CDC 2A374N A-10/F-15/F-16/U-2 Avionic Systems Craftsman Volume 2. F-15A–D/F-15E Avionic Systems

___________

Air Force Career Development Academy

Air University Air Education and Training Command

2A374N 02 1310, Edit Code 01 AFSC 2A374

Material in this volume is reviewed annually for technical accuracy, adequacy, and currency. For SKT purposes the examinee should check the Weighted Airman Promotion System Catalog to determine the correct references to study.

Author: MSgt Michael R. Ratliff 365th Training Squadron USAF Technical Training School (AECT) 365TRS/TRR 609 9th Avenue Sheppard Air Force Base, Texas 76311-2335 DSN: 736-6054 E-mail address: cdc.2a354@sheppard.af.mil

Instructional Systems Specialist:

Evangeline K. Walmsley

Editor: Nelva J. Brown

Air Force Career Development Academy (AFCDA) Air University (AETC) Maxwell-Gunter Air Force Base, Alabama 36118–5643

Preface ___________________________________________________________________ i

This second volume of CDC 2A374N, A-10/F-15/F–16/U-2 Avionic Systems Craftsman, contains the basic knowledge required to upgrade to the 2A374N Air Force specialty code (AFSC). The subjects covered in this volume range from the F-15 attack control systems and instrument and flight control systems to communications, navigation, and penetration aids systems. Unit 1 of this volume covers F-15 radar, integration, and display systems. Unit 2 covers F-15 instrument, flight control, engine air intake, and air data systems. Unit 3 covers F-15 data link, electronic warfare systems and PODS used on the F-15E aircraft. A glossary of abbreviations and acronyms used in this course is included at the end of each volume. Code numbers on figures are for preparing agency identification only.

The use of a name of any specific manufacturer, commercial product, commodity, or service in this publication does not imply endorsement by the Air Force.

To get a response to your questions concerning subject matter in this course, or to point out technical errors in the text, unit review exercises, or course examination, call or write the author using the contact information on the inside front cover of this volume.

NOTE: Do not use the IDEA Program to submit corrections for printing or typographical errors.

Consult your education officer, training officer, or NCOIC if you have questions on course enrollment, administration, or irregularities (possible scoring errors, printing errors, etc.) on unit review exercises or course examination. For these and other administrative issues, you may email the Air University e-Campus Support (helpdesk) at au.ecampussupport@maxwell.af.mil.You should receive a response in four days or less.

This volume is valued at 27 hours and 9 points.

ii __________________________________________________________________ Preface

NOTE:

In this volume, the subject matter is divided into self-contained units. A unit menu begins each unit, identifying the lesson headings and numbers. After reading the unit menu page and unit introduction, study the section, answer the self-test questions, and compare your answers with those given at the end of the unit. Then, do the unit review exercises.

Contents _________________________________________________________________ iii

Page

Unit 1. Attack Control Systems ..........................................................................................1–1 1–1. Radar Systems ............................................................................................................ 1–1 1–2. Integration Systems .................................................................................................. 1–32 1–3. Display Systems ....................................................................................................... 1–49

Unit 2. Instrument and Flight Control Systems ................................................................2–1 2–1. Instrument Systems .................................................................................................... 2–1 2–2. Primary Flight Control Systems ............................................................................... 2–16 2–3. Automatic Flight Control Systems ........................................................................... 2–26 2–4. Engine Air Intake System ........................................................................................ 2–60 2–5. Air Data Systems...................................................................................................... 2–68

Unit 3. Communications/Navigation/Penetration Aids Systems .....................................3–1 3–1. Data Link Systems ..................................................................................................... 3–1 3–2. Electronic Warfare Systems ..................................................................................... 3–20 3–3. PODS ....................................................................................................................... 3–35

Glossary............................................................................................................................................. G–1

Unit 1. Attack Control Systems 1–1. Radar Systems .......................................................................................................................... 1–1

201. APG-63 radar ..................................................................................................................................... 1–1 202. F-15C/D APG-70 radar set .............................................................................................................. 1–16 203. F-15E APG-70 radar set .................................................................................................................. 1–20 204. APG-63 (V)1 radar .......................................................................................................................... 1–24

1–2. Integration Systems ................................................................................................................ 1–32 205. Overload warning system ................................................................................................................ 1–32 206. Central computer complex ............................................................................................................... 1–38 207. Advanced display core processor .................................................................................................... 1–43

1–3. Display Systems ...................................................................................................................... 1–49 208. Video tape recording system ........................................................................................................... 1–49 209. F-15E digital map system ................................................................................................................ 1–54

N THIS unit, you will cover information on the F-15 attack control systems that you maintain. This information is not classified, and classified information is not discussed in this career development course (CDC). You will start off learning about the various radar systems on the F-

15 C/D and E model aircraft. Then we will discuss the integration systems on the F-15 C/D and E aircraft. Finally we will wrap up this section by learning about a few of the display systems you will troubleshoot on the flightline.

1–1. Radar Systems We begin this section by giving you a description of the APG-63 radar. Then we will discuss the more advanced APG-63(V)1 radar. We conclude the section with a discussion on the APG-70 C/D and E radar sets.

201. APG-63 radar As a 7-level you will spend a great deal of your time troubleshooting radar if you work the APG-63 system. It is important that you know all of the system components and understand what they do and how they communicate with each other. In this lesson you will learn the system components and basic theory of operation.

System components The APG-63 radar set is made up of nine major line replaceable units (LRU), several minor LRUs, and waveguides. See figure 1–1 for the names and locations of the major LRUs. Each part has a numerical nickname, which comes from the part number, plus the descriptive name. As a 7-level you need to know the numerical nickname and the descriptive name of each radar part.

Major radar LRUs We will discuss and describe the APG-63 radar system’s major LRUs in the following paragraphs.

Radar control panel (541) The radar control panel located in the cockpit left console provides manual control of radar set power, operating modes, and mode parameters. All outputs from the various control switches are sent to the data processor (081) except for the POWER switch, which also sends power control commands to the power supply (610) and transmitter (011).

I

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Figure 1–1. APG-63 radar control panel (541) and general system layout.

Power supply (610) The power supply located in door 3L converts and conditions aircraft power to the low-voltage forms required by the various LRUs of the radar set. The 610 also contains the antenna (031) servo-electronics that respond to commands coming from the data processor (081). These servo-electronics drive the corresponding hydraulic and electrical mechanisms in the 031 to produce movement.

Radio frequency oscillator (001) The radio frequency (RF) oscillator (RFO) located in door 3L establishes the basic operating radio frequencies for the transmitter based on data processor (081) commands. These are the “channels” which are selected either manually on the radar control panel (541), or automatically by the central computer (CC). The RFO provides the frequency to the transmitter (011). It also provides the local oscillator signal to the receiver (022). The local oscillator, or LO, is always set at 30 MHz (megahertz) above the transmitted frequency.

Transmitter (011) The transmitter, located in door 3L, uses a gridded traveling wave tube (GTWT) to amplify the low power RF signal from the RFO (the GTWT drive signal). The transmitter then couples the high power RF to either the antenna (031), flood horn (019), or into a dummy load for ground operation. The dummy load is inside the 011. The primary output source is the antenna (031). Additionally, the transmitter provides a low-power sample of the radar signal to the missile auxiliaries for missile tuning during radar time-in and at the start of an AIM-7 Sparrow missile launch. The transmitter contains a duplexer (four-port circulator) that couples low power return signals to the receiver. It also has a multipactor, which prevents high power RF damage to the receiver (022).

Receiver (022) The receiver, located in door 3L, amplifies the RF returns, converts them to 30 MHz intermediate frequency (IF) signals, amplifies them again, and sends them to the analog processor (039). The receiver processes signals in two separate channels: main and guard. RF energy from the main and guard channels are processed in search modes. If the radar is in a track mode, the receiver provides the initial processing of the sum and difference signals. These signals are also passed along to the 039.

Analog processor (039) The analog processor, located in door 3L, receives two 30 MHz IF signals from the receiver (022) and frequency-shifts them to put clutter at a specific frequency for processing (clutter is discussed later).

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Both signals are filtered and processed in one of four IF channels, converted from analog-to-digital, and sent to the programmable signal processor (042) for digital processing.

Data processor (081) The data processor, located in door 3L, is a stored program digital computer that performs radar set management, radar control, selected radar data processing, and radar set performance monitoring and measurement. It stores the radar operational flight program (OFP). The 081 also provides 031 positioning commands to the 610. The data processor is the only radar LRU connected to the H009 or 1553 multiplex (MUX) bus, and is the only radar LRU that communicates with the CC. It acts as the interface between the radar and other avionic systems.

Programmable signal processor (042) The programmable signal processor, located in door 3L, is a digital processor. The 042 performs target detection, filtering, range measurement, target parameter calculations, clutter canceling, and identification friend-or-foe (IFF) correlation. It also provides radar display parameters to the indicator group (IG).

Antenna (031) The antenna “planar array assembly” (fig. 1–2), located in the nose radome, receives high-power, X-band, RF energy from the transmitter and radiates a narrow, vertically polarized pencil beam for target illumination. The planar array is the flat, circular portion of the antenna that makes up the ends of a large network of waveguides. Radar echoes received by the antenna are routed through the microwave circuitry (waveguides) to the receiver. The 031 moves on a three-axis system with motion in azimuth, elevation, and roll. The azimuth and elevation axes are hydraulically controlled, while the roll axis is electrically controlled. The data processor (081) sends antenna positioning commands to the servo electronics in the power supply (610). These commands produce the azimuth and elevation hydraulic valve control signals and roll torquer drive signals for the antenna. Also located on the 031 are the guard horn and null horn (discussed later). Dipoles, mounted on the planar array, radiate and receive energy for air-to-air interrogations/identification friend or foe (AAI/IFF). The radome that houses the antenna is a ceramic-fiber, sandwich-type construction that does not disrupt the radar transmission.

Figure 1–2. Antenna (031) (back view).

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Non major LRUs Now let’s discuss and describe other LRUs of the APG-63 radar system.

Flood horn (019) The flood horn (fig. 1–3) is essentially the termination of a waveguide that radiates RF energy in a large fan-shaped pattern in front of the aircraft. If the radar is unable to angle-track due to a malfunctioning antenna, the pilot can select FLOOD on the radar control panel (541). This will cause the flood horn to radiate, providing a backup, range-only tracking ability for gun attack. (The wide pattern can enhance rapid detection of targets within the beam at close range.) The 019 is located near the top of the bulkhead behind the 031, under the radome. A plastic-type material called Mylar covers its opening and prevents loss of waveguide system pressurization. Mylar also covers the planar array surface as well as the guard and null horns.

Figure 1–3. The flood horn (019).

Waveguides RF energy at microwave frequencies using standard RF cables produces unacceptable amounts of power loss. Waveguides reduce this problem (fig. 1–4). Waveguides connect microwave LRUs (the 031, 011, 022, and RFO).

Waveguides are susceptible to dents, cracks, corrosion, contaminants, and misuse of seals. Many radar problems are inadvertently caused by “maintenance-induced” faults, like those resulting from misuse of waveguides. Cracks, corrosion, and incorrect seals can cause changes in the impedance of a waveguide. This causes problems such as weak detection, noise, or even birds. Birds is a term used to describe false targets on the display, usually in an arc or a straight line.

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Figure 1–4. Waveguides.

Indicator group The IG is a video display system (fig. 1–5) made up of the programmable signal data processor (PSDP) and the multiple indicator control panel (MICP). The IG converts both analog and digital input signals for displays. It is capable of displaying radar, TV weapons, or miscellaneous displays as selected by the CC.

Programmable signal data processor The PSDP (fig. 1–5) receives serial digital data from the CC and radar set, analog and discrete signals from the radar set, and a composite video signal from the programmable armament control set (PACS). The PSDP processes the information and transmits it to the MICP. On the F-15D, the information also goes to the rear cockpit MICP.

Multiple indicator control panel The MICP (fig. 1–5) receives deflection and symbology data from the PSDP for display on a cathode ray tube (CRT). The MICP also contains the controls for operating the IG. The F-15D has a front and rear cockpit MICP. The two MICPs are identical. On the flightline, the MICP is usually called a VSD, or vertical situation display.

Figure 1–5. MICP and PSDP.

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In addition to the radar components and the indicator group components, there are several LRUs that are associated with the radar. Some of these are: stick grip, throttle grips, avionics status panel, built-in test (BIT) control panel, lights panels, engine control panel, multipurpose color display (MPCD), head-up display (HUD), video tape recording system (VTRS) components, and the joint tactical information display processor group.

System operation Now that you understand the components of the APG-63 radar system, let’s dive into the system theory of operation.

Transmitting function The RFO (001), transmitter (011), antenna (031), flood horn (019), and missile auxiliaries perform the transmit function (fig. 1–6). The radar control panel CHAN switch sends signals to the 081 for frequency selection. Frequencies can be selected either manually by choosing 1 through 6 or automatically by choosing A.

The CC controls what frequency is selected during automatic frequency selection. The 081 sends the frequency selection to the RFO, where it is generated. The RFO sends this GTWT drive signal to the transmitter through a waveguide. The GTWT amplifies and pulses the low-power RF from the RFO. The high-power RF is then sent to a four-port circulator. The circulator, with the multipactor, provides isolation of the receiver from the GTWT source.

During transmission, the multipactor is activated to prevent the high-power RF from leaking into the receiver. If it were not for the multipactor, the 022 would burn out due to radar main bang, which is too much power entering the receiver. In addition, the signal travels from the circulator to a three-position WAVEGUIDE switch. This switch directs the high-power RF to the antenna (031), flood horn (019), or into a dummy load (labeled “D.L.” on fig. 1–6).

Figure 1–6. Transmitting function.

Under most circumstances, the antenna is used for radiation of the energy. If the radar is unable to angle track a target (the antenna mechanics go bad), the flood horn radiates to provide range-only tracking ability for a gun attack.

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The final place the high-power RF can go is the dummy load. It is used during ground operation to prevent radiating the area around the aircraft. The dummy load is inside the 011 and it is controlled by the position of the PROXIMITY switch (weight-on-wheels switch) on the right main landing gear.

Sidelobes and ground returns When a radar system emits RF energy from an antenna, sidelobes are always produced (fig. 1–7). Sidelobes radiate in all directions from the F-15 antenna, but at lower power than the main beam. If a sidelobe reaches the ground and produces a return, we might receive it. We call this type of return ground clutter. Of course, this poses a problem for the radar, because now the returns from true targets may become lost in the ground clutter. Using the guard horn, the radar is designed to reduce this effect.

Figure 1–7. Sidelobes and the main beam.

The antenna has a planar array (the flat, round part), a guard horn, and a null horn (fig. 1–8). The guard horn is a small, wide-angle horn located at the bottom of the antenna, angled down slightly. It is a receive-only horn designed to intercept sidelobe echoes. With the ability to measure Doppler shift, we can use these ground returns to measure our ground speed. Knowing this, the radar can filter out all main-beam returns that come back with the same Doppler signature, as they must be just returns from the ground ahead. This process is called clutter rejection and occurs in the 042.

Figure 1–8. Antenna (031) planar array.

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Receiving function The antenna (031), transmitter (011), RFO (001), and receiver (022) perform the receiving function. Radar operations are divided into two broad categories; search and track. Search is when the radar is sweeping the skies, looking for targets. Track is when the radar has found a likely target and is pointing directly at it, following the target along its flight path.

Search modes In search modes, the antenna receives main beam returns and routes them to the 011(fig. 1–9). They travel through the deactivated multipactor and into the 022. The sidelobe energy received through the guard horn is sent directly to the 022. The 022 takes the two signals and creates two 30 MHz IF signals using the LO signal from the RFO (heterodyne process). The 30 MHz IF signals retain all of the information of the received signals, but at a frequency that is easier to process. Two triaxial cables transfer the IF signals to the 039 for processing.

Figure 1–9. Receiving function during search.

Track modes In track modes the antenna tracks (follows) the target by comparing the levels of energy received through the four quadrants of the 031 planar array (fig. 1–10). Reception in the top and bottom pairs determines the elevation-difference signal, while reception in the left and right pairs determines the azimuth-difference signal. The 031 points directly at a target when all four quadrants receive equal amounts of energy. The azimuth (AZ) and elevation (EL) difference signals are sent directly to the 022. Additionally, all four quadrants are added together and sent to the 011. This signal goes through the deactivated multipactor to get to the 022. From here, the receiving function is about the same. The 022 creates two 30 MHz IF signals to send to the 039 for processing.

Signal processing function The analog processor (039) and the programmable signal processor (042) perform the signal processing function. First, the 039 takes in the 30 MHz IF signals from the 022, filters them, and converts them to even lower frequencies. Then it converts them to digital and passes the information to the 042. The 042 is responsible for processing the data to obtain useful information, such as the target’s position, range, and Doppler shift. It also eliminates clutter and false targets. Additional 042 functions include providing the timing for the radar package, developing the radar display data that is

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sent to the indicator group, and interfacing with the IFF reply evaluator, a part that is used to help identify targets as friend or foe.

Figure 1–10. Receiving function during track.

Data processing function The data processor (081) is the only LRU involved in the data processing function. It does the radar set functions through digital, analog and discrete interfaces with the radar units. With a few exceptions, all communications between the radar set and other aircraft systems are through the data processor. The 081 stores the radar OFP, downloading portions of it into the 042 as needed. It also controls processing for the BIT function. Finally, to efficiently process the large volume of data that’s continually transferred between the radar, head-up-display, and navigation (NAV) systems (inertial navigation system (INS) and attitude heading reference system (AHRS)), the 081 controls to the high-speed radar data bus. This bus is completely separate from the H009 and 1553 buses.

Antenna control function Servo electronics within the 610 receive azimuth, elevation, and roll drive commands from the 081 to produce the azimuth and elevation hydraulic control signals and roll torquer drive signals for the antenna. The 031 azimuth and elevation are hydraulically driven, while roll is driven electrically. The 031 takes in hydraulic fluid from the utility system and restricts it to 1500 pounds per square inch (psi) for 031 use.

Cooling function Two aircraft cooling inputs control the temperature of the radar set: air and liquid. The 001, 011, 022, 042, 610, 081, and 039 all require cooling air. (Hint: An easy way to remember this is that all radar LRUs under door 3L require air.) The antenna and radar control panel do not require cooling air. During cold weather, the environmental control system (ECS) can supply warm air.

The ECS light on the caution lights display panel (fig. 1–11) will come on when cooling airflow or temperature problems occur. Inadequate cooling air will cause the radar to turn off. The transmitter (011) is the only unit that requires liquid cooling.

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Figure 1–11. ECS and AV BIT lights.

Waveguide pressurization function Pressurization prevents high voltage component arcing at high altitudes and keeps moisture and contamination out of the air used in pressurized RF components. The transmitter, receiver, antenna, and flood horn are pressurized. The source of pressure is bleed air from the ECS system. If waveguide pressure falls below 12 psia (pounds per square inch absolute) in flight, avionics status panel (ASP) 25 will latch (turn orange).

Built-in test The purpose of radar BIT is to detect a radar system fault and isolate it to an LRU. The types of BIT are described below.

The 081 stores the BIT program as well as the BIT matrices. A matrix is simply a database of failures. It helps maintainers decide which LRU needs replacement. There are six types of APG-63 radar BITs:

Power up-tests The primary purpose of the power-up test sequence is to do system tests and calibrations on the radar set. This establishes a high level of confidence that the radar is operational. Switching the radar control panel directly from OFF to OPER starts an operational readiness test (ORT) that performs most BIT checks prior to calibrating the system. If failures are detected during ORT, then the pilot can call for a technician to troubleshoot it. Failures during power-up are recorded in the continuously-monitored built-in test (CM BIT) matrix.

Continuous monitoring Radar set performance is continuously monitored (CM) by BIT. The RDR (RADAR) light on the BIT control panel (BCP) and the AV BIT (AVIONICS BIT) light on the main instrument panel indicate failures. Failure data is stored in the CM BIT matrix. The CM BIT is done in all modes of operation except at power up. This test is non-interruptive.

Automatic calibration The radar set requires various calibrations to continue functioning at peak performance. These calibrations are performed automatically by the radar set and eliminate the requirement for scheduled maintenance. They occur periodically or at major mode changes.

Operate initiated BIT Initiating a BIT, while in operate (operate initiated BIT (IBIT)), discontinues normal radar operations and starts all possible tests of the radar set. Operate IBIT is divided into two categories: ground and

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airborne. The type of test done depends upon the right main landing gear (MLG) weight-on-wheels switch position. When IBIT is running on the ground, G-TEST is displayed in the BIT window. If TEST appears, the radar thinks it is airborne and transmits an obvious hazard. IBIT also includes a track-test option. It generates synthetic targets that can be displayed, acquired, and tracked in various modes. See figure 1–12 for a view of the BIT control panel and the RDR and VSD lights/system selections.

Figure 1–12. BIT control panel.

Standby IBIT Faults detected by the radar are displayed during standby IBIT (fig. 1–13). It displays a plain English readout of the IBIT matrix and CM-BIT matrix. The matrices tell the maintainer what the primary failed LRU is and exactly which tests failed. A secondary LRU is listed for most faults.

Figure 1–13. Typical STBY IBIT display.

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Power-off tests Power-off tests are simply tests done during normal shutdown of the radar. From the time the radar POWER knob is turned to OFF, the radar performs its final tests and will latch the appropriate fault indicators. If abnormal power-down occurs, such as premature engine shutdown, the LRU fault indicators are considered suspect because the radar shutdown sequence did not have time to complete.

Pulse repetition frequency operation The F-15 radar uses high pulse repetition frequency (HPRF), medium (MPRF), and interleaved (HPRF/MPRF) PRF operation. The type used depends on the radar mode selected. In long-range search mode, the PRF can be selected manually. The BIT window displays which PRF is in use. The different PRFs have advantages and disadvantages. The 541 allows selection of 10, 20, 40, 80, and 160 NMI (nautical miles) range scales. HPRF is not allowed at the 10 NMI range, and MPRF is not allowed at the 160 NMI range, except as part of the interleaved PRF waveform.

HPRF operation A HPRF radar mode provides maximum detection against nose-aspect targets. Its high average power results in long-range detection and tracking capability. (The close spacing of the pulses means the transmitter is ON a lot, compared to its listening time (fig. 1–14).) A problem with HPRF is that it has poor performance against tail aspect targets and near zero capability against targets with no closure rate. HPRF is available in all ranges except 10 NMI.

Figure 1–14. Low, medium, and high PRF.

MPRF operation The radar uses MPRF for all-aspect-angle detection. It has longer listening time than HPRF and determines range-to-target very accurately (fig. 1–14). However, it has lower average power than HPRF and cannot be used by itself at the 160 NMI range. It can be used as part of the interleaved PRF waveform at the 160 NMI range. The ranges available to all-MPRF are 10 through 80 NMI.

Interleaved HPRF/MPRF operation This is actually a dual-mechanization type of scan that switches between HPRF and MPRF during the scan pattern (see fig. 1–15). It provides good performance balance with the combination of MPRF and HPRF operation. Each mode has strengths that complement the other’s weaknesses. Targets not picked up by one PRF, will be picked up in the other. All range selections are available. However, the radar will automatically switch to all-MPRF at the 10 NMI range. If 160 NMI is selected, the radar will use either all HPRF or interleaved PRF at the pilot’s discretion.

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Figure 1–15. Interleaved PRF.

Low pulse repetition frequency operation The radar only uses low pulse repetition frequency (LPRF) during the air-to-ground (A/G) ranging mode (discussed later).

Air-to-air mode of operation Radar modes are categorized into four basic mode groups: search, acquisition, track, and track-while-scan (TWS). Typically, a pilot sweeps the skies in a search mode. When the pilot finds a target to track, he/she acquires it using either automatic or manual acquisition (the 031 points directly at the target and tries to get a couple of good returns from it). If that happens, the radar automatically enters track. Finally, if the pilot wants to continue searching for other targets while tracking the first one, he/she can enter track-while-scan or TWS. Below describes the air-to-air (A/A) mode groups.

Search Let’s discuss the basic search modes used during radar set maintenance (fig. 1–16):

Long-range search Long-range search (LRS) is the most versatile of the search modes and is used in most tactical situations. It is the primary A/A search mode and is used for detection and acquisition of both closing- and opening-rate airborne targets. PRF is selectable using the PRF menu, and all ranges are available.

Vector scan Vector scan mode detects small targets by slowing the 031 scan-rate. A greater number of pulses are sent out in a particular direction, increasing the number of returns from a target. All HPRF is used, and all ranges are selectable.

Velocity search Velocity search allows the radar to detect high closing-rate targets by using only HPRF without ranging. All range scales are selectable.

Short range search Short-range search (SRS) uses all MPRF to detect all-aspect targets from 10 to 80 NMI ranges.

Pulse mode The pulse mode operates the same as SRS look down conditions. In look up conditions, when ground clutter is not a factor, the clutter rejection circuitry is bypassed. This allows detection of very small targets. It uses all MPRF and is good out to 80 NMI.

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Figure 1–16. Typical search display (B-scan format).

Acquisition Acquisition is the process of picking out a target to track. The two ways to acquire a target are manually and automatically. During manual acquisition, the operator selects which target to track (fig. 1–17). The target designator control (TDC) on the right throttle grip moves the acquisition symbol around on the radar display; similar to a computer mouse cursor. The acquisition symbol looks like a pair of captain’s rank insignia (fig. 1–16). Pushing and releasing the TDC will cause the radar to scan the area around the acquisition symbol in a 6 degree mini-raster pattern. If it gets two live hits, the radar enters a track mode, storing the information in a track-file. The auto acquisition group of modes does just what the name implies; it automatically acquires a target, the first one it hits. There are six types of modes, and each is tested during the checkout except flood. Most are selected using the AUTO ACQ/REJECT switch on the control stick grip (fig. 1–17).

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Figure 1–17. Acquisition switchology.

Boresight The boresight (BST) mode causes the antenna to cease scanning, move to the aircraft boresight line (straight ahead), and stop. The first target (shortest range) to cross through the main beam is locked-on to (acquired), and the track function begins. Boresight uses all MPRF and is good out to 10 NMI. Pressing the AUTO ACQ/REJECT switch forward twice in less than one second selects boresight.

Long range boresight Long range boresight (LRBST) acts exactly the same as boresight does. The only difference is that long-range boresight uses interleaved PRF and has a range of 40 NMI. Holding the AUTO ACQ/REJECT switch forward for longer than one second selects long-range boresight.

Supersearch Supersearch (SS) has a scan pattern that approximates the HUD field of view: about 20 degrees by 20 degrees. The radar locks-on to the first target that falls within the scan pattern out to a range of 10 NMI. Supersearch uses all MPRF and is considered the primary automatic acquisition mode. Pressing the AUTO ACQ/REJECT switch forward once in less than one second selects supersearch.

Vertical scan Vertical scan (VTS) causes the antenna to scan up and down in a very narrow and very tall pattern (+5 to +55 degrees of elevation and only two bars wide). The first target to fall into the pattern is locked on to. Vertical scan uses all MPRF and is good out to 10 NMI. Pressing the AUTO ACQ/REJECT switch aft for less than one second selects vertical scan.

Gun scan The gun scan mode automatically acquires targets detected within a 60 degree by 6 bar scan, from 0.5 to 15 NMI. It uses all MPRF.

Flood mode The flood auto acquisition mode is a backup mode that provides range-only ability for guns support. It is selectable by placing the 541 SPL MODE knob to FLOOD and pressing the AUTO ACQ/REJECT switch forward.

Single target track Single target track (STT) is entered after a target is acquired, either manually or automatically. The target data is kept in a track file, which contains the target’s range, angles, and range-rate.

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Track-while-scan TWS modes provide a wide-angle coverage, multi-target detection and track capability. The TWS modes maintain up to 10 target track files while continuing to detect and display up to 18 more (half-intensity) observation targets.

Air-to-ground mode of operation A/G modes are selected automatically by the CC or manually by the 541 MODE SEL switch. Although there are three A/G selections available (DPLR (Doppler), RNG (ranging), and MAP), only two (DPLR and RNG) are operational. Selection of MAP will put the radar in A/A LRS.

A/G ranging mode The A/G ranging mode uses LPRF to determine the slant range from the aircraft to a ground point. It is used primarily for bombing/gun strafing. The display is a 10 NMI PPI (pixels per inch) format (fig. 1–18).

Figure 1–18. The A/G ranging display (PPI format).

DPLR mode The Doppler mode performs a navigational update function. It provides north and east velocity corrections to the CC to correct long-term velocity errors developed by the INS. It is recommended that pilots use this function prior to making bombing runs. This selection has no display.

202. F-15C/D APG-70 radar set The APG-70 radar operates similarly to the APG-63 radar. There are only a few differences in the LRUs, modes, and the BIT system. It performs sophisticated A/G functions like standard and high-resolution ground mapping. Currently, the majority of the APG-70 packages are in F-15E aircraft. In this lesson, we will highlight the major differences between the F-15C/D model APG-63 and APG-70 radar systems. We will then discuss the F-15C/D model APG-70 radar set operation.

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System components The APG-70 radar set is made up of eight major LRUs, several minor LRUs, and connecting waveguides. Most of the units are located under door 3L in the forward left nose equipment bay (fig. 1–19).

Figure 1–19. APG-70 radar control panel (542) and general radar layout.

Radar control panel (542) The 542 performs the same role as the APG-63’s 541 did, but it is designed to work with the new modes available in the APG-70. It is located on the left console of the front cockpit.

Receiver-exciter (025) The receiver-exciter combines the functions of the APG-63 receiver and RFO into one LRU. It establishes the basic operating radio frequencies for the transmitter based on radar data processor (082) commands. When returns are received, the 025 amplifies them with very low noise. It also produces an internal LOCAL OSCILLATOR signal set at 30 MHz above the transmitted frequency. This is heterodyned with the incoming returns to develop the two 30 MHz IF signals (main and guard in search modes, and sum and difference in track modes.) The IF signals are sent to the analog signal converter (038) for processing.

Other APG-70 parts The APG-70 parts shown in the table have the same function as their counterparts in the APG-63 system:

PART APG-70 NUMBER APG-63 NUMBER Transmitter 111 011 Analog Signal Converter 038 039 Radar Data Processor 082 081 Programmable Signal Processor 044 042 Antenna (same part) 031 031 Flood Horn (same part) 019 019 Power Supply (same part) 610 610

System operation The following paragraphs describe the APG-70 radar system theory of operation.

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Setup Look at the waveguide setup for the APG-70 radar system illustrated in figure 1–20. The waveguide that carries the GTWT drive signal is between LRUs 025 and the 111. The microwave LRUs are now the 031, 111, and 025.

Figure 1–20. APG-70 waveguide routing.

Range gated high PRF The APG-70 uses HPRF, MPRF, and LPRF waveforms for transmissions. It also uses range gated high PRF. Range gated high PRF is an intermediate PRF that falls between HPRF and MPRF. The pulses are spaced a little farther apart than in HPRF, but not quite as far as in MPRF. This produces some of the desirable effects of both the old waveforms.

APG-70 operating modes The APG-70 has several operating modes which we’ll discuss here.

Air-to-air search modes The following are available A/A search modes:

1. HI. This selects the HPRF range-while-search mode, which will search in the 20, 40, 80, and 160 NMI ranges. Its use of all high PRF means that tail aspect targets are very unlikely to be picked up, but performance against nose-aspect, long-range targets is outstanding.

2. MED. This selects the MPRF range-while-search mode. It uses all MPRF to detect all-aspect targets from 10 to 80 NMI. Selection of 160 NMI is not allowed.

3. INLV. This selects the interleaved PRF range-while-search mode. It is available in all range scales and has a PRF menu available for the operator to change PRF at his or her discretion.

4. RGH. This selects the range gated high PRF range-while-search mode. The RGH mode is used in intermediate and short range, lookdown, and clutter environments to detect both opening and closing targets.

Manual and automatic acquisition Acquiring targets with the APG-70 is the same as with the APG-63. During manual acquisition, the operator uses TDC to position the captain’s bars over a target, then presses and releases the TDC to

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acquire the target and establish a track file. AUTO acquisition still includes the boresight, long-range boresight, supersearch, vertical scan, gun scan, and flood modes found in the APG-63 section. Each auto-acquisition mode uses MPRF, except for long-range boresight, which uses interleaved PRF.

Track The APG-70 still tracks in the single-target-track format when performing the radar operational checkout on the ground. It is initially entered from either manual or automatic acquisition. TWS may be entered from STT during actual operation.

Track-while-scan The TWS modes in the APG-70 are similar to those found in the APG-63, in that they can maintain 10 target track files plus 18 observation targets.

Air-to-ground modes The APG-70 radar system makes four A/G modes available to the pilot. All of them use LPRF waveforms. Only A/G ranging, A/G beacon (BCN), precision velocity update (PVU), and real beam map (RBM) are discussed.

A/G ranging This mode determines the slant range from the aircraft to the ground for bombing. The display is a 10 NMI PPI scan like the APG-63 A/G ranging mode uses (fig. 1–18).

A/G beacon This mode is a navigational mode that transmits interrogations, processes replies, and displays positions of ground-based radar beacons. It operates in all ranges and has a one-bar, 100 degrees wide PPI-type scan (10 degrees is allocated to either side for drift compensation).

Precision velocity update The PVU mode provides radar-measured velocity that compensates for errors in INS velocities. There is no display and range is not applicable, the antenna looks down at different points on the ground to gather velocity data. Pilots utilize this function prior to making a bomb run.

Real beam map The RBM mode provides conventional A/G functions of large-area mapping and general navigation. The presentation is in a 100 degree PPI scan format, with 10 degree on each side for drift compensation. All range scales are available. All LPRFs are used.

Built-in test The APG-70 BIT has a persistence counter and an enhanced matrix of detectable failure modes.

Automatic modes of BIT There are two automatic modes of BIT that require no operator action. They are the operational readiness test and the continuous monitor.

Operational readiness test The ORT BIT mode does system tests and automatic calibrations to establish a high level of pilot confidence that the radar will perform its mission. It happens during cold-start power-ups either on the ground or in the air. As the test occurs, the LRU under test will be displayed in the BIT window on the lower left of the display. If the test stops on a particular LRU, then it is likely that LRU has failed a test.

Continuous monitor The CM BIT runs in the background when the radar is ON. It is non-interruptive. Failures are stored in the CM BIT matrix.

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Operator initiated BIT The operator initiated BIT is divided into G-BIT and F-BIT. Each of these may be divided into standby or operate modes depending on the position of the RADAR POWER switch. The radar BIT type (G- or F-BIT) run depends on the state of the right MLG weight-on-wheels (WOW) switch.

Operator initiated G-BIT The operator initiates this BIT on the ground. If the RADAR POWER switch is in OPR when BIT is initiated, the radar ceases all operations and performs every possible test. If the POWER switch is in STBY when BIT is initiated, the radar ceases all other functions, latches fault indicators, and displays the BIT matrices on the MICP.

Operator initiated F-BIT The operator initiates this BIT in the air. If the RADAR POWER switch is in OPR when BIT is initiated, the radar ceases all operations and performs every possible test. If the POWER switch is in STBY when BIT is initiated, the radar again ceases all other functions, sets fault indicators, and displays the BIT matrix on the MICP.

BIT history matrices The APG-70 BIT has five BIT history matrices available:

1. ORT matrix. 2. GBIT matrix. 3. FBIT matrix. 4. CM matrix. 5. FRESH matrix.

The FRESH matrix contains the latest known test results. Performing standby- IBIT will display the matrices. The AUTO-ACQ/REJ switch cycles between the five matrices.

A persistence counter tracks the number of CM BIT failures. The failures are not used in fault isolation if their persistence count is less than seven in one minute. Not all CM BITS will be disregarded with less than seven failures. You must refer to the Fault Isolation TO for more information. Seven or more failures are considered high persistence, and these failures are used in fault isolation.

203. F-15E APG-70 radar set The F-15E is equipped with the APG-70 radar set. The system searches for, acquires, and tracks airborne targets while providing a clutter-free display of all radar information. The system provides A/A the same as the F-15C. It also provides air to ground mapping, ranging, and a radar beacon mode in both air-to-air and air-to-ground operation.

System components The radar set is made up of seven major LRUs and connecting waveguides.

Transmitter (111) The transmitter (111) amplifies the low power RF signal from the receiver-exciter (025) and couples high-power RF through the waveguide sum channel to the antenna (031) or flood horn (019) for radiating into space, or into a dummy load for ground operation.

Receive- exciter (025) Amplifiers in the receiver section provide low noise amplification of the returns in the main/sum and guard/difference channels. The receiver-exciter (025) mixes the returns with a local oscillator frequency to produce an IF frequency which it amplifies.

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The exciter section of the receiver-exciter (025) establishes the basic operating RF for the transmitter (111) using radar data processor (082) commands. The exciter section provides the transmitter (111) drive signal, drive signal modulation and receiver section local oscillator signal.

Power supply (610) The power supply (610) performs the same functions as the F-15C APG-63 APG-70 power supply.

Analog signal converter (038) The analog signal converter (038) receives the main/track 1 IF and guard/track 2 IF signals from the receiver-exciter (025) and processes the IF inputs. It converts the IF analog data to digital information which is then sent to the programmable signal processor (044) for processing.

Antenna (031) The antenna (031) planar array assembly performs the same functions as the F-15C APG-63/APG-70.

Radar data processor (082) The radar data processor (082) is a stored program digital computer which does radar set management, radar control, selected radar data processing, and radar set performance monitoring and measurement. Functions of the radar data processor (082) include input/output signals handling and conditioning of interface signals between LRUs of the radar set and other avionic systems.

Programmable signal processor (044) The programmable signal processor (PSP) (044) receives digital data and controls from the analog signal converter and the radar data processor (RDP). The PSP also does data preprocessing, altitude line and main beam clutter rejection, subdivision of data into discrete frequency bands, range/velocity and angle tracking, high resolution map or real-beam map processing, and BCN return processing.

Air-to-air mode of operation The AN/APG-70 radar mechanization makes use of three basic waveforms, MPRF, HPRF, and range gated high (RGH), plus fallout interleaved HPRF and MPRF waveforms. A/A modes are manually selectable by switch activation on the multipurpose display/multipurpose color display (MPD/MPCD) along with the throttle, stick grip, and hand controllers, or automatically by the advanced display core processor (ADCP). The A/A modes are made up of the mode groups search, manual and auto acquisition, track, and multi-target track.

Air-to-ground mode of operation The A/G modes of operation are the primary operating modes of the F-15E. We will begin our discussion with a brief description of the following A/G modes:

• RBM mode. • Ground moving target (GMT) mode. • High resolution map (HRM) mode. • PVU mode. • Air-to-ground ranging (AGR) mode. • A/G BCN mode. • A/G backup mode.

Real beam map mode RBM mode is used to provide a conventional mapping mode for low-resolution ground mapping, weather returns or for HRM cueing.

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Ground moving target mode Refer to figure 1–21 as you read the following. The GMT mode is used for the detection of ground moving targets (trucks, convoys) out to a maximum of 32 NMI. Targets are displayed as the antenna sweeps across them and are erased at the start of the next GMT frame.

Figure 1–21. MPD/MPCD RDR display.

High resolution map mode The HRM mode is used to provide a high-resolution map for A/G weapons delivery, high range resolution, and azimuth resolutions by employing synthetic aperture radar (SAR) techniques.

Precision velocity update mode PVU mode is used to update the system velocity in mission navigator (MN) and/or INS. INS velocity update is only performed if errors are suspected.

Air-to-ground ranging mode AGR mode provides a slant range measurement used for target altitude determination, for target designation, or for position updates. AGR is not directly selectable. The mode is entered automatically when the system determines that a slant range measurement is required, unless the radar is supporting a missile in flight.

Air-to-ground beacon mode In A/G BCN, the radar locates a beacon transmitter on the ground enabling direct interpretation of relative bearing for tracking, homing or other navigation using a ground BCN reference. The beacon returns are displayed as several short horizontal lines in range order. The horizontal line closest in range is the actual position of the beacon.

Air-to-ground backup mode The radar goes into backup mode when the ADCP has failed. During A/G radar backup mode, the displays and functions are limited to the left MPD in both cockpits.

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Built-in test BIT modes are an integral part of the radar that uses hardware and software to test, calibrate, detect faults, and fault isolate to an LRU. BIT matrix buffers are used to store the faults resulting from tests to be used for fault isolation and determination of radar performance. LRU and system calibration is done during BIT. The BIT program is stored in the radar data processor (082) and the 082 controls BIT.

BIT operates in an automatic or initiated mode or a combination of both types of modes. There are two automatic modes that require no operator action. These are ORT and CM.

Operational readiness test The primary purpose of the ORT test sequence is to do system tests and calibrations on individual LRUs in the entire radar set to establish a high level of confidence that the radar will do its mission satisfactorily. ORT starts during a cold start power-up on the ground or in the air.

Continuous monitor CM is an automatic sequence of tests run in a non-interference manner any time the radar is ON. CM detects failures in the selected radar operating mode and channel. The detected failures are recorded in the CM BIT matrix, which is used to fault isolate to the defective LRU.

Initiated BIT IBIT, when commanded by the operator, runs all possible tests of the radar. On the ground, the FLOOD WAVEGUIDE switch is not tested because of radiation hazard protection interlock. In the air, the antenna gyro tests are not done because of aircraft motion.

IBIT is divided into two categories, FBIT and GBIT. FBIT is the in-flight BIT (performed in the air) and GBIT is the maintenance BIT (performed on the ground). When BIT is started, the state of the WOW switch determines which type of BIT is selected. If WOW is false, FBIT is selected. If WOW is true, GBIT is selected. GND or AIR is displayed in the lower right hand corner of the radar display to indicate the position of the WOW switch. IBIT failures are displayed in the order of most serious failure first.

System integration The following components interface with the F-15E APG-70 radar set.

Sensor control panel The sensor control panel contains the RADAR POWER switch and the INS mode switch. The RADAR POWER switch is a four-position switch consisting of the OFF, STBY, ON, and EMERG selections. This panel replaces the 541 and the 542 on APG-63/APG-70 on the F-15C.

Multipurpose display system The ADCP processes all radar video and display symbology for display on the MPD/MPCD. Any of the MPD/MPCDs are used to display the A/A radar format. The A/G radar format is not available on the MPCD. Switches on the outer edge of the displays are used to control mode selection, video for recording, RF channel selections and many mode functions.

Left and right hand controller The left and right hand controllers contain switches used in A/A radar to control acquisition symbol position and antenna (031) elevation scan, for mode reject, and to undesignate targets and select highlight search. In A/G radar, the controllers switches control cursor position, control antenna (031) elevation angle and position and display window size, mode reject, designate and undesignate targets, select sequence points, select cursor functions and allow HRM map expansion.

Control stick grip In A/A radar, the control stick grip contains an auto ACQ/REJECT/IFR DISENGAGE switch which enables selection of boresight, long range boresight, supersearch, and vertical scan. The switch also

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enables missile tuning off, track-while-scan, high data rate track-while-scan and, when locked on, enables return to search (break lock). In A/G radar, this switch controls display window size and mode reject.

Throttle grips The throttles contain controls used with the radar system. In A/A radar, these controls enable antenna (031) elevation control and azimuth scan, target acquisition and lock-on, weapon selection, and a variable supersearch scan pattern. In A/G radar, these controls enable antenna (031) elevation control to target designate, select sequence points, control the cursor, and freeze and unfreeze the display.

Avionics interface unit 1 and 2 The avionics interface unit (AIU) 1 and AIU 2 work together processing discrete signals from controlling switches and analog signals from transducers into 16 bit data words which are sent to the ADCP on the avionics 1553 MUX bus. The AIU 1 functions as primary bus controller for the up-front control (UFC), while the AIU 2 functions as backup bus controller for the UFC.

Blanker When the radar set is transmitting, the blanker causes the radar warning receiver (RWR) high band receiver, the LANTIRN (low altitude navigation and targeting infrared for night) NAV pod, and the LANTIRN targeting pod to be blanked. Internal countermeasures system (ICMS) transmissions cause the radar receiver (025) to be blanked.

Environmental control system Circulating air at a predetermined temperature from the ECS cools the transmitter (111), receiver-exciter (025), programmable signal processor (044), power supply (610), radar data processor (082), and analog signal converter (038). Radar cooling air is provided by the ECS with the aircraft engine operating. During ground maintenance, cooling air is supplied by the ground air conditioning unit (C10) through the ECS ground-cooling receptacle. When the ECS is shut down in flight, cooling is automatically provided by ram air.

Liquid coolant system The transmitter’s (111) GTWT, microwave processor and high voltage power supply are cooled by circulating liquid at a predetermined temperature through the transmitter (111) from the aircraft ECS.

IFF reply evaluator The IFF reply evaluator (IRE) provides the IFF target information when commanded on the A/A mode during AAIs. The targets are displayed as diamonds or circles indicating the confidence level of a target being a friend. The diamond indicates the lowest confidence level and the circle indicates the highest.

Avionics status panel The indicators related to the radar system are ASP numbers 2, 4, 25, 34, 61, and 62.

204. APG-63 (V)1 radar If you have worked the APG-63 radar in the past and now are working the APG-63(V)1 radar you will be pleasantly surprised at how much less the (V)1 system fails. The APG-63(V)1 radar is installed on F-15C/D aircraft. The (V)1 is projected to have a mean time between maintenance actions (MTBMA) of 120 hours. This means each individual radar system should operate for 120 hours before any part goes bad. This is much better than the 13 hours MTBMA for the APG-63 and APG-70 radar systems. Reduction of malfunctions that cannot be duplicated is also guaranteed. This guarantee is made possible through the use of new components and an upgraded CFRS/CFI system tied directly to the data transfer module (DTM) data downloaded during pilots’ debriefs. (CFRS/CFI stands for computerized fault reporting system and computerized fault isolation.) CFRS is already established in the field. CFI is simply CFRS software that provides maintenance diagnostic capability for the APG-63(V)1 radar.

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New maintenance concepts The approach to maintenance on the APG-63(V)1 radar system is different than most technicians are used to on the flightline. Traditionally, the radar BIT was just one piece of information available for troubleshooting. Technicians considered BIT indications when troubleshooting, but they were not completely dependent on them. The APG-63(V)1 dictates how maintenance is to be accomplished. The BIT functions of the radar have been drastically improved. It is capable of troubleshooting the jet in 98 percent of all conditions, and has measures to grow as idiosyncrasies are discovered. Stipulations in the contract with the manufacturer require USAF personnel to strictly adhere to the BIT and CFRS/CFI fault isolation process. For example, if CFRS/CFI tells you to remove and replace the transmitter, but you choose to change the power supply instead, then you have voided the warranty. The squadron will be liable for transportation and repair costs of the LRU. Squadron commanders are going to be very upset if money gets wasted in this manner.

Importance of CFRS/CFI There is no fault isolation manual for the APG-63(V)1 radar. Faults are isolated exclusively by the BIT tests within the radar and CFRS/CFI. The process works because all of the failures that BIT can identify are stored within the CFRS/CFI system at your work center. The combination of data from the pilot’s DTM and inputs from the pilot about what he/she observed in-flight will lead CFRS/CFI to the probable faulty LRU. CFRS/CFI always gives one of three types of positive maintenance actions:

1. Job control number generated and a maintenance action identified. After three maintenance actions have been tried and the problem still exists, the manufacturer will get involved and provide further assistance.

2. Informational write-up generated with no maintenance action required. 3. False alarm report generated, indicating normal operation under reported flight conditions.

Data transfer module The APG-63(V)1 uses the DTM for troubleshooting. During the mission, radar stores BIT faults and electronic serial numbers of all the radar LRUs. When the pilot writes to the DTM, all the BIT faults and radar LRU serial numbers are transferred to the DTM. Then, the pilot loads the data into CFRS/CFI where fault isolations take place. Now, keep in mind that all of the data in CFRS/CFI is shared between the USAF and the manufacturer. If you decide to swap LRUs with another aircraft instead of following the proper supply procedures, you again risk voiding the warranty. (Remember that the serial numbers of each part are recorded on the DTM.) Simply follow the CFRS/CFI instructions without deviation!

System components There are seven major LRUs in the APG-63(V)1 radar system, shown in figure 1–22. Notice that the power supply has moved up to the top-right of door 3L and that there is now only one digital processor, the 385. The 031 is unchanged and interchangeable with other radar systems.

Radar control panel (342) The 342 performs the same role as the 541 and 542 did, but it’s a new unit designed to work with the modes and BIT available in the APG-63(V)1. Most outputs go directly to the radar data processor (385). One visible change is the omission of the emergency flag. Now, if EMERG is selected, the pilot will see an E in the lower left corner of his display and the event will be recorded on the DTM.

Receiver-exciter (325) The receiver-exciter performs the same functions as the APG-70 receiver-exciter.

Transmitter (311) The transmitter performs the same functions as the 011 and 111. The manufacturer did add coffin handles to the sides of the transmitter to assist you in transporting it from the B-4 stand to the container.

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Figure 1–22. APG-63(V)1 RSC and general LRU layout.

Analog signal converter (338) The 338 performs the same functions as the 039 and 038 did.

Radar data processor (385) The 385 provides control, data processing, and high-speed digital signal processing functions that were formerly done by two processors (081/042 or 082/044). The RDP stores the radar OFP and protects the radar from overheating.

Power supply (310) The 310 performs the same functions as the 610 did for the APG-63 and APG-70. It meets the APG-63(V)1 specifications for BIT testability and reliability.

Antenna (031) The antenna remains unchanged. Funding was not allocated for 031 improvements in conjunction with the APG-63(V)1 installation.

System operation There are only a couple of mode changes in the APG-63(V)1 radar system, primarily in the A/G area. Look at 342 in figure 1–22 as we discuss. A/A modes remain the same as APG-70 radar.

A/G modes There are three modes selectable with the A/G master mode switch pressed:

1. A/G ranging - This mode provides a 10 NMI PPI display for determining slant range to the ground for bombing. All LPRF is used.

2. RBM - The RBM mode provides conventional A/G functions of large-area mapping and general navigation. The presentation is in a 100 degrees PPI scan format, with 10 degrees on each side for drift compensation. All range scales are available. All LPRF is used.

3. PVU - This is the same as PVU on the previously discussed systems. It provides radar-measured velocity estimates to correct INS errors and has no display. PVU uses all LPRF.

Built-in test As mentioned earlier, the APG-63(V)1 has a much-improved BIT. It runs in the same manner as the other radar BITs, but it has increased number of test points and uses CFRS/CFI.The new BIT is designed to detect at least 95 percent of all possible failures and out-of-tolerance conditions that degrade tactical operation. It is designed to isolate the fault to one failed LRU 98 percent of the time; to an ambiguity group of two or fewer LRUs 99 percent of the time; or to an ambiguity group of three or fewer LRUs 100 percent of the time. Ambiguity group refers to a situation where the fault may lie

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within more than one LRU and the radar cannot isolate it any further. In addition, BIT is required to reduce the amount of could-not-duplicate (CND) solutions to an 8 percent maximum rate. The CND solution is used when a maintainer cannot reproduce the failure reported by a pilot. Since the system is not failing regularly, maintainers will “let it fly” and wait to see if it breaks again. BIT operates either automatically or in an initiated mode.There are two automatic BIT modes that require no operator action. They are power-on self-test (POST) and periodic BIT (PBIT).

Power-on self-test The POST BIT is executed automatically at either “standby cold start” or “operate/emergency cold start.”

POST when executed at “standby cold start” This occurs when the POWER switch goes from OFF to STBY and runs virtually all tests run by an IBIT. LRU names are displayed as each is tested.

POST when executed at “operate/emergency cold start” This occurs when the POWER switch goes from OFF to OPR or EMERG. It runs only the core tests necessary to get the radar operational.

Periodic BIT The periodic BIT (PBIT) automatically checks radar function in a non-interference manner during tactical operation of the radar.Think of it as being very similar to continuous monitor BIT in older radars.

Initiated BIT In addition to the automatic BIT modes, the APG-63(V)1 has an operator IBIT. IBIT is the complete set of tests, including PBIT tests, which the operator must choose to run either on the ground or in flight. An operator initiated BIT will cease all operations and cause the radar to perform its test. During the initial stages of IBIT, the radar determines whether or not it is in the air by using the right MLG WOW switch position. If it is in the air, the radar performs the flight BIT (FBIT) functions of IBIT. If it is on the ground, the radar performs the ground BIT (GBIT) functions. During IBIT, the radar will give the operator the option to run a track-test and/or a switch test that checks the functionality of the 342 control knobs and switches. An initiated BIT with the radar in STBY will cause the BIT matrices to be displayed. The BIT displays include the following (fig. 1–23):

1. Pilot BIT matrix. 2. Software version ID display. 3. Maintenance BIT matrix (includes results of most recent IBIT, PBIT, persistence count, and

FBIT). 4. BIT window (lower left corner).

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Figure 1–23. Typical APG-63(V)1 BIT displays.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

201. APG-63 radar 1. List the major LRUs of the APG-63 radar set.

2. Which LRU generates the basic RF operating frequencies for the radar set?

3. Which LRU amplifies low power RF (GTWT drive) from the RFO?

4. The receiver amplifies the RF returns, converts them to 30 MHz IF signals, amplifies them again, and sends them to which radar component?

5. What type of commands does the data processor (081) send to the power supply (610)?

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6. What does the programmable signal processor (042) provide?

7. Which LRU receives serial digital data from the CC and radar set, analog and discrete data from the radar set, and a composite video signal from PACS?

8. During the transmitting function, which component sends the frequency selection to the RFO?

9. During the transmitting function, how does the RFO send the GTWT drive signal and to which component is it sent?

10. How does the APG-63 radar set filter out ground clutter?

11. The receiving function of the APG-63 radar is divided into what two broad categories?

12. Which component is the only LRU involved in the data processing function?

13. Which LRU is the only radar component that requires liquid cooling?

14. When running an operate IBIT on the ground, what should be displayed in the BIT window?

15. During which BIT are the faults detected by the radar displayed?

16. Explain the HPRF operation of the APG-63 radar set.

17. What detection ranges are available only during MPRF operation?

18. If the pilot wants to continue searching for other targets while tracking another target, he can accomplish this by using which basic mode group?

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19. What is the primary A/A search mode used for detection and acquisition of both closing- and opening-rate airborne targets?

20. What is acquisition and how can the pilot acquire a target?

21. During manual acquisition, what control is used to move the acquisition symbol around the radar display?

22. Explain the A/G ranging mode.

202. F-15C/D APG-70 radar set 1. The F-15 C/D APG-70 radar set is made up of how many major LRUs?

2. The receiver-exciter (025) combines the functions of which APG-63 LRUs?

3. Which APG-70 radar set major components use waveguides or are termed microwave LRUs?

4. Explain range gated high PRF.

5. When in the HI A/A search mode, what range will the radar search?

6. When in the MED A/A search mode, what range will the radar search?

7. How many A/G modes are available and what waveforms do they use?

8. Pilots will use the precision velocity update function prior to making a bomb run to compensate for what?

9. What happens when a G-BIT is initiated with the RADAR POWER switch in STBY?

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10. List the five BIT history matrices available.

11. What action will display the BIT history matrices?

12. What control will cycle between the BIT history matrices?

13. Explain how the persistence counter is used for fault isolation.

203. F-15E APG-70 radar set 1. Name the seven major LRUs of the APG-70.

2. Which LRU receives digital data and controls from the analog signal converter and the RDP?

3. Explain the purpose of the RBM mode.

4. Which A/G mode is used for detection of trucks or convoys out to 32 NMI?

5. What does the HRM mode provide?

6. Explain the purpose for the PVU mode.

7. What is the purpose of the ORT?

8. What is the difference between an FBIT and a GBIT?

9. What determines whether an FBIT or GBIT will be performed?

10. What portion of the initiated BIT is performed only in the air?

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204. APG-63 (V)1 radar 1. What computer system is used in addition to BIT to help achieve the contracted requirements for

fault isolation in the APG-63(V)1?

2. How will the BIT results from the radar get to the CFRS/CFI workstation in debrief?

3. Aside from pilot directed maintenance, are you allowed to deviate from the BIT and CFRS/CFI instructions?

4. What are the seven major LRUs in the APG-63(V)1 radar?

5. What new 311 feature makes transportation of the part easier?

6. Where is the APG-63(V)1 radar OFP stored?

7. BIT for the APG-63(V)1 is divided into what two modes?

8. What will be displayed when running an IBIT with the POWER switch in STBY?

9. What type of BIT is started when the radar is switched from OFF to OPR or EMERG?

10. What will be displayed when running an IBIT with the POWER switch in STBY?

1–2. Integration Systems In this section, we will concentrate on A through D models of the F-15 overload warning system. We will then discuss the F-15 central computer complex. And finally, we’ll cover the F-15E model advanced core display processor. 205. Overload warning system The overload warning system (OWS) determines structural loads (stress) being applied to the airframe during airborne operations. This alerts the pilot to a potential over-G of the jet. The pilot can change or ease up on the maneuver being executed. Preventing over-Gs on the aircraft reduces the

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amount of stress placed on the aircraft; thus reducing maintenance and maximizes aircraft performance.

Definitions These definitions will be useful to understanding OWS theory:

G-force A unit of force equal to the gravity exerted on a body at rest. Measured at one G, this is the amount of force you are experiencing as you sit and read this CDC.

Positive G Any force on a body that is greater than one G. A positive G is what you experience on a roller coaster while being pressed down into your seat. During a two-G maneuver, a 200-pound man would exert 400 pounds of force on the seat.

Negative G Any force on a body that is less than one G. A negative G is what you experience on a roller coaster when it makes its initial drop and you are pressed upward against the shoulder straps. A zero-G maneuver is weightlessness.

Over-G Over-G is a condition that occurs when the maximum allowable force on a structure has been exceeded. This condition causes structural damage if the force is great enough.

System components The OWS system has only one LRU, the OWS RESET switch. The OWS is a software program that is associated with two other LRUs: the CC and PSDP. There are several other LRUs that provide various inputs for OWS calculations.

OVERLOAD WARNING RESET switch The OVLD WRN RESET SW is a two position (norm and reset) momentary toggle switch mounted in the nose landing gear wheel well, left of the ASP. The switch is used with the navigation control indicator (NCI) panel to clear overload conditions from the CC memory and erase the OWS matrix display from the MICP.

Central computer The CC is a digital computer that contains the OWS software and performs the calculations, using inputs from various components.

Programmable signal data processor The PSDP is a display processor that receives OWS data from outside systems, conditions them, and sends them to the CC to be used in the OWS calculations. The PSDP also processes the OWS information displayed on the MICP, HUD, and heard in the headset.

System operation The following paragraphs describe the OWS system operation.

Inputs signals The OWS is a software system, which draws on existing aircraft configurations (i.e. weapons, fuel) and flight data (i.e. air speed, angle of attack) to calculate the structural loads being applied to the aircraft. These computations show the maximum allowable G force that the aircraft can withstand without causing structural damage. The OWS computation is done by comparing the OWS computed maximum allowable G force with the G force currently being applied to the aircraft. The CC receives digital inputs across the H009 MUX from three LRUs: the air data computer (ADC), PACS, and PSDP. Let’s look at the different inputs coming from each LRU. Figure 1–24 is a block diagram of the inputs and output of the OWS.

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Figure 1–24. Overload warning system block diagram.

Air data computer The ADC sends flight data consisting of angle of attack (AOA), Mach number, pressure ratio, and barometric corrected altitude across MUX bus channels 1 and 3.

Programmable armament control system The PACS sends signals telling the CC what is loaded on the aircraft. The OWS program needs to know what pylons, racks (different bombs require certain mounting racks), weapons type (500, 1000, 2000 lbs and missiles), and wing tanks and/or conformal fuel tanks (CFT) are loaded for accurate computations.

Programmable signal data processor The PSDP receives analog data from several different components. It takes this data and converts it into a digital signal and sends it to the CC on MUX bus channels 2 and 4. Here is a list of the components and the signals sent to the PSDP.

• Multipurpose color display - The MPCD is used to manually input information on aircraft stores (i.e. bombs) loaded on the aircraft. This information is sent to the PACS via the PSDP.

• Fuel quantity indicator - Sends the total fuel quantity for the aircraft to include the all internal fuel cells and external wing tanks and centerline tank. It will NOT send fuel quantity for CFTs.

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• Fuel quantity signal conditioner - Sends CFT fuel quantity to the PSDP.

• Roll/Yaw computer - Sends lateral stick force and roll rate to the PSDP.

• Pitch computer - Sends stabilator (stab) position to the PSDP.

• Accelerometer - Measures the current G-force on the aircraft and sends it to the CC. The accelerometer is located in the right main landing gear wheel well. If the PSDP and/or OWS are not working, the INS sends current Gs to the CC as backup so the pilot still knows how many Gs are being pulled.

• OVERLOAD WARNING RESET switch - Sends a signal to clear existing OWS information from the CC.

Results of inputs The OWS program uses the data provided to it to compute the G force being applied to seven different areas of the aircraft. These seven areas are the fuselage (FUS), wings (WNG), left tail stabilator (LTS), right tail stabilator (RTS), pylons (PYL), CFT, and mass items (MIT) (i.e. engines, jet fuel starter airframe mounted accessories drive, etc.). After the computations are complete, the CC compares the OWS information with the stored structural data on the seven different areas to determine if an overload condition exists. The structural load or stress placed on the aircraft is measured and displayed in percentages. The percentages are then grouped in ranges and the ranges are assigned a number called the severity code. See the below table.

Percent of Design Load Limit Severity Code a85 – 100 0 100 – 110 1 110 – 120 2 120 – 130 3 130 – 140 4 140 and higher 5

Output signals The CC sends OWS information to two LRUs, the PSDP and the HUD data processor.

Programmable signal data processor The PSDP receives several signals from the CC. They include the OWS matrix display information, a tone control or voice warning command, and the overload latch signal.

OWS matrix display The CC transmits the matrix information to the PSDP. The PSDP sends this display information to the MICP for display once it is requested using the NCI panel.

Tone control/voice warning The tone control/voice warning lets the pilot know the level of stress that is being put on the aircraft. Different percentage rates give different tones or voice warnings (see below table). Each signal is commanded by the CC and sent to the PSDP. The PSDP sends these signals to the integrated communications control panel (ICCP).

Percent of maximum allowable load Signal 85 to 92 900HZ tone interrupted at 4HZ rate 92 to 96 900HZ tone interrupted at 10HZ rate 96 to 100 900HZ solid tone above 100 Voice Warning (OVER-G, OVER-G)

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Overload latch Once the overload condition is above 100 percent, a ground is sent to the ASP to latch fault indicator 72. This indicates that an overload condition exists.

HUD data processor The HUD data processor (DP) receives the normal acceleration (current G) and maximum allowable acceleration data (maximum allowable Gs) or NOWS (OWS not operational) to be displayed in window 8 of the HUD.

Displays There are three displays for the OWS. They are the MICP OWS matrix display, MICP test pattern and HUD window 8. Look at figure 1–25 for an explanation of how to read the MICP OWS matrix.

Figure 1–25. MICP OWS matrix display.

MICP OWS matrix display The three worst overload conditions for each component are stored in memory for an OWS matrix display on the MICP. The CC also computes the overload percentage; the worst overload conditions and latest overload condition that happened on each flight. These, along with the G measurement during each overload, are stored in memory.

What is done with overload information? Depending on the severity of the over-G, various panels may have to be removed to inspect the aircraft frame for cracks. A severe over-G will require an extensive inspection of much of the airframe.

MICP test pattern The MICP test pattern is used during the operational checkout of the OWS. It is used to check the integration signals for the stick force and fuel quantity systems (fig. 1–26).

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Figure 1–26. MICP test pattern.

HUD window 8 display The HUD window 8 will show the current Gs and the maximum allowable Gs (fig. 1–27). The maximum allowable Gs are computed by the CC and sent to the HUD processor. If the maximum allowable Gs cannot be computed for any reason, then window 8 will display the current Gs from the accelerometer and NOWS. NOWS lets the pilot know that OWS warnings are not available at that time. During normal operation, the HUD window 8 display is the only OWS display the pilot will see.

Figure 1–27. HUD window 8 display.

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System operation The CC/OWS is constantly working. The ADC is sending AOA, speed and pressure readings. PACS is telling the CC what is loaded on the exterior of the aircraft: one tank on each wing, a centerline tank, and four missiles. The CC knows exactly how much each item weighs from its memory of stored data. However, since it does not know how much fuel is in the tanks, the fuel quantity indicator tells the CC how much fuel is on board. The CC takes the amount of fuel and computes its weight. The CC now knows the exact configuration of the aircraft. Next, it looks at the amount of commanded pitch by the pilot (stick force) and compares that with the stab position (actual pitch of aircraft). The CC takes the speed, pitch, and configuration to compute the maximum allowable Gs on each of the seven areas of the aircraft. It then compares the maximum allowable G with the current G force from the accelerometer to determine if an overload condition exists.

The CC sends the computed maximum allowable and current Gs, from the accelerometer to the HUD. As the overload condition increases, the CC sends a signal to the ICCP to enable the 4 hertz (Hz)-interrupt rate warning tone for the pilot, when the condition goes between 85–92 percent. Then the 900 Hz tone would shift to the 10 Hz-interrupt rate when the condition went from 92 to 96 percent and a solid tone for conditions from 96 to 100 percent. Once the overload condition goes above 100 percent, a voice warning of “OVER-G, OVER-G” would be heard. A signal would also be sent by the CC through the PSDP to latch ASP 72. Once the overload condition peaks out, the CC records the overload condition in memory if it ranks in the top three for each component or if it is the last overload condition.

206. Central computer complex The heart of the F-15 A through D model avionics is the central computer complex (CCC). The CC does detailed computations for aircraft navigation, weapons delivery and control, and display systems. It also performs computations for the various avionics displays and indicators. After studying this lesson, you’ll have a good idea how the CC functions and the importance of the system to the F-15 A through D model aircraft.

Major components The CCC consists of five major components: the CC, two MUX buses (H009 and 1553), DTM receptacle, and memory loader verifier (MLV) receptacle. Central computer The CC (fig. 1–28) is a high-speed, general-purpose, stored program, digital computer. Through MUX buses, it integrates the various avionics subsystems into a single weapon system. By using information from various avionics systems and sensors, the CC calculates navigational, display, weapon delivery, and weapon control information. Under guidance of the OFP, the CC computes aircraft position, velocity, altitude, and attitude for navigation, weapons delivery, and display purposes.

The development of the CC gave aircraft avionics enormous flexibility and capability. The CC totally defines the term integrated avionics. Here’s an example of that flexibility and capability. The CC performs different functions, depending on the aircraft avionics master mode selected on the mode beacon light panel. There are three aircraft avionics master modes:

1. A/A. 2. A/G. 3. Attitude director indicator (ADI).

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Figure 1–28. Location of the central computer.

Data transfer module receptacle The DTM receptacle is simply a device that accepts the data transfer module. The receptacle provides an interface between the DTM and PSDP in order to load mission data into the CC. The DTM and DTM receptacle will be covered in more detail later.

Memory loader verifier receptacle The MLV receptacle provides a connection for the programmable loader verifier (PLV) to the aircraft. The PLV (AN/ASM-700) is commonly referred to as the PLV/NT (PLV/new technology). The MLV receptacle is used to load OFPs and other instructional code required for avionic components that are linked through the 1553 MUX bus.

Multiplex buses The MUX buses are super highways for digital data between the CC and its peripherals (avionics systems/equipment). The CC is the master terminal and the peripherals are the remote terminals (RT). Each RT communicates with the CC only after the CC sends a select word to that RT. RTs can communicate with the CC on the MUX buses, but they can’t communicate with each other. There are two types of MUX buses—H009 and 1553.

Major sections of the central computer and built-in test The very high-speed integrated circuit (VHSIC) central computer contains five different types of modules and a battery. The modules of the VHSIC CC are listed:

1. Three data processing modules (DPM). 2. Three input/output (I/O) modules (IOM).

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3. Bulk memory module (BMM). 4. Bulk storage module (BSM). 5. Timing and discrete module (TDM).

The battery is used to retain variable data such as time-of-day, navigation data and so on. The CC is rack mounted and secured by two ratchet-type-knurled fasteners.

CC built-in test The CC BIT reports failure in any of the CC shop replaceable units (SRU). The failure is then stored in the CC nonvolatile memory. SRU data can be displayed on the MPCD using the BIT menu. The SRU failures also can be cleared through the BIT menu. After an SRU failure has been cleared on the MPCD, press and release the CC RESET on BCP. If CC failures reappear, the CC should be replaced.

CC operation and integration The CC requires three-phase, 115 volts alternating current (VAC), 400-Hz power input. The CC switch on the ground (GND) power (PWR) control panel applies power to the CC. The CC communicates directly with the DTM.

F-15 data transfer module The DTM is a programmable, battery-powered, nonvolatile memory device used to transfer flight operations mission data to the CC. Figure 1–29 shows the DTM and the DTM receptacle.The DTM receptacle allows the DTM to interface with the CC.

Figure 1–29. Data transfer module and receptacle.

Data transfer module types Two types of DTMs—maintenance (MAINT) and operations (OPS)—can be used by the DTM receptacle. The OPS DTM (used by pilots) contains more memory storage than the MAINT DTM because the OPS DTM requires a larger memory to hold a wide variety of mission data. The MAINT DTM is used only to read bit-type data for maintenance purposes. The DTM transfers A/A and A/G operations data, as well as maintenance data, from the CC to the DTM by way of the PSDP.

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Data transfer module READ function The READ function is a process of the CC reading data from the DTM. Selecting READ on the MPCD causes the CC to lock out keyboard entries on the NCI and generates a READ IN PROGRESS legend to be displayed on the MPCD.

Data transfer module WRITE function The WRITE function is used to record mission data on the DTM. The DTM receives mission data from the CC, by way of the PSDP.

CC operational flight program The OFP is the program loaded into the VHSIC CC. It contains instructions necessary to do radar, general navigation, and weapon delivery computations. The VHSIC CC OFP consists of seven program functions. Although each function is a complete set of instructions for doing specific duties, other functions may help. In this regard, they all work together, as a team, to get the job done.

The functions that make up the total program are the executive, A/A, A/G, navigation, flight director, controls and displays, and computer self-test functions.

Executive function The executive function interfaces with all other program functions. The executive function has five program subfunctions, which are identified in the table below.

Subfunction Explanation

Initialization Controls all power-up or reset parameters along with setup for processing of all executive sub-functions.

Interrupt processing Services “armament interrupt” routines by giving memory and registering priority to the armament system for data processing; thus interrupting any normal routine in work.

External I/O processing Interfaces with the CC peripherals in testing all MUX capability.

Program scheduling Sequences the functions timing and calls the functions at the required scheduling rates.

Fault processing Provides OFP fault tolerances along with collecting and storing data for the computer self-test function.

A/A function The A/A function assists in the management and delivery of air-to-air weapons and receives radar data to assist in target tracking and display of target data. This function also provides data for steering and displays, does computations, displays HUD cues for gun mode operations (providing lead angle), and provides missile prelaunch and post-launch management. It interfaces with the executive function, navigation function, and the controls and displays function. The A/A function also interfaces, through the I/O section, with CC peripherals.

A/G function The A/G function operates when the A/G master mode is selected and valid attitude, velocity, and heading data are available. This function provides weapons and delivery control of air-to-ground weapons, delivery parameter setup, and steering and release computations.

Navigation function The navigation function of the CC collects data from the INU and other peripherals, analyzes and computes the information for use by the pilot and other peripheral systems. The navigation function performs the following tasks:

1. Computes the best available navigation data (attitude, heading, velocity, altitude, and present position).

2. Updates position and velocity.

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3. Computes range and bearing to a waypoint or tactical air navigation (TACAN) station destinations.

4. Processes joint tactical information distribution system (JTIDS) position transfer data (target information) and vectoring data from the JTIDS terminal.

5. Designates an A/G target and identification point (IP) upon aircrew selection.

6. Performs NCI data interface, computes sideslip for the engine diagnostic unit (EDU), and initiates spin recovery displays.

Flight director function The flight director function provides TACAN, NAV, and ILS steer mode processing (ground track, command heading, range, and bearing) for display on the horizontal situation indicator (HSI). It also performs range and bearing or steering computations for display on the ADI and HUD in TACAN, NAV, and ILS steering modes.

Controls and displays function The controls and displays function provides the CC (also PACS, DTM, and the VTRS) with the latest configuration of aircraft system controls. This function also updates control panel indicators, manages symbology for the HUD and MICP, processes data for the signal data recorder, and handles processing for the overload warning system.

Computer self-test function The computer self-test function provides initialization self-test, CPU self-test, memory self-test, CC I/O and multiplex test, and error handling functions.

Inspecting the CC memory You can troubleshoot virtually any avionics system by examining the CC I/O memory data and the internal memory parameters of the CC database to find malfunctions. To do this, you enter all inputs and requests to the CC through the MPCD or the NCI keyboard. You’ll start with the NCI.

Using the navigation control panel to inspect central computer memory The NCI right DRD window displays the data in CC memory in the form of a five- or six-digit octal code. This code is meaningless in its original form; therefore, you must decode it. Generally speaking, this involves converting the octal number to a 16-digit binary word and using data tables to extract its full meaning. Detailed procedures are in TO 1F-15C-2-31GS-00-1, General System- Indicating/Recording Systems.

The information contained in the 16-digit data words, of course, is the thousands of parameters the CC needs to get its job done. Antenna position, true airspeed, AOA, Mach number validity, and true airspeed validity are some examples. So you see, the information can be a very precise value requiring many digits, or it may be a simple valid or not valid signal requiring only one digit.

Using the multipurpose color display to inspect central computer memory TO 1F-15A-2-31GS-42-1 provides the information you need to audit the VHSIC CC using the MPCD. The MPCD provides a much easier method of inspecting the CC memory. You can access the MPCD AUDIT display directly from the BIT menu.

The NCI uses an octal readout, where the MPCD uses numerous readout formats. Hexadecimal (HEX) is the preferred readout format. This is a process of converting the four-digit HEX number to a 16-BIT binary number. Remember, a binary number is counted from left to right from 0 to 15. You need to convert the HEX readout in order to read the most significant binary number and the least significant binary number. Once you find the applicable MPCD AUDIT addresses and readout table, you can start examining the CC memory. Figure 1–30 shows a typical MPCD audit display.

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Figure 1–30. Typical MPCD audit display format.

207. Advanced display core processor Time Compliance Technical Order (TCTO) 1F-15E-807, Installation of Advanced Display Core Processor (ADCP) P/N 8525470-950 into F-15E Aircraft, 1 Jun 06, dramatically changed the configuration of the F-15E aircraft. This TCTO saw the removal of both the CC and the multipurpose display processor (MPDP) and in their place, the installation of the ADCP. The ADCP combines all the functions of both units and is the bus controller and display processor for the entire aircraft.

System description The F-15E ADCP combines central computer processing, analog to digital conversion, discrete processing, and video display processing in a single LRU. The ADCP is located where the MPDP used to be, on the top shelf, left corner, in door 3L (fig. 1–31). The ADCP is made up of the following components:

General purpose processor The general purpose processor (GPP), including the intelligent serial module (ISM) (main and essential), accomplishes all mission data processing within the ADCP. The ISM supports loading LRU software by way of the 1553 MUX bus using a PLV. The GPP interfaces and controls the avionics and fighter data link (FDL) MUX buses. The GPP stores the OFP that controls the tactical operation of the ADCP system.

Image processor modules The multipurpose display processing is performed by the image processor module (IPM) section of the ADCP. The IPM does display related processing (for example, generation of graphical and symbology overlays) within the ADCP. The IPM stores part of the OFP and controls all cockpit displays (MPD/MPCD/HUD) and secondary HUD and VTRS.

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Figure 1–31. ADCP location.

VERSA module Eurocard 64 main and essential busses The VERSA module Eurocard (VME) 64 main and essential buses service the multipurpose display processing and central computer processing functions. During the normal mode of operation (no detected faults), these buses are interconnected and the system is in full operation. In degraded mode (fault detected), the buses are split and the system operates on the bus without detected faults. The displays connected to the functioning bus will operate, and since the buses are redundant, many of the core processing functions will be available.

Principles of operation The ADCP system has four modes of operation: normal, degraded, emergency, and test.

Normal mode The normal mode of operation is enabled when aircraft power is applied and there are no detected failures. Normal mode produces the complete set of displays available for the MPD, MPCD, and HUD. The ADCP produces the displays as directed by the ADCP OFP. The ADCP OFP defines the displays to be produced. The ADCP operates as the primary display controller for the MPD system.

Degraded mode The degraded mode is enabled when a simple failure is detected that does not reduce functionality or when a severe failure occurs and causes a main and essentials bus split. All failures are stored in the

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bridge modules fault logs. If a bus split occurs, the functioning bus (main/essential) will automatically be selected. The displays associated with the failed bus will not be available with the exception of the functions that are redundant on the buses. A bridge module failure will force ADCP operation to the main bus. A GPP failure will be flagged, but the other GPP will support all systems. If a module fails, the displays supported by that module will not be available for display.

Emergency mode (dual generator failure) When the ADCP detects a loss of power, the ADCP checks the generator no go discrete signal input. If the discrete is a ground, only one generator is off line and the ADCP operates in the degraded mode. If the no go discrete is open, both generators are off line and the ADCP operates in the emergency power mode. When both generators are off line, the essential bus is powered by the aircraft emergency power unit. The HUD, all MPD, and both rear MPCDs go blank and only the electronic attitude director indicator (EADI) format is displayed on the MPCD and the push buttons (PB) are disabled.

Test mode The ADCP system does four types of BIT (power-up, background, maintenance, and initiated). Power-up BIT is done when power is applied to make sure the ADCP is operational. Background BIT is done at intervals during normal system operation and does not interrupt system operation. Initiated and maintenance BIT are manually initiated, and interrupts normal system operation. Equipment failures are displayed at the lower center of the BIT format displays.

Avionics master modes of operation The following paragraphs describe the ADCP avionics master modes of operation.

Air-to-air The A/A master mode provides for visual short range, high-g performance during A/A combat. It also provides for medium range, head down attack ability. When a weapon type is selected, the ADCP automatically initiates display and radar operating modes. The ADCP also sequences and initiates pre-launch signals for the selected weapon.

Air-to-ground The A/G master mode provides for visual attack ability for delivering bombs, dispensers, and for firing the gun. The conventional weapon delivery modes are continuously displayed impact point (CDIP) and automatic (AUTO). In the occurrence of an ADCP failure, a manual delivery mode is also available. The ADCP controls the attack displays displayed on the HUD and the required weapon delivery computations.

Navigation The NAV master mode provides for attitude and steering display information on the HUD for general navigation. The ADCP controls the NAV displays on the HUD. Steering mode selection is done using the integrated multifunction display system.

Instrument (INST) The INST master mode of operation provides a one-step procedure to quickly call up basic flight instrument displays. The HUD displays for the INST master mode of operation are the same as the HUD displays during NAV master mode of operation. The basic flight instrument displays that appear on the cockpit multipurpose displays are as follows:

• A/A radar format on the left MPD. • EADI on the MPCD. • Electronic horizontal situation indicator (EHSI) on the right MPD.

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Multiplex bus integration The ADCP interfaces with other avionics systems over the multiplex bus. Let’s discuss these interfaces.

Avionics 1553 MUX bus 5A/5B The ADCP interfaces with several peripherals over the 5A/5B bus. These include the automatic flight control system (AFCS), AIU 1, AIU 2, ICMS, INS, LANTIRN pods, and the radar system.

Avionics 1553 MUX bus 7A/7B The FDL interfaces with the ADCP over buses 7A/7B.

Avionics 1553 MUX bus 8A/8B The embedded GPS/INS (EGI) receives time of day, date, almanac data, initialization data, wander angle, barometric altitude, and position updates. The PACS system also communicates on bus 8A/8B, providing weapon selection/control and launch parameters.

ADCP operational flight program The ADCP OFP is made up of seven program functions. These functions are identical to the ones in the CC system previously mentioned. The program functions that make up the ADCP OFP are listed:

1. Executive. 2. A/A. 3. A/G. 4. Navigation. 5. FDL. 6. Controls and display. 7. Computer self-test.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

205. Overload warning system 1. What is the purpose of the OWS?

2. What is the OWS RESET switch and what does it do?

3. List and give a brief description of the two LRUs associated with OWS.

4. How many inputs does the ADC give the CC?

5. What does PACS provide to OWS?

6. Explain the purpose of the accelerometer for the OWS?

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7. When will the pilot hear overstress warning tones? Describe these tones.

8. Use figure 1–25 to answer this question. What row of the OWS matrix shows the worst overload condition for each column?

9. During normal operation, what display will the pilot see for OWS and what does it display?

206. Central computer complex 1. List the five major components of the CC complex.

2. What are the three aircraft avionics master modes?

3. What is the purpose of the DTM receptacle?

4. What are the two types of MUX buses?

5. What are the five different types of modules contained in the VHSIC CC?

6. What is the input power requirement for the CC?

7. Describe the DTM.

8. What are the two types of DTMs?

9. With which component does the DTM interface?

10. What is the DTM READ function?

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11. What information does the OFP contain?

12. What program functions are included in the OFP?

13. The executive function interfaces with what program functions?

14. What is the purpose of the A/A function?

15. What tasks are performed by the flight director function?

16. Explain how the CC can be used for troubleshooting the avionics systems.

17. How do you decode an octal word?

207. Advanced display core processor 1. What did TCTO 1F-15E-807 change in the F-15E aircraft?

2. Where is the ADCP located?

3. What components make up the ADCP?

4. What are the four modes of ADCP operation?

5. Explain the degraded mode of operation.

6. What are the four types of BIT tests?

7. Which BIT format does not interrupt normal operation?

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8. On which MUX bus does the FDL operate?

9. The EGI receives time of day, date, almanac data, initialization data, wander angle, barometric altitude, and position updates on which MUX bus?

10. List the program functions making up the ADCP OFP.

1–3. Display Systems The F-15C/D/E aircraft has several different types of display systems. In this section we will discuss two of them. We will start with the F-15C/D/E VTRS. Then we will wrap up this section with the digital map system (DMS). The video tape recording system does not display, but records the display systems; thus making it fit into this section.

208. Video tape recording system The VTRS is used to record in-flight mission data that is displayed on the HUD, MICP, and the MPCD. Upon landing, the aircrew will carry the cassette into debrief where they will be able to replay it.

The ability to record has allowed aircrew (pilots) to analyze the actual effectiveness of each mission flown. In addition to analyzing mission effectiveness, the VTRS will aid in troubleshooting display problems.

There have been many modifications to the VTRS over the last 10 years. Currently the F-15 is being upgraded to the digital video recording system (DVRS). In this lesson, we will cover the VTRS; though different than the DVRS, much of the theory is the same.

System description The 8 millimeter (mm) VTRS records inflight color video of the MICP, HUD, and MPCD. Voice communications to and from the pilot are recorded on the audio track of the MICP, HUD and MPCD video to aid correlation of inflight occurrences. The recorded video and audio can be reviewed when the video tape cassettes are played back using monitoring equipment.

Component description The 8 mm VTRS is made up of the following parts:

• HMD/VIDEO CONTROL panel. • Power converter. • Video event marker generator. • HUD camera. • HUD camera control unit. • HUD video tape recorder. • Beamsplitter assembly. • MICP camera. • MICP camera control unit. • MICP video tape recorder. • MICP video sensor head assembly.

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• MPCD camera. • MPCD camera control unit. • MPCD video tape recorder.

Refer to figures 1–32a, b, and c for the VTRS system components. To use these figures simply find the common name on figure 1–32a and note the index number for that name. Next, refer to figures 1–32b or c to find the index number with a circle around it pointing to the component. Also note that the arrows describe which TCTO is applicable for a given component. As long as you know which TCTOs are complete on your aircraft, you will be able to use the chart to know which components are installed.

Figure 1–32a. 8 mm video tape recording system component location.

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Figure 1–32b. 8 mm video tape recording system component location.

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Figure 1–32c. 8 mm video tape recording system component location.

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HMD/video control panel The HMD/video control panel (fig. 1–32b) is mounted on the right side of the HUD unit and uses the OFF/STBY/AUTO/HUD switch to provide aircraft power to the 8 mm video tape recording system.

Power converter The power converter (fig. 1–32c) is mounted to the interior lights power supply and is located below the right console under the map case. The power converter converts aircraft 28 volts direct current (VDC) to system operating voltages.

Video event marker generator The video event marker generator (fig. 1–32c) is used to electronically put an event marker on the MICP and HUD video signal when either the TRIGGER switch is pressed to the second detent or the WEAPON RELEASE switch is pressed. This causes a rectangle to be displayed on the MICP and HUD recorded video displays.

MICP/HUD/MPCD cameras The MICP, HUD and MPCD cameras (fig. 1–32b) are charge coupled devices used to provide a color display of the HUD, MICP, and MPCD to the applicable MICP, HUD, and MPCD video tape recorders. The MICP camera is mounted to the beamsplitter assembly located above the MICP. The HUD camera is mounted in a fixture located on the center glareshield. The MPCD camera is mounted on the lower left windscreen.

NOTE: Depending on which TCTO has been accomplished on your aircraft there may not be a MPCD camera. If the aircraft is upgraded then the MPCD is recorded from within and does not have an external camera to do the recording.

MICP/HUD camera control unit The camera control units (fig. 1–32c) provide power, control, and processing of the video signal from the applicable MICP, HUD camera. The MICP/HUD camera control units are located in a cockpit housing, forward of the HUD unit.

MPCD camera control unit The MPCD camera control unit (fig. 1–32c) provides power, control and processing of the video signal from the MPCD camera. The MPCD camera control unit is mounted to the bottom of the map case located in the right console.

MICP/HUD/MPCD video tape recorders The MICP, HUD and MPCD video tape recorders (fig. 1–32c) are identical color recorders. The recorders are located in the cockpit in the map case. The recorders use an 8 mm video tape cassette that provides 120 minutes of recording time.

MICP video sensor head assembly The MICP video sensor head assembly (fig. 1–32b) is used to mount the MICP video camera to the beamsplitter. The MICP video sensor head assembly allows the MICP display image, directed up by the beamsplitter, to be directed forward to the MICP camera.

Beamsplitter assembly The beamsplitter assembly (fig. 1–32b) is attached to the left main instrument panel, in front of the MICP. The polarized beamsplitter glass, set at a 45 degree angle to the MICP cathode ray tube, reflects the MICP display up to the MICP video sensor head, then to the MICP camera. A polarized filter, at the top of the beamsplitter assembly in front of the MICP camera, prevents surrounding area light from reaching the MICP camera.

Related component description The following units are not part of the 8 mm video tape recording system, but work directly with it:

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Helmet display unit camera The helmet display unit (HDU) camera (fig. 1–32b) is a component of the joint helmet mounted cueing system (JHMCS). The HDU camera has a 20 degree field of view for recording video along the helmet line of sight.

Electronics unit The electronics unit (EU) (fig. 1–32c) is also a component of the JHMCS system. The EU receives HDU video from the HDU camera and HUD video from the HUD video control unit. The EU overlays the video from the HDU camera with the symbology that makes up the HMD. With the HMD/video control panel OFF/STBY/AUTO/HUD switch set to AUTO, the EU outputs HUD or HDU video for recording based on commands from the CC.

System operation There are two lights on each of the video tape recorders that dictate most of the maintenance required on the VTRS system. The REC (record) light is red and comes on when the recorders are in record mode. The other is the CAUTION light. This light is amber and comes on/flashes when any of the following exists:

• An input power problem exists. • The cassette tape is near end or has run out. • Moisture is detected inside video tape recorder. • The video heads are contaminated. • Some other malfunction exists.

Or with record mode enabled and the following conditions: • The red tab on tape cassette is out. • No cassette is inserted.

The operation of the VTRS is controlled by the OFF/STBY/AUTO/HUD switch on the HMD/video control panel. Setting the OFF/STBY/AUTO/HUD switch to STBY causes 28 VDC to be applied to the power converter. The power converter outputs system operating voltages to the video tape recording system components. In addition, the power converter applies the standby signal to the HUD, MICP and MPCD video tape recorders. This applies power to the video tape recorders and enables the eject function. The eject function allows for opening the cassette compartment to remove or install a video tape cassette. Setting the OFF/STBY/AUTO/HUD switch to AUTO or HUD causes the power converter to output the LANC (local application control bus) signal to the recorders. This signal puts the recorders in the record mode.

209. F-15E digital map system The DMS provides a continuous moving digital map image for the aircrews. The digital map is an aeronautical chart created from compressed digitized raster graphics and represents the area being flown over. The DMS supplies this information to the MPD system for the tactical situation display (TSD) format. The map display forms the background against which the current flight situation is shown complete with targets, threats, and steer points.

System components The digital map system consists of three main components: the digital map processor (DMP), the theater cartridge (TC), and the mission cartridge (MC).

Digital map processor The DMP is located in the rear cockpit on the right console. It contains two removable cartridges, a TC and an MC (fig. 1–33). Both cartridges interface with the Air Force mission support system (AFMSS) which the pilots use for DTM mission data loading or unloading. They contain several map segments and each location can have maps at different scales; they could be 10, 20, 40, 80, or 160

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nautical mile maps. The DMP uses present position information from the ADCP and scale select information from the MPD system to determine which map to display. Then, it generates a color video representation of the selected size and sends it to the MPDP. The MPD system takes the map and situational information from several avionics systems, and displays these as the TSD. The DMP has a unit fault indicator that turns black and white when the DMP fails. DMP must be manually reset when this happens. A door on the DMP provides access to the cartridges.

Figure 1–33. DMP and cartridges.

Theater cartridge The TC stores the digital map, digital terrain elevation data (DTED), and static data frames. The TC data is loaded using the AFMSS and is reloaded after aircraft theater reassignment, map updates, or cartridge failures. Opening the lid of the DMP provides access for TC removal and installation and removes power to the TC.

Mission cartridge The MC stores data loaded from the AFMSS in battery-backed static random access memory. The type data stored includes mission specific points, digital map, DTED, and data frame information. Flight data is written to the MC during flight and downloaded using AFMSS. The locking slide on the DMP removes power from the DMP and secures the MC in the DMP.

System operation The DMP uses 3 phase, 115 VAC from the right circuit breaker (CB) panel #3 through miscellaneous relay panel #8. The DMP power supply converts the aircraft power to low voltage DC required by the DMS circuits. The DMS is initialized immediately following the application of aircraft power to the DMP. The DMS will only initialize when both cartridges are installed and the access doors are secured. The DMS automatically enters standby mode after initialization. The DMS remains in standby until a TSD or data frame is selected on any MPD/MPCD. Once aircraft power is on, the TSD format is selected and the cassette is installed, the DMS is operational.

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Modes of operation The DMS operates in two modes: TSD or data frame. TSD mode is entered when TSD is selected from menu 1 on the MPD/MPCD. Data frame mode is entered when DATA FRAME is selected from menu 2 on the MPD/MPCD. It is possible to have a TSD format on one display and a data frame format displayed on another at the same time.

Tactical situation display The DMP provides a moving map by transmitting map data from the MC and TC to the MPCD in response to a command from the ADCP. Navigational data is superimposed over the map display. TSD display enhancements are available and can be selected individually or in combinations using the UFCs, HOTAS (hands on throttle and stick) functions, and MPD/MPCD switches. Figure 1–34 shows a typical TSD. Here are some of the many TSD symbols, controls, and indicators.

Figure 1–34. Typical TSD.

Inverse video An inverse video (INV) option is similar to viewing a photographic negative and is easier for viewing at night. INV is available while the TSD is on a MPD. Inverse video is not an option (blank) on the MPCD. The ADCP commands the DMS to invert the monochrome video (black to white or white to black). A repeated inverse selection toggles the display between inverted and noninverted video.

Zoom The operator can zoom (ZM) or magnify the map by factor of 10/7. For example, a 10 NMI scope shows 7 NMI and an 80 NMI scope shows 56 NMI. Zoom is not an option in 160 NMI range.

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Scale selection The DMS allows the aircrew to select five scales on the TSD format when an aeronautical chart is displayed. The available map scales are 10, 20, 40, 80, and 160 NMI.

Map positioning The aircrew has the capability to select a moving map image that corresponds to the current aircraft position or look ahead to a sequence point on the mission route loaded during the mission planning. The present position format provides a display of the aircraft’s present position with respect to a planned course. The display is oriented with the aircraft heading pointing toward the top of the display. The look-ahead mode shows the route points at the center of the map. The pilot can step through his mission route to see what is coming up. The look-ahead mode orients the display with north at the top of the display (north-up).

TSD declutter When the aircrew selects declutter (DCL), a selected group of information is removed from the TSD. If the DCL is pressed again, another level of information is removed. When pressed again, the TSD is returned to its initial state.

Various symbologies The TSD also displays various symbologies to orient the pilot to the flight path or targets. The table helps define the various symbols found on figure 1–34.

Tactical Situational Display Symbology Type Display Description

Steer points Route lines These points make up the basic route to be flown; they indicate where steering maneuvers will take place. Route lines connect these sequence points. They show where the pilot should be during his or her mission. The pilot can select the desired airspeed and bank angle so the route lines accurately show the path during turns. The TSD also shows the aircraft symbol. The pilot can easily see if the aircraft is off course.

Initial points A square These points show the place that a specific maneuver must be initiated. An initial point is the last steer point before a target point. Each steer or initial point can have up to seven aim points tied to them. They are used to fine tune guidance systems to steer points or initial points. They are displayed as dashed circles or dashed squares, depending on the point with which they are associated.

Target points

A triangle Target points are points to attack. Target data can also be displayed as bearing and range from the target offset point.

Offset points A broken-line triangle

Offset points are always associated with a target point and used for offset bombing. Up to seven offsets per target point can be tied to a target. The bearing and range to the target from the offset point are also displayed.

Bulls eye points

The bull’s eye of an archery or rifle target

These are points from which bearing/range calls can be made. All aircraft on a specific mission will have the same bulls-eye point data. Up to 10 bulls eye points may be loaded either manually using the UFC or automatically by way of the DTM.

Cursor A cross symbol Cursor is a cross symbol that can be moved by the TDC. Static ring threat

Circles centered on a point

Static ring threat (lower left portion of fig. 1–34) is displayed as circles centered on a point where the aircraft is potentially vulnerable to a defined threat. There may be up to four rings associated with a threat point.

Dynamic threat masking

Transparent tint overlays

Threat masks (not shown in fig. 1–34) are transparent tint overlays that represent areas where the aircraft is potentially vulnerable to a defined threat.

Dynamic elevation banding

Two transparent bands

Dynamic elevation banding (not shown in fig. 1–34) compares the aircraft’s altitude and position to surrounding terrain elevations and provides two transparent bands identifying corresponding terrain elevation bands. The DMS uses DTED information stored on the cartridges and the

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Tactical Situational Display Symbology Type Display Description

aircraft’s latitude, longitude, and altitude data from the ADCP to create the bands.

Ownship position

Aircraft symbol The ADCP commands the DMS to display the (middle lower portion of fig. 1–34) either centered on the TSD display or on the bottom of the display by aircrew toggling S6 (CTR/BOT).

TSD sensor cueing

The pilot can command the radar or LANTIRN TGT (target) pod to point at specific positions on the TSD. They do this to develop radar maps or to aim the TGT pod. By positioning the cue point over the selected area, the operator is telling the sensing system to paint data on that area. The data will be painted on the radar or LANTIRN TGT format. The TDC positions the cursor over the desired area. The radar can map a 1.3, 3.3, 4.7, or 10 NMI area.

Data frame mode Data frames are digital files stored within the DMS theater and mission cartridges that contain a variety of imagery information. Selecting DATA FRAME from menu 2 on any MPD/MPCD displays a data frame directory display. A combined maximum of 50 data frames are stored and the MPDP provides a selection box around the data frame number and title. The MPCD/MPD switches are used to page and move the selection box to the desired data frame.

DMS built-in test The three elements that perform BIT are the DMP, MC, and TC. When a BIT failure occurs, the fault is logged and the test repeated. A test is considered failed after five consecutive attempts. A BIT failure is then reported to the avionics MUX-bus and the DMS no-go discrete is set. The cartridge BIT is the same as the DMP except its failure is reported on the DMP pixel bus A.

BITs There are three types of BITs for the DMS; they are power-up, periodic, and initiated BIT.

Built-in-tests Types Description

Power-up BIT Power-up BIT executes automatically when the system powers up. Partial power failures set the no-go discrete and the ASP 59.

Periodic BIT PBIT executes 14 tests during normal operation. One test for each pass through the software executive. If test results are abnormal, the test repeats during the next pass. Test repeats until five failures are recorded.

Initiated BIT IBIT occurs in response to an avionics 1553 MUX bus message. IBIT suspends normal DMP operations to test all circuits. The IBIT may take up to 30 seconds to perform. The DMP stays in IBIT until a stop signal is received from the ADCP. At the end of IBIT, a set of failure codes are written to the MC and relayed to the ADCP. The DMP executes power up initialization and is ready for normal operation within three seconds after receiving the stop command.

Fault indications During failures in any test, the following fail indications happen: ASP 59 latches and the AV BIT caution is displayed on the MPD/MPCD. Additionally, a DMP detail BIT is provided which displays SRU failures.

System integration The following describe the system integration for the DMS.

Advanced display core processor The ADCP controls the interchange of information between the DMS and other avionics components. Command signals from the MPD/MPCD and the DMS processor to the ADCP control the displays in

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the cockpit. The ADCP also combines mono and color video from the DMS with inputs from the MUX bus to produce the map displays on the MPD/MPCD.

Multi-purpose display system The MPD/MPCD displays TSD symbology superimposed over a moving map. Switches on MPD/MPCD provide a means to select options and edit the displays.

Left and right hand controllers The left and right hand controllers are used to take command and position sensors.

Avionics status panel The ASP latches ASP 59 during failures.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

208. Video tape recording system 1. What does the VTRS system record?

2. Where is the HMD/video control panel located?

3. Which switch on the HMD/video control panel is used to provide aircraft power to the VTRS?

4. The power converter provides what functions?

5. When is an event marker put on the MICP and HUD video signal?

6. What does an event marker symbol look like?

7. Where are the MICP, HUD, and MPCD cameras mounted?

8. Where are the MICP, HUD, and MPCD video tape recorders located?

9. What is the record time of the 8 mm video tape cassette?

10. Explain the purpose of the EU in the video tape recording system?

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11. The VTR REC light comes ON under what condition(s)?

12. The VTR CAUTION light comes on under what condition(s)?

209. F-15E digital map system 1. List the DMS components.

2. Where is the DMP located?

3. Where does the DMP obtain present position?

4. What is stored on the TC?

5. What type of data is stored on the MC?

6. What are the power requirements for the DMS?

7. Where are the DMP circuit breakers located on the aircraft?

8. When will the DMS come out of standby mode?

9. Name the two operating modes of the DMS.

10. Describe the TSD format.

11. What permits improved viewing of the TSD display at night?

12. What are the available map scales?

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13. What does declutter do for the pilot?

14. Describe the symbol and purpose of a steer point.

15. Describe the static ring threat.

16. What is sensor cueing?

17. What is data frame mode?

18. When is a test considered failed?

19. Which BIT suspends normal DMP operations?

20. Which DMS BIT will display a failed SRU?

Answers to Self-Test Questions 201 1. Radar control panel (541), power supply (610), RF oscillator (001), transmitter (011), receiver (022),

analog processor (039), data processor (081), programmable signal processor (042), and antenna (031). 2. The RFO (001). 3. Transmitter (011). 4. Analog processor (039). 5. 031 positioning commands. 6. The 042 performs target detection, filtering, range measurement, target parameter calculations, clutter

canceling, and IFF correlation. It also provides radar display parameters to the IG. 7. PSDP. 8. The 081. 9. The transmitter through a waveguide. 10. By using the guard horn, a small, wide-angle horn located at the bottom of the antenna, angled down

slightly. It is a receive-only horn designed to intercept sidelobe echoes. With the ability to measure Doppler shift, we can use these ground returns to measure our ground speed. Knowing this, we can filter out all main-beam returns that come back to us with the same Doppler signature, as they must be just returns from the ground ahead. This process is called clutter rejection and occurs in the 042.

11. Search and track.

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12. The data processor (081). 13. The transmitter (011). 14. G-TEST is displayed in the BIT window. If TEST appears, the radar thinks it is airborne and will transmit. 15. Standby IBIT. 16. An HPRF radar mode provides maximum detection against nose-aspect targets. Its high average power

results in long-range detection and tracking capability. (The close spacing of the pulses means the transmitter is ON a lot, compared to its listening time.) A problem with HPRF is that it has poor performance against tail aspect targets and near zero capability against targets with no closure rate. HPRF is available in all ranges except 10 NMI.

17. 10 through 80 NMI. 18. Track-while-scan. 19. LRS. 20. It is the process of picking out a target to track. You may acquire a target manually or automatically. 21. The TDC on the right throttle grip. 22. It uses LPRF to determine the slant range from the aircraft to a ground point. It is used primarily for

bombing/gun strafing.

202 1. Eight. 2. The receiver and RFO. 3. 031, 111, and 025. 4. It is an intermediate PRF that falls between HPRF and MPRF. The pulses are spaced a little farther apart

than in HPRF, but not quite as far as in MPRF. This produces some of the desirable effects of both the old waveforms.

5. It will search in the 20, 40, 80, and 160 NMI ranges and it uses all high PRF. 6. It uses all MPRF to detect all-aspect targets from 10 to 80 NMI. 7. There are four A/G modes available to the pilot and all of them use LPRF waveforms. 8. To compensate for errors in INS velocities. 9. The radar ceases all other functions, latches fault indicators, and displays the BIT matrices on the MICP. 10. ORT matrix, G-BIT matrix, F-BIT matrix, CM matrix, and FRESH matrix. 11. Performing standby-initiated BIT will display the matrices. 12. The AUTO-ACQ/REJ switch cycles between the five matrices. 13. A persistence counter tracks the number of CM-BIT failures. The failures are not used in fault isolation if

their persistence count is less than seven in one minute.

203 1. Transmitter (111), receiver exciter (025), power supply (610), analog signal converter (038), antenna (031),

radar data processor (082), and PSP (044). 2. 044. 3. It is used to provide a conventional mapping mode for low resolution ground mapping, weather returns or

for HRM cueing. 4. GMT mode. 5. It is used to provide a high-resolution map for A/G weapons delivery, high range resolution, and azimuth

resolutions by employing SAR techniques. 6. Used to update the MN or INS. 7. Performs system tests and calibration on individual LRU and the entire radar set to establish a high level of

confidence that the radar will do its mission satisfactorily. 8. The GBIT is the maintenance BIT and FBIT is the in-flight BIT. FBIT is performed in the air, while the

GBIT is performed on the ground. 9. The state of the WOW switch.

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10. FBIT.

204 1. CFRS/CFI. 2. DTM. 3. NO! Any deviation could void the equipment warranty. 4. Radar control panel (342), receiver/exciter (325), transmitter (311), analog signal converter (338), radar

data processor (385), power supply (310), and antenna (031). 5. Coffin handles. 6. RDP (385). 7. Automatic and initiated. 8. POST when executed at standby cold start. 9. POST when executed at operate/emergency cold start. 10. It will cause the BIT matrices to be displayed.

205 1. To determine the structural loads being applied to the aircraft. 2. A two-position momentary toggle switch in the NLG wheel well. The switch is used with the NCI to clear

overload conditions from the CC memory and erase the OWS matrix display from the MICP. 3. CC—Digital computer that contains OWS software and performs the calculations. PSDP— is a display

processor that receives OWS data from outside systems, conditions them, and sends them to the CC to be used in the OWS calculations. The PSDP also processes the OWS information displayed on the MICP, HUD, and heard in the headset.

4. 4 inputs: AOA, Mach #, pressure ratio, barometric corrected altitude. 5. The PACS sends signals telling the CC what is loaded on the aircraft. 6. Measures the current G force on the aircraft and sends it to the CC. 7. When a certain level of stress is being put on the aircraft. A 900 Hz tone interrupted @ 4 Hz rate when 85–

92% max allowable load. A 900 Hz tone interrupted @ 10 Hz rate when 92–96% max allowable load. A 900 Hz solid tone when 96–100% max allowable load. Voice warning when 100+% max allowable load.

8. 1st row. 9. The HUD window 8 display is the only OWS display the pilot will see and it will show current Gs and the

maximum allowable Gs. If the maximum allowable Gs cannot be computed for any reason, then window 8 will display the current Gs and NOWS.

206 1. (1) CC.

(2) H009 MUX bus. (3) 1553 MUX bus. (4) DTM receptacle. (5) MLV receptacle.

2. (1) A/A. (2) A/G. (3) ADI.

3. To provide an interface between the DTM and the PSDP in order to load mission data into the CC. 4. H009 and 1553. 5. (1) DPM.

(2) IOM. (3) BMM. (4) BSM. (5) TDM.

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6. Three-phase, 115 VAC, 400-Hz. 7. It’s a programmable, battery powered, nonvolatile memory device that transfers flight operations mission

data to the CC. 8. MAINT and OPS. 9. PSDP. 10. The process of the CC reading data from the DTM. 11. Instructions necessary to do radar, general navigation, and weapon delivery computations. 12. Executive, A/A, A/G, navigation, flight director, controls and displays, and self-test. 13. All other program functions. 14. To assist in management and delivery of air-to-air weapons; receive radar data to assist in target tracking

and display of target data; provide data for steering and displays; do computations; display HUD cues for gun mode operations (providing lead angle); and provide missile prelaunch and post-launch management.

15. Provides TACAN, NAV, and ILS steer mode processing for display on the HSI. It also performs range and bearing or steering computations for display on the ADI and HUD in TACAN, NAV, and ILS steering modes.

16. By examining the CC I/O memory data and the internal memory parameters of the CC database. 17. Convert it to a 16-digit binary word and then use tables from TO 1F-15A-2-31GS-42-1 to extract its

information.

207 1. Removed the CC and MPDP, replaced with the ADCP. 2. Top left shelf of door 3L. 3. The GPP, the IPM, and the VME64 main and essential busses. 4. Normal, degraded, emergency power, and rest. 5. The degraded mode is enabled when a simple failure is detected that does not reduce functionality or when

a severe failure occurs and causes a main and essentials bus split. All failures are stored in the bridge modules fault logs. If a bus split occurs, the functioning bus (main/essential) will automatically be selected. The displays associated with the failed bus will not be available with the exception of the functions that are redundant on the buses. A bridge module failure will force ADCP operation to the main bus. A GPP failure will be flagged, but the other GPP will support all systems. If a module fails, the displays supported by that module will not be available for display.

6. Power-up, background, maintenance, initiated. 7. Background BIT. 8. MUX 7A/7B. 9. MUX 8A/8B. 10. Executive, A/A, A/G, NAV, FDL, controls and displays and computer self test functions.

208 1. It records inflight color video of the MICP, HUD and MPCD. 2. It is mounted on the right side of the head-up display unit. 3. The OFF/STBY/AUTO/HUD switch. 4. It converts aircraft 28 VDC to system operating voltages. 5. When either the TRIGGER switch is pressed to the second detent or the WEAPON RELEASE switch is

pressed. 6. A rectangle. 7. The MICP camera is mounted to the beamsplitter assembly located above the MICP. The HUD camera is

mounted in a fixture located on the center glareshield. The MPCD camera is mounted on the lower left windscreen.

8. The recorders are located in the cockpit in the map case. 9. It provides 120 minutes of recording time.

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10. The EU receives HDU video from the HDU camera and HUD video from the HUD video control unit. The EU overlays the video from the HDU camera with the symbology that makes up the HMD. With the HMD/video control panel OFF/STBY/AUTO/HUD switch set to AUTO, the EU outputs HUD or HDU video for recording based on commands from the CC.

11. The red REC (record) light comes on when the recorders are in record mode. 12. The CAUTION light comes on/flashes when any of the following exists: an input power problem exists,

cassette tape is near end or has run out, moisture is detected inside video tape recorder, video heads are contaminated, some other malfunction exists or with record mode enabled red tab on tape cassette is out, no cassette is inserted.

209 1. DMP, MC, and TC. 2. Rear cockpit on the right console. 3. It uses present position information from the ADCP. 4. Digital map, DTED, and static data frames. 5. Mission specific points, digital map, DTED, and data frame information. 6. 3 phase, 115 VAC. 7. Right, CB panel #3. 8. The DMS remains in standby until a TSD or data frame is selected on any MPD/MPCD. 9. TSD or data frame. 10. Navigational data is superimposed over the map display. 11. Inverse video (INV). 12. 10, 20, 40, 80, and 160 NMI. 13. Removes a selected group of information from the TSD. If the DCL is pressed again, another level of

information is removed. When pressed again, the TSD is returned to its initial state. 14. Indicate where steering maneuvers will take place. 15. Displayed as circles centered around a point where the aircraft is potentially vulnerable to a defined threat. 16. Commanding the radar or targeting pod to look at specific points on a TSD. 17. Digital files stored within the DMS theater and mission cartridge that contains a variety of imagery

information. 18. After five consecutive attempts. 19. IBIT. 20. Detail BIT.

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Unit Review Exercises Note to Student: Consider all choices carefully, select the best answer to each question, and circle the corresponding letter. When you have completed all unit review exercises, transfer your answers to the Field-Scoring Answer Sheet.

Do not return your answer sheet to the Air Force Career Development Academy (AFCDA).

1. (201) On the F-15 APG-63 radar system, which component provides manual control of radar set power, operating modes, and mode parameters? a. Radio frequency (RF) oscillator. b. Radar control panel. c. Analog processor. d. Power supply.

2. (201) Which F-15 APG-63 radar system component sends antenna positioning commands to the servo electronics in the power supply? a. Data processor. b. Analog processor. c. Radio frequency (RF) oscillator d. Programmable signal processor.

3. (201) On the F-15 APG-63 radar system, which component process clutter rejection? a. Data processor. b. Analog processor. c. Radio frequency (RF) oscillator. d. Programmable signal processor.

4. (201) Which F-15 APG-63 radar system component stores the radar operational flight program (OFP)? a. Data processor. b. Analog processor. c. Radio frequency (RF) oscillator. d. Programmable signal processor.

5. (201) On the F-15 APG-63 radar system, ASP 25 will latch when the waveguide pressure falls below how many pounds per square inch absolute (PSIA)? a. 12. b. 16. c. 20. d. 24.

6. (201) Which F-15 APG-63 radar search mode locks on to the first target that falls within the scan pattern out to a range of 10 nautical miles (NMI) and is considered the primary automatic acquisition mode? a. Supersearch. b. Velocity search. c. Long-range search. d. Short-range search.

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7. (201) Which F-15 APG-63 radar scan mode can maintain up to 10 target track files while continuing to detect and display up to 18 more (half-intensity) observation targets? a. Gun scan. b. Vertical scan. c. Single-track-scan. d. Track-while-scan.

8. (202) Which F-15C/D APG-70 radar set component combines the functions of the APG–63 receiver and radio frequency oscillator (RFO) into one line replaceable unit (LRU)? a. Power supply. b. Receiver-exciter. c. Radar control panel. d. Radar data processor.

9. (202) Which F-15C/D APG-70 radar air-to-air (A/A) search mode uses the “intermediate” pulse repetition frequency (PRF) to detect targets in intermediate-range and short-range? a. HI. b. MED. c. RGH. d. INLV.

10. (202) What F-15C/D APG-70 radar air-to-ground (A/G) mode does the pilot utilize prior to making a bomb run? a. Air-to-ground (A/G) beacon. b. Precision velocity update. c. Real beam map. d. A/G ranging.

11. (202) On the F-15C/D APG-70 radar set, which built-in test (BIT) history matrix contains the latest known test results? a. CM. b. F-BIT. c. G-BIT. d. FRESH.

12. (203) Which F-15E APG-70 radar set component establishes the basic operating radio frequency (RF) for the transmitter? a. Power supply. b. Receiver-exciter. c. Radar data processor. d. Analog signal converter.

13. (203) Which F-15E APG-70 radar set component converts the intermediate frequency (IF) analog data to digital information which is then sent to the programmable signal processor (044) for processing? a. Power supply. b. Receiver-exciter. c. Radar control panel. d. Analog signal converter.

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14. (203) Which F-15E APG-70 air-to-ground (A/G) radar mode provides a slant range measurement used for target altitude determination, for target designation, or for position updates? a. Real beam map mode. b. High resolution map mode. c. Air-to-ground ranging mode. d. Air-to-ground beacon mode.

15. (204) On the F-15 APG-63(V)1 radar, which component stores the radar operational flight program (OFP) and protects the radar from overheating? a. Power supply. b. Receiver-exciter. c. Radar data processor. d. Analog signal converter.

16. (204) Which F-15 APG-63(V)1 radar air-to-ground (A/G) mode provides general navigation? a. A/G ranging. b. Real beam map (RBM). c. Ground moving target (GMT). d. Precision velocity update (PVU).

17. (205) The F-15 overload warning reset switch is used with the navigation control indicator (NCI) to clear overload conditions from the memory of which component? a. Central computer (CC). b. Multiple indicator control panel (MICP). c. Head-up display (HUD) data processor. d. Programmable signal data processor (PSDP).

18. (205) Which F-15 overload warning system (OWS) component processes the OWS information displayed on the multiple indicator control panel (MICP), head-up display (HUD), and is heard in the headset? a. Central computer. b. Radar data processor. c. Multiple purpose display processor. d. Programmable signal data processor.

19. (205) Which F-15 component sends flight data consisting of angle of attack (AOA), Mach number, pressure ratio, and barometric corrected altitude across MUX bus channels 1 and 3? a. Central computer (CC). b. Air data computer (ADC). c. Multiple indicator control panel (MICP). d. Programmable signal data processor (PSDP).

20. (205) Which F-15 component sends signals telling the central computer (CC) whether wing tanks and/or conformal fuel tanks (CFT) are loaded on the aircraft? a. Air data computer (ADC). b. Fuel quantity signal conditioner. c. OVERLOAD WARNING RESET switch. d. Programmable armament control system (PACS).

21. (205) The F-15 central computer (CC) sends overload warning system (OWS) information to which line replaceable unit (LRU)? a. Air data computer. b. Inertial navigation unit. c. Head-up display (HUD) data processor. d. OVERLOAD WARNING RESET switch.

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22. (205) At what F-15 overload warning system (OWS) percentage of maximum allowable load will the voice warning (OVER-G, OVER-G) be heard? a. 85 to 92. b. 92 to 96. c. 96 to 100. d. Above 100.

23. (205) What is displayed in window 8 on the F-15 head-up display (HUD) if the overload warning system (OWS) is not operational? a. NOWS. b. OWS OFF. c. OWS FAIL. d. Display is blank.

24. (206) On the F-15, what types of data transfer modules (DTM) can be used by the DTM receptacle? a. Communications (COMM) and maintenance (MAINT). b. COMM and control (CONT). c. Operations (OPS) and MAINT. d. OPS and CONT.

25. (206) Which operational flight program (OFP) function interfaces with all other program functions? a. Executive. b. Navigation. c. Flight director. d. Controls and displays.

26. (206) Which operational flight program (OFP) function handles tactical air navigation (TACAN), navigation (NAV), and instrument landing system (ILS) steer mode processing for display on the horizontal situation indicator (HSI)? a. Executive. b. Navigation. c. Flight director. d. Controls and displays.

27. (207) On the F-15E aircraft, which component within the advanced display core processor (ADCP) services the multipurpose display processing and central computer processing functions? a. Intelligent serial module (ISM). b. Image processor modules (IPM). c. General purpose processor (GPP). d. VME64 main and essential busses.

28. (207) On the F-15E aircraft, which system does the advanced display core processor (ADCP) communicate with over the avionics 1553 MUX bus 5A/5B? a. Fighter data link (FDL). b. Internal countermeasures set (ICMS). c. Embedded global positioning system (EGPS). d. Programmable armament control system (PACS).

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29. (208) Which F-15 video tape recording system component causes a rectangle to be displayed on the multiple indicator control panel (MICP) and head-up display (HUD) recorded video displays? a. Power converter. b. Electronics unit (EU). c. Video event marker generator. d. Helmet mounted display (HMD)/video control panel.

30. (208) Which F-15 video tape recording system light will come on/flash when moisture is detected inside video tape recorder? a. FAIL. b. CAUTION. c. MOISTURE. d. Record (REC).

31. (209) Which F-15E digital map system (DMS) component uses present position information and scale select information to determine which map to display? a. Mission cartridge (MC). b. Theater cartridge (TC). c. Digital map processor (DMP). d. Advanced display core processor (ADCP).

32. (209) Which F-15E digital map system (DMS) component stores data loaded from the Air Force mission support system (AFMSS) in battery-backed static random access memory? a. Mission cartridge (MC). b. Theater cartridge (TC). c. Digital map processor (DMP). d. Advanced display core processor (ADCP).

33. (209) What F-15E digital map system (DMS) tactical situation display is displayed as circles centered on a point where the aircraft is potentially vulnerable to a defined threat? a. Bulls eye points. b. Static ring threat. c. Dynamic threat masking. d. Dynamic elevation banding.

34. (209) Which F-15E digital map system (DMS) built-in test (BIT) suspends normal digital map processor operations to test all circuits? a. Periodic. b. Initiated. c. Power-up. d. Background.

Please read the unit menu for unit 2 and continue

Unit 2. Instrument and Flight Control Systems 2–1. Instrument Systems .................................................................................................................. 2–1

210. F-15 fuel quantity indicating system ................................................................................................. 2–1 211. F-15E engine instrument system ........................................................................................................ 2–6 212. Acceleration indicating/G-exceedance system ................................................................................ 2–11

2–2. Primary Flight Control Systems ........................................................................................... 2–16 213. Lateral flight controls ...................................................................................................................... 2–16 214. Longitudinal flight controls ............................................................................................................. 2–19 215. Directional flight controls ................................................................................................................ 2–21

2–3. Automatic Flight Control Systems ........................................................................................ 2–26 216. F-15A through D control augmentation system ............................................................................... 2–26 217. F-15 trim systems ............................................................................................................................ 2–38 218. F-15E automatic flight control system ............................................................................................. 2–45

2–4. Engine Air Intake System ...................................................................................................... 2–60 219. Engine air intake system components, inputs, and outputs .............................................................. 2–60 220. Auto and emergency modes and ground operation .......................................................................... 2–65

2–5. Air Data Systems .................................................................................................................... 2–68 221. Pitot-static system ............................................................................................................................ 2–68 222. F-15 A through D air data computer ................................................................................................ 2–73 223. F-15E air data processor .................................................................................................................. 2–82

E WILL BEGIN THIS UNIT by providing you with the operating theory of the F-15 instrument systems. Afterwards, you will learn how the F-15 primary flight controls keep the aircraft inflight using the lateral, longitudinal, and directional flight controls. We will then build on

that knowledge as we dive into the automatic flight controls. Then we will tackle the ever complex engine air intake system components and operation. We will conclude the unit with a discussion on the air data systems of the F-15.

2–1. Instrument Systems In order for any aircraft to successfully take off, fly to and accomplish its mission and then return home safely, it requires well designed and properly operating instrumentation systems. As an avionics system craftsman, your job will be to ensure these systems are properly maintained to enable your F-15 pilot to complete a successful mission.

210. F-15 fuel quantity indicating system The fuel quantity indicating system consists of the tank units (commonly called fuel probes), an indicator, and tank aboard relays. A fuel quantity signal conditioner is also installed on the F-15C, D, and E model aircraft. The F-15 D and E models also have a rear cockpit indicator.

System components The following paragraphs describe the fuel quantity indicating system components.

Tank units The tank units are commonly referred to as fuel probes. Look at the bottom of the fuel probe illustrated in figure 2–1. The drawing shows the fuel probe consists of a skinny metal tube inserted into a larger tube. The outside surface of the inner tube and the inside surface of the outer tube act as the capacitor’s plates. The probe is designed and mounted so fuel easily passes into the tubes, filling the space between them to the same level as the surrounding tank. This causes the probe’s capacitance

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to change proportionally to changes in the tank’s fuel level. The fuel probes also contain two diodes, which rectify the AC (alternating current) input into positive and negative DC (direct current) pulses. There are 13 internal tank units (this includes fuselage and wings), one in each 600-gallon external tank and three in each conformal fuel tank, for a maximum of 22 tank units.

Figure 2–1. Typical fuel probe.

Fuel quantity indicator The fuel quantity indicator continuously displays the total internal fuel on the internal pointer, and internal and external fuel on the total pounds counter. In addition, the fuel quantity indicator contains a fuel quantity selector, which allows monitoring of each individual internal and external fuel tank on the left and right pounds counters. Because of the increased number of internal fuel tanks, tank probes, and conformal fuel tanks on the F-15C, D, and E model aircraft, along with the addition of the signal conditioner, a new indicator (fig. 2–2) had to be designed.

Figure 2–2. F-15 fuel quantity indicator (front cockpit).

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The fuel quantity selector also has a built-in test (BIT) position for troubleshooting purposes. Selection of the BIT on the indicator commands the internal (INTL) pointer to 6,000 ± 200 pounds (lbs) TOTAL LBS COUNTER to 6,000 ± 200 lbs, and LEFT and RIGHT LBS COUNTERS to 600 ± 50 lbs. The F-15D and E rear cockpit INTL pointers indicate to within 100 pounds of the main fuel quantity indicator INTL, and the rear TOTAL LBS COUNTER indicates to within ± 200 pounds of the main fuel quantity TOTAL LBS COUNTER.

Fuel quantity signal conditioner The signal conditioner is a separate line replaceable unit (LRU) installed on F-15C, D, and E model aircraft. The unit has a manually resettable fault indicator. The signal conditioner receives pulsed DC inputs from the tank units and supplies signals to drive the fuel quantity indicator. It provides the power to tank units, senses the output (capacitance) of the tank units, compares it to the reference capacitance, and drives the appropriate pointer or counter. It also contains the adjustment screws for calibrating the system.

Tank aboard relays These relays tell the fuel quantity system that a 600-gallon external fuel tank is installed. When the tank is installed, the relay allows the signal conditioner to receive the fuel signal from the tank’s probe. When the tank is not installed, the relay activates a zero reference capacitor in the signal conditioner; thus forcing that indication to zero. This ensures the tank indication is zero when the tank is not installed.

Fuel quantity indicator operation principles The value of the tank capacitor (probe or tank unit) varies with the amount of fuel. The fuel, acting as a dielectric, controls the variable capacitance of the tank unit. Tank unit diodes rectify input AC signals and provide two pulsed DC signals to the signal conditioner. All positive DC inputs are applied to control individual tank indications (LEFT lbs and RIGHT lbs counter) and all negative DC inputs are applied to control total indications (INTL pointer and TOTAL lbs counter).

Figure 2–3 shows a basic representation of a DC system. Only one tank unit is depicted. All the tank units for each tank are connected in parallel. Power is provided to the tank unit by the signal conditioner on the F-15C, D, and E aircraft. After the tank unit produces an AC signal proportional to the probe capacitance, the diodes in the tank unit rectify that signal (make it into a pulsating DC). The resulting DC signal then goes to the signal conditioner on the F-15C, D, and E.

Figure 2–3. DC operation in a tank unit.

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Internal pointer The internal pointer displays the total fuel in the internal fuselage and internal wing tanks. It uses the negative DC output pulse from the tank units. Otherwise, the system operation is the same as what we covered in the previous paragraph.

Left and right counters Individual tank amounts are displayed on counters controlled by the FUEL QUANTITY SELECTOR switch on the fuel quantity indicator. The switch positions are as follows:

• With the switch in the FEED position, the left counter displays the fuel in tanks 3A and 3B that feeds the left engine and the right counter displays tank 2 that feeds the right engine.

• In the INTL WING position, the left counter displays the fuel in the left-wing tank, while the right counter displays fuel in the right-wing tank.

• In the TANK 1 position, the left counter displays the fuel in fuselage tank 1. The right counter reads zero.

• In the EXT WING position, the left and right counters display the fuel in the external wing tanks.

• In the EXT CENTER position, the left counter displays the fuel in the centerline tank. The right counter reads zero.

• In the CONF TANK position, the left and right counters display the fuel in the conformal fuel tanks.

Total counter and internal pointer (totalizer) The total counter and internal pointer (totalizer) use the negative DC pulse from the tank units, but display the fuel in all the internal and external tanks installed on the aircraft. If no external tanks are installed, the total counter should display the same amount of fuel as the internal pointer.

BINGO The BINGO fuel indicator on the F-15C, D, and E fuel quantity indicator provides an indication when aircraft fuel reaches a preset level. You or the pilot can set the BINGO indicator to any position by rotating the BINGO knob, located on the upper right hand corner of the fuel quantity indicator. When the internal pointer reaches the preset level, the BINGO FUEL light on the caution lights display panel illuminates and a voice warning of BINGO FUEL is initiated, repeated once with a two-second delay. On the F-15D, and E, the rear BINGO FUEL light also illuminates. When fuel DUMP is selected, the aircraft forces fuel out the dump mast. Once the BINGO fuel level is reached, the fuel system automatically stops fuel dump operations.

Calibrating the F-15 fuel quantity system To ensure the accuracy of the fuel quantity indicating system, the system needs to be calibrated after replacing the signal conditioner (F-15C, D, and E), or, occasionally, when the fuel quantity system fails the operational check.

Prepping the system Adjusting the fuel quantity indicating system is quite simple. First, make sure the aircraft is defueled. It’s important that the tanks are not depuddled (completely drained) because the engines do not consume every drop of fuel in the aircraft. The system leaves minor amounts of residual fuel in each fuel tank. With the fuel tanks empty of fuel, perform the empty calibration. The empty calibration sets a zero fuel reference needed to perform a full calibration. The adjustment screws are rotated for each tank until a zero indication is reached on the internal pointer and all pounds counters.

Adjustment screws The adjustment screws for the F-15C/D/E are located on the signal conditioner (fig. 2–4).

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Figure 2–4. F-15C/D/E fuel quantity adjustment screws.

Hydrometer testing Once you’ve adjusted the empty potentiometer screws, have the aircraft refueled and adjust the full adjustment screws. The procedure has you use a hydrometer before making your adjustments. A hydrometer is a devise used to measure fuel density. Fuel density varies with temperature in a range of 6.0 to 7.0 lbs per gallon. The hydrometer allows you to draw fuel, from a sample taken from the aircraft, into its two fluid chambers where density can be measured. The measurement consists of observing the fuel level on a float device etched with a pounds scale. The fuel sample is taken from the tank 1 drain, just above the nose of the centerline tank. The full indication to which you adjust the pointer and counters depends on the density of the fuel. Fuel density can make quite a difference in the values you must adjust to. Make sure you follow applicable technical data when performing these procedures.

Performing maintenance on the F-15 fuel quantity indicating system Now, let’s talk about two areas you’ll encounter on almost any system—operational checks and troubleshooting.

System checkout Before a suspected malfunction can be analyzed, a system operational check must be performed. The operational check of the fuel quantity indicating system consists of BIT, and empty, full, and fuel low-level warning checkouts.

BIT check When the SELECTOR switch on the fuel quantity indicator is placed to the BIT position, the internal pointer and counters drive to predetermined values. Ensure the pointer and counter drive smoothly and are within technical order (TO) tolerances. The BINGO function is also checked by BIT.

Empty and full checkouts The empty checkout must be done after the aircraft is defueled. Simply look for indications of zero ± TO tolerances. The full checkout is performed after the aircraft has been fully fueled and is done to ensure the indications are within tolerance of the full fuel indications. If some doubt exists about the displays, troubleshoot the system. Since compensation isn’t provided by the system, changes in fuel density cause the system to read somewhat higher or lower.

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Typical malfunctions Uncovering problems with the fuel quantity indicating system is normally quite simple. There are generally three types of problems—above or below normal, at the mechanical stop, and fluctuations.

Above or below normal The above normal malfunction may be caused by a faulty tank unit or a short in the wiring, and the below normal malfunction is caused by a faulty tank unit or an open in the wiring.

At the mechanical stop When the fuel quantity indicator pounds counter is driven to its mechanical stop it is a probable indication of a faulty tank unit.

Fluctuations The cause of indication fluctuations can be loose or faulty wiring, or an improperly installed tank unit. The troubleshooting procedures are very basic. Use the fuel quantity gauging test set to isolate a faulty tank unit in the internal wing tanks. To isolate a faulty internal fuselage tank unit, disconnect one tank at a time and observe the respective indicator. If the indicator does not decrease its indication when the tank is disconnected, the wiring is faulty.

NOTE: The gauging box for the F-15C, D, and E simulates capacitance and uses utility receptacle power. Troubleshooting with the gauging box is the same for all F-15 aircraft. Remember to always use the fault isolation TO to isolate and correct any malfunction.

211. F-15E engine instrument system In this lesson we will discuss how the F-15E engine instrument system functions and how it differs from the F-15C/D models. The primary difference is how the information is displayed to the pilot.

Purpose The F-15E uses the engine monitor display (EMD) as the primary engine display and the advanced display core processor (ADCP) as a backup source for engine information. The EMD in the F-15E, along with the multipurpose displays (MPD) and multipurpose color displays (MPCD) of the ADCP system, are major improvements over the old synchro/servo indicators developed in the 1920s and still used on many aircraft, like the F-15C/D. The BIT capabilities of this digital display simplify maintenance and increase the reliability of the system.

System components and operation The F-15E engine instrument system main component is the EMD. It interfaces with the engine diagnostic unit (EDU), avionics interface units (AIU), and fuel flow transmitter for engine indications. The MPDs/MPCDs are the backup and weapons system officer’s (WSO) display. They display data provided by the EDU and digital engine electronic control (DEEC) on each engine through the engine multiplex (MUX) bus transformers, AIUs, and the ADCP.

Engine diagnostic unit The EDU is a self-contained microcomputer located in panel 95L/R; it replaces the engine interconnect box found on the C–D model while adding new features. The EDU receives power from the aircraft electrical system, receives data from the DEEC, and various engine sensors and the airframe. The data is stored and processed to indicate displays, faults, cautions, and maintenance requirements. Signals from the engine-mounted sensors (revolutions per minute (RPM), temperature, nozzle position, and oil pounds per square inch (psi)) are fed to the engine-mounted EDU. The EDU sends this information directly to the EMD and to the ADCP, by way of the 1553 MUX bus. The fuel flow (FF) signal is fed from the aircraft-mounted fuel flow transmitter directly to the EMD and AIU No. 1. The AIU No. 1 digitizes the FF signal and sends it to the ADCP over the 1553 MUX bus to be displayed on the MPD/MPCD.

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Two engine parameters, fan turbine inlet temperature (FTIT) and OIL PRESS, are monitored by ADCP for engine operating limitations and are part of the F-15E caution and warning system. An over-temperature condition will activate the voice warning system and the minimum oil pressure condition will be shown on the display along with an L/R OIL PRESS light. In addition, the ENGINE caution lamp and the MASTER caution lamp will illuminate.

Engine monitor display The EMD is internally lit, and located on the right side of the main instrument panel. It is a liquid crystal display (LCD) with identical, divided, digital windows. These dedicated windows display their respective engine RPM, FTIT, fuel flow, nozzle position, and oil pressure. The analog readings from various sensors are converted to digital format in the indicator and are presented as one display, rather than on 10 individual analog indicators. Refer to figure 2–5 for an illustration of the EMD.

Figure 2–5. EMD.

The EMD has three BIT functions; continuous, periodic, and initiated. During continuous BIT, the EMD inspects internal signals without interruption, to ensure they are within normal signal ranges. The periodic BIT (PBIT) combines with normal data transfer and does not interfere with equipment operation. Initiated BIT (IBIT) is the same as the PBIT; however, the IBIT adds a test display, program test, and read/write test.

If the EMD fails BIT, a failure signal is sent to the AIU No. 1 and the avionics status panel (ASP). Additionally, a shop replaceable unit (SRU) failure signal is sent to AIU No. 2. The LRU failure signal will cause ASP indicator 61 to latch, the avionics (AV) BIT light on the caution lights display panel to illuminate, and AIU No. 1 will send a signal to the ADCP over the 1553 MUX bus. The ADCP will display EMD* on the MPD/MPCD BIT display. AIU 2 receives the SRU failure data and the exact SRU failure can be determined by viewing the DETAIL page via the MPD/MPCD.

If one of the signals for the engine parameters is out of range, no LRU or SRU failure signal is sent from the EMD. The window for the failed parameter will go blank, but the AV BIT light will not come on and the ASP 61 will not latch. The one exception to this is if the RPM input exceeds its maximum 110 percent range. In that case, the EMD will send the signal to cause the AV BIT light to illuminate and ASP 61 to latch. All engine parameters may be displayed on any MPD or MPCD as a backup for the EMD (EMD repeater), for troubleshooting purposes or as rear cockpit readout. Refer to figure 2–6 for an illustration of the MPD with the engine parameters displayed.

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Figure 2–6. MPD.

N2 RPM indication system (0–110 percent) The auxiliary winding of the engine-driven alternator produces a single phase alternating signal with voltage and frequency directly proportional to N2 RPM. This signal is sent to the DEEC, and the EDU. From the EDU, it is sent to the EMD for display. The EDU also converts the alternating signal to a digital signal and sends it to the ADCP over the avionics 1553 MUX bus. The ADCP system uses this signal for the EMD repeater. Refer to figure 2–7 for an illustration of the RPM signal flow.

Figure 2–7. RPM signal flow.

Once received, the EMD converts the RPM signal to a digital signal for display on the LCD. The display range is 0 to 110 percent in increments of 1 percent. The EMD will blank that specific RPM window if the engine parameters are out of the EMD indicating range during flight and trigger an AV BIT light and ASP 61 if the engine RPM reaches above the 110 percent range.

Fan turbine inlet temperature indicating system Seven alumel-chromel sensors connected in parallel determine fan turbine inlet temperature. The exposed sensors develop a DC voltage direct proportion to FTIT. This voltage is transmitted from the sensors by way of alumel-chromel leads and passes through the EDU on its way to the EMD. The EMD receives the FTIT signal, converts it to digital, and displays FTIT in 10º increments from 200º to 1,400º C (degrees Celsius). The EMD will blank the applicable FTIT window if an engine’s temperature indication goes beyond EMD range. Refer to figure 2–8 for an illustration of the FTIT signal flow.

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Figure 2–8. FTIT signal flow.

As this figure indicates, the EDU also converts the received FTIT voltage to a digital signal and sends it to the ADCP over the avionic 1553 MUX bus. The digital FTIT information is used by the ADCP system for the EMD repeater and by the ADCP to establish “overtemp” conditions. If the ADCP determines a high turbine temperature (greater than 1,000º C), it commands the AIU 1 to send FTIT L/R warning signals to the intercommunications set control panel (ICSCP). The ICSCP then sends the audio signal “WARNING-FTIT OVERTEMP LEFT/WARNING-FTIT OVERTEMP RIGHT” to the head set.

Fuel flow indicating system The FF transmitters (accessed in 95 L/R & 113 L/R) provide a synchro signal to the EMD and AIU 1. Each transmitter is made up of two fuel-tight compartments, one containing a drive motor and the other, a synchro transmitter. Refer to figure 2–9 for an illustration of the F-15E fuel flow signal development. As the drive motor operates, it magnetically turns the momentum wheel at a constant rate in opposition to the fuel flow. As fuel flow increases, the torque required to turn the momentum wheel also increases. This energy is magnetically transferred to the synchro transmitter, which sends a fuel flow synchro signal to the EMD and AIU 1. The fuel flow signal received by the EMD is converted to a digital signal to drive the LCD window FF pounds per hour (PPH). The EMD will blank that specific FF window if the indication goes beyond the range of the EMD. AIU 1 converts the signal to a digital signal and sends it to the ADCP through the avionics 1553 MUX bus for display in the MPD system.

Figure 2–9. Fuel flow signal flow.

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Nozzle position indicating system The convergent exhaust nozzle control (CENC) drives the engine exhaust nozzle position transmitter through a mechanical interface. The transmitter senses nozzle position and converts this measurement into a synchro type electrical signal. Related wiring transmits this signal through the engine EDU to the EMD. The signal is applied to the EMD where the signal is changed to a digital signal for display on the LCD; the range is 0–100 percent open. In the event the indication goes beyond the range of the EMD, it will blank that specific window. The EDU also converts the analog signal to digital and sends it to the AIU 1 for EMD repeater. Refer to figure 2–10 for an illustration of the nozzle position signal flow.

Figure 2–10. Nozzle position signal flow.

Oil pressure indicating system Breather pressure and main oil pressure are connected to the oil pressure transmitter. The transmitter senses the difference between the two pressures and converts this measurement into a synchro signal. The signal is transmitted by related wiring to the EDU and EMD. The EDU also converts the signal to digital and sends it to the AIU 1 and ADCP over the avionic 1553 MUX-bus. This signal is used by the MPD system for the EMD repeater.

The EMD receives the synchro signal and converts it to a digital signal for display on the LCD. Oil pressure is displayed from 0–100 psi in 5-psi increments. The EMD will blank that specific window if the indication goes beyond the range of the EMD.

The ADCP determines low oil pressure (less than 8 psi). During a low oil pressure condition the ADCP sends a command to AIU 1 through the avionics 1553 MUX bus instructing it to turn on the L/R OIL PRESS, ENGINE and MASTER CAUTION lights. See figure 2–11.

Figure 2–11. Oil pressure signal flow.

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212. Acceleration indicating/G-exceedance system Although we don’t normally think much about it, each of us live out our lives under a constant force of gravity (G). This is the force that attracts all bodies toward the center of the earth. The amount of force we live under is a gravitational force of 1 G. Have you ever been in an elevator and felt heavier as the elevator started to go up? That was not merely a feeling. Had you been standing on a scale you would have seen your weight increase on the scale as the elevator began to ascend.

When the G-forces are combined with speed and acceleration they become more dangerous. As a matter of fact, when we talk of aircraft maneuvers during flight, the G-force becomes critical. It’s critical because the fatigue the G-forces may cause to the aircraft structure can eventually cause an aircraft catastrophe. In addition to fatigue, G-forces place stress upon the aircraft. When this stress exceeds the design limits, the aircraft can literally disintegrate in the air. That’s why monitoring acceleration and G-force data is vital for maintaining the integrity of an aircraft’s structure due to stress. To be useful, the G-force on the aircraft must be displayed to the pilot during flight. In this lesson we’ll discuss how this is accomplished on the F-15.

Purpose The recording of aircraft stresses in flight helps engineers and maintainers predict and repair stress induced failures before they occur. The acceleration indicating systems of the F-15 provides the means to record these stresses. During this lesson we will be covering the components and operation of the F-15 acceleration indicating system. This system is further broken into the accelerometer counter set and the acceleration indicator (G meter).

System components The following paragraphs describe the acceleration indicating/G-exceedance system components.

Accelerometer counter set Made up of the counter accelerometer unit (CAU) and counter display unit (CDU), the accelerometer set measures the aircraft vertical-axis acceleration forces and counts when the force exceeds –2G, –1G, 0G, +3G, +4.5G, +6G, or +7.5G, with +1G being the reference. Acceleration force in the negative direction with respect to +1G is defined by aircraft acceleration in the downward direction. Acceleration in the positive direction is caused by aircraft acceleration in the upward direction.

Counter accelerometer unit The CAU or accelerometer as it is often called is located in the right main landing gear wheel well (fig. 2–12). The accelerometer contains a vertical acceleration sensor in a hermetically sealed unit. This means that if the unit fails, you can’t repair it on the line; thus, it must be turned in for repairs. The accelerometer is mounted near the aircraft center of gravity in the right main wheel well.

The accelerometer sensor works under the concept of piezoresistivity. The sensor consists of two seismic masses supported by metallic flexures connected by two silicon beams. Two piezoresistors, a resistor whose resistance changes with a strain applied, are diffused into each silicon beam. The sensor is oriented so that a vertical acceleration force causes the seismic masses to flex the silicon beams and the resulting stress changes the resistance of the piezoresistors. This generates a signal proportional to the acceleration force. The signal is .75 volts for each G-force.

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Figure 2–12. Counter accelerometer unit.

Counter display unit The counter is shown in figure 2–13. Rack-mounted under door 6R, the counter has two BIT fault indicators, one for the CDU and one for the CAU. There are also seven G display counters (–2G, –1G, 0G, +3G, +4.5G, +6G, and +7.5G).

A front panel connector is used for input power and connection with the accelerometer, signal data recording system (SDRS), BIT control panel (BCP), and ASP. There are two elapsed-time indicators (ETI), one will record the counter’s operating time and the other records flight time.

The accelerometer provides an output of 0.75 volt per G over the range –4G to +12G to the counter. The circuit then advances the appropriate counter on the front panel. The counter also routes BIT status to the BIT system and G-exceedance counts to the SDRS. We’ll cover those signals later, for now we need to look at the acceleration indicator.

Figure 2–13. Counter display unit.

Acceleration indicator The acceleration indicator shown in figure 2–14 is commonly called the G-meter. It’s a combined accelerometer and indicator that’s self-contained within the unit. It’s designed to provide a visual

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indication of −5 to +10 G-forces imposed on the cockpit of the aircraft during climbs, dives, and turns. The main pointer continually registers the aircraft G-units and two auxiliary pointers (one for positive G-units; one for negative G-units) indicate the maximum G-units attained during any flight or maneuver. These pointers remain fixed at maximum until manually reset. The acceleration indicator is on the cockpit main instrument panel left side and has the push-to-set knob on the lower left of the indicator face. The F-15D model aircraft have an identical indicator on the rear main instrument panel left side.

Figure 2–14. Accelerometer indicator.

System integration The accelerometer provides G-force data to the programmable signal data processor (PSDP) while the counter provides BIT status to the BIT system and G-exceedance counts to the SDRS.

The accelerometer routes .75 volts for each G-force exceeded to the PSDP. The PSDP converts this data into digital format and routes it to the central computer (CC). The CC will compute the amount of G-forces. The CC routes G-data directly to the head-up display (HUD) system for display. OVERLOAD data is sent through the PSDP to the integrated communications control panel (ICCP) for generation of warning tones and voice warning messages. Also, the PSDP converts and routes the OVERLOAD data into a usable format for display of the overload matrix on the multiple indicator control panel (MICP). Finally, the CC routes OVERLOAD data to the SDRS for recording purposes.

There are two operational checks used to test the acceleration indicating system. These operational checks are the accelerometer counter set checkout and the acceleration indicator checkout.

Accelerometer counter set checkout The accelerometer counter set is operationally checked by using the BIT system. When initiated, the BIT automatically checks both the counter and accelerometer; but, it doesn’t cause the counters to advance.

Acceleration indicator checkout The operational checkout of the acceleration indicator is accomplished by tapping the indicator and watching its movement. The indicator should indicate 1G when no external force is applied. As you tap the indicator in an up direction, one pointer indicates maximum deflection of the main pointer, and as you tap the indicator down, the other pointer indicates the maximum deflection in the other direction.

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Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

210. F-15 fuel quantity indicating system 1. What type input voltage is supplied to the tank units and what type signal does the tank unit

output?

2. What does the fuel quantity indicator continuously display?

3. If the FUEL QUANTITY SELECTOR switch is in the TANK 1 position, what will the left and right counter display?

4. With the FUEL INDICATOR switch in the FEED position, what does the left and right pounds counters display?

5. What can the pilot set to a predetermined position to automatically stop fuel dumping?

6. What is the purpose of the BINGO function?

7. What LRU requires calibration after removal and replacement?

8. Where are the fuel quantity system adjustment screws located on the F-15C, D, and E aircraft?

9. Why do you check the fuel density before performing a full adjustment?

10. How is a BIT initiated on the fuel quantity indicating system?

11. While performing a fuel quantity operation check, you notice the total pounds counter reads 35,000 pounds above what the reading should be. What is the probable cause?

12. After completion of a mission, the pilot writes up that the fuel quantity indications were fluctuating during maneuvers. What is the probable cause of this malfunction?

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211. F-15E engine instrument system 1. Which engine indications are displayed on the EMD?

2. What happens to the EMD if one of the signals for the engine parameters is out of range?

3. Which ASP will latch if the EMD fails the BIT?

4. The EMD displays engine RPM in what range?

5. In what range does the EMD display FTIT?

6. Which LRU determines if an FTIT overtemp condition exists?

7. How is the fuel flow signal routed to the EMD?

8. The EMD nozzle position is displayed in what range?

9. Which LRU drives the engine nozzle position transmitter?

10. Which LRU ultimately determines a low oil pressure condition?

11. At what pressure will the L/R OIL PRESS light illuminate?

212. Acceleration indicating/G-exceedance system 1. What is the constant force of gravity that we all live under?

2. The accelerometer counter set is composed of what two LRUs?

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3. What does the accelerometer counter set measure?

4. What flight line repairs can you perform on the accelerometer?

5. Where’s the counter accelerometer unit located?

6. The accelerometer generates how much voltage for each G-force encountered?

7. Where is the counter display unit located?

8. What are the exceedances, in Gs, that the CDU displays?

9. If an aircraft performs a 7G maneuver, which counter(s) will advance? Explain?

10. Explain the purpose of each pointer on the acceleration indicator?

2–2. Primary Flight Control Systems The F-15 has a flight control system designed to give exceptional maneuverability and control at both subsonic and supersonic speeds. In this section we will cover the three subsystems of the primary flight control system; lateral, longitudinal, and directional flight controls.

213. Lateral flight controls The lateral control subsystem is a hydro-mechanical system and provides a method of controlling the aircraft around its longitudinal axis. It uses the differential movements of the ailerons and stabilators for roll maneuvering. Refer to figure 2–15 for LRU locations on the aircraft.

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Figure 2–15. Lateral flight control subsystem.

System components The following paragraphs describe the components of the lateral control subsystem.

Control stick The control stick is located in the cockpit (also in the rear cockpit on the F-15D and F-15E models). It transmits pilot inputs to the control stick boost and pitch compensator (CSBPC). The CSBPC is made up of the pitch and roll channel assembly (PRCA), the aileron rudder interconnect (ARI) and the lateral control stick damper in lateral.

ROLL RATIO switch The ROLL RATIO switch is located on the miscellaneous control panel. This two-position switch is used to hydraulically shut down the lateral system of the PRCA. Its positions are AUTO and EMERG. When the switch is placed to EMERGENCY, the hydraulic boost to the lateral flight control subsystem is shut down. This will cause the roll ratio lamp to illuminate, alerting the pilot that the system is shut down and to use caution when inputting lateral commands.

Control system damper The control system damper is used to reduce oscillations in the control stick. It works by using permanent bar magnets and a drag cup that rotates through the bar magnet when the control stick is moved. As the stick attempts to oscillate during flight, friction is applied to dampen the control stick.

Lateral feel trim actuator The lateral feel trim actuator is located under door 6L alongside the longitudinal feel trim actuator. The feel trim actuator has built-in centering springs to provide artificial feel.

Pitch and roll channel assembly The PRCA is located under door 10L. It is a hydro-mechanical device that basically functions as a hydraulic computer. It receives the control stick inputs from the pilot, adds hydraulic boost and varies the ratio between lateral input and output, depending on airspeed, longitudinal (pitch) input, landing gear position, and rate of change in the yaw axis.

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Roll ratio controller The roll ratio controller is a replaceable subcomponent of the PRCA. The roll ratio controller increases and decreases lateral control based on variations in airspeed, longitudinal position, landing gear position, and yaw rate.

Mode select assembly The mode select assembly is another subcomponent of the PRCA. The mode select assembly is used to control the flow of hydraulic fluid to the PRCA. When energized, hydraulic pressure is blocked and a pressure switch will open, causing the ROLL RATIO light to illuminate.

Aileron rudder interconnect The ARI is located under door 10R. It works as a solid link in the directional subsystem mechanical linkage and couples directional movement in series with lateral and longitudinal movement as one common output for turn coordination. The mechanical signal is transmitted from the ARI output rod through the breakout assembly and push-pull cables to both right and left rudder rotary hydraulic servocylinders. The ARI is hydraulically shut down by the PRCA at Mach 1 and above.

Mechanical mixer The mixer assembly is located under panel 60 on the top center fuselage. The mechanical mixer separates the single PRCA output into two mechanical outputs. One goes to the ailerons, through safety spring cartridges, and the other to the stabilators, which aid in lateral control by operating differently.

Aileron safety spring cartridges The safety spring cartridges are located under panels 62L and 158L, for the left, and 62R and 158R for the right, on the top center fuselage. They are used aft of the mechanical mixer to allow for one wing operation should the linkage to the other aileron become jammed or inoperative.

Aileron hydraulic servocylinder The aileron hydraulic servocylinders are located under panels 143L, for the left, and 143R for the right. They are hydro-mechanical actuators that control the up and down movement of the ailerons.

System operation Pilot input to the lateral control subsystem is mechanically linked to the PRCA, ARI, lateral feel trim actuator, and a control system damper. The lateral input to the PRCA is modified by the roll ratio controller taking into effect the variations in airspeed, longitudinal subsystem position, and landing gear handle position. With increasing airspeed or longitudinal control stick displacement from neutral, the ratio changer decreases the amount of roll output to the ailerons and stabilators; thereby decreasing the amount of their deflection. This is known as aileron washout. When the gear handle is placed in the down position or when the roll/yaw computer senses a high rate of yaw with the handle up, a gear down solenoid is energized in the PRCA. This gear down solenoid moves the ratio changer to the maximum output position, giving the ailerons and stabilators full roll deflection for better aircraft control.

The ROLL RATIO switch on the miscellaneous control panel in the cockpit controls an electrical shutdown solenoid in the PRCA that shuts down hydraulic pressure to the roll ratio portion of the PRCA. The PRCA then works as a fixed ratio bellcrank allowing unboosted pilot commands to the servocylinders. Whenever the roll ratio portion of the PRCA is shutdown by the ROLL RATIO switch, the ROLL RATIO light on the caution lights panel is turned on by an internal pressure switch. The ROLL RATIO light is also turned on by a switch inside the PRCA whenever roll ratio scheduling is not at the proper setting as determined by the airspeed and the landing gear position, warning the pilot of a problem in the system.

The ROLL RATIO light plays an important part of the lateral control subsystem by warning the pilot when the output ratio is not correct. If the PRCA is shutdown (ROLL RATIO switch is put to EMER)

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or if the rudder limiter system is activated (at approximately 1.5 Mach) and the roll ratio airspeed scheduling valve has not shifted to minimum, the ROLL RATIO light will come on. If you look back at figure 2–15, you can see that as the roll input is applied to the PRCA it is also applied to the ARI by a control rod that runs across to the ARI input linkage. The ARI receives roll input and varies the ratio between the input and output depending on forward or aft stick inputs. This assists the pilot in turning the aircraft, also known as turn coordination. Longitudinal input comes from the PRCA through an interconnect cable. If the pilot elects to apply direct rudder input by using the rudder pedals, then that rudder pedal input is combined in series with the lateral output to form a single output to the rudder system.

The lateral output from the PRCA is divided into four outputs by the mechanical mixer, which operates the ailerons with differential stabilator deflection to increase lateral control of the aircraft. Mechanical linkage to the ailerons has a safety spring cartridge to allow the controls in one wing to operate if controls in the other wing are jammed.

Input linkage to the aileron and stabilator hydraulic servocylinder moves a spool, which controls pressure to the servocylinder. When hydraulic pressure to the servocylinders is lost, the servocylinder acts as a flutter damper by allowing fluid to flow from one side of the piston to the other side.

214. Longitudinal flight controls Now that you know the basic operation of the lateral flight control subsystem, we will discuss the longitudinal flight control subsystem. In this lesson we will discuss the components of the longitudinal flight control subsystem, as well as its operation.

System components The longitudinal control subsystem is a hydro-mechanical system and provides a method of controlling the aircraft around the lateral axis. It uses the symmetrical up and down movement of the horizontal stabilators for the pitch maneuvering of the aircraft. Figure 2–16 shows the longitudinal flight control subsystem. Refer to this figure as we discuss the components of the longitudinal control subsystem.

Figure 2–16. Longitudinal flight control subsystem.

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Control stick The control stick is located in the cockpit (also in the rear cockpit on the F-15D and F-15E models). It transmits pilot inputs to the CSBPC. The CSBPC is made up of the PRCA and the ARI.

PITCH RATIO switch The PITCH RATIO switch is located on the lower left main instrument panel directly below the pitch ratio indicator. This two-position switch is used to hydraulically shut down the longitudinal system of the PRCA. Its positions are AUTO and EMERG.

Pitch ratio indicator The pitch ratio indicator is located on the lower left main instrument panel directly above the PITCH RATIO switch. The pitch ratio indicator gives the pilot a visual indication of the output ratio in the longitudinal system. It has a range from zero to one.

Longitudinal mass balance A longitudinal mass balance is mounted on the linkage between the control stick and the PRCA to help neutralize the effect of acceleration on the control stick during maneuvers.

Longitudinal feel trim actuator The longitudinal feel trim actuator (LFTA) is located under door 6L in the top compartment. The feel trim actuator has built-in centering springs which resist stick movement giving the pilot a stick feel force.

Pitch and roll channel assembly Located in door 10L, the PRCA is a hydro-mechanical device used to vary the ratio between the control stick input and output to the control surfaces, depending on airspeed and aircraft response. The PRCA also boosts the input from the control stick to reduce the force required by the pilot to move the control surfaces. The PRCA contains an electrical shutdown solenoid and an internal switch used to turn on the PITCH RATIO light if the unit is shut down or if the ratio scheduling is not at the correct position according to the input conditions.

Pitch ratio controller The pitch ratio controller (PRC) is also a subcomponent on the PRCA. It contains an airspeed scheduling valve (for determining the required pitch ratio), a Mach factor computer, and an ARI Mach valve (for controlled shutoff of the ARI at speeds of Mach 1 and greater). The PRC receives left pitot and left S1 to read airspeed.

Mode select assembly The mode select assembly is another subcomponent of the PRCA. The mode select assembly is used to control the flow of hydraulic fluid to the PRCA. When energized, hydraulic pressure is blocked and a pressure switch will open, causing the PITCH RATIO light to illuminate.

Pitch trim controller The pitch trim controller (PTC) is also a subcomponent of the PRCA. The PTC automatically compensates for trim changes caused by accelerating from subsonic to supersonic flight, operating flaps or speed brake, or stores separation. The PTC travel is limited as the PRC nears minimum ratio. When the pitch control augmentation system (CAS) is engaged, the CAS interconnect (CASI) servo controls the PTC.

Mechanical mixer assembly The mixer assembly is located under panel 60, on the top center fuselage. The mixer assembly receives a single input from the PRCA and divides it into two outputs, one for each stabilator.

Stabilator hydraulic servocylinder The stabilator hydraulic servocylinders are located under panels 118L, 123L, 126L, 118R, 123R, and 126R in the tail cone area. The stabilator hydraulic servocylinders are hydro-mechanical actuators

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that control the synchronous and differential up and down movement of the stabilators. The servocylinders also contain the CAS servovalves used in the automatic flight control system (AFCS).

System operation Pilot input to the longitudinal control subsystem is mechanically linked to the PRCA, ARI, and the feel trim actuator. The ratio changer modifies the input to the PRCA for various airspeeds. The ratio changer decreases the amount of the output with increasing airspeed. The pitch ratio indicator in the cockpit shows the ratio between input and output of the PRCA.

The control stick supplies mechanical inputs to the PTC. A load force sensor inside the PTC measures the aircraft’s response. If the aircraft response does not equal the input from the control stick, the PTC sends hydraulic inputs to the pitch trim compensator. The pitch trim compensator will then send the correct pitch output to both stabilators through mechanical linkage. Failure of the pitch trim compensator shuts down pitch CAS and causes the pitch ratio changer actuator to lock in a predetermined position (fixed ratio bellcrank) allowing continued operation with limited pitch operation.

The pilot can shutdown the pitch portion of the PRCA by setting the PITCH RATIO switch to EMERG, which energizes the mode select assembly, a subcomponent of the PRCA. The mode select assembly shuts off hydraulic pressure in the pitch portion of the PRCA. This causes the PRCA to act as a fixed ratio bellcrank. With the PRCA shutdown, the pitch ratio will drive to .4 and the output from the PRCA is unboosted, causing increased stick friction.

When the PRCA is shut down, an integral pressure switch turns on the PITCH RATIO light. The PITCH RATIO light also comes on if the air data computer (ADC) senses that the altitude is less than 20,000 feet, airspeed above 330 knots and pitch ratio is more than .9, or if the pitch ratio is less than .9 with the landing gear down. The PRCA contains a PRC. The pitch ratio controller increases and decreases longitudinal control based on airspeed. The PRC also houses the airspeed sensor that controls the ARI by positioning the ARI shutoff valve in the roll ratio controller, blocking hydraulic pressure to the ARI at Mach 1 and higher.

Once the PRCA has boosted the input from the pilot, the signal goes to the mechanical mixer. From the mixer, the one longitudinal input from the PRCA is transmitted as two inputs to the stabilator hydraulic servocylinders through the bellcranks, control rods, and cables. The input linkage to the stabilator hydraulic servocylinder positions a spool, which controls pressure to the servocylinders, moving the stabilators up and down symmetrically. When hydraulic pressure to the servocylinders is lost, the servocylinder acts as a flutter damper by allowing fluid to flow from one side of the piston to the other side.

215. Directional flight controls Thus far we have discussed the operation of the lateral and longitudinal flight control subsystems. In this lesson we will be discussing the directional control subsystem; this will include the components that make up the system as well as the operation of the system.

System components The directional control subsystem is a hydro-mechanical system and provides a method of controlling the aircraft around its vertical axis, also known as yaw. Unlike the lateral and longitudinal control subsystems, which consist mainly of rods, the directional control subsystem consists mostly of push-pull cables. The rudder pedals provide a method for the pilot to control the directional control subsystem by applying force on the left or right pedal pivot points. Figure 2–17 shows a simplified diagram of the directional flight control subsystem. Refer to the figure as we discuss the components of the directional flight control subsystem.

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Figure 2–17. Directional flight control subsystem.

Rudder pedals The rudder pedals provide a means for the pilot to apply inputs to the directional (yaw) channel of the flight controls. There are two pedals (left and right). These are connected in such a manner that pushing forward on one pedal moves the other pedal aft. During ground operation (while taxiing) the rudder pedals are used to control the nose wheel steering system and steer the aircraft. The aircraft brakes also utilize the rudder pedals. Braking is initiated by applying pressure to the top of the rudder pedals to slow and stop the aircraft while on the ground.

Safety spring cartridge The safety spring cartridge is located in the cockpit on the rudder pedal assembly. It usually acts as a solid link, but spring tension in both directions allows rudder/brake movement to provide nose wheel steering and CAS operation if there is a jam in the directional control subsystem mechanical linkage.

Directional feel trim actuator Located in front of the rudder pedal assembly, the directional feel trim actuator (DFTA) contains centering springs, which provides an artificial feel used to keep the pedals centered.

Aileron rudder interconnect The ARI is located under door 10R. It works as a solid link in the directional subsystem mechanical linkage and couples directional movement in series with lateral and longitudinal movement as one common output. The mechanical signal is transmitted from the ARI output rod through push-pull cables to both right and left rudder rotary hydraulic servoactuators. It is a hydro-mechanical device that combines four inputs to provide turn coordination below Mach 1. It receives inputs from the following:

1. Pilot command – rudder pedals provide inputs for the directional control subsystem.

2. Pitch and roll inputs – the pitch input from the PRCA is used for turn coordination to determine direction and amount of rudder deflection. The roll input from the control stick is also used in turn coordination to determine the amount of rudder deflection.

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3. Hydraulic pressure – the PRCA controls hydraulic pressure to the ARI and above Mach 1 the PRCA removes hydraulic pressure from the ARI; thus disabling turn coordination. When pressure is removed from the ARI, it acts as a fixed linkage to transmit rudder pedal inputs to the rudder actuators.

NOTE: The ARI is also deenergized by the antiskid system during landings. This occurs when the ANTISKID switch on the miscellaneous control panel is placed to either the pulser or OFF position, wheel speed greater than 50 knots (such as landing), or when a malfunction has occurred in the antiskid system.

4. Flap signal – the flaps system provides a position signal to the ARI. This signal is required enabling the ARI to increase the amount of rudder deflection for lateral stick movement.

Rudder control breakout assembly The breakout assembly transmits the input force from the ARI into two separate push-pull cables, which provide inputs to the left and right rudder rotary hydraulic servoactuators. The rudder breakout assembly contains two shear rivets. If a malfunction occurs aft of the breakout assembly that requires approximately 80 pounds or more of force to move the left or right rudder push-pull cable or prevent movement of the cable, the rivet for that cable will shear. This allows the other cable to continue normal operation.

Rudder actuator The rudder actuators are located in the vertical stabilizers under panels 126L (left rudder) and 126R (right rudder). The rudder actuators serve the same purpose as the stabilator and aileron actuators, except that it uses a rotary motion instead of extending or retracting. The input from the mechanical linkage directs hydraulic pressure to the chambers of the rotary actuator.

Rudder limiter actuator Excessive rudder application at high airspeeds can cause structural damage or loss of control of the aircraft. Therefore, the ARI is hydraulically deactivated after the aircraft exceeds Mach 1. This eliminates turn coordination. The rudder limiter actuator is used to prevent commanded excessive rudder travel when the aircraft is traveling above Mach 1.5. It’s a linear actuator that is located in the cockpit on the rudder pedal assembly.

System operation If force is applied to the rudder pedals, a mechanical linkage transmits the input force through the DFTA (artificial feel). From the directional feel trim actuator, the signal is sent through the safety spring cartridge, the output torque tube assembly, and push-pull cables to the ARI. When the ARI is energized, it operates the rudders with lateral and longitudinal stick movement. When the ARI is deenergized, above Mach 1, the input force is transmitted as a solid link to the rudder control breakout assembly. The breakout assembly transmits the input force to two separate push-pull cables that provide inputs to the left and right rudder rotary hydraulic servoactuator input rods.

The rudder control breakout assembly contains two shear rivets. If a malfunction occurs aft of the breakout assembly that requires approximately 80 pounds or more of force to move the left or right rudder push-pull cable or prevents movement of the cables, the rivet for that cable will shear. This allows the other cable to continue normal operation.

A rudder travel limiter system in the directional control subsystem prevents excessive rudder deflection during high airspeeds. As airspeed exceeds Mach 1.5, a MACH switch in the right air inlet controller (AIC) closes and completes an electrical circuit, which energizes the rudder travel control relay. The travel relay then completes an electrical circuit to extend the rudder travel limiter actuator, which limits the movement of the rudder pedals. As airspeed decreases below Mach 1.5, the MACH switch in the right AIC opens, breaking the electrical circuit and de-energizing the rudder travel control relay. With the travel relay deenergized, another electrical circuit is completed, driving the rudder travel limiter actuator to the retracted position, returning the rudder pedals to full movement.

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A RUD LMTR (rudder limiter) caution light on the caution lights display panel warns the pilot when a malfunction exists in the rudder limiter system. If the airspeed is above Mach 1.5 and the rudder limiter actuator has not engaged the stop mechanism or if the airspeed is below Mach 1.5 and the rudder limiter has not disengaged the stop mechanism, the RUD LMTR light is turned on through a series of switches and relays. The left AIC supplies the Mach signal for the RUD LMTR light.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

213. Lateral flight controls 1. What hydro-mechanical system provides a method of controlling the aircraft around its

longitudinal axis?

2. What flight control surface does the lateral control subsystem use for roll maneuvering?

3. What components make up the CSBPC in the lateral control subsystem?

4. What is the purpose of the control stick damper, and how does it perform this function?

5. What is the purpose of the PRCA?

6. What modifies the lateral input to the PRCA for variations in airspeed, longitudinal linkage position, and landing gear position?

7. What component of the flight control system controls the rudders during lateral and longitudinal stick movement? When is this component shut down?

8. What is the purpose of the safety spring cartridges?

9. Under what conditions does the ROLL RATIO light illuminate for the lateral control subsystem?

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214. Longitudinal flight controls 1. What does the longitudinal control subsystem provide?

2. What gives the pilot a visual indication of the output ratio in the longitudinal system and has a range from zero to one?

3. What component is mounted on the linkage between the control stick and the PRCA to help neutralize the effect of acceleration on the control stick during maneuvers?

4. What is the purpose of the PRCA?

5. Explain the function of the pitch trim controller.

6. What system component divides the longitudinal input into two separate outputs?

7. Setting the PITCH RATIO switch to which position will shut down the PRCA?

8. When the PITCH RATIO switch shuts down the PRCA, it causes the PRCA to function as what?

9. When the PITCH RATIO switch is put to EMERGENCY, what does the pitch ratio indicator drive to?

10. When will the pitch ration light illuminate?

215. Directional flight controls 1. What does the directional control subsystem provide?

2. The directional control subsystem consists mostly of what?

3. What is the purpose of the safety spring cartridges?

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4. Where is the ARI located?

5. What four inputs does the ARI combine to provide turn coordination below Mach 1?

6. Explain why the ARI require the flap position signal.

7. What component prevents commanded excessive rudder travel when the aircraft exceeds 1.5 Mach?

8. Which AIC MACH switch drives the rudder travel limiter actuator and which AIC supplies the Mach signal for the RUD LMTR light?

2–3. Automatic Flight Control Systems Now that you understand the primary flight control system, let’s learn about the automatic flight control system on the F-15 A through D and E model aircraft. In this section we will begin with the F-15 C/D CAS. Then we will discuss the F-15 trim system. The section concludes with a lesson covering the F-15E automatic flight control system.

216. F-15A through D control augmentation system In this lesson we’ll focus on the components of the flight control system and AFCS. In our discussions, we’ll see what function(s) they provide for the CAS. After this, we’ll discuss the signals received by the CAS from other systems or components.

Flight control component functions as they apply to the control augmentation system The CAS components should be somewhat familiar to you. That’s because they’re used in many of the systems we’ve previously discussed. Keep in mind that in the CAS, the components have different functions than previously discussed. Although some of the components will be new to you, these are used primarily for CAS operation. The components we’ll be discussing are as follow:

• Roll yaw computer. • Pitch computer. • Rate gyro assembly. • Accelerometer assembly. • CAS control panel. • Dynamic pressure sensor. • Stick force sensor. • Directional feel trim actuator. • Pitch trim controller (CAS interconnect). • Stabilator servocylinders.

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• Rudder servocylinders. • Caution lights display panel.

Roll/yaw computer The roll/yaw computer (fig. 2–18) contains the engage logic for roll and yaw CAS. It receives roll and yaw CAS sensor inputs, processes those inputs, and generates output drive signals that eventually control surface deflection in the roll and yaw channels. The roll/yaw computer also contains the logic used to control the caution lights that inform the pilot how the CAS channels are operating.

Figure 2–18. F-15A thru D roll/yaw computer.

Pitch computer The pitch computer (fig. 2–19) is also located next to the roll/yaw computer. The pitch computer contains the engage logic and receives the sensor inputs for the pitch channel of CAS. It then processes the signals into a drive signal for the stabilator actuators. In addition, the pitch computer contains the differential stabilator servocylinder (DSS) amplifier, which is used to combine pitch and roll CAS inputs into a common output to drive the stabilators according to the pitch and roll inputs. The pitch computer also controls the CAS PITCH caution light that indicates the status of the pitch channel.

Figure 2–19. F-15A thru D pitch computer.

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Rate gyro assembly The rate gyro assembly (fig. 2–20) contains six rate gyros in a single housing. Their functions are as follows: two rate gyros provide pitch rate signals proportional to angular motion about the pitch axis. The pitch rate gyro outputs are applied to the pitch computer to monitor aircraft response to pilot commands in pitch. Roll and yaw are identical in function, but are applied to the roll/yaw computer.

Figure 2–20. F-15A thru D rate gyro assembly.

Accelerometer assembly The accelerometer assembly (fig. 2–21) contains four accelerometers. Each accelerometer sensor provides an electrical output signal that’s proportional to aircraft acceleration along the two sensitive axes of pitch and yaw. Two sensors are oriented to sense normal acceleration and two are oriented to sense lateral acceleration. The normal acceleration outputs are applied to the pitch computer and are used in pitch CAS to monitor aircraft response to pilot commands. Similarly, the lateral acceleration outputs are applied to the yaw CAS to monitor aircraft response to pilot commands.

Figure 2–21. F-15A thru D accelerometer assembly.

Dynamic pressure sensor The dynamic pressure sensor (fig. 2–22) contains two differential pressure transducers driven by bellows. The bellows are connected to pitot and static pressure lines. The transducers supply dual-electrical signals that are proportional to dynamic pressure (Qc). The dynamic pressure signals are

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applied to the roll/yaw computer and are used to limit the total differential stabilator deflection in roll CAS at high airspeeds.

Figure 2–22. F-15A thru D dynamic pressure sensor.

CAS control panel The CAS control panel (fig. 2–23) is located in the forward portion of the left console. It contains three dual-pole, lever-lock toggle switches, one pushbutton switch, and one indicating light. The three dual-pole, lever-lock switches are the CAS YAW, ROLL, and PITCH switches. Each switch has three positions: OFF, RESET, and ON. The CAS is operational in each axis when the switch for that axis is set to ON; that is, if the engaging parameters are satisfied. The RESET position initiates a re-engagement (reset) signal to the applicable computer following a malfunction. If the malfunction was momentary, the CAS for the failed axis can be reset and engaged again.

Figure 2–23. F-15A thru D CAS control panel.

Stick force sensor As you can see, the stick force sensor (fig. 2–24) is installed between the control stick grip and the control stick column. It contains strain gage elements that are designed to measure the forces applied to the control stick. The strain gage elements generate signals proportional to applied forces. The amplified signals are then applied to the pitch computer and roll/yaw computer for use as pilot input for pitch and roll CAS operation.

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Figure 2–24. Stick force sensor.

Directional feel trim actuator The directional feel trim actuator is located between the rudder pedals. You should remember that it contains the rudder pedal position linear variable differential transformers (LVDT) that supply the roll/yaw computer with pilot inputs for yaw CAS operation.

Pitch trim controller (CAS interconnect) The PTC is a component of the PRCA. The PTC is electrically connected to the pitch computer for CASI operation. CASI is designed to provide a tracking function. To do this, the PTC’s CASI servo receives pitch CAS commands and drives the PRCA’s pitch trim compensator. This forces the mechanical system to track pitch CAS so that if pitch CAS fails, the mechanical system can take over at the precise point of failure without any large uncommanded pitch-up or pitch-down maneuvers.

Stabilator servocylinders The stabilator servocylinders contain the electrohydraulic valve (EHV) that controls the stabilator displacement. There are two EHVs, one for each CAS channel. Pressure to the EHVs is controlled by shutoff valves. When the CAS is energized by turning the control switch to ON, the shutoff valves energize and allow pressure to the EHVs. There’s a differential pressure sensor (DPS) that monitors hydraulic pressure to and between the EHVs. If pressure isn’t available or a difference exists between the EHVs, the DPS shuts them down. There are two servo valve position LVDTs and two main ram position LVDTs that supply servo valve and ram position information back to the pitch computer.

Rudder servocylinders In the rudder servocylinders, yaw CAS commands are applied to reposition the rudders according to the command from the roll/yaw computer. Each actuator (located in the vertical fin) contains only one EHV and shutoff valve. The EHV controls pressure to the master control valve (MCV) which controls the rotary shaft of the actuator. External feedback is used to reposition the MCV and stop the actuator. The MCV position LVDT sends a signal back to the roll/yaw computer that’s proportional to the MCV position.

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Caution lights display panel The caution lights display panel (fig. 2–25) contains the CAS warning lights: CAS YAW, CAS ROLL, and CAS PITCH. The lights are controlled by logic circuits in the roll/yaw computer for the CAS YAW and CAS ROLL lights, and logic circuits from the pitch computer for the CAS PITCH light. CAS lights illuminate when the respective channel of CAS is manually turned off by using the switch, or when a malfunction has been detected in the respective CAS channel.

Figure 2–25. Caution lights display panel.

Signals applied to the control augmentation system Earlier we said that the CAS receives signals from many sources. The signals are used to provide proper CAS operation. Although the signal sources aren’t considered to be part of the CAS, they’re very important to the CAS operation. To be proficient in your job, you should be able to identify these signals and their origin. In addition, you should know their function in the CAS. To assist you in troubleshooting the CAS in the future, we’ll discuss these subjects: trim system; modular relay panels; pitot-static system; angle-of-attack (AOA) transmitters; and engine air intake system.

Trim system Each CAS axis receives an input from the respective trim actuator to show its position. This is important to the CAS because as the trim actuator drives, it establishes/updates a new zero reference for the CAS. As an example, let’s say the pilot trimmed 2° nose down. This would be the zero reference for the pitch CAS axis. The roll and yaw axis operate the same way.

Modular relay panels The #1 and #2 modular relay panels provide logic signals to the pitch computer for changing conditions such as landing gear and flap position, and weight off wheels to control CAS limits and functions. Remember, they’re applied only to the pitch computer.

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Pitot-static system The pitot-static system supplies pneumatic pressures (S2 static and right pitot) to the dynamic pressure transducers in the dynamic pressure sensor.

Angle-of-attack transmitters The left and right AOA information is supplied to the pitch computer for pitch limiting. The roll/yaw computer receives AOA from the pitch computer for use in calculating turn coordination (CAS ARI).

Engine air intake system At 1.5 Mach, the left and right AIC send a logic signal to the roll/yaw computer. This signal disables the turn coordination (CAS ARI) function of yaw CAS.

Control augmentation system operation Since the operation of the yaw, roll, and pitch CAS channels are similar, we’ll only discuss the operation of the yaw CAS channel. We’ll then highlight the differences with the other channels of operation.

Yaw control augmentation system operation The yaw CAS is a dual-channel, self-monitoring system, incorporating automatic channel shutdown in case of an axis failure. Pushing the left or right rudder pedal causes the mechanical controls to move the rudders. The CAS adds to or subtracts from the rudder positions commanded by the mechanical system. It does this until the sum of the lateral acceleration and yaw rate signals is equal to the sum of the rudder pedal position and ARI command signals. In other words, yaw CAS compares the pilot input (rudder pedal position) to aircraft response (lateral acceleration and yaw rate) and either adds to or subtracts from surface deflection to achieve the desired aircraft response. In the following pages we discuss these areas of yaw CAS operation: engagement, command augmentation, (ARI), and AC interlock and yaw rate limit.

Engagement Engaging a CAS axis isn’t as simple as placing the appropriate switch in the ON position. Instead, certain requirements and conditions must be met before the channel will engage. Figure 2–26 shows the CAS engage diagram. This diagram shows all three axes. Refer to the diagram as we discuss each axis separately.

Figure 2–26. F-15A thru D CAS engage logic diagram.

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We’ll start with the yaw CAS switch, shown in figure 2–23. When you move the switch from OFF to ON, yaw CAS ON logic signals are supplied to the roll/yaw computer’s engage logic starting the engagement process. Remember, the CAS axes have dual channels and are referred to as channels A and B.

The engage logic is contained inside the roll/yaw computer. If the rudder servo isn’t failed and the yaw rate is less than 41.5° per second, the logic will activate two transistor switches. One transistor is for channel A, providing 28 volts direct current (VDC) to the left rudder servocylinder shutoff valve (SOV). At the same time, the channel B transistor switch provides a ground to the right servocylinder SOV. The SOVs are wired in series, so the 28 VDC signal, and the single ground signal, energize both SOVs. When the shutoff valves energize, hydraulic pressure is supplied to the left and right rudder EHVs.

In short, the yaw CAS engagement requirements are as follows: 1. Yaw CAS switch placed to ON. 2. Rudder servo not failed. 3. Yaw rate less than 41.5° per second.

Control augmentation Once yaw CAS is engaged, it can perform its primary function of control augmentation. To do this, aircraft response is compared to the manual commanded input (rudder pedal displacement) to compute the yaw CAS signal. The yaw CAS axis will add rudder deflection if the aircraft is under responsive. Likewise, it will subtract rudder deflection if the aircraft is over responsive.

The primary directional flight control system can produce ±15° of rudder deflection. Yaw CAS can add an additional ±15° of rudder deflection, for a maximum rudder deflection of ±30°.

Figure 2–27 shows a simplified diagram of the yaw CAS axis. Pilot commands are applied to the roll/yaw computer through the rudder pedal position LVDTs in the directional feel trim actuator. At the same time; the mechanical system drives the rudders and causes the aircraft to yaw. The response of the aircraft is monitored by the accelerometer and rate gyro assemblies. Remember, there is both an A and B yaw rate gyro output and an A and B lateral accelerometer output. These outputs are applied to the A and B channels of the yaw CAS inside the roll/yaw computer. If the response of the aircraft is proper, the signals from the accelerometers and rate gyros cancel the pedal position input and the yaw CAS doesn’t have an output. However, if the aircraft is under responsive, the computer will command additional rudder deflection. Likewise, if the aircraft is over responsive, the computer will command less rudder deflection. Once the inputs are received by the yaw axis of the roll/yaw computer, they’re processed by the yaw computation circuits and applied to the rudder servocylinder EHV. The left rudder servocylinder is controlled by yaw channel A, and the right by a channel B. Rudder servocylinder position is monitored in each channel, respectively, by a position LVDT signal from the actuator to the roll/yaw computer.

Inside the computation circuits are balance amplifiers which will attempt to balance both channels for small signal differences. Both channels are monitored for large signal differences, or small long-standing signal differences. Also, the rudder servocylinder is monitored for large or long-standing differences. When these conditions exist, the yaw channel is automatically shut down and the CAS YAW light illuminates. If it was a temporary condition, the channel may be reset.

NOTE: The balance function is disabled during landing. To do this, a landing gear signal from the pitch computer disables the balance circuits.

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Figure 2–27. F-15A thru D yaw CAS logic diagram.

ARI The yaw axis of CAS must compute a turn coordination command much like the function performed by the mechanical ARI. The roll/yaw computer contains a pulse width multiplier that provides the outputs required for coordinated turns. The pulse width multiplier uses roll rate and AOA to compute the yaw CAS ARI signal. Just like the mechanical ARI, the CAS ARI has no input to the rudders if the aircraft is flying straight and level in pitch. Figure 2–27 shows that the roll rate signal is received from inside the roll/yaw computer. Also, notice that the AOA signal is received from the pitch computer. The CAS ARI works very much like the mechanical ARI, except instead of having a longitudinal (pitch) flight control input, CAS ARI uses AOA. Also, instead of having a lateral (roll) flight control input, CAS ARI uses roll rate. Both roll rate and AOA can affect the amount of rudder deflection that is achieved through CAS ARI. Notice in figure 2–27 that there’s a 1.5 Mach signal applied to the roll/yaw computer from the left and right AICs. The left is applied to channel A and the right to channel B. This signal is used to disable the CAS ARI function of yaw CAS at 1.5 Mach.

Power applied (26 volts alternating current (VAC)) to the rate gyro is compared to the power applied to the roll/yaw computer. If the power of the rate gyro is interrupted, the yaw channel shuts down. On F-15C/D aircraft, the 26 VAC is routed through the stick force sensor (also the rear cockpit in the F-15D) to prevent yaw and pitch CAS operation if the stick force sensor(s) is disconnected. Since yaw CAS engagement is a requirement for roll CAS engagement (discussed in more detail later) a disconnected stick force sensor will prevent all three axes of CAS from engaging. The yaw rate level detectors are used to sense yaw rate. If the yaw rate is exceeding 41.5° per second, the yaw CAS axis is automatically shut down.

Roll control augmentation system operation The roll CAS is designed to augment the roll primary flight control channel to maintain a consistent aircraft response. Much of the roll CAS channel operation is similar to the yaw CAS channel. These differences are discussed below.

Engagement The roll CAS is engaged in a manner that’s almost the same as the yaw CAS. In fact, the roll CAS switch functions the same as the yaw CAS switch. Moving the switch from OFF, through RESET, to

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the ON position causes the same RESET and ON logic signals to be applied to the roll CAS engage logic. Looking at figure 2–26, we see that the logic signals from the switch are applied to the roll CAS logic circuits inside the roll/yaw computer. Three additional requirements must be met before the roll channel will engage:

1. Roll CAS not failed. 2. Yaw CAS engaged. 3. The stabilators must be engaged (DSS not failed).

If all these conditions are met, the roll engage logic will energize two transistor switches. One switch controls the power to the roll channel computer circuits, while the other removes the ground from the roll channel output. The transistor switch that applies power also applies the excitation voltage to the dynamic pressure sensor. As you can see from figure 2–26, the roll CAS engage logic sends a signal to the pitch CAS engage logic. This is necessary because the pitch computer sends the surface commands to the stabilator servocylinders. So when roll CAS is engaged, it must send the engage signal to the pitch engage logic to energize the SOVs of the stabilator servocylinders. This happens even though pitch CAS isn’t engaged.

NOTE: A very important fact to remember is that the yaw CAS must be engaged prior to roll CAS being engaged.

Once all the requirements are met and the roll engage logic is proper, the channel functions and the CAS ROLL caution light is extinguished.

Command augmentation Earlier, we said the purpose of roll CAS is to augment roll control of the aircraft just as yaw CAS was to augment yaw control. Figure 2–28 shows the simplified roll CAS diagram. Again, there are two channels (A and B) for each sensor and computation.

Figure 2–28. F-15A thru D roll CAS logic diagram.

During flight the lateral stick force (pilot input) is measured by the stick force sensor while the aircraft response is measured by the roll rate gyros. The signals from the rate gyro will null each other, or add to, or subtract from the commanded input. Since the mechanical roll ratio is varied as a function of airspeed, so must the roll CAS signals be varied according to airspeed. To do this, the dynamic pressure sensor and the AOA signal from the pitch computer are used to schedule the roll CAS signal. When the aircraft is flying at less than 544 knots, roll stabilator control is ±5° of authority for each stabilator. At speeds above 544 knots, the authority is gradually decreased until it’s reduced to 1.1° at 800 knots. The AOA controls the roll CAS authority so that at 23° and –1.0° the

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roll CAS authority is 0°. Also at 1.5 Mach, a signal from the yaw CAS circuit is applied to the roll CAS circuit to reduce (attenuate) the input (pilot command) signals by about 50 percent.

Remember, the roll trim signal applied to the roll computation circuits is used to establish the zero reference point for the roll CAS.

The roll CAS A and B channels are monitored for large signal differences and long-term small signal differences. Either of these will cause the roll CAS axis to shut down and illuminate the CAS ROLL light. Note (fig. 2–28), that the output of the roll computation channels is applied to an amplifier which has its output applied to the DSS in the pitch computer. At this point the roll and pitch CAS commands are summed into a single drive command for the stabilator servocylinders. Each stabilator receives a channel A and a channel B command. Don’t forget that each stabilator servocylinder has two EHVs instead of one like the rudder servocylinders. The stabilator servos are monitored in the DSS circuits, and a failure of either causes both the roll and pitch CAS to shut down. Because the DSS monitor circuits are in the pitch computer, a pitch CAS failure causes the roll CAS to also shut down. If the failure doesn’t include the stabilator circuits, the roll CAS may be reset. Again, it’s very important to remember that the roll CAS has no input to the ailerons, only the stabilators.

Pitch control augmentation system and stall inhibit operation Pitch CAS performs a variety of functions. It’s designed to augment the primary pitch flight control channel, provide limited oscillation damping, and stall inhibit for the pitch axis. The pitch CAS operates much like the yaw and roll CAS axis. Here we’ll discuss the following subjects: engagement, command augmentation, and stall inhibit function.

Engagement You probably guessed that pitch CAS is engaged very similar to yaw and roll CAS. You’re correct; however, there are some different requirements that we’ll explain as we progress through the lesson. Refer again to figure 2–26 as we see how pitch CAS is engaged.

The RESET and ON logic from the pitch CAS switch is applied to the pitch CAS engage logic; however, this time the stabilator servocylinder servo shutoff valves and the CASI shutoff valve must be energized. In addition to the pitch CAS switch logic, the pitch CAS must not be failed for the CASI shutoff valve to energize. The stabilator servocylinder shutoff valves are energized when the following conditions are met:

• The pitch CAS switch logic is present (pitch CAS ON). • DSS not failed. • Yaw rate is less than 41.5° per second. • CASI servo not failed (PTC).

Pitch CAS can be reset and brought back on-line in the event the CASI servo fail by placing the PITCH RATIO switch to the EMERG position.

Remember, since the stabilator DSS is located in the pitch computer, a pitch CAS failure also shuts down roll CAS; however, the roll CAS resets if the failure only affects the pitch CAS axis.

Command augmentation The pitch axis of the CAS serves the same purpose as the yaw and roll CAS axis. During operation, the pitch CAS can add or subtract 10° of stabilator symmetrical deflection. Figure 2–29 shows a simplified pitch CAS diagram. As you can see, the pitch axis is a dual-channel system with monitoring and automatic axis shutdown just like the yaw and roll axis.

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Figure 2–29. F-15A thru D pitch logic diagram.

The stick force sensor senses forward and aft stick forces and sends it to the pitch computer. Aircraft response is measured by the accelerometer assembly (normal acceleration) and the rate gyro assembly. But, when the gear is down, normal acceleration is removed from the pitch computer to avoid producing adverse signals during possible hard landing. The switching is controlled inside the pitch computation circuits by a gear signal from the modular relay panel.

The output drive signal from the pitch computer is applied to the stabilator servocylinder and the PTC. The signal applied to the stabilator servocylinders causes the stabilators to either add to or subtract from the manually commanded input, depending on the aircraft response. The signal applied to the PTC’s CASI servo repositions the manual linkage to track the pitch CAS commanded input. Remember, this is necessary so that if pitch CAS fails, the mechanical flight control system will take command where the pitch CAS failed.

The pitch trim signal received from the roll/yaw computer establishes the zero reference point just as it did in the yaw and roll CAS axes. The pitch and roll CAS commands are summed at the DSS circuitry inside the pitch computer. The feedback from the stabilator servocylinder LVDTs is a combination of pitch and roll commands, and the feedback from the LVDTs must be separated back into separate pitch and roll CAS feedback signals. This is accomplished in the DSS circuits as well. The reason for separating the pitch and roll feedback signals is to cancel only the proper amount of input signal to provide proper control surface movement for each axis.

NOTE: The pitch axis applies redundant channel A and B inputs to the DSS just as roll CAS does. This method of applying a signal from the A and B channel provides for system failure monitoring. The computer performs this dual-channel failure detection, where if the two channels don’t match, the computer shuts the axis down.

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Stall inhibit function The stall inhibit function warns the pilot that a pending stall condition exists and increases the effort necessary by the pilot to increase the aircraft’s nose-up attitude. AOA, gear position, and flap position control the stall inhibit function circuits inside the pitch CAS computation circuits. At an AOA of 13° with the gear up, negative pitch commands are feed the pitch CAS; so, any attempt to increase the nose-up attitude requires extra force by the pilot. Placing the flaps down allows 10° of additional nose-up attitude before additional stick is required. The pitch rate from the pitch rate gyros allows the circuit to prepare for a stall condition. Once the aircraft has landed and the weight is on the wheels, the stall inhibit function is deactivated.

Control augmentation system troubleshooting The flight control system test set (FCSTS) is used to verify/fault isolate the automatic flight control system components on the F-15 A through D aircraft. This is done using interconnect cables from the FCSTS to the pitch and roll/yaw computers, rate sensor assembly (RSA), acceleration sensor assembly (ASA), and both left and right rudder and stabilator servocylinders. The FCSTS used in conjunction with the job guide and fault isolation TO will be critical to your ability to troubleshoot automatic flight control system malfunctions.

217. F-15 trim systems We begin this lesson by explaining the function and operation of the components and switches of the trim system. Although some of the components were previously mentioned when discussing flight controls; they also serve functions as part of the trim system. The components we’ll be covering are:

• Longitudinal (pitch) trim actuator. • Lateral (roll) trim actuator. • Directional (yaw) trim actuator. • CONTROL STICK TRIM switch. • RUDDER (YAW) TRIM switch. • TAKEOFF-TRIM (TOT) switch and light. • Roll yaw computer. • Pitch computer.

Longitudinal (pitch) feel/trim actuator A closeup view of the actuator is shown in figure 2–30. The actuator contains a single-phase, 115 VAC, reversible motor that’s connected to a screw jack. When the motor is activated in either direction, it repositions the mechanical pitch linkage and the control stick. In order for the motor to drive, it must receive 115 VAC and a ground through two separate relays. The pitch trim actuator is controlled by relays in the pitch computer.

Figure 2–30. The longitudinal (pitch) feel/trim actuator.

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Lateral (roll) feel/ trim actuator Figure 2–31 shows a closeup of the lateral feel trim actuator. In design, the roll trim actuator is identical to the pitch trim actuator. The major difference is that the roll trim actuator is controlled by relays in the roll/yaw computer whereas the pitch trim actuator is controlled through relays in the pitch computer. The roll trim actuator repositions the roll channel linkage and control stick.

Figure 2–31. The lateral (roll) feel/trim actuator.

Directional (yaw) feel/trim actuator Figure 2–32 shows a closeup of the directional feel trim actuator. This actuator is a little different from the pitch and roll trim actuators. It contains a single-phase, 115 VAC, reversible motor, which is connected to a screw jack like the pitch and roll feel trim actuators. But in addition, it contains the relays that control the voltage and ground for the drive motor. Like the pitch and roll trim actuators, the yaw trim actuator has two trim position LVDTs. In addition, it has two additional rudder pedal position LVDTs.

Figure 2–32. The directional (yaw) feel/trim actuator.

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NOTE: The F-15E is triple redundant; that is it uses three pedal position and trim position LVDTs.

The outputs of these LVDTs are used in the AFCS. The trim position LVDT’s supply trim actuator position information to the roll/yaw computer and are repositioned any time the motor is driven. The relays for all three feel trim actuators are controlled by the roll/yaw computer. The pedal position LVDTs are repositioned when the rudder pedals are moved, and supply pedal position information to the roll/yaw computer. Since the trim actuator repositions the yaw channel linkage and rudder pedals when it’s actuated, the rudder pedal position LVDTs are also repositioned.

CONTROL STICK TRIM switch This switch provides pitch and roll trim commands. It’s a five-position switch, which is spring-loaded to the center (off) position. A closeup view is shown in figure 2–33. Forward switch movement commands nose down pitch control, aft commands nose up pitch control; left switch movement commands left wing down roll control, while right switch movement commands right wing down roll control.

Figure 2–33. CONTROL STICK TRIM switch.

When the CONTROL STICK TRIM switch is placed in either the NOSE-UP, NOSE-DOWN, LEFT WING DOWN or RIGHT WING DOWN positions, two sets of contacts are closed. These two sets of contacts apply two grounds (or logic lows) to the roll/yaw computer. Later in the unit, you’ll see that both logic lows are necessary in order to provide the proper logic data flow to energize the trim motor.

NOTE: The F-15E combined the roll/yaw and pitch computers into a single computer called the flight control computer (FCC).

RUDDER (YAW) TRIM switch The RUDDER TRIM switch is located on the throttle quadrant. It’s a three-position switch that’s spring-loaded to the center (OFF) position. Moving the switch to the left causes a NOSE-LEFT COMMAND, and moving the switch to the right causes a NOSE-RIGHT COMMAND. The YAW TRIM switch has two sets of contacts. When it’s placed in either position, two logic lows are applied to the roll/yaw computer.

TOT switch (button) and light Both the TOT switch and light are located on the CAS control panel. The TOT switch is a pushbutton-type switch. Depressing the button applies two 28 VDC signals (or logic highs) to the roll/yaw computer. The two logic highs start the TOT sequence of events. When all the requirements have been completed, the roll/yaw computer energizes a transistor switch which illuminates the TOT

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light. The TOT light is an indication to the pilot that the aircraft control surfaces are at the TOT position.

Roll/yaw computer The roll/yaw computer contains the control circuits for the trim system. It receives the logic inputs from all the trim switches. In addition, it contains all the logic circuits, level detectors, transistor switches, and time delays that are used to control the trim systems. The roll/yaw computer also contains the relays used to control the roll trim system.

Pitch computer The last component having an effect on the trim system is the pitch computer whose relays are controlled by the roll/yaw computer. The pitch computer contains the relays used to control the voltage and ground for the longitudinal feel trim actuator.

Pitch trim operation As you study this material, keep in mind that pitch is referenced to the nose of the aircraft. When we talk about pitch we’re using the term in reference to aircraft nose up and nose down. From this, we can say that the pitch trim function is used to compensate for nose/tail heavy conditions. Figure 2–34 is a simplified block diagram of the pitch trim function. Refer to the figure as we look at the following subjects: CONTROL STICK TRIM switch, trim command processing, trim motor operations, and LVDTs.

Figure 2–34. Pitch trim functional block diagram.

CONTROL STICK TRIM switch We’ll start with the PITCH TRIM switch in the center (OFF) position. In this position there’s no input to the roll/yaw computer. Holding the PITCH TRIM switch aft (NOSE-UP) applies trim commands (two separate grounds) or logic lows to the roll/yaw computer. One logic low is called the nose-up command, and the other logic low is called nose-up/nose-down command. Holding the PITCH TRIM switch in the forward (NOSE-DOWN) position sends a nose-down trim command logic low, and the same nose-up/nose-down logic low to the roll/yaw computer. The nose-up/nose-down circuit operates the same as before.

Trim command processing The roll/yaw computer contains all the pitch trim logic. Except for the addition of logic circuits that sense that the pitch trim actuator is not at its limit, the nose-up/nose-down command is processed in the same way as the nose-up command. If the longitudinal feel trim actuator isn’t at its nose-up limit, the logic from the nose-up/nose-down command activates circuitry sending 28 VDC to the nose-

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up/nose-down relay in the pitch computer. When the nose-up/nose-down relay is energized, contacts are closed to apply the 115 VAC to the trim motor. When the 115 VAC is applied, the trim motor moves.

Remember, the pitch trim logic circuits are located in the roll/yaw computer and the pitch trim relays are in the pitch computer. The relays in the pitch computer provide power and ground to the longitudinal feel trim actuator trim motor for it to drive nose up or nose down.

Trim motor operation When 115 VAC and a path to ground are provided, the trim motor moves the control linkage so the aircraft’s nose moves up. As the trim motor is driving, its repositioning is monitored by two pitch trim position LVDTs (A and B). LVDT A sends a signal back to the roll/yaw computer as an input to the nose-up limit-level detector. Once the nose-up (trim) limit-level detector determines that the pitch trim motor is at the nose-up limit, it changes its logic output. This stops the pitch computer from continuing to drive the actuator. This action removes the 115 VAC from the pitch trim motor. As long as the nose-up level detector senses that the nose-up limit is reached, it inhibits the logic circuit for the nose-up/nose-down relay.

Linear variable differential transformers Remember, there are two LVDTs in the trim motor. The output of LVDT A determines the nose-up and nose-down limits. Both LVDTs A and B are used later for the TOT system. On the F-15E there are three LVDTs; an A, B, and C LVDT.

Take-off trim will be covered later in this lesson.

NOTE: Another very important fact to remember is that when the longitudinal or lateral feel trim actuator motor drives, it repositions the control stick. So any time the pitch or roll trim system operates, the control stick is repositioned to correspond to surface displacement.

Roll trim operation Roll is referenced to the downward motion of the left and right wingtips. A left roll is called left wing down, and right roll is called right wing down. At times, roll actions are planned and deliberately executed by the pilot. At other times, the aircraft can become unbalanced or wing heavy to one side or the other due to changing flight conditions. The purpose of the roll trim function is to compensate for these wing-heavy conditions. Figure 2–35 shows a simplified block diagram of the roll trim function. Refer to the figure as we discuss the following subjects: CONTROL STICK TRIM switch, energizing the roll trim motor, roll trim relays, and trim motor operations.

Figure 2–35. Roll trim functional block diagram.

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CONTROL STICK TRIM switch Holding the CONTROL STICK TRIM switch to the left (ROLL LEFT) generates two logic low trim commands (two separate grounds) to the roll/yaw computer. One command is the left roll trim command; the other is the left/right roll trim command. These two commands work just like the pitch trim commands.

Energizing the roll trim motor One of the roll left trim commands causes the logic inside the roll/yaw computer to energize the left roll relay, thus providing a ground to the lateral feel trim actuator trim motor. The second trim command enables power to be supplied to the trim motor.

Moving the trim switch to the right does the same as before except a right roll command is provided to the roll/yaw computer and the roll right relay provides the ground for the trim motor. The trim motor drives in the opposite direction.

Roll trim relays One major item to notice is that the roll trim relays are located inside the roll/yaw computer; otherwise the pitch and roll trim operate the same.

Trim motor operations The trim motor drives downstream lateral linkage that repositions the control stick and the roll trim position LVDTs A and B. LVDT A provides trim motor position to the limit-level detectors in the roll/yaw computer. The limit-level detectors work identical to the limit-level detector in the pitch trim system.

Yaw trim operation Like pitch, yaw is also referenced to the nose of the F-15. When we talk about yaw, we’re doing so in terms of nose-left and nose-right. The yaw trim function is designed to compensate for left or right yaw conditions, and is controlled by the RUDDER TRIM switch located on the throttle quadrant. Figure 2–36 is a simplified diagram of the yaw trim system. Refer to the figure as we discuss the yaw trim motor operation and LVDTs.

Figure 2–36. Yaw trim functional block diagram.

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Yaw trim motor operations The yaw trim function is very similar to the pitch and roll functions. Moving the RUDDER TRIM switch to the NOSE-LEFT position will generate two logic low commands. The yaw trim commands are sent to the roll/yaw computer. Logic inside the computer energizes trim relays which provide power and ground to the directional feel trim actuator. Yaw trim is unique in that the trim relays are part of the directional feel trim actuator. Moving the trim switch to the right will generate nose-right commands, and the nose-right relay in the actuator drives and the motor turns in the opposite direction.

LVDTs Once the yaw trim actuator drives in either direction, it repositions the yaw channel linkage, the rudder pedals, and the rudder trim position LVDTs. Again LVDT A is used to supply trim actuator position to the limit level detectors in the roll/yaw computer. The limit-level detectors prevent the actuator from driving beyond its limit just as it did in the pitch and roll functions. Both trim position LVDTs A and B are used later for the TOT system. On the F-15E there is an A, B, and C LVDT.

Unlike the longitudinal and lateral feel trim actuators, the directional feel trim actuator contains two sets of LVDTs. You should be familiar with the trim position LVDTs, but the actuator also contains rudder pedal position LVDTs. The pedal position LVDTs provide pilot inputs to the yaw CAS system.

Takeoff-trim function The purpose of the TOT function is to automatically drive the control surfaces to a predetermined position for takeoff. At takeoff, the rudders and ailerons are positioned at neutral, and the longitudinal feel trim actuator positions the pitch channel linkage so the control stick is 1° aft which positions the trailing edge of both stabilators 5° trailing-edge up. Figure 2–37 shows a simplified diagram of the TOT function. Refer to this figure as we cover these areas of the TOT function: TOT switch, trim-level detectors, moving to TOT, stopping TOT movement, and TOT light.

Figure 2–37. Take-off trim functional block diagram.

TOT switch The TOT switch and light are located on the CAS control panel (left console). It sends TOT commands to the roll/yaw computer. All the level detectors and logic circuits are contained in the roll/yaw computer. The trim actuators and relays function the same as they did for each individual trim system.

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Trim-level detectors Inside the roll/yaw computer (in addition to the limit-level detectors) there are trim level detectors. There are four detectors for each trim channel. There’s a trim level detector A and B for one direction of trim motor travel, and detector A and B for the other direction. Both A detectors are supplied trim motor position by the trim motor position LVDT A. In addition, both B detectors are supplied trim motor position by LVDT B. When any of the three trim actuators are out of the TOT position, and the TOT button is pressed, the TOT logic energizes trim relays to bring the trimmed system back to TOT. The pitch channel has a bias applied to the trim level detectors. This bias is used to provide the 1° of aft stick deflection required for pitch TOT.

Moving to TOT Depressing the TOT switch provides two initiate signals (28 VDC) to the roll/yaw computer. The logics are TOT command A, and TOT command B. Commands A and B are provided to TOT logic circuits. If the logic circuits sense an out of trim condition from any of the three systems trim level detectors, a drive sequence is started. This allows the trim actuator(s) to drive in the TOT position.

Stopping TOT movement Once the trim level detectors sense that the trim actuator is in the TOT position, the TOT logic circuits stop the applicable trim relays from driving the trim motors.

TOT light The TOT light illuminates when all three trim systems have reached the take-off trim position. This process is activated by a logic output from the A trim level detectors in the roll/yaw computer for each channel of trim. This output is used as an input to a four input AND gate which initiates a 28-VDC signal to illuminate the TOT light. When the TOT switch is depressed, it provides a high logic to the AND gate, and when each channel is in the TOT position, a logic high is sent to the AND gate. So, when the AND gate has four highs applied to the inputs (one from the switch, one from each of the pitch, roll, and yaw channels), the output causes the TOT light to illuminate. With the TOT switch providing the high input to the AND gate, releasing the switch will turn OFF the light. The light signals to the pilot that TOT has been accomplished.

218. F-15E automatic flight control system Although the F-15E model aircraft contains a similar AFCS, there are some notable differences. The biggest difference is that the F-15E AFCS is triple redundant. With this, it has the ability to operate after a single channel failure. Triple redundant means there are three CAS channels: channel A, channel B, and channel C, instead of the older system’s two CAS channels. If the FCC determines that a single channel has failed, the system design allows for continued AFCS operation with the remaining functional channels. There’s no doubt that system design such as this provides additional mission flexibility. Because the F-15E system is failure-tolerant, the pilot can experience a greater number of failures before the CAS becomes completely inoperative.

F-15E automatic flight control system components and differences The F-15A through D AFCS contains two flight control computers. These are the pitch and roll/yaw computers. The F-15E AFCS combined the pitch and roll/yaw computer into a single unit called the FCC. The FCC is the heart of the F-15E AFCS.

Flight control computer The FCC is a triple-redundant digital computer which controls the operation of the F-15E automatic flight control system. With the top cover removed (fig. 2–38), you can see the computer is divided into three distinct channels; channel A, channel B, and channel C. Each channel contains six different SRUs. The SRUs are interchangeable between the three channels.

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Figure 2–38. FCC.

Control augmentation system control panel The CAS control panel is located on the forward section of the front cockpit left console. The F-15E CAS control panel, shown in figure 2–39, is basically the same as that of F-15A through D, yet there is a difference unique to the F-15E AFCS. In addition to the pitch, roll, and yaw engagement switches, and the TOT button, there are two additional switches (BIT CONSENT and TF COUPLE (terrain following couple)) unique to the F-15E AFCS. The BIT CONSENT switch is the square button which enables the AFCS to enter a PBIT or a maintenance (MBIT). Setting the TF COUPLE switch to ON engages the terrain following (TF) function. This function allows the AFCS to control the aircraft during auto TF (automatic terrain following) operation.

NOTE: The ALT HOLD and ATT HOLD enable switches have been removed on the E-model control panel. Those functions of pilot relief are now selected on the up-front controller (UFC).

Figure 2–39. F-15E AFCS CAS control panel.

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Pressure sensor assembly The pressure sensor assembly (PSA) (fig. 2–40), with a few differences, performs essentially the same function as the F-15C AFCS dynamic pressure sensor. It’s a two channel device containing two pressure sensors for each channel. In each channel, one sensor measures pitot pressure and the other measures static pressure. A major difference from the C-model AFCS is that the PSA receives pitot and static pneumatic pressures from both the left and right pitot static probes. The resulting pitot and static signals produced by these sensors are hardwired into the channel A and B of the AFCS. Also, these signals are cross-channel communicated so that each of the three channels in the FCC has access to all three signals (left and right pitot and S2 static) produced in the PSA.

Figure 2–40. F-15E pressure sensor assembly.

Accelerometer assembly The accelerometer assembly (fig. 2–41) performs the exact same function as in the F-15A through D AFCS; but, there are channel A, B, and C acceleration signals produced instead of the channel A and B used in the older system. During operation, normal and lateral accelerations are sensed using six accelerometers. The accelerometer output is sent to the FCC.

Figure 2–41. F-15E accelerometer assembly.

Rate sensor assembly The RSA (fig. 2–42) performs the exact same function as in the older AFCS system. It contains nine rate gyros that sense pitch, roll, and yaw rate signals that are proportional to the angular motion about

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each axis. Again, there are three gyros for channel A, three for channel B, and three for channel C. The RSA output is sent to the FCC.

Figure 2–42. F-15E rate sensor assembly.

Stick force sensor The F-15E stick force sensor functions identically to the F-15A through D unit, except the F-15E stick force sensor produces three signals to facilitate outputs to channel A, channel B, and channel C within the FCC. The strain gauge elements output is sent to the FCC. Figure 2–43 shows the stick grip mounted on the stick force sensor.

Figure 2–43. F-15E stick grip and stick force sensor.

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Stabilator servocylinders The stabilator actuators operate exactly as they did in the F-15A through D AFCS; however, there are differences in the F-15E AFCS servocylinders. The F-15E AFCS requires a three-channel input for its operation; therefore, these actuators have a channel A, channel B, and channel C command and feedback lines instead of two. The forward and aft shut-off valves, as well as the rotary force motor, has a three-channel input to accommodate the triple redundant system design. Only two channels at a time are required for proper CAS operation. The rotary force motor is exclusive to the F-15E. The rotary force motor is an electric motor designed to move the shuttle valve of the EHVs.

Rudder servocylinders The rudder actuators are identical to those used on the older systems. In fact, each actuator has only one channel for operation. The outputs of each channel are fed to the FCC. Channel A is used for the left rudder servoloop operation while channel B is used for right rudder servoloop operation. Channel C is used for failure monitoring and shutoff valve control. Whenever a channel failure occurs, that channel is disengaged. Unlike the older yaw CAS, the F-15E yaw CAS continues to function with the remaining good rudder servo operating at twice the normal gain on the command. A rudder servo is disengaged whenever two out of three channels agree to disengage the rudder, or when the drive channel A or B determines that the rudder should be disengaged.

Trim actuators The trim actuators in the F-15E AFCS operates the same as in the older AFCS system. The difference is that the position feedback LVDTs on each trim motor contains three channels; A, B, and C. The LVDTs send longitudinal, lateral, and directional feel trim actuator position signals to the FCC, which monitors changes in pilot trim commands as well as stick and rudder pedal position. In addition, the rudder pedal position LVDTs are used for yaw CAS commands.

AFCS circuit breakers The AFCS AC and DC circuit breakers are located in the center main instrument panel of the front cockpit. Figure 2–44 shows the specific location of the AFCS AC and DC circuit breakers in the front cockpit. Three AFCS DC circuit breakers are located to the right of the air conditioning vent, and five are located above the vent between the EMERG BK/STEER handle and the rudder pedal adjustment knob. You may be directed by the TO to pull these circuit breakers during troubleshooting procedures.

System operation The F-15E AFCS operation is essentially the same as the older F-15A through D AFCS. As mentioned before, the biggest difference is the ability of the newer system to continue to operate after a channel failure. The purpose of the F-15E AFCS is the same as the older system; that is, to provide a uniform aircraft response to pilot input regardless of changing in-flight conditions.

Yaw CAS Engagement of yaw CAS is the same as in the older system. The reasons that yaw CAS won’t engage or disengage are reported to the operator by way of BIT codes. Also, to engage yaw CAS, the yaw rate must not exceed 41.5° per second. The validity of the rudder pedal position signal is checked, and if invalid is set to zero. When this occurs, the yaw CAS provides damping with no rudder pedal CAS command.

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Figure 2–44. F-15E AFCS circuit breakers.

The yaw CAS provides command augmentation adding or subtracting a maximum of ±15° rudder travel. This is in addition to the mechanical rudder travel of 15° for a total of ±30° maximum rudder travel. All three channels of the yaw CAS compute a rudder command and perform failure monitoring. Unlike the stabilator servocylinder, which contains three channel LVDT outputs, the rudder servocylinder contains only a single channel LVDT output. However, channel A drives only the left rudder servo, channel B drives only the right rudder servo, and channel C provides SOV control and failure monitoring. As mentioned before, in the event of a single channel failure, that channel is disengaged, and yaw CAS operation is continued with the good rudder servo. The yaw CAS command gain is doubled on the remaining functioning rudder servo in order to assure proper yaw CAS control. The pilot is alerted of the failure by display of the lateral stick limit (LAT STK LMT) caution and one rudder CAS functional fail advisory. During one rudder yaw CAS operation, the remaining rudder command is compared to the LVDT feedback position. If the compared positions don’t agree, a BIT code is generated. In addition, that rudder actuator is shut down, and yaw CAS is also shut down. During normal yaw CAS operation, the left and right rudder servocylinder position is also compared to each other. If the left and right rudder mistrack by 4° or more, a BIT code is generated and one of the actuators will shut down.

Outputs from the lateral accelerometer and yaw rate gyro control yaw CAS gain and response to aircraft movements in yaw. The outputs of these are compared with pedal position in order to detect aircraft over- and under responsive conditions. Yaw CAS then adds or subtracts rudder deflection to

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provide uniform response during those conditions. If a single channel from either the yaw rate gyro or the accelerometer occurs, that failure is isolated and yaw CAS continues to operate with the other two good channels. The pilot is notified of the failure by a generated BIT code and associated first functional failure advisory. Only when the system experiences a second channel failure of the same function does the yaw CAS then shut down. This is an example of the benefits gained from moving to a triple redundant system.

Another important feature of the yaw CAS is to assist in improving turn coordination. This is known as CAS ARI. It’s just like that found on the older aircraft and is used together with the mechanical ARI function. The AOA signal is used to control the gain on the yaw CAS ARI function. At higher AOA, greater rudder deflection is needed for turn coordination. Above 1.5 Mach, the rudder deflections from crossfeed inputs to the yaw CAS are reduced to zero and remain there for any AOA below +18°. This is to minimize roll/yaw coupling tendencies.

Roll CAS Engagement of roll CAS is the same as in the F-15A through D model AFCS. Keep in mind that the yaw CAS must be engaged first before roll CAS can be engaged. Again, if yaw rate exceeds 41.5° per second, yaw CAS disengages, which also disengages roll CAS. Just like the older system, if pitch CAS disengages the roll CAS will disengages as well. If the problem causing disengagement was in the pitch CAS circuits only, roll CAS can be reset and operated normally. As with yaw CAS, the reasons the roll CAS won’t engage or disengage are reported by means of generated BIT codes.

Roll CAS provides command augmentation which adds or subtracts a maximum of ±5° of differential stabilator travel. Several factors work to reduce that 5° authority toward zero. First, airspeed is used to reduce roll CAS authority by the function of the pressure sensor assembly. At about 544 knots, the roll CAS authority is at ±5°. Above 544 knots, the higher airspeed changes roll CAS output so that at 800 knots, roll CAS authority is reduced to 1.1°. Also, AOA plays a part in roll CAS output reduction. Between +7° and –1° AOA, roll CAS authority is at the maximum ±5°. Between 7° and 23° AOA the roll CAS is reduced so that at 23° AOA, the roll CAS authority is 0°. Also, below –1° AOA, roll CAS authority is at 0°. The AOA signal is fixed at 1.44° with weight-on-wheels (WOW)/wheel spin-up condition in effect. This is done to ensure full roll CAS authority immediately after touchdown as well as minimizing the roll-to-yaw crossfeed gain. Such design features aid the pilot during crosswind landings.

The roll CAS command is generated from the summed front and rear stick force sensor (SFS) pilot commands to control the aircraft in roll. If the rear cockpit SFS signal is invalid, the FCC recognizes that and sets that SFS output to zero. The total SFS command is fed through a deadband and gradient network in order to desensitize the roll commands around the neutral point. The reduced sensitivity diminishes roll ratcheting and lateral pilot induced oscillation. The filtered roll pilot command (from the force sensor) is combined with the roll rate signal (from the rate sensor assembly) and roll trim position signal, from the roll trim actuator to produce the roll CAS command. The unlimited roll CAS command is fed through a roll CAS authority fader and safety limiter before reaching the stabilator series servo. After a roll CAS command is computed, it’s fed into the pitch-roll servoloop, or DSS, where the pitch and roll CAS commands are combined to produce a single command. This is identical to the F-15 A through D AFCS, in that the DSS modifies the pure pitch command into a “differential pitch” to represent the roll CAS command to the servocylinder.

The pitch-roll servoloop software monitors the actuator interface for proper AFCS operation. BIT codes are generated if the FCC detects differences in predetermined values among channels. The CAS ram position, servoamp voltage, and command wraparound are monitored in this fashion. The command versus position monitor compares the stabilator command to the CAS ram position feedback. These two are compared and if these values disagree, a BIT code is generated to reflect this. If two out of three channels disagree (in addition to a BIT code produced) both pitch and roll CAS will disengage. If a failure is incurred with the pressure sensor assembly, such as pitot or static

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signal from a single channel, the roll CAS authority will be calculated on the last known valid signal in the failed channel. The pilot is notified of the failure with a LAT STK LMT caution and a ROLL LIMIT functional failure. As before, if a single channel fails, the failure is reported as a BIT code to the pilot and the roll CAS will continue to operate with reduced redundancy.

Pitch CAS Engagement of the pitch CAS is similar to the older system. The stabilator servocylinder and the CASI (pitch trim controller) solenoid valves and servoamplifiers must engage in order to engage pitch CAS. Again, a yaw rate at 41.5° per second will cause the pitch CAS to disengage. This is considered normal and won’t generate any associated BIT codes. Like yaw and roll CAS, generated BIT codes report the disengagement and the reason for disengagement to the pilot.

The front and rear stick force sensor signals are summed to produce one pitch CAS pilot command. Again, if the rear SFS signal is invalid, the FCC detects that and sets the rear SFS signal to zero. As in roll CAS, pitch CAS SFS signals are fed into a deadband and dual-gradient network to prevent oversensitivity near the null and to reduce the stick force required during high-g maneuvers. Deadband breakout force in basic CAS is 1 pound (3.5 pounds when altitude hold is used) and +10.64 pounds/–4.5 pounds when auto terrain following is in control.

During normal flight, with the landing gear handle up, pitch CAS feedback is a combination of normal acceleration signals from the accelerometer and filtered pitch rate signals from the rate sensor assembly. With the gear handle down, only unfiltered pitch rate is used for feedback to the pitch CAS. This is because during landing and touchdown, acceleration signals would fight the pilot commands required to provide stable control of the aircraft.

Just as in the older AFCS system, AOA is used to reduce gain on the pitch CAS. At an AOA of 13°, the pitch CAS gained is reduced to implement the stall inhibit function. The gain reduction also aids in matching the pitch CAS stabilator commands with the mechanical flight control system during high-AOA maneuvers. The pitch CAS authority is also reduced with large nose-up servocylinder deflection exceeding 18°. This is to ensure that roll CAS inputs to the stabilator command can still be executed during large aft stick maneuvers. The AOA signal is fixed at 1.44° during WOW/wheels spin up because the AOA sensor output is unstable at low taxi speeds. The pitch CAS authority is controlled by the pitch CAS authority fader, which monitors the previously mentioned flight conditions.

Pitch CAS commands sent to both stabilator actuators are of the same polarity, resulting in collective stabilator pitch movements. Pitch commands are also applied to the pitch trim compensator, so the CASI servo can force the mechanical pitch axis to track the pitch CAS axis in case of CAS failure. This allows the mechanical pitch axis to take over at the point of failure.

Since there are three channels in the F-15E AFCS pitch CAS, it has a high-failure tolerance, as compared to the older system. In fact, as with the roll and yaw CAS, you can have a single channel failure and pitch CAS will still operate, although with reduced redundancy. Again, the pilot is notified of the failure, by way of BIT codes, and the system uses the remaining two functional channels of pitch CAS.

Input signal management The philosophy of input signal management is the ability to maintain three independent computers within the FCC. These are channel A, channel B, and channel C. They provide for continued CAS operation when a failure exists. As stated before, this triple redundant, fail-tolerant design provides greater system reliability and flexibility to the pilot, who can count on the continued operation of the AFCS to complete the mission. Only a subsequent second failure of the same function will cause the CAS to shut down. This system design supplies the desired qualities of fail-operate/fail-safe operation.

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Each FCC channel is hardwired to a respective triplex sensor. Refer to figure 2–45 for the following discussion. The triplex signal is sent to the FCC where it first passes through a digital-to-analog converter within the input output processor (IOP). A channel A input is considered the local signal for channel A, channel B input signal is considered the local signal for channel B, and the channel C input signal is considered the local signal for channel C. The raw signal is bias compensated to reduce transients by attempting to make the local (L) signal the same as the selected or median (middle) signal. The triplex signals within the IOP are cross-connected to the other channels by what’s termed the cross-channel data link. By using this arrangement, all three FCC channels have access to the same signal. For example, a channel A forward cockpit roll SFS signal is also available to channel B and C. The cross channel data link (CCDL) makes it possible to have a failure in one channel, yet enables the other two channels to operate with information supplied to that failed channel. In instances where the input sensors provide duplex outputs (such as the pressure sensor assembly), input system management enables operation of the roll CAS in the event of a single channel failure. Software within the input system management computes the roll CAS authority based on the last known valid signal in the failed channel. The pilot will be notified of the failure with a LAT STK LMT caution and a ROLL LIMIT which warns him/her to use care when moving the control stick in the roll axis.

Figure 2–45. F-15E triplex signal inputs to the FCC.

Another function of the input system management is failure detection. To do this, the local signal is compared with the selected signal. If the two signals differ by a predetermined amount, the FCC considers the local signal failed. The failure detection process considers only two signals: (1) selected signal and (2) local (L) signal.

Therefore, a channel can only fail its local signal. Although a channel B roll rate signal failure would be seen by all three channels, only the channel B roll rate signal will be reported as a failure generating a B channel BIT code. It would be displayed to the aircrew as a BIT code XXX, Roll Rate Fail 1st local in the B channel only.

Discrete signal processing The input system management processes discrete signals for use by the FCC. A discrete signal is either in one state or another. On or off, engaged or not engaged, and up or down are examples of discrete signals. Discrete signal processing is categorized by the type of signal being processed. This can be either triplex or three state discretes.

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Triplex discrete signals are produced by switches or relay contacts and are brought into the FCC one discrete per channel. Each channel receives a hardwired discrete from each triplex relay or switch, and two discretes from the CCDL, for a total of three discretes of that input to each FCC channel. Triplex discretes are used for the following:

• Pitch CAS engage. • Roll CAS engage. • Yaw CAS engage. • Pitch CAS reset. • BIT consent. • Manual terrain-following (MTF) armed. • FORWARD AUTOPILOT DISENGAGE switch (ADS).

There are three state discretes: 1. WOW. 2. Wheel-spin-up. 3. Landing gear handle signals.

The discretes are ground, 28 VDC, or –15 VDC. The FCC monitors these discretes for an out-of-range condition.

Output discrete management The FCC produces the output discrete signals to control the following:

• Automatic speed brake retract. • Trim actuator extend and retract. • Pitch CAS status. • Roll CAS status. • Yaw CAS status.

Extend and retract signals are determined in each FCC channel. The pitch trim extend and retract signals are driven by channel A, while the roll trim extend and retract are driven by channel B. Yaw trim extend and retract are driven by channel C. The CCDL routine sends each channel’s data to the other channels. The output signals are majority voted and used to control the trim actuator drivers. The AFCS READY discrete signal indicates to the AIU 1 the status of the AFCS, in turn lighting the AV BIT light if necessary. The AFCS READY signal is majority voted through CCDL and is driven by channel B. The AUTOPILOT OFF discrete signal tells the AIU 1 the status of the autopilot, which, in turn, lights the AUTOPILOT CAUTION if necessary. This signal is also majority voted and is driven by channel C.

The FCC provides the AIU 1 with status signals of the CAS in each axis. These status signals power the caution drivers if the CAS disengages. Channel A drives the PCAS caution display, channel B drives the RCAS caution display, and channel C drives the YCAS caution display. In addition to these, the FCC provides three auto TF discretes to drive their respective cockpit caution lights. The FCC also provides a ground to the avionics status panel if any one of the following AFCS LRUs fails:

• FCC. • Force sensor. • Rate sensor assembly. • Pressure sensor assembly.

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• Acceleration assembly. • CAS control panel.

The ground latches ASP fault indicator 1. All three channels of the FCC can cause an output to the ASP circuit.

External data management The external data management (EDM) function allows the FCC to communicate with other avionic components on the MIL STD (Military Standard) 1553MUX bus. Although the FCC can communicate on two avionics MUX buses, only one MUX bus is active at any time. The central computer is the normal bus controller (BC) and treats the FCC as only one peripheral unit (with one address) even though the FCC has three identical internal computers. The channels within the FCC have the capability to transmit and receive messages over the 1553 MUX bus on either the 5A or 5B MUX buses. The EDM performs three important functions within the FCC:

1. Selects an active transmitter from the three potential transmitters. 2. Manages data from other FCC functions for transmission to other aircraft systems. 3. Monitors “received data” and prepares it for use by other FCC functions.

The typical systems the FCC must communicate with over the MUX bus include the following: • Inertial navigation unit (INU). • LANTIRN Nav Pod. • AIU 1. • AIU 2.

This communication link enables several capabilities the aircraft offers such as: autopilot, terrain following displays, FCC status displays, BIT display, and BIT initiate.

The validity of the transmitted data is checked and compared to the other data sources. If two of the channels agree with the data being presently transmitted from the selected source, that channel continues to transmit data. If at least two channels don’t agree on which source to use, the FCC turns off all data sources to reset them. It then selects channel A as the initial selected channel source. Channel A is also the initial channel used after a cold start power up.

Communication over the 1553 MUX bus is critical to successful AFCS operation. That’s because it’s the process used to (1) engage the autopilot, (2) select the autopilot modes, (3) select the autopilot steering modes and TF and autopilot (A/P) signal data, (4) initiate BIT, (5) monitor the results of BIT, and (6) audit the information stored in the FCC.

Upon AFCS IBIT, the pilot selects AFCS on the BIT menu on the MPCD/MPD together with the BIT consent switch on the CAS control panel. This switched information is sent via the MUX bus to the FCC to start the initiation process. If the FCC BIT requirements are met, the FCC will send (also via the MUX bus) an AFCS IN TEST message to be displayed on the MPCD/MPD. This gives the pilot confirmation that the AFCS system BIT has begun.

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Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

216. F-15A through D control augmentation system 1. What unit combines the pitch and roll CAS inputs into a common output to drive the stabilators?

2. What’s the purpose of the rate gyro assembly?

3. What function does the dynamic pressure sensor serve?

4. Why is the CASI function of the PRCA necessary?

5. Why do the servocylinders contain LVDTs?

6. Why does the CAS need an input from the trim system?

7. What does the pitot-static system provide to the AFCS?

8. When do the left and right AICs send a logic signal to the roll/yaw computer? What is the purpose of the logic signal?

9. What is the result of moving the yaw CAS switch from the OFF to ON position?

10. What requirements must be met prior to the engage logic activating the transistor switches?

11. How are the rudder servocylinder SOVs energized?

12. What signals are compared to compute the yaw CAS signal?

13. How will the yaw CAS compensate for an under responsive aircraft?

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14. What signal differences are monitored during yaw CAS?

15. What’s the purpose of the AC interlock?

16. Why does the yaw rate limit detectors sense yaw rate?

17. In addition to the RESET and ON logics, what are the requirements needed to engage roll CAS?

18. Why must an engage logic be applied to the pitch computer to engage roll CAS?

19. Which signals are used to vary the roll CAS output?

20. What’s the purpose of the roll trim signal that’s applied to the roll CAS computation circuits?

21. Why does a failure of the pitch CAS cause the roll CAS to shut down? When can roll CAS be reset?

22. State the additional requirements to the RESET and ON logics from the pitch CAS switch necessary to engage pitch CAS.

23. What signals are used to control the stall-inhibit function of pitch CAS?

217. F-15 trim systems 1. What’s the main difference between the pitch and roll trim actuator system?

2. How is the yaw trim actuator different from the pitch and roll trim actuators?

3. What does the TOT light indicate?

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4. Which LRU contains the control circuits for the trim system?

5. What two commands are generated when the PITCH TRIM switch is placed in the NOSE-UP position?

6. Which pitch trim motor LVDT determines trim limits?

7. Explain how the control stick is repositioned when the pitch trim motor is driven in either direction.

8. What is the purpose of the roll trim function?

9. Where are the roll trim relays located?

10. What commands are generated by the RUDDER TRIM switch when it’s placed in the NOSE-LEFT position?

11. What is unique to the yaw trim system as compared to the other trim systems?

12. What’s repositioned when the yaw trim actuator drives?

13. To what position are the control surfaces positioned when in the TOT position?

14. How many trim level detectors are used in each channel?

15. Explain when and in what direction the trim motors drive for TOT.

16. What causes the trim motors to stop driving once they reach the TOT position?

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218. F-15E automatic flight control system 1. How many channels does the F-15E AFCS contain?

2. What switch allows the AFCS to control the aircraft during auto TF operation?

3. For AFCS to operate, how many channels on the stabilator servocylinders on the F-15E must be operating?

4. How many channels are required for CAS operation?

5. What’s the purpose of the F-15E AFCS?

6. Which yaw CAS channel provides SOV control and failure monitoring?

7. What happens to the yaw CAS command during one rudder CAS operation?

8. What happens when the left and right rudder mistrack by 4° or more?

9. What’s employed to reduce pilot induced oscillations in roll CAS?

10. Where is the roll CAS command sent after it’s computed?

11. What happens to the pitch CAS authority during nose-up servocylinder deflections that exceed 18°?

12. What happens when the pitch CAS experiences a single channel failure?

13. What’s the purpose of the input signal management within the F-15E AFCS?

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14. What two signals are compared by the input signal management in order to provide failure detection?

15. What produces the triplex discrete signals on the aircraft?

16. If one of the AFCS LRUs fails, what does the FCC provide to the ASP?

17. What’s the function of the EDM within the FCC?

18. How does the EDM react when at least two channels don’t agree on which data source to use?

19. What message does the FCC send over the MUX bus to alert the pilot that the AFCS system has begun a test?

2–4. Engine Air Intake System As a 7-level you must understand the engine air intake system. This system often produces redballs and grounds the aircraft when it breaks. If you have great knowledge of the components and how they work you will be very valuable to your unit by enabling the aircraft to return to a flyable status when it may otherwise sit broken. On the flip side if you can’t troubleshoot this system you will spend many hours scratching your head while the mission is waiting on you.

219. Engine air intake system components, inputs, and outputs In this lesson we will discuss the components of the engine air intake system along with their inputs and outputs. The components we’ll be discussing are the variable inlet ramps, two AICs, actuators, and associated controls.

Variable inlet ramps There are four ramps and a bypass door that are controlled by the ramp system. Figure 2–46 shows a side view of the engine inlet and you can see all four ramps. (The fourth ramp is called the diffuser ramp.) You can also see the position of the bypass door. Additionally, the figure shows the exit louvers, bleed air holes, and protective screens that protect the ramp system and engine from damage due to excessive boundary layer air and entry of foreign objects.

Now, let’s see how shock waves are established to reduce airflow to the engine. The thin arrowed lines in figure 2–46 show how the ramps control the flow of air into the intake. The air exit louvers, on the first ramp (inlet) and aft of the bypass door, are for the removal of boundary layer air. Boundary layer air is bled off the first ramp (inlet), second ramp, and third ramp through bleed air holes in the ramps and is dumped overboard through the air exit louvers.

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Figure 2–46. F-15 inlet ramps.

Air-inlet controllers For the ramp system to operate properly the ramps must be appropriately scheduled. To do this, AICs are used. An AIC (fig. 2–47) is a digital computer which contains monitoring circuits, BIT circuits, and a fault indicator. Each engine has its own independent intake system, so there are two AICs—left and right—one for each engine. The left and right AICs are identical. The AICs are used to control the position of the inlet ramps and bypass door for their respective engine.

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Figure 2–47. Air inlet controller.

Actuators The ramp system uses three actuators. These are shown along with the ramps in figure 2–46. The first ramp, diffuser, and bypass door actuators are hydraulically powered and controlled by the AIC. The utility hydraulic system supplies hydraulic pressure to the actuators. The actuators receive electrical signals from the AIC and convert them into a hydraulic signal which extends or retracts the actuator according to the command. The first ramp and diffuser actuators provide LVDT position feedback signals to the AIC. These signals are used to verify that the actuators have moved to the position commanded by the AIC. The actuators have a locking device which prevents their actuation until certain conditions are met.

Associated controls The controls for the ramp system are the INLET RAMP switches located on the miscellaneous control panel on the cockpit left console, and the AIC TEST switches located in doors 6L and 6R.

INLET RAMP switches These switches are located on the miscellaneous control panel, shown in figure 2–48. There are two positions—AUTO and EMERGENCY. In the AUTO position, the AIC schedules the ramps and bypass door according to the inputs. Should an emergency occur or the left or right AIC fail, the EMERGENCY position allows that side to be locked in a predetermined position. A system failure could be critical to engine operation, so the EMERGENCY mode raises and locks all ramps (except above 1.5 Mach) so the engine isn’t starved of air.

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Figure 2–48. INLET RAMP switches.

AIC TEST switch This three-position switch (fig. 2–49) is used during ground checkout of the ramp system. Position A and B are used during ground checkout to initiate a series of tests on the AIC and to move the ramps when hydraulic pressure is applied. The center position is OFF and the switch is naturally spring loaded to this position. Remember, there’s a switch for each side.

Figure 2–49. AIR INLET CONTROL TEST switch.

AIC inputs and outputs People are often awed by the ability of electronic systems to maintain critical tolerances. The systems aren’t so awesome if you think of them as decision makers. Most of the modern electronic wonders

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we find on our aircraft are basically decision makers. For you to make a decision, you normally use inputs (or things you sense) to decide on the outputs (actions to take). The AICs work in the same manner; that is, they use inputs from other systems to produce outputs to carry out decisions.

Inputs For ramp positioning, the AICs receive inputs of pitot-static pressure, total temperature, diffuser and first ramp actuator position, AOA, and a diffuser actuator lock/unlock signal from the ADC.

Pitot-static pressures The AIC is supplied pitot and static pressures from the freestream pitot/static probes, the duct pitot probes, and the duct static ports. Figure 2–50 shows the pitot-static pressures supplied to each AIC. These pressures are used to compute the desired position of the first ramp, diffuser, and bypass door actuators.

Figure 2–50. Pitot-static pressures.

Total temperature Total temperature probes supply a total temperature signal to the AIC. Why do we want to monitor temperature? We monitor temperature because it affects the density of air which has a direct effect on the operation of the engines. The sensed temperature is used to compute the relative air density for positioning the first ramp, bypass door, and diffuser ramp.

AOA The AIC needs an AOA input to compute and control the amount of freestream air entering the inlet. The left and right AOA transmitters supply local AOA to the respective AIC. The AIC uses the AOA data to schedule the first ramp position.

ADC The ADC supplies a discrete Mach signal (ground) to the engine air inlet system. This signal unlocks the diffuser ramp actuator at 0.84 Mach and locks it at 0.80 Mach.

Actuator position The first and diffuser ramp actuators have an LVDT that supplies position signals back to the AIC. The AIC compares the position of the actuators to the commanded position. If there’s a disagreement, the (L or R) INLET light on the caution lights display panel will illuminate. Upon this warning, the

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pilot may select EMERGENCY on the INLET RAMP switch and continue to fly the aircraft, but only under certain flight restrictions.

Outputs The ramp system provides various signals to the engine electronic control (EEC, F100-PW–100 engine) or the digital engine electronic control (DEEC, F100-PW–220/229 engine), ASP, and caution lights display panel.

EEC/DEEC The AIC supplies a digital Mach signal ranging from 0.15 to 2.80 Mach to the EEC/DEEC. The EEC/DEEC uses the Mach signal to control engine operations at various speeds within the noted range. Since the EEC/DEEC and AIC are both digital computers, a clock pulse is also used to synchronize the Mach signal.

ASP The ASP has fault indicators that latch from black to orange in the event of an AIC failure. ASP fault indicator No. 6 is for the left AIC, while ASP fault indicator No. 7 is for the right AIC. ASP fault indicator 6 and 7 indicate a failure exist under doors 10L and 10R respectively.

Caution lights display panel The AICs control illumination of two lights on the caution lights display panel. These are the L INLET and the R INLET lights. If either the left or right system fails, the appropriate caution light illuminates to warn the pilot of a malfunction. The MASTER CAUTION light will also illuminate with either L or R INLET light.

220. Auto and emergency modes and ground operation Now that you know the components of the engine air intake system you need to know how it operates. In this lesson, we’ll discuss the two modes of operation for the ramp system: auto mode and emergency mode. We’ll conclude with a discussion of ground operation.

Auto mode To obtain the most efficient aircraft performance at speeds greater than approximately Mach 1.4 it’s necessary to use oblique shock waves created by the vari-ramps to reduce the speed of the supersonic airflow before it gets to the engines. Supersonic airflow is reduced by using external oblique shock waves at the first ramp (inlet), second ramp, and third ramp. These oblique shock waves are generated by extending the ramps ahead of the inlet entrance. The final reduction of airflow to subsonic flow takes place through a normal shock wave positioned at the inlet duct lip.

NOTE: 50 RPM is used to explain inlet system operation. Depending on what F-15 airframe you are working on, and even which switch is installed (50 percent switch) on the airframe, the percentage may differ slightly.

Prior to engine start, the ramps and diffuser are up and locked and the bypass door is closed. If for some reason, the ramps or door actuators aren’t in the proper position as soon as sufficient hydraulic pressure is supplied to the system after engine start, the actuators move to the proper position. After the applicable engine is started and reaches 50 percent RPM, the first ramp is unlocked through the engine run relay and moves to the full DOWN position. Once the aircraft is airborne, a WOW relay ties both left and right systems together. This combining of the systems allows either system to use both 50 percent switches for continued ramp operation in the event of an engine shutdown or switch failure. Remember, the second ramp, third ramp, and diffuser remain locked in the full UP position until the aircraft speed reaches 0.84 Mach. Although the bypass door isn’t locked closed, hydraulic pressure is shutoff to the actuator until after the engine has started.

After engine start, the first ramp position is computed using AOA, total temperature, and pitot-static inputs. The bypass door position is computed using total temperature and duct pitot-static inputs. Once the diffuser actuator is unlocked by the 0.84 Mach signal from the ADC, the AIC computes its

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position using AOA, total temperature, and pitot-static inputs. The AIC receives and computes the inputs into a signal to position the first ramp, diffuser, and bypass door actuators. This signal is applied to the servo valve of the appropriate actuator. The actuator converts this signal into a hydraulic signal that extends or retracts the actuator. A rate limiter is incorporated into the return port of the actuators and provides a limited flow rate of hydraulic fluid to ensure the piston in the actuator doesn’t exceed its designed operating velocity. The actuators generate and route a position signal of the first ramp and diffuser actuators to the AIC. The AIC uses this signal to drive the actuator to its proper position.

The AIC continuously monitors the entire engine air intake system to detect any failure within the system. If a system fails, an output from the AIC turns on the L or R INLET caution light in the cockpit and latches the LRU fault indicator on the face of the AIC. Fault indicators 6 (left) and 7 (right) on the ASP in the nose wheel well provide normal (black) or failed (orange) indications for the left or right system.

Monitoring the inlet duct area for unstable inlet air and for proper throat pressure ratio is also accomplished by the AIC. When the throat pressure ratio exceeds a supercritical condition or when the inlet air becomes unstable, the AIC provides an output to illuminate the L or R INLET caution light in the cockpit as well as the MASTER CAUTION light.

Emergency mode An emergency mode of operation is entered by placing the L and/or R INLET RAMP switches to the EMERG position.

NOTE: Individual switches for each system allow one system to operate automatically while the other system is in emergency.

Placing either the L or R INLET RAMP switch to EMERG commands the ramps and bypass door (for that system) to the full UP position. This allows hydraulic pressure to be ported to the retract side of the piston in both the first ramp and diffuser ramp actuators and to the extend side of the piston in the bypass door actuator. However, if aircraft speed is Mach 1.5 or greater, the AIC routes a signal to the Mach blocking valve solenoid in the diffuser ramp. This locks the diffuser ramp at its present position until the speed decreases below Mach 1.5. At airspeeds of less than Mach 1.5, a rate limiter in the return port of the actuator limits the flow of hydraulic fluid and prevents rapid retraction of the diffuser ramp.

If utility hydraulic pressure is lost in-flight, air loads push the first ramps and diffuser ramps to the full UP position until the locks are automatically engaged. The bypass door may close or it may remain open, depending on the air pressure inside the inlet duct.

Ground operations You can use the AIC for ground operation to visually observe ramp movement. You do this by placing the L or R AIC TEST switch in door 6L or 6R to position A or position B. Of course, external electrical power (ground power switch 3 to ON), utility hydraulic pressure, and cooling air must be applied to the aircraft.

The AIC TEST switches apply a ground to the applicable AIC to drive the ramp actuators to a fixed preset test position. The following table shows the position of the ramps for position A and B.

AIC TEST Switches Switch Position A Position B

First Ramp +3.43° (scale position) (nearly full up) –2° (scale position) (nearly full down) Diffuser Ramp Approx. 24 inches from bottom of duct Approx. 20 inches from bottom of duct Bypass Door Closed Open

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AIC static BIT The AIC static BIT uses the monitor circuits in the AIC to ensure a failure doesn’t exist. To perform the checkout you must introduce a failure into the system, ensure it is detected, and make certain the fault indicators on the appropriate AIC and ASP are latched. The static BIT also ensures the fault indicators can be reset after the failure is corrected.

Test equipment The TTU-205D/F pressure temperature test set is designed to assist you in troubleshooting the ramp system. It’s used to check for either pitot/static leaks in the system fittings and associated hoses and tubing, or for an air induction system operational checkout. An operational check may be required for troubleshooting engine stalls and stagnations. During the engine air induction systems operational checkout, the TTU-205 test set simulates altitude and airspeed.

F-15E differences On the E-model, the engine air intake system GO/NO-GO signals flow through the AIU to the ADCP via the 1553 MUX bus. The ADCP drives the MPD and the MPCD. The MPD/MPCD will display either L INLET or R INLET on the cautions warning display format during system failures. Otherwise, the BIT and operations are essentially the same as on the F-15A through D model aircraft. One other difference is that the AIC will route data to a DEEC only.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

219. Engine air intake system components, inputs, and outputs 1. How many F-15 AICs are there? What purpose do they serve?

2. How do the first ramp and diffuser ramp actuators provide feedback position to the AIC?

3. Why is the emergency mode for the inlet ramp system necessary?

4. What signals are provided to the AIC for ramp positioning?

5. What type signal unlocks the diffuser ramp actuator at speeds greater than .84 Mach?

6. What type signal do the AICs supply to the EEC or DEEC?

7. Where will a failure of the left AIC be displayed?

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220. Auto and emergency modes and ground operation 1. In the AUTO mode, how does the ramp system obtain the most performance at speeds greater

than approximately Mach 1.4?

2. At what percent of RPM does the first ramp go to the full DOWN position at engine start?

3. What inputs are used to compute the first ramp and diffuser positions?

4. How will the ramps react when the RAMP switch is placed in EMERG?

5. How are the diffuser ramp actuators prevented from being driven too rapidly?

6. What happens if utility hydraulic pressure is lost in-flight?

7. What provisions are provided to check the AICs on the ground?

8. What test equipment is used to assist in the troubleshooting the ramp system?

2–5. Air Data Systems Pitot-static is another important system you must understand as a 7-level. You will deal with these air data systems often and like the engine air intake system they will ground the aircraft when they break. In this section we will discuss the pitot-static system. We will then learn about F-15 A through D air data computer. To conclude this section we will dive into the F-15E air data processor. Let’s begin with the pitot-static system.

221. Pitot-static system The pitot-static system senses and delivers pitot and static pressures to numerous systems and indicators. These pressures are used to generate displays such as altitude and airspeed. The pressures that the aircraft is experiencing dictate what the other aircraft systems do to keep the aircraft flying properly and the engines running at top performance.

Components The pitot-static system is designed to sense pitot and static pressures and supply these pressures to the LRU of several systems. The LRUs of those systems are as follows:

1. ADC. 2. AIC. 3. ADP (F-15E only).

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4. PRCA. 5. DPS (F-15C and D). 6. PSA (F-15E only). 7. FLAP AIRSPEED switch. 8. Standby airspeed and altitude indicators.

NOTE: The ADP is an F-15E component that replaces the ADC, the PSA, and both left and right AICs. The PSA on the F-15E serves the same purpose as the DPS on the F-15C and D model aircraft.

The F-15 pitot-static system consists of these LRUs:

1. Two electrically heated pitot-static probes. 2. Two electrically heated duct pitot probes. 3. Two flush-mounted duct static ports. 4. Eleven pitot drains. 5. Thirteen static drains. 6. Tubing to transmit the pressure inputs to the various units.

The illustration on figure 2–51 shows the entire pitot-static system and instruments. Refer to this figure as you complete the text.

Figure 2–51. F-15 pitot-static system.

Pitot-static probes There are pitot-static probes mounted on each side of the aircraft, one each on doors 3L and 3R. They are designated as “left” and “right” probes. These probes are mounted aerodynamically to compensate for flight-induced errors. The probes contain internal heaters to prevent icing conditions from blocking the openings that sense the pitot and static pressures.

Each probe (pitot tube) has a pitot port (the opening at the forward end of the tube) and static ports (fig. 2–51). Each probe has ports for two static systems— S1 and S2. The pitot port gathers ram air for airspeed calculations. In contrast, the static ports sense static pressure surrounding the probe for airspeed and altitude calculations. The static pressure enters the static holes around the middle of the probe. Impact pressure is derived by subtracting the static pressure from pitot pressure. Figure 2–51

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shows that the pitot and static pressures are connected to the system through separate lines. Notice that S2 is connected between the left and right side pitot probes. This gives us a total of 5 pitot/static systems from the two free stream pitot static probes— left pitot, right pitot, left S1, right S1, and S2.

Each pitot static probe has two heaters used to prevent ice buildup. One heater is in the head of the probe (the part of the probe that contains the pitot and static ports) and the other heater is in the mast of the probe (the part of the probe between the mounting flange and the head of the probe). These heaters are controlled by the ANTI-ICE PITOT HEAT switch on the ECS (environmental control system) control panel.

Duct pitot probes/static ports Each engine inlet duct has a pitot probe and a flush-mounted static port. The pressures sensed by these devices are used to schedule the ramps in the inlet system. With these four duct pitot/static systems, that brings the total to nine pitot static systems on the F-15.

Each duct pitot probe contains a heater that is controlled by the ANTI-ICE PITOT HEAT switch on the ECS control panel. Like the free stream pitot static probes, the heaters are used to prevent ice buildup which could clog the ports on pitot probes.

Pitot and static drains An additional hole is located on the bottom of each of the four probes to allow moisture to drain out. Moisture in the probes can block the pitot or static ports and disrupt their functions. Some water does occasionally get past the probes. For this reason, several drains are placed into the tubing. These drains are located at the lowest point in the pitot-static lines. The drains allow you to get rid of any moisture that may have accumulated in the lines.

Tubing The F-15 uses rigid and flexible types of pitot-static tubing. Flexible tubing is used in areas of high vibration or where the line needs to bend, such as a door hinge or at the point of connection to an LRU. The rigid tubing is used anywhere the flexible tubing isn’t required.

Features of the F-15 pitot-static indicating system The pitot-static instruments are designed to back up the altitude indicator and airspeed Mach indicator. They include the standby altimeter and the standby airspeed indicator (SAI). All F-15s, including the F-15E, use these instruments.

Standby altimeter The standby altimeter measures and displays aircraft altitude above sea level using S2 (static 2) pressure from the pitot-static system. As you can see in figure 2–52, the dial is graduated in feet. There are two indications on the altimeter. The first is a barometric correction scale that shows a barometric setting in inches of mercury (Hg). The second is altitude as presented by the three pointers shown in figure 2–52. This altitude indication is directly affected by atmospheric pressure change because the change acts on an aneroid inside the indicator.

An aneroid is the pressure sensing element used in the standby altimeter. It is evacuated, sealed, and mounted inside the altimeter case, which is also sealed. Static pressure is vented into the case through a small inlet. At higher pressures (on the ground), the aneroid is compressed. As the aircraft climbs, static pressure decreases, allowing the aneroid to expand. The movement of the aneroid is mechanically connected through a series of multiplying gears to the altitude pointers.

Figure 2–52 shows three pointers, indicating altitude in feet. The long pointer indicates hundreds of feet, the intermediate pointer indicates thousands of feet, and the short pointer indicates tens of thousands of feet. Also notice that the shortest pointer has a small extension connected to a triangle that moves along the outer edge of the standby altimeter. This allows for more accurate readings.

Barometric pressure changes constantly. To allow the standby altimeter to measure this changing pressure and deliver an accurate altitude readout, the instrument is designed so that you or the pilot

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can adjust barometric correction to the current barometric pressure. This is done by turning the BARO SET knob to set the barometric scale to the current pressure. Check with the expediter, maintenance operation center (MOC), or control tower for current barometric pressure (baro set). Repositioning the BARO SET knob also drives the pointers to the correct altitude in feet. The barometric scale range is from 28.10 to 31.00 inches of mercury.

Figure 2–52. Standby altimeter.

A problem still arises when two aircraft take off from two different locations and cross paths with two different baro settings. They could be at the same physical altitude, though they have two different indicated altitudes. For this reason, above 18,000 feet the pilot adjusts the baro setting to a standard barometric setting of 29.92 inches of mercury. This helps avoid mid-air collisions because all aircraft above 18,000 feet have the same identical baro setting. This policy increases the accuracy of altitude indication in relation to other aircraft, rather than sea level. As an aircraft descends through 18,000 feet to its destination, the pilot requests the local altimeter setting and adjusts the barometric setting to that pressure. This gives the aircraft an accurate altitude indication compared to the other aircraft flying below 18,000 feet at its destination.

The atmospheric pressure at a given altitude isn’t always the same; instead, changing weather conditions cause the pressure to vary. For this reason the pilot may have to adjust the standby altimeter for variations in barometric pressures. The terms in the table are related to instruments used in determining barometric pressure.

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Term Definition

Pressure altitude Is the indicated height using 29.92 Hg as the basic reference only. Using this reference only, the altimeter displays the altitude as indicated by the atmospheric pressure around the aircraft.

Altimeter setting Indicates altitude above sea-level. This is accomplished by setting the altimeter to the current barometric pressure at sea level. The station (base) altimeter is always set to this reference.

Field elevation pressure

Indicates altitude above field level. The existing atmospheric pressure in Hg at 10 feet above the mean (average) elevation of the runway. An elevation of 10 feet above the runway is used because the standby altimeter is in the cockpit and we assume the cockpit is 10 feet high. This is also known as station pressure.

The use of altimeter setting is universal. However, if a pilot desires to land with the pointers reading zero rather than the altitude of the field above sea level, he or she may do so. The pilot does this by contacting the tower and requesting the observed station pressure and then setting this on the instrument scale, using the BARO SET knob. The instrument then reads zero altitude upon landing.

The case of the altimeter is airtight. Static pressure from the pitot-static probe is fed inside the case through an inlet. The changing pressure around the aneroid causes them to move, driving the altimeter pointers. At higher pressures, the aneroid is compressed. The higher pressures occur the closer the aircraft gets to the ground. As the aircraft climbs, the surrounding pressure decreases and the aneroid expands. The movement of the aneroid is small, but the mechanism multiplies this movement and causes the following:

Pointer Makes: Larger 10 revolutions for a change of 10,000 feet. Intermediate One revolution for a change of 10,000 feet. Small One-tenth of a revolution for a change of 10,000 feet.

The movement of these pointers indicates changes in pressure altitude above a known reference, which may be sea level or the local field elevation. This reference is normally sea level.

Standby airspeed indicator The SAI is shown in figure 2–53. The indicated airspeed is displayed by a moving pointer that’s read against a fixed dial graduated in knots × 100. Notice the nonlinear graduations on the dial. This allows for greater accuracy at lower airspeeds where the pilot has to worry about aircraft stalls.

Figure 2–53. Standby airspeed indicator.

The SAI basically operates in four steps: 1. Pitot pressure is applied to a differential pressure diaphragm. 2. Static pressure is supplied to the inside of the indicator case. 3. Since the static pressure inside the case and the static pressure portion of the pitot pressure

cancel each other, only the impact pressure causes the diaphragm to expand.

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4. The expansion of the diaphragm is mechanically connected to the pointer of the indicator.

The pointer moves only when there’s a change in impact pressure. This pointer indicates the airspeed of the aircraft.

222. F-15 A through D air data computer There are seven components that make up the F-15 A through D ADC system:

1. ADC. 2. AOA transmitters. 3. Total temperature probes. 4. Altimeter indicator. 5. Airspeed Mach indicator (AMI). 6. Vertical speed indicator (VSI). 7. AOA indicator.

NOTE: The VSI is often referred to as the vertical velocity indicator or VVI.

On the F-15D model, there’s an additional set of indicators in the rear cockpit that are controlled by the same signals from the ADC. Air data computer The ADC, shown in figure 2–54, is a solid-state digital computer that uses pneumatic and electrical inputs to compute environmental data for using systems. The computer also continuously monitors its own performance and provides ADC status signals to the appropriate indicators.

Figure 2–54. F-15 ADC and AOA transmitter.

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AOA transmitters The F-15 uses two AOA transmitters (see fig. 2–54), one under panel 5R, and one under panel 5L. These AOA transmitters provide local AOA to the following:

• ADC. • AOA indicator. • AFCS. • Engine air intake system. • ICCP.

The AOA transmitter probes extend into the airstream and contain a pair of parallel slots that are positioned 90° apart. The slots allow the probe to rotate with a change in airstream direction. The air pressure exerted on paddles inside the transmitters will position the probes to maintain equal pressure in each slot. The movement of the probes also moves four potentiometers in each AOA transmitter. The potentiometers generate electrical signals representing aircraft AOA. The output of the potentiometers from each AOA transmitter is broken down in the table below.

Output of the Potentiometers Left AOA transmitter Right AOA transmitter

ADC ADC AFCS (pitch computer) AFCS (pitch computer) AOA indicator ICCP Left AIC Right AIC

Each AOA probe contains a heating element to prevent icing. In addition, each case contains a thermostatically controlled heater to reduce condensation. Operation of the probe heaters is controlled by the main landing gear (MLG) WOW switches. These heaters have a thermostatic overheat protection built into them.

Total temperature probes The left and right total temperature probes are located to the left and right of panel 15 (see fig. 2–55). The left total temperature probe supplies an indicated total temperature signal to the ADC and the left AIC. The right total temperature probe supplies an indicated total temperature signal to the right AIC. The total temperature probes are cylindrical tubes that are open to the airstream at the front end and form a nozzle at the rear. The probe contains a resistive temperature element. In the resistive temperature element, the line resistance of the temperature element and the line resistance leads are part of a bridge circuit. The heating element in the total temperature probe is operated by the WOW switch and relays in the same manner as the heating element in the AOA probe. The heater prevents ice from forming on the probe. The output of this bridge is fed to the ADC and AICs.

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Figure 2–55. F-15 total temperature probe.

Altimeter The altimeter is located on the right main instrument panel and receives correction inputs of altitude and barometric pressure from the ADC. The face of the altimeter is shown in figure 2–56. Aircraft altitude is displayed on a four-digit counter indicating thousands of feet and a single rotating pointer, indicating hundreds of feet in 20-foot increments. Barometric correction is supplied from the altimeter barometric potentiometer (baro pot) to the ADC and is displayed on the altimeter in inches of mercury. The altimeter has a BIT mode where an OFF flag is displayed to warn the pilot of a malfunction.

Figure 2–56. F-15 altimeter.

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Airspeed Mach indicator The AMI is located on the left main instrument panel. The AMI is an electromechanical device that displays airspeed and Mach number. An illustration is shown in figure 2–57. Airspeed is displayed on a pointer and Mach is displayed on a rotating Mach card. The AMI also provides an adjustable airspeed index pointer used for visual reference only. The input signals, used to control the airspeed and Mach displays, are provided by the ADC.

Figure 2–57. F-15 airspeed Mach indicator.

Vertical speed indicator The VSI, shown in figure 2–58, is an electromechanical device that accepts an altitude rate signal from the ADC and converts it into a pointer indication of vertical speed. The VSI indicates the aircraft rate of climb or dive in thousands of feet per minute. The VSI also has a BIT mode where an OFF flag can be displayed to warn the pilot of a malfunction.

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Figure 2–58. F-15 vertical speed indicator.

Angle-of-attack indicator The AOA indicator is mounted in the main instrument panel (see fig. 2–59) and provides a visual display of local AOA. The pointer shows local AOA against a fixed scale graduated in units. The optimum AOA function is not used. This pointer is driven to 45 units to keep it out of the way. A small T-shaped indexer is fixed at 21 units to indicate optimum approach AOA. In addition, an off warning flag indicates loss of power or a failure inside the indicator. The left AOA transmitter supplies a varying voltage, which represents local AOA, to the indicator. The indicator is an electromechanical device that converts the voltage into a corresponding pointer indication. The ADC provides the signal to keep the optimum AOA triangle pointer at 45 units.

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Figure 2–59. F-15A through D model AOA indicator.

Related systems and the F-15 ADC system interface The ADC is interfaced with various systems, providing digital, analog, or discrete inputs. Figure 2–60 is a block diagram that shows the systems that interface with the ADC.

Central computer The ADC provides digital airspeed, AOA, Mach number, altitude, pressure ratio, and relative air density data when requested by the CC. The CC uses this information to compute aircraft steering and weapon release data. The CC also provides air data information to the HUD for airspeed scale, altitude scale, and AOA scale indications. Airspeed and altitude scales are displayed in all modes A/A, air-to-ground (A/G), indicated velocity (VI), and attitude director indicator (ADI)), but the AOA scale is only displayed in the ADI mode with landing gear DOWN.

Identification friend-or-foe transponder The ADC provides digital altitude information to the identification friend-or-foe (IFF) transponder. This altitude information is encoded into mode C and used for automatic altitude reporting.

Engine air induction system The ADC provides a Mach discrete signal to unlock the left and right diffuser ramp actuators as airspeed increases to 0.84 Mach and to lock the diffuser ramp actuators as airspeed decreases to 0.80 Mach.

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Figure 2–60. ADC interface.

Environmental control system The ADC provides Mach and airspeed discretes that control ram air and engine bleed air to the primary heat exchanger. At airspeeds above 0.84 Mach, the ADC Mach discrete energizes a relay that closes the primary heat exchanger ram air door and ejector valve. At airspeeds below 180 knots and with weight-on-wheels, the ADC airspeed discrete deenergizes a relay that opens the secondary heat exchanger ejector valve. Engine bleed air is then used for cooling during ground operations.

Integrated communication control panel The ADC routes a signal to the ICCP initiating an unsafe landing warning tone if all the following conditions exist:

1. Altitude is less than 10,000 feet. 2. Rate of descent is greater than 250 ft/min. 3. Indicated airspeed is less than 200 knots. 4. The landing gear control handle is in the UP position.

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Unsafe landing warning tone The unsafe landing warning tone will also come on if the ADC power is removed or the ADC fails. The ADC unsafe landing warning initiate signal is applied to the tone generators in the ICCP generating a 250-Hz tone in the headset, interrupted at a 5-Hz rate, alerting the pilot to danger.

AOA stall warning tone The ICCP also receives an input directly from the right AOA transmitter. This signal initiates the AOA stall warning tone when aircraft AOA reaches 28.4 and increases in frequency as AOA goes to 33.1 units. The tone is a 1600-Hz tone in the headset, interrupted at a 1 to 20 Hz rate, depending on the AOA, and alerts the pilot of a stall condition. This tone is inhibited when weight is on wheels or the landing gear is UP.

AOA unit limit warning tone The ICCP also generates an AOA unit limit warning tone. The tone is a 900-Hz doublet (two 900-Hz tones interrupted at a 10 Hz rate with the doublet separated by a 0.3-second pause). The warning tone sounds whenever the AOA units are exceeded and the landing gear is UP. The tone initiation sequence is programmable or determined by the programmable armament control system (PACS). The 900-Hz warning tone is used for several different warnings. These are prioritized as follows:

1. The departure warning tone (yaw limit) has first priority. 2. The overload warning system (OWS) has second priority. 3. The AOA unit limit has third priority.

The AOA unit limit (warning tone) parameters are controlled by the PACS according to the following aircraft configuration:

• Defaulted to OFF when the aircraft configuration is clean, or A/A missiles are loaded. • Defaults to 30 AOA units when conformal fuel tanks (CFT), A/G stores or miscellaneous

stores are loaded.

The HUD will display (AOA XX) if the AOA reaches 18 units or greater. The pilot also has the option of changing (programming) the AOA limit through the NAV CONTROL indicator (NCI).

Automatic flight control system The ADC system integrates with the AFCS to provide a pilot relief mode called altitude hold.The ADC provides three signals to the pitch computer for the altitude hold function:

1. Analog altitude error signal. 2. Discrete altitude error valid signal. 3. Discrete altitude hold engage signal.

Flight control system The flight environmental data system integrates with the longitudinal flight control system to provide warnings to the pilot when the aircraft pitch ratio function isn’t operating properly. The ADC monitors pitch ratio using the following signals:

• A varying LVDT signal from the PRCA representing the pitch ratio changer position. • 28 VDC if the landing gear is down.

With computed airspeed and altitude, the ADC is able to determine unsafe flight control conditions. The following are considered unsafe flying conditions:

• A pitch ratio indication greater than 0.9 with an altitude less than 20,000 feet and airspeed greater than 330 knots.

• A pitch ratio indication less than 0.9 with the landing gear DOWN.

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When any of the previous conditions occur, the ADC routes a signal through the PRCA to the caution lights logic unit, illuminating the PITCH RATIO and MASTER CAUTION lights. This lets the pilot know that the aircraft will probably under or over react to aggressive stick inputs.

Air data computer built-in test operation The ADC has automatic and initiated BIT modes.

Automatic BIT The ADC BIT function operates in automatic mode when equipment power is applied and the BIT initiate mode is not activated. In this mode the ADC automatically monitors its own operation and the operations of many other components of the system. It does this without interfering with the normal operation of the system. A failure detected by the monitor illuminates the ADC light on the BCP and AV BIT light on the caution lights display panel. Also, a failure of the ADC will latch the ASP fault indicator number 3. In addition, the TOT TEMP HI light warns of high temperatures of the ram air in the intake area or a failure in the total temperature circuits.

Initiated BIT The ADC and instruments have BIT capabilities that can be initiated from the BCP. When a BIT is initiated, the pitot-static pressures, total temperature signal, AOA signal, and baro pot signal are replaced by an internal signal in the ADC to verify system operations. During the BIT, the TO requires you to pull the indicators and ADC circuit breakers and verify proper indications and OFF warning flag operation. Be sure to check the ASP before and after the BIT to see if fault indicator number 3 is latched. Fault number 3 indicates a failure of a LRU under door 3R, where the ADC is located. In addition, the ADC has its own fault indicator that latches when an internal failure of the ADC occurs.

Functional characteristics of the F-15 air data computer The operation of the ADC is divided into five sections:

1. Power supply. 2. Pressure sensor module (pitot and static inputs). 3. Input data converter (IDC). 4. Digital processor (DP). 5. Output data converter (ODC).

During operations, the ADC accepts inputs from the pitot-static system, AOA transmitters, total temperature probe, and baro pot.

Power supply The power supply uses 115-VAC, single-phase, 400-Hz aircraft power, and generates all internally required power.

Pressure sensor module There’s a pressure sensor module in the ADC for pitot and static inputs. Each module contains the following:

• A pressure sensor. • Sensor driving electronics. • Sensor temperature sensing circuitry. • Sensor calibration memory.

During operation, the pressure sensor module receives, converts, and routes vibration frequency and a voltage equivalent of temperature from the pressure sensor to the IDC.

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Input data converter section The IDC performs data conversion and provides the required interface between the air data sensors and the DP. The pressure sensor frequencies are converted into digital data by a frequency-to-digital (F/D) converter. In addition, the other inputs are converted from DC analog format into digital data using an analog-to-digital (A/D) converter. These signals are stored in registers; and upon request, the data is multiplexed to the DP. The IDC also generates required voltage for A/D conversions and the input networks associated with the total temperature probes.

Digital processor section The DP does three tasks: (1) receives digital inputs from the IDC; (2) compensates for input errors; and (3) monitors input, output, and intermodule signals among the ADC sections.

The DP contains the operational program memory, aircraft calibration memory, scratch-pad memory, arithmetic element, and the control logic for computing input data

The control logic synchronizes the operation of the IDC and ODC with the DP. The computed input digital data is processed by the DP then sent to the ODC.

Output data converter section The ODC provides required signal conditioning for interfacing the ADC with the indicators and other using systems. The ODC receives and stores the data from the DP into appropriate storage registers. Upon request, the data control retrieves and processes the data for operations. Serial digital data is provided to the altitude (ALT) indicator(s) and the CC, which may request air data outputs at any time. Linear DC analog voltages are converted into an absolute or a ratio of reference voltages. If required, the signals are buffered before being supplied to the using system.

223. F-15E air data processor The F-15E aircraft has two ADPs installed in place of the AICs with the ADC removed. In this lesson we’ll discuss how the ADP functions differ from the ADC/AIC operation and the ADP’s built-in diagnostics or BIT. Most of the related systems that interface with the ADP are identical to those of ADC/AIC equipped aircraft.

Air data processor system description The ADP system performs the functions of the ADC, electronic air inlet controller (EAIC), PSA, and FLAP BLOWUP switch (FBS).The two identical and independent ADPs perform these functions. The left performs the functions of the left EAIC and a single PSA channel (B). The right ADP performs the functions of the ADC, the right EAIC, FBS, and a single PSA channel (A). The air data processor system receives inputs from the same systems as the ADC/AIC equipped aircraft discussed in the previous lesson.

ADC function The following are the different ADC function output signals developed by the ADP.

• Altitude rate/validity to both VSIs. • Excitation to AOA transmitter potentiometers. • Optimum AOA/validity to AOA indicator. • Altitude to IFF transponder. • Mach warning to ECS. • Airspeed to ECS • Total temperature probe excitation. • Total Temp (TT) high caution to AIU for MPD/MPCD. • Pitch ratio caution to AIU 1 for MPD/MPCD display.

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• Cross channel data link to opposite ADP. • ADP BIT acknowledge to AIU for ADCP. • Right ADP fault to ASP. • Unsafe landing warning to ICSCP. • 1553 MUX data to ADCP.

EAIC function The EAIC function develops the following output signals:

• Ramp positioning commands to inlet ramp actuators. • Mach number to the DEEC. • Mach block to AFCS. • Left ADP fault to ASP. • Inlet caution to AIU for MPD/MPCD display.

FBS function The FBS function develops two output signals:

1. Airspeed high to flap control circuits. 2. Airspeed low to flap control circuits.

PSA function The PSA function develops an uncorrected pitot/static pressure signal to FCC for roll CAS limiting.

Components The ADP system contains 5 components. They are: (1) ADPs; (2) VSIs; (3) AOA transmitters; (4) TT probes; and (5) AOA indicator.

The ADP system exchanges information with 4 others components. They are: (1) up-front control/rear up-front control; (2) multipurpose display system; (3) standby ALT indicator/rear standby ALT indicator; and (4) data transfer module (DTM).

Air data processors The ADPs are solid state digital computers. The left and right ADPs are rack-mounted in doors 10L and 10R respectively. The ADP uses pneumatic and electrical inputs to compute required outputs for using systems. Each processor also continuously monitors its own performance and provides ADP status signals to the applicable indicators. Figure 2–61 shows the air data processor.

Figure 2–61. Air data processor.

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Vertical speed indicators The VSIs receive an altitude rate signal from the right ADP.

AOA transmitters Both the left and right ADPs receive input from the left AOA transmitter. The right ADP also receives input from the right AOA transmitter.

Total temperature probes The left and right total temperature probes provide indicated total temperature signal to the left and right ADPs respectively.

Angle-of-attack indicator The ANGLE-OF-ATTACK indicator receives a varying voltage ratio from the left AOA transmitter and converts it into a corresponding pointer indication of local AOA. Optimum AOA and AOA validity are received from the right ADP.

Up-front control/rear up-front control The UFC and rear UFC display altitude and airspeed from the ADP with data 1 menu selected.

Multipurpose display system The multipurpose display system displays the same information as in ADC/AIC equipped aircraft.

Standby ALT indicator/rear standby ALT indicator The baro set value from the standby ALT indicator is sent to the right ADP to provide an indication of local barometric pressure. Data transfer module The DTM is a programmable, nonvolatile battery powered memory device used to transfer flight operations mission data to/from the ADCP. The DTM has both read and write functions. The read function allows for preprogramming of flight data. The write function allows mission collected flight data and failure data to be copied from the ADCP to the DTM. The DTM cartridge can be read on the computerized fault reporting system (CFRS) for use in debriefing in-flight malfunctions.

Principles of operation ADP principles of operation will be discussed in three separate categories: (1) ADP operation; (2) BIT operation; and (3) related systems interface.

ADP operation APD operation can be broken down into the following functions:

• Pressure sensors (pitot and static). • Power supply. • IDC section. • DP section. • ODC section. • CCDL. • Avionics 1553 MUX bus.

All inputs the ADP receives are conditioned and converted in the IDC section. The resulting digital data is transmitted to the DP section and there corrected for source (probe) errors. The required ADP outputs are processed and transferred to the ODC section. The ODC reformats the digital data from the DP to the correct output formats for the using system pressure sensors.

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Pressure sensors Each sensor module contains a pressure sensor, sensor driving electronics, sensor temperature sensing circuitry, and sensor calibration memory. The pressure sensor oscillator frequency and the analog temperature signal are digitized and sent to the digital processor in the ADP. The digital processor uses calibration constants stored in the sensor module to compute actual pressure.

Power supply The power supply converts 115 VAC, B phase, 400 Hz aircraft power into internally required system power.

Input data converter section The IDC provides the conversion and required interface between the air data sensors and the DP. All analog inputs are converted into digital data in the A/D converter. Discrete signals are stored in the input status register. The register output is multiplexed with the outputs of the F/D and A/D modules and transmitted to the DP.

Digital processor section The DP is a digital computer with an operating speed of 20 Hz. The DP receives digital inputs from the IDC and provides them with source error corrections. The DP monitors input, output, and intermodule signals between all the ADP sections. The DP contains the operational program memory, aircraft calibration memory, scratch pad memory, arithmetic element, and control logic. The control logic synchronizes the operation of the IDC and ODC with the DP.

Output data converter section The ODC conditions the ADP signals for interfacing with the indicators and other onboard systems. ODC operation is controlled by the DP. The output data is transferred from the DP to the ODC whenever a data transfer signal is received. The data control routes the data from the output data bus to the parallel to serial (P/S) converter shift register, digital to analog (D/A) converter, or discrete buffer storage registers as required. The data control also provides applicable output devices with required data processing signals. Serial digital data is provided for communication with the ADCP.

CCDL transmit and receive The purpose of the CCDL is to transmit free stream pitot and static electrical values from one ADP to the other. This allows the ADP to compute an averaged static value and to use the higher of the left and right side pitot values. Since the left and right side static lines (S2 ports) are not plumbed together, the static is averaged electrically by the ADP for the ADC, FBS, and PSA functions. By choosing the maximum value of the right or left pitot pressure, the ADP can compensate for a plugged pitot probe.

Avionics 1553 MUX bus operation Both the left and right ADP are connected to the avionics 1553 MUX bus 5A/5B but only the right ADP performs operational communications on the bus. The connection to the left ADP is used for operational flight program (OFP) loading only.

The ADCP uses a MUX bus test message to determine if communications can be established over the bus. If communications cannot be established, the ADCP directs the MPD/MPCD to display the A/D* legend on the MPD/MPCD. If bus communications is established and then lost, the ADCP will check the status of the right ADP no go discrete to AIU 1 to see if the ADP has failed. If the ADP has failed, the A/D* legend will be displayed. If the ADP has not failed, the ADP 5A and ADP 5B legends will be displayed indicating a bus failure to the ADP.

BIT operation The DP section performs most of the BIT function. The following paragraphs describe the types of BIT modes and stored BIT codes.

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Periodic mode The BIT function operates automatically when ADP power is applied and the BIT initiate mode is not activated. PBIT is internally activated and is made up of periodic tests interleaved with normal operations and does not interfere with equipment operation.

Initiated mode (ADC) IBIT is performed on the right ADP only. ADP IBIT is activated by pressing S5 (A/D/EXCS) on the MPD/MPCD displaying the BIT menu. During IBIT, the right ADP simulates static pressure, pitot pressure, AOA, and TT inputs. The simulated inputs are processed as normal inputs with the results transmitted as test indications. The static and pitot pressure sensor output signals are also compared for agreement within set limits. When the BIT initiate signal is removed, the right ADP reverts to periodic BIT mode. IBIT is provided for ADP on the ground only. It takes approximately 20 seconds for ADP IBIT to run.

ADP IBIT operation starts with a 28-VDC BIT initiate signal from AIU l. The ADP responds with a BIT acknowledge signal, which causes an A/D IN TEST legend to be displayed on the MPD/MPCD. When BIT is complete, the ADP removes the BIT acknowledge signal and the IN TEST legend is removed. If the ADP does not respond with a BIT acknowledge signal, the A/D IN TEST legend will not be displayed and the ADCP continues processing as if an IBIT was never requested.

During ADP IBIT, the ADP outputs test air data for observation by the maintenance crew on the HUD, EADI, and VSI. The HUD and EADI will both display 11,500 feet altitude and 250 knots airspeed. The EADI will also display 19.4 units of AOA. The VSI will display 500 ft/min up rate.

Initiated mode (PSA) When AFCS BIT is initiated, the FCC sends a BIT request signal to the ADPs. When the BIT request is received, the ADP sets the uncorrected pitot and static signals to a BIT value, to be read and interpreted by the FCC. If a failure is detected, the FCC causes an L/R ADP legend to be displayed on the AFCS detail maintenance BIT display. Initiated mode (EAIC) The EAIC function of the ADP is tested using the AIR INL CONT TEST SW located in doors 6L and 6R the same as in AIC equipped aircraft.

ADP stored BIT codes When the ADP BIT program tests a function, and that function fails, the ADP generates a BIT code and stores that code in nonvolatile memory. There are 318 possible BIT codes. If the same BIT code is generated on a subsequent test, it is not written into the BIT log a second time. This keeps the BIT log from being filled with codes all representing the same failure. The failure must be absent for six or more seconds before it is again entered in the BIT log. When the BIT log is full, no further BIT codes may be entered.

Related systems interface The ADP provides and receives digital, analog, or discrete signals for the ADCP, IFF transponder, engine air inlet system, ECS, and ICSCP, flight control system, landing gear, and the flap system.

The ADC, PSA and FBS functions all use the electrically averaged S2. The greater of left or right pitot pressure is used in the ADC and FBS functions. Using the pitot source that provides the greatest pressure will compensate for plugged probes and pitot system leaks. The EAIC function also uses the local side duct pitot and static pressures. The PSA function uses the electrically averaged S2 and the local side pitot pressure.

Backup altitude to the inertial navigation system (INS) During normal operation, the ADC function of the right ADP computes the pressure altitude signal by averaging the raw static pressure from both left and right ADP (via CCDL), and using local AOA. This pressure altitude is sent to the INS via the ADCP to stabilize the INS vertical loop. The ADP

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declares this signal invalid for any critical ADP fault, if both AOA inputs failed or static pressure failed. Both pressure altitude and validity is sent to the ADCP over the MUX bus to be used by the AFCS to calculate a backup pressure altitude.

The AFCS receives an uncorrected static pressure signal from the ADP. By comparing the uncorrected static pressure with the pressure altitude it receives over the MUX bus, the AFCS calculates a difference between the signals called a bias. It uses this bias to compute a backup pressure altitude from the uncorrected static pressure. This backup pressure altitude is not used as long as the ADP pressure altitude is good. If the ADP pressure altitude is declared invalid, the AFCS freezes the bias at the last known good value and uses this value to continue computing the backup pressure altitude. This backup pressure altitude is then used by the INS.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

221. Pitot-static system 1. Where are the pitot-static probes located on the F-15 aircraft?

2. What’s the purpose of the probe heaters?

3. What’s the purpose of the opening in the forward end of the pitot tube?

4. What’s the purpose of the hole on the bottom of the pitot tube?

5. Where are the system drains located?

6. Why are two types of pitot-static tubing used?

7. Explain how the pointers are moved in the standby altimeter.

8. What happens to the aneroid in the altimeter at higher altitudes?

9. What does the small pointer on the standby altimeter indicate?

10. When flying long distances at altitudes over 18,000 feet, what do pilots do with their baro settings to help avoid midair collisions?

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11. When the pilot descends below 18,000 feet at his or her destination, what baro setting is used?

12. How does the pilot adjust the altimeter so that it indicates zero on landing?

13. Why does the SAI have nonlinear markings on the dial?

14. What pressures are applied to the SAI?

15. When will the SAI pointer move?

222. F-15 A through D air data computer 1. What type of inputs does the ADC use to compute the outputs to the using system?

2. What output does the AOA transmitter provide?

3. When do the AOA probe heaters operate?

4. What displays are indicated on the altimeter?

5. What does the VSI indicate?

6. What inputs are provided to the IFF transponder by the ADC? Why are they used?

7. How does the ADC tie in with the ECS?

8. Where is the stall warning tone generated?

9. Describe the AOA stall warning.

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10. Why is the ADC integrated with the AFCS?

11. Why is the ADC integrated with the longitudinal flight control system?

12. What inputs are simulated by the ADC initiated BIT?

13. What LRU is used to initiate an ADC BIT?

14. What are the five distinctive sections of the ADC?

15. What section of the ADC performs data conversion and provides the required interface between the air data sensors and the DP?

16. What kind of signal is processed by the DP?

17. Which ADC section provides signal conditioning for associated system interface?

223. F-15E air data processor 1. The F-15E air data processor performs the same functions as which LRUs in non-ADP equipped

aircraft?

2. What functions does the right ADP perform?

3. Why is the DTM cartridge used?

4. Which section of the ADP provides the conversion and required interface between the air data sensors and the DP section of the ADP?

5. Which section of the ADP contains the operational program memory?

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6. Which ADP section conditions ADP signals for interfacing with the indicators and other onboard systems?

7. What is the purpose of the CCDL?

8. What is the advantage of choosing the maximum value of the right or left pitot pressure?

9. Which ADP performs operational communications on the 1553 MUX bus?

10. Which ADP section performs most of the BIT function?

11. Where does the ADP store BIT codes for failed test functions?

12. Under which conditions will the ADP declare the pressure altitude signal invalid?

Answers to Self-Test Questions 210 1. AC; positive and negative DC pulses. 2. Total internal fuel on the internal pointer; internal plus external fuel on the total pounds counter. The

indicator also displays selected individual tank amounts on the left and right pounds counters. 3. The left counter displays the fuel in fuselage tank 1 and the right counter reads zero. 4. The left counter displays the fuel in tanks 3A and 3B, while the right counter displays fuel in tank 2. 5. The BINGO indicator. 6. The indicator will monitor fuel consumption and give the pilot a BINGO FUEL light when the internal fuel

pointer reaches any preset valve. If fuel DUMP is selected, the indicator will shut off the fuel dump system when the desired fuel level is reached.

7. The signal conditioner. 8. On the signal conditioner. 9. The full indication to which you adjust the pointer and counters depends on the density of the fuel. 10. Placing the SELECTOR switch on the fuel quantity indicator to the BIT position. 11. Shorted wiring or a faulty tank unit. 12. Loose or faulty wiring, or an improperly installed tank unit.

211 1. RPM, FTIT, fuel flow, nozzle position, and oil pressure. 2. The window for the failed parameter will go blank. 3. ASP 61.

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4. 0–110 % in increments of 1%. 5. 200º to 1,400º C in 10º increments. 6. The ADCP. 7. It is routed directly from the transmitter to EMD. 8. 0–100 percent open. 9. The CENC. 10. ADCP. 11. When the oil pressure drops below 8 psi.

212 1. 1g. 2. CAU and CDU. 3. Aircraft vertical-axis acceleration forces (G-forces). 4. None, LRU is hermetically sealed. 5. Rt Main Landing Gear Wheel well. 6. 0.75V. 7. Under door 6R. 8. –2, –1, 0, 3, 4.5, 6, 7.5. 9. 6 G counter because it registers the maximum G units attained during any flight or maneuver. 10. Main pointer = Current G. 2, aux pointers = Max positive and negative.

213 1. Lateral control subsystem. 2. Ailerons and stabilators. 3. PRCA, the ARI and the lateral control stick damper. 4. Reduce stick oscillations. By using permanent bar magnets and a drag cup that rotates through the bar

magnet when the control stick is moved. 5. It receives control stick inputs and varies the output ratio depending on airspeed, longitudinal input, landing

gear position, and rate of change in yaw. 6. Roll ratio controller. 7. ARI. Shut down at 1 Mach by PRCA (also by antiskid). 8. Allows single aileron operation if linkage becomes jammed on other side. 9. If the PRCA is shutdown (ROLL RATIO switch is put to EMER) or if the rudder limiter system is activated

(at approximately 1.5 Mach) and the roll ratio airspeed scheduling valve has not shifted to minimum.

214 1. A method of controlling the aircraft around the lateral axis. 2. Pitch ratio indicator. 3. Longitudinal mass balance. 4. It varies the ratio between the control stick input and output to the control surfaces, depending on airspeed

and aircraft response. The PRCA also boosts the input from the control stick to reduce the force required by the pilot to move the control surfaces.

5. Automatically compensates for trim changes caused by accelerating to supersonic, flaps, speed brake, or stores separations.

6. Mixer assembly. 7. Emergency. 8. Fixed ratio bellcrank. 9. .4. 10. When the PRCA is shut down, if the ADC senses that the altitude is less than 20,000 feet, airspeed above

330 knots and pitch ratio is more than .9, or if the pitch ratio is less than .9 with the landing gear down.

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215 1. A method of controlling the aircraft around its vertical axis. 2. Push-pull cables. 3. It usually acts as a solid link, but spring tension in both directions allows rudder/brake movement to

provide nose wheel steering and CAS operation if there is a jam in the directional control subsystem mechanical linkage.

4. Under door 10R. 5. Pilot command, pitch and roll inputs, hydraulic pressure, and a flap signal. 6. It enables the ARI to increase the amount of rudder deflection for lateral stick movement. 7. Rudder limiter actuator. 8. The right AIC drives the rudder travel limiter actuator and the left AIC supplies the Mach signal for the

RUD LMTR light.

216 1. DSS section of the pitch computer. 2. To monitor the aircraft’s pitch, roll, and yaw response to pilot commands. 3. Monitors pitot and static pressures to reduce differential stabilator movement in roll CAS at high speed. 4. To allow the PRCA mechanical system to track the CAS so if the CAS fails, the mechanical system can

take over at the point of failure. 5. To provide the pitch computer with ram and servo valve position. 6. To update the CAS zero reference. 7. Pitot and static pressures to the dynamic pressure sensor. 8. At 1.5 Mach. This signal disables the turn coordination (CAS ARI) function of yaw CAS. 9. Yaw CAS ON logic signals are supplied to the roll/yaw computer’s engage logic starting the engagement

process. 10. YAW CAS switch placed in the ON position, rudder servo not failed, yaw rate less than 41.5° per second. 11. The yaw CAS engage logic energizes one transistor that 28 VDC to the left rudder servocylinder SOV. At

the same time, another transistor switch provides a ground to the right servocylinder SOV. The SOVs are wired in series, so the single 28 VDC signal and the single ground signal energize both SOVs.

12. Commanded input (rudder pedal displacement) and aircraft response. 13. YAW CAS will add rudder deflection to increase response. 14. Rudder actuator position LVDT signals. Channel A and B signals are monitored for large differences or

small long-standing differences. 15. Prevent all three axis of CAS from engaging if a stick force sensor is disconnected. 16. Automatically shut down the yaw CAS channel if the yaw rate exceeds 41.5° per second. 17. Roll CAS not failed, yaw CAS engaged, and stabilators engaged (not failed). 18. The pitch computer engages the shutoff solenoids of the stabilator servos. 19. Dynamic pressure and AOA. 20. Establish a zero reference for roll CAS. 21. Since the stabilator control circuits are in the pitch computer, a failure of the pitch channel also causes the

roll channel to also shut down. Roll CAS can be reset if the failure only affects the pitch CAS circuits and not the stabilator circuits.

22. The DSS not failed, yaw rate less than 41.5° per second, CASI not failed. 23. AOA, gear position, and flap position.

217 1. The pitch trim actuator is controlled through relays in the pitch computer, while the roll trim actuator is

controlled through relays in the roll/yaw computer. 2. It contains the relays which control the operation of the drive motor, plus two additional position LVDTs. 3. The aircraft control surfaces are at the TOT position.

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4. Roll/yaw computer. 5. Nose-up trim command and the nose-up/nose-down trim command. 6. LVDT A. 7. When the longitudinal or lateral feel trim actuator motor drives, it repositions the control stick. So any time

the trim system operates, the control stick is repositioned to correspond to surface displacement. 8. To compensate for a wing heavy condition. 9. Inside the roll/yaw computer. 10. Two logic low commands. 11. The trim relays are part of the directional feel trim actuator. 12. Yaw channel linkage, rudder pedals, and rudder trim LVDTs. 13. Rudders and ailerons at neutral, and the control stick at 1° aft which positions the trailing edge of both

stabilators 5° up. 14. Four. 15. When any of the three trim actuators are out of the TOT position, and the TOT button is pressed, the TOT

logic energizes trim relays to bring the trimmed system back to TOT. 16. Once the trim level detectors sense that the trim actuators are in the TOT position, the TOT logic circuits

stop the applicable trim relays from driving the trim motors.

218 1. 3. 2. TF COUPLE switch. 3. 3. 4. 2. 5. Provide a uniform response to pilot input regardless of changing in-flight conditions. 6. Channel C. 7. Gain to the functioning rudder actuator is doubled. 8. A BIT code is generated and one of the actuators is shut down. 9. The SFS roll command is fed through a deadband and gradient network in order to desensitize the roll

commands around the neutral point. 10. Fed to the pitch-roll servoloop, or DSS. 11. It is reduced. 12. The pilot is notified of the failure through displayed BIT codes. Pitch CAS operation continues with the

two remaining functional CAS channels. 13. It maintains three independent computers within the FCC; these allow continued CAS operation when a

failure exists. 14. Local and selected signals. 15. Switches or relay contacts. 16. A ground. 17. Allows the FCC to communicate with the other avionic components on the MIL STD 1553 MUX bus. 18. Turns off all data sources to reset them; then, selects channel A as the initial source. 19. AFCS IN TEST message to be displayed on the MPCD/MPD.

219 1. Two - one AIC for each inlet. They’re used to control the position of the inlet ramps and bypass door for

their respective engine. 2. Through the use of LVDTs. 3. A system failure could be critical to engine operation, so the EMERGENCY mode raises and locks all

ramps (except above 1.5 Mach) so the engine isn’t starved of air.

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4. Pitot-static pressure, total temperature, diffuser and first ramp actuator position, AOA, and a diffuser actuator lock/unlock signal from the air data computer.

5. Discrete (ground) from the ADC. 6. A digital Mach signal ranging from 0.15 to 2.80 Mach. 7. On the caution lights display panel and avionic status panel (fault indicator No.6 will be orange).

220 1. The system uses oblique shock waves created by the vari-ramps to reduce the speed of the supersonic

airflow before it gets to the engines. 2. 50 percent. 3. AOA, total temperature, and pitot static inputs. 4. The ramps and bypass door for that system will go to the full UP position once aircraft speed goes below

1.5 Mach. 5. They contain rate limiters. 6. Air loads force the first ramps and diffuser ramps up, and they automatically lock. The bypass door may or

may not close. 7. Each system has an AIC GROUND TEST switch which provides a preset test output from the AIC to the

system actuators. 8. The TTU–205D/F.

221 1. One on door 3R and one on door 3L. 2. To prevent icing of the probe. 3. Provides the means of gathering ram air for airspeed calculations. 4. To allow moisture to drain. 5. At the lowest point in the pitot-static lines. 6. Rigid tubing is normally used. Flexible tubing is used in areas of high vibration or where the line needs to

bend, such as a door hinge or at the point of connection to an LRU. 7. Static pressure is applied to an aneroid sensor which compresses or expands the aneroid. The movement of

the aneroid is mechanically connected, through a series of multiplying gears, to the pointers. 8. It expands. 9. Altitude × 10,000 feet. 10. The baro set on all altimeters are to be set to 29.92. 11. The local altimeter setting. 12. By using the observed station pressure as a baro setting. 13. To show greater accuracy at lower airspeeds. 14. Pitot and static pressure. 15. Only when there’s a change in impact pressure.

222 1. Pneumatic and electrical inputs. 2. Local AOA. 3. When weight is off the wheels. 4. Aircraft altitude in 20-foot increments and barometric corrections in inches of mercury. 5. The rate of climb or dive in thousands of feet per minute. 6. Altitude information encoded into mode C. Used for automatic altitude reporting. 7. The ADC provides Mach and airspeed discretes to control ram air and bleed air to the primary heat

exchanger. 8. The ICCP. 9. A 1600-Hz tone in the headset, interrupted at a 1 to 20 Hz rate.

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10. To provide a function of pilot relief called altitude hold. 11. To monitor pitch ratio operations and provide warnings to the pilot if a malfunction exist. 12. Static pressure, pitot pressure, AOA, barometric pressure, and total temperature inputs to verify system

operations. 13. BIT control panel. 14. (1) Pressure sensors.

(2) Power supply. (3) Input data converter. (4) Digital processor. (5) Output data converter.

15. IDC. 16. Digital data. 17. Output data converter.

223 1. ADC, EAIC, PSA, and FBS. 2. The functions of the ADC, the right EAIC, FBS, and a single PSA channel (A). 3. To transfer flight operations mission data to/from the ADCP and it is used by the CFRS for use in

debriefing in-flight malfunctions. 4. IDC section. 5. DP section. 6. ODC section. 7. Transmit free stream pitot and static electrical values from one ADP to the other. 8. Allows the ADP to compensate for a plugged pitot probe. 9. Right ADP. 10. DP section. 11. Nonvolatile memory. 12. Any critical ADP fault, if both AOA inputs failed or static pressure failed.

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Unit Review Exercises Note to Student: Consider all choices carefully, select the best answer to each question, and circle the corresponding letter. When you have completed all unit review exercises, transfer your answers to the Field-Scoring Answer Sheet.

Do not return your answer sheet to the Air Force Career Development Academy (AFCDA).

35. (210) Which F-15C, D, and E fuel quantity indicating system line replaceable unit (LRU) provides power to the tank units? a. Indicator. b. Thermistor. c. Tank aboard relays. d. Fuel quantity signal conditioner.

36. (210) On the F-15C/D/E aircraft, the fuel quantity system requires calibration after replacement of the a. external tank. b. signal conditioner. c. tank unit number 2. d. low-level control unit.

37. (211) Which F-15E fuel flow indicating system component provides a synchro signal to the engine monitor display (EMD) to drive the liquid crystal display (LCD) window FF PPH (fuel flow pounds per hour)? a. Fuel flow transmitter. b. Avionics interface unit no. 1. c. Engine diagnostic unit (EDU). d. Advanced display core processor (ADCP).

38. (211) Which F-15E nozzle position indicating system component drives the engine exhaust nozzle position transmitter through a mechanical interface? a. Engine diagnostic unit (EDU). b. Digital electronic engine control (DEEC). c. Advanced display core processor (ADCP). d. Convergent exhaust nozzle control (CENC).

39. (211) Which component determines if a low oil pressure condition exist on the F-15E? a. Engine diagnostic unit (EDU). b. Avionics interface unit (AIU) no. 1. c. Digital electronic engine control (DEEC). d. Advanced display core processor (ADCP).

40. (212) Which F-15 acceleration indicating/G exceedance system component is designed to provide a visual indication of −5 to +10 g-forces imposed on the cockpit of the aircraft during climbs, dives, and turns? a. Accelerometer. b. Counter display unit. c. Acceleration indicator. d. Counter accelerometer unit.

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41. (213) Which F-15 component increases and decreases lateral control based on variations in airspeed, longitudinal position, landing gear position, and yaw rate? a. Roll ratio controller. b. Mode select assembly. c. Lateral feel trim actuator. d. Aileron rudder interconnect.

42. (213) Which F-15 component is used to control the flow of hydraulic fluid to the pitch and roll channel assembly (PRCA)? a. Roll ratio controller. b. Mode select assembly. c. Lateral feel trim actuator. d. Aileron rudder interconnect.

43. (214) Which F-15 longitudinal flight control subsystem component is used to help neutralize the effect of acceleration on the control stick during maneuvers? a. Pitch ratio switch. b. Longitudinal mass balance. c. Pitch and roll channel assembly. d. Longitudinal feel trim actuator (FTA).

44. (214) Which F-15 longitudinal flight control subsystem component has built-in centering springs which resist stick movement giving the pilot a stick feel force? a. PITCH RATIO switch. b. Pitch ratio controller. c. Longitudinal mass balance. d. Longitudinal feel trim actuator (LFTA).

45. (214) Which F-15 longitudinal flight control subsystem component contains an airspeed scheduling valve, a Mach factor computer, and an aileron rudder interconnect (ARI) Mach valve? a. PITCH RATIO switch. b. Pitch ratio controller. c. Longitudinal mass balance. d. Longitudinal feel trim actuator (FTA).

46. (214) Which F-15 longitudinal flight control subsystem component automatically compensates for trim changes caused by accelerating from subsonic to supersonic flight? a. Mechanical mixer. b. Pitch trim controller (PTC). c. Pitch ratio controller (PRC). d. Longitudinal feel trim actuator (FTA).

47. (215) Which F-15 directional flight control subsystem component contains centering springs, providing an artificial feel used to keep the rudder pedals centered? a. Rudder actuator. b. Rudder limiter actuator. c. Directional feel trim actuator. d. Directional control breakout assembly.

48. (215) Which F-15 directional flight control subsystem component transmits the input force from the aileron rudder interconnect (ARI) into two separate push-pull cables? a. Rudder actuator. b. Rudder limiter actuator. c. Directional feel trim actuator. d. Rudder control breakout assembly.

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49. (215) The F-15 rudder limiter actuator is used to prevent commanded excessive rudder travel when the aircraft is traveling above what Mach? a. 1. b. 2. c. 1.5. d. 2.5.

50. (216) What F-15 component provides pitch rate signals proportional to angular motion about the pitch axis to the pitch computer? a. Stick force sensor. b. Rate gyro assembly. c. Accelerometer assembly. d. Control surface actuator linear variable differential transducer (LVDT).

51. (216) How many accelerometers are in the F-15A through D automatic flight control system (AFCS) accelerometer assembly? a. 1. b. 3. c. 4 d. 6.

52. (216) The F-15 #1 and #2 modular relay panels provide logic signals to the pitch computer for changing conditions such as a. CAS engage, weight off wheels, and angle of attach (AOA) signals. b. CAS engage, speed brake, and flap down signals. c. Landing gear and flap position and weight off wheels. d. Gear position, speed brake, and AOA signals.

53. (216) Which is required for engaging the F-15 roll control augmentation system? a. Yaw CAS engaged. b. Pitch CAS engaged. c. Yaw CAS disengaged. d. Pitch CAS disengaged.

54. (216) On an F-15, how much stabilator deflection can the pitch control augmentation system (CAS) add to or subtract from the manual flight controls? a. ± 5° differential movement. b. ± 5° symmetrical movement. c. ± 10° differential movement. d. ± 10° symmetrical movement.

55. (217) What supplies the F-15 directional feel trim actuator position to the limit level detectors to prevent the actuator from driving beyond its limit? a. Linear variable differential transducer (LVDT) A. b. RUDDER STOP SENSOR switch. c. Rudder limit transducer. d. Rudder synchro.

56. (218) Which F-15E line replaceable unit (LRU) is basically a combination of pitch and roll/yaw computers? a. Avionics interface unit (AIU). b. Flight control computer (FCC). c. Central computer complex (CC). d. Multipurpose display processor (MPDP).

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57. (218) What is the maximum amount of rudder travel the F-15E yaw control augmentation system (CAS) can add or subtract? a. 5°. b. 10°. c. 15°. d. 30°.

58. (218) What yaw rate in degrees per second will cause the F-15E pitch control augmentation system (CAS) to disengage? a. 41.5. b. 45.1. c. 51.4. d. 54.1.

59. (218) What angle-of-attack (AOA) in degrees is the pitch control augmentation system (CAS) gain reduced to implement the stall inhibit function? a. 9. b. 13. c. 17. d. 21.

60. (218) The F-15E flight control computer (FCC) produces the output discrete signals to control what? a. Roll engage command. b. Pitch engage command. c. Trim actuator position signals. d. Automatic speed brake retract.

61. (219) How many variable inlet ramps are there in the F-15 air intake system? a. 2. b. 3. c. 4. d. 5.

62. (219) Which F-15 ramps, doors, or actuators are hydraulically powered and controlled by the air inlet controllers (AIC)? a. First, diffuser, and bypass. b. Second, third, and diffuser. c. Diffuser, bypass, and third. d. First, second, and third ramps.

63. (219) At what airspeed does the F-15’s air data computer (ADC) unlock the diffuser ramp? a. 0.80 Mach. b. 0.84 Mach. c. 1.0 Mach. d. 1.5 Mach.

64. (220) The F-15 diffuser ramp actuator’s Mach blocking valve solenoid prevents the actuator from retracting when aircraft speed is above Mach a. 1.0. b. 1.2. c. 1.4. d. 1.5.

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65. (221) The F-15 pitot-static probes are located on the a. radome. b. inlet duct walls. c. doors 6L and 6R. d. doors 3L and 3R.

66. (221) The additional hole in the bottom of the F-15 pitot-static probes is designed to a. drain moisture. b. relieve pressure. c. sense static pressure. d. sense impact pressure.

67. (221) What is the standard barometric setting in inches of mercury (Hg) used above altitudes of 18,000 feet? a. 2.92. b. 29.2. c. 22.92. d. 29.92.

68. (221) The altimeter setting is defined as a setting on the altimeter to indicate a. altitude above sea level. b. altitude above the airfield. c. a standard pressure setting, worldwide. d. a standard pressure setting, below sea level.

69. (221) The pointer of the F-15 standby airspeed indicator moves when what pressure is present? a. Stall. b. Pitot. c. Static. d. Impact.

70. (222) Which is not part of the F-15 air data computer (ADC) system? a. Altimeter. b. Head-up display (HUD). c. Vertical speed indicator (VSI). d. Airspeed Mach indicator (AMI).

71. (222) What type inputs are provided to the F-15 air data computer (ADC)? a. Electrical only. b. Electrical and pneumatic. c. Electrical, pneumatic, and mechanical. d. Electrical, pneumatic, mechanical, and hydraulic.

72. (222) On an F-15, operation of the angle-of-attack (AOA) probe heaters is controlled by the a. thermostat in the probe. b. thermocouple in the air data computer (ADC). c. main landing gear weight-off-wheels (MLG WOW) switches. d. pitot heat switch on the environmental control system (ECS) control panel.

73. (222) On an F-15, what information does the air data computer (ADC) provide to the identify friend-or-foe (IFF) transponder? a. Mach number. b. Mode 1 interrogation code. c. Digital altitude information. d. Aircraft identification number.

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74. (222) On an F-15, the operation of the air data computer (ADC) is divided into a. three sections: input data converter (IDC), digital processor (DP), output data processor (ODC). b. five sections: IDC, DP, ODC, pitot and static pressure sensor inputs, and power supply. c. six sections: IDC, DP, ODC, power supply, calibrator, and built-in test (BIT). d. four sections: IDC, DP, ODC, pitot and static pressure sensor inputs.

75. (223) How are the left and right F-15E air data processors mounted and where are they located? a. Hard mounted in doors 6L and 6R. b. Hard mounted in doors 10L and 6R. c. Rack mounted in doors 3L and 10R. d. Rack mounted in doors 10L and 10R.

76. (223) Which section of the F-15E air data processor (ADP) conditions and converts all received inputs to the ADP? a. Digital processor (DP). b. Input data converter (IDC). c. Output data converter (ODC). d. Electronic air inlet controller (EAIC).

77. (223) Which section on the F-15E air data processor (ADP) conditions the ADP signals for interfacing with the indicators and other onboard systems? a. Digital processor. b. Output data converter. c. Cross channel data link. d. Electronic air inlet controller.

78. (223) Why is the 1553 multiplex (MUX) connection to the left F-15E air data processor (ADP) used? a. As backup for the right ADP. b. For operational flight program (OFP) loading only. c. For data transfer module (DTM) information transfer. d. For line replaceable unit (LRU) status for the central computer (CC).

79. (223) Which F-15E air data processor section performs most of the built-in test (BIT) function? a. Digital processor (DP). b. Input/output (I/O) section. c. Input data converter (IDC). d. Electronic air inlet controller (EAIC).

Please read the unit menu for unit 3 and continue

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Student Notes

Unit 3. Communications/Navigation/Penetration Aids Systems Page 3–1. Data Link Systems .................................................................................................................... 3–1

224. Fighter data link ................................................................................................................................. 3–1 225. F-15E avionics interface unit ........................................................................................................... 3–15

3–2. Electronic Warfare Systems .................................................................................................. 3–20 226. Air-to-air interrogator ...................................................................................................................... 3–20 227. F-15C AN/ALQ-135 internal countermeasures set components ..................................................... 3–24 228. F-15C internal countermeasures set operation ................................................................................. 3–29

3–3. PODS ....................................................................................................................................... 3–35 229. F-15E LANTIRN navigation pod .................................................................................................... 3–35 230. F-15E data link pod ......................................................................................................................... 3–45

OMMUNICATIONS/NAVIGATION/PENETRATION AID SYSTEMS provide the pilot with a myriad of useful information. Ensuring proper operation of these various systems is a job you are tasked to do and, like all aircraft maintenance, should be taken seriously. In this unit, you’ll

learn about the F-15 data link system along with the F-15E avionics interface unit operation. Then you will learn about the electronic warfare systems used on the F-15 aircraft. The unit will finish off by giving you a look at the F-15E low altitude navigation targeting infrared for night (LANTIRN) navigation (NAV) and data link pod.

3–1. Data Link Systems The data link system provides the pilots with all sorts of vital information that enables them to communicate and share this information during the fight. Data link systems use aircraft integration like the avionics interface unit (AIU) and other systems to comprise the information required to make the mission a success. As you will see, the data link system is much like a home computer and the Internet. We will discuss how the fighter data link system operates in this way and will conclude with a lesson on the AIU operation.

224. Fighter data link The fighter data link (FDL) is a digital data link system that allows aircraft to exchange and display real-time tactical information. A secure, jam-resistant system operates in the L band radio frequency range. FDL functions as a terminal in the Link-16 network (fig. 3–1).

FDL and Link-16 can be compared to the Internet. Think about a home computer and the Internet for a moment. The home computer is similar to a single F-15 FDL system while the Internet is similar to the Link-16 network. From that single home computer, you can access the information on any of the thousands of computers connected to the Internet. Likewise, one FDL terminal can access multiple FDL systems information operating on Link-16 and see the information gathered from those FDL systems. Using FDL, pilots can transfer or display information with members in their own flights, airborne warning and control system (AWACS), or joint surveillance and target attack radar system (J-STARS) command and control aircraft, sensor aircraft, ground units or even ships equipped with the FDL capability. The major difference between the home computer/Internet scenario and FDL/Link-16 is that FDL is a line-of-sight system with a limited range.

C

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Figure 3–1. Flight data link/Link-16 network.

System description FDL is a data link system that transfers information at high rates. The data is encrypted to provide security, and is jam resistant to yield high reliability in a hostile radio frequency (RF) environment. It also provides the ability to interconnect many widespread sources and users of information.

The FDL system transmitter allows effective use of the system for up to 100 nautical miles (NMI) for fighter-to-fighter aircraft operations and up to 200 NMI for fighter-to-command and control aircraft operations. The effective range can be increased by using relay platforms between participants that are not within line of sight.

System components FDL components include a JTIDS (joint tactical information distribution system) mode control panel, sensor control panel, FDL radio/transmitter (RT), FDL remote power supply, and batteries. The following paragraphs describe these components.

F-15C/D joint tactical information distribution system mode control panel The FDL system on the F-15C/D is operated by using the JTIDS mode control panel (MCP). The JTIDS MCP is located in the cockpit on the left console (fig. 3–2). The JTIDS MCP contains the controls used for FDL operation. These include FDL operating mode, battery support mode, mission channel selection, FDL crypto code zeroizing, and the TACAN (tactical air navigation) /FDL master reset.

Figure 3–2. Joint tactical information distribution system mode control panel.

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F-15E sensor control panel The FDL system on the F-15E is operated by using the sensor control panel. The sensor control panel is located in the cockpit on the left console (fig. 3–3). The JTIDS mode switch controls power and mode selection for FDL.

Figure 3–3. Sensor control panel.

Flight data link receiver/transmitter The FDL RT is located in door 3R (fig. 3–4). It contains the receiver, transmitter, processor, antenna control, and interface circuits for data link operations. Also, all TACAN signals are passed through the FDL RT.

Figure 3–4. Flight data link RT.

Flight data link remote power supply The FDL remote power supply (RPS) is located in door 3R (fig. 3–5). It provides operating power for the FDL RT and TACAN RT. It also provides data shift clocks to the FDL RT for TACAN interface and data shift clocks and logic power to the instrument landing system (ILS)/TCN (TACAN) control panel.

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Figure 3–5. Flight data link remote power supply.

Batteries The battery compartment is located in the FDL RT (fig. 3–6). It holds three 3.0 VDC (volts direct current) lithium batteries. The batteries provide enough power to store the crypto codes and run the internal clock when system power is off (hold mode).

Figure 3–6. Flight data link RT battery compartment.

Related system description (non-model specific components) Related system components include the upper TACAN antenna, TACAN/very high frequency (VHF)/ultra high frequency (UHF) antenna, control stick grip, right throttle grip, interface blanker system, data transfer module (DTM), and the avionics status panel (ASP). The following describes these non-model specific components.

Upper tactical air navigation antenna The upper TACAN antenna is located aft of the canopy on door 30 (fig. 3–7). This is a flush-mounted antenna that transmits and receives RF in the L band for both the FDL and TACAN systems.

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Figure 3–7. Upper tactical air navigation antenna.

Lower tactical air navigation antenna The lower TACAN antenna is located on the lower fuselage near the radome (fig. 3–8). This is a dual element blade antenna that transmits and receives RF for the UHF and TACAN/FDL systems.

Figure 3–8. Lower tactical air navigation antenna.

Control stick grip The control stick is located in the cockpit (fig. 3–9). It contains the castle switch which is used for FDL control.

Figure 3–9. Control stick.

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Right throttle grip The right throttle grip is located on the forward cockpit left console (fig. 3–10). The target designator control (TDC) switch is mounted on the right throttle grip, and controls the acquisition symbol on the FDL display. The TDC switch acts like the mouse pad on a laptop computer.

Figure 3–10. Right throttle target designator control switch.

Interference blanker system The interference blanker is located in door 3R. It prevents interference between active aircraft RF transmitters and receivers. The FDL RT provides a blanking pulse to the blanker to inhibit other receivers when the FDL transmits and receives a blanking pulse to inhibit the FDL receiver when other systems transmit.

Data transfer module The DTM (fig. 3–11) is a programmable battery powered memory device. It is similar to a removable flash drive or memory card for a computer. There are two types of DTMs that are used for FDL. One is the operations (OPS) DTM used by the pilots to load mission data to the central computer (CC)/advanced display core processor (ADCP) and FDL. The other is the MAINT (maintenance) DTM which is used to transfer built-in test (BIT) data to the computerized fault reporting system (CFRS) for both debrief and computerized fault isolation.

Figure 3–11. Data transfer module and aircraft receptacle.

Avionics status panel The ASP is located in the nose wheel well. The ASP shows the status of the avionic systems. Indicator 3 includes systems that are located under door 3R such as FDL. When any avionic system under door 3R fails, the ASP indicator 3 will turn orange.

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F-15C/D model specific components Now let’s discuss the F-15C/D model specific FDL components.

Programmable signal data processor The programmable signal data processor (PSDP) is located in the avionics equipment bay behind access door 6L. It communicates with the CC over the 1553 multiplex (MUX) bus, channels 8A and 8B. This communication includes interface for the FDL pushbutton controls (on the multipurpose color display (MPCD)) to the CC as well as communication from the CC to the integrated communications control panel (ICCP) for audio FDL messages and threat voice warnings. The PSDP provides an interface path for the DTM to and from the CC for FDL initialization data and flight maintenance data. The PSDP provides interface for the stick grip castle switch control of situation displays.

The PSDP also provides the majority of the information that is displayed on the MPCD. It generates the FDL display symbology, provides display drives for the situation and FDL mode displays on the MPCD. The PSDP receives direction from the CC as to the positioning of FDL symbology.

Central computer The CC is a digital computer that provides most of the processing requirements and is a distribution center for information from many systems in the aircraft. Nearly all of the processing related to FDL is contained within the CC. The CC controls the data interchange with the FDL terminal on the 1553 MUX bus, channels 7A and 7B.

Multipurpose color display The MPCD is located on the cockpit, main instrument panel. The MPCD displays include armament, situation FDL mode, DTM and BIT displays.

Built-in test control panel The BIT control panel (BCP) is located in the cockpit on the left console. The BCP provides indications for both continuous monitor and initiated BIT (IBIT). The FDL system is tested by selecting JTIDS with the system select switch and pressing INITIATE. The JTIDS light indicates a malfunction with the FDL units. If both TACAN and FDL are powered on, initiating BIT on either system will run BIT on both systems.

Integrated communications control panel The ICCP is located in the cockpit on the left console. There is an additional ICCP in the rear cockpit of F-15Ds. The ICCP supplies FDL voice warnings to the aircrew. Volume control is accomplished by a knob on the ICCP.

F-15E specific components The specific F-15E FDL components include the ADCP, MPDs and MPCDs, AIU 1 and 2, left and right hand controllers, remote intercommunication control panel (RICP), and the intercommunication set control panel (ICSCP). The following paragraphs will provide a clearer description of each component.

Advanced display core processor The ADCP, in door 3L, is a digital computer that provides most of the processing requirements for FDL and is a distribution center for information from many systems in the aircraft. Information obtained by the ADCP for FDL processing is passed to the display function for display. Nearly all of the processing related to FDL is contained within the ADCP. The ADCP is the bus controller for the data interchange with the FDL RT on the 1553 MUX bus, channels 7A and 7B.

Multipurpose display/multipurpose color display There is a combination of seven multipurpose displays and multipurpose color displays in the F-15E. The front cockpit contains two MPDs and one MPCD. The rear cockpit contains two MPDs and two

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MPCDs. The seven displays provide the operator with situation displays, a JTIDS display for FDL control, and BIT displays for testing FDL.

Avionics interface units 1 and 2 The AIUs are located in the avionics equipment bays. AIU 1 is located behind access door 6L while AIU 2 is located behind access door 3R (fig. 3–12). The AIUs provide an interface from the up-front controller (UFC) and hands on throttle and stick (HOTAS) to the ADCP for control of aircraft systems. HOTAS controls for the FDL include the control stick grip castle switch and auto acquisition (ACQ) switch, throttle control TDC switch, and hand controller switches. AIU 1 also receives an FDL line replaceable unit (LRU) fail discrete that drives ASP fault indicator 3 and the AV BIT fail light.

Figure 3–12. Avionics interface unit set locations.

Left and right hand controllers The left and right hand controllers are located in the rear cockpit (fig. 3–13). The hand controllers provide control of the acquisition symbol using the TDC switches.

Figure 3–13. Left and right hand controllers.

Remote intercommunication control panel The RICP is located in the front cockpit on the left console. It contains the CRYPTO switch which may be used to zeroize FDL crypto codes.

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Intercommunications set control panel The ICSCP is located in the rear cockpit on the left console. It produces the message voice warning for FDL and has a CRYPTO switch which may be used to zero FDL crypto codes.

System operation The FDL system receives power from the right bus circuit breaker panel. When power is applied to the FDL RPS, the POWER APPLIED light on the FDL RPS illuminates.

Placing the JTIDS MCP MODE switch (F-15C/D) or SENSOR CONTROL PANEL MODE switch (F-15E) in the NORM, or SIL position causes the FDL RPS to route a ground signal to the FDL RT. This signal then returns to the FDL RPS which turns on the FDL system. The POWER ON light on the FDL RT will illuminate when the FDL is turned on.

F-15C/D joint tactical information distribution system mode control panel The following table describes the F-15C/D JTIDS mode control panel control/indicator and functions (fig. 3–14).

Figure 3–14. F-15C/D joint tactical information distribution system mode control panel.

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F-15C/D JTIDS Mode Control Panel Control Control/Indicator Position Function

MASTER RESET Switch Press and Release

When pressed, power is removed from the FDL. At the same time, battery power is applied to the FDL RT to maintain initialization data, crypto variables, and existing operating modes during reset. When the MASTER RESET switch is released, power is restored, causing both systems to restart.

CIPHER Switch Zero

All crypto variables within the FDL are zeroed. The variables cannot be zeroed if the MCP MODE selector is in HOLD.

NORM Normal operating position. MISSION CHAN Selector Turn Dial (000 to 127) Selects any one of 127 different subnets (controller

operating channels 000 to 126).

MODE Selector

(PULL) OFF All power is removed. All data, including crypto codes, in the FDL is lost.

POLL Not used on FDL. SIL All transmissions are prohibited.

NORM Normal operating position. Allows full participation in the Link-16 community. Provides full transmit and receive ability.

HOLD

Battery power allows retention of the initialization data and crypto variables. In this position, the crypto variables cannot be zeroed using the CIPHER switch action. To load crypto variables, the MODE switch must be in HOLD or OFF.

F-15E sensor control panel The following table describes the F-15E sensor control panel control/indicator and functions relating to the FDL system (fig. 3–15).

Figure 3–15. F-15E Sensor control panel.

F-15E Sensor control panel Control/Indicator Position Function

MODE Switch

(PULL) OFF All power is removed. All data, including crypto codes, in the FDL is lost.

POLL Not used by FDL. SIL All transmissions are prohibited.

NORM Normal operating position. Allows full participation in the Link-16 community. Provides full transmit and receive ability.

HOLD Battery power allows retention of the initialization data and crypto variables with the MODE switch in the HOLD position.

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F-15E RICP/ICSCP The following table describes the F-15E RICP/ICSCP control/indicator and functions relating to the FDL system (fig. 3–16).

Figure 3–16. F-15E RICP/ICSCP.

F-15E RICP/ICSCP Control/Indicator Position Function

CRYPTO Switch HOLD Not used for FDL. NORM Normal operations. ZERO Zeros all crypto memory locations in FDL.

Menu displays The FDL displays are very similar throughout all models of F-15s. FDL is controlled by accessing a series of menus (fig. 3–17) from the main menu. The MPD/MPCD multifunction switches are used to access menu displays used with FDL which are described in the below table.

FDL Menu Displays

Menu Description

BIT Selects the BIT display for monitoring JTIDS testing.

SIT The situation display (SIT) is the main screen used with FDL. The SIT screen presents situational data, giving the pilot a broad view of what is going on around the aircraft.

DTM Control the exchange of information between the DTM and the FDL system.

FLIGHT DATA FLIGHT DATA calls up information which is used to monitor other members of the flight.

OWN DATA The OWN DATA display allows the aircrew to set up or modify the DTM data for their own aircraft. This data is normally loaded by the DTM during initialization.

RESET IPF

FDL shares frequency spectrum with multiple other avionics systems. To ensure FDL is not interfering with more flight essential avionics the FDL RT has a built-in Interference Protection Feature (IPF) monitor. This monitor circuit ensures that the FDL is not interfering with other systems in the same frequency band.

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FDL Menu Displays

Menu Description

ENTER NET

Provides the status of FDL’s attempt to logon to the Link-16 network. It can either be blank, or display PENDING, COARSE, or FINE. When the display after ENTER NET is blank, it indicates that no status is reported and the FDL is not logged onto the Link-16 network. PENDING indicates that time synchronization is in progress. COARSE indicates that the FDL has refined the time well enough to receive messages, but is still unable to transmit. FINE indicates that time synchronization is complete and the FDL can both send and receive messages.

Figure 3–17. Flight data link menu access displays.

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Flight data link radio frequency characteristics FDL operates using pulsed transmissions in the Lx band. The Lx band itself is a portion of the larger UHF. The selection of this wide band of operation complicates enemy jamming attempts because of huge bandwidth coverage and enormous power required to override the Link-16 RF.

Accessing the Link-16 network We have previously established that in order for FDL to transmit and receive all its required information along the Link-16 network, it needs to establish a FINE synchronization with the network. Next, we are going to break down exactly what has to happen for FINE synchronization to occur.

For starters, because of the classified information transmitting along the Line-16 network, FDL needs to ensure all its information is secure. In order to keep the transmission secure and jam resistant, the information is spread out over a wide frequency spectrum. The specific frequency in use needs to be constantly changing, using a frequency-hopping pattern. Imagine listening to a fast song on the radio. Every fraction of a second, the station changes and you have to go to the next station for the next part of the song. The radio is essentially hopping frequencies. FDL automatically hops frequencies 76,923 times a second.

The FDL RT will hop around on the 51 possible FDL channels in a pattern controlled by the crypto code key loaded in the FDL system. A crypto code key is a classified set of electronic instructions. This key determines the frequency hopping sequences.

It is possible to have more than one group of operators use the system without interfering with each other. This is accomplished by changing the channel (mission channel or fighter channel). The channel settings are used to select the network that the FDL will operate on. The fighter channel selects a fighter-to-fighter network and a mission channel selects the command and control network. They can be the same network. The selection does not select a specific frequency, but gives an indicator so the FDL software will know which code keying pattern and which of the 51 channels to use at the current time to communicate in the selected network. Theoretically, up to 20 different FDL systems can work together with common keys. The channel determines a starting point so none of the 20 different systems interrupts each other.

This massive amount of data constantly changing frequencies is possible because FDL architecture is based on concept called time division multiple access (TDMA). TDMA is where small intervals of time are allocated to each FDL user for transmission of digital data. TDMA theory can get overly complex for a technician’s required level of understanding. Figure 3–18 is a detailed breakdown of FDL’s time division slots. What to take away from figure 3–18 is that accurate time is extremely important to proper FDL operation. In order for pilots to transmit and receive data they need to know what time slot they are allocated.

To enter onto the Link-16 net the pilot needs to load FDL TERM data to the CC/ADCP. TERM data can be downloaded to the aircraft via the DTM or mission cart. Most importantly, term data contains the pilot’s time slot allocation.

Once FDL term data is loaded the pilot needs access the Link-16 network. This is done by choosing one of the terminals participating in the network to act as the net time reference (NTR). The NTR is the first terminal, which is transmitting on the network, and its clock establishes system time.

All other network participants enter the network via the NTR. After the NTR is activated, the users synchronize their clocks with the NTR clock. Think of the NTR as the person hosting a game on X-box live. X-box live is the Link-16 network. The people wanting to play need to access the settings determined by the person hosting the game. Just like X-box live, when the host (or NTR in the case of FDL) leaves, all other users get kicked off the game.

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Figure 3–18. Flight data link time division slots.

To fully complete fine synchronization and decode the data in the Link-16 messages, the FDL RT must also have valid crypto codes. These are the same crypto codes used to determine the frequency hopping sequence. Four pairs of crypto codes can be loaded into the FDL RT. Each pair allows operation in a different network. Pairs are loaded to allow operation during the current day and the next day, which allows operation after midnight Zulu time.

To summarize, in order for a group of pilots to sync up, achieve FINE synchronization and fully communicate along the Link-16 network, they need 5 things:

1. Properly loaded FDL TERM data. 2. A common channel number (either mission or fighter channel). 3. A host to establish the NTR. 4. Time sync with the NTRs established time. 5. Common crypto codes.

Built-in test There are three types of BITs used with FDL: startup, operational (continuous), and initiated.

Failure indications As with any complicated system, faults will occur. When that happens in the FDL system, there are several ways to identify those faults. If a fault exists, ASP 3 will indicate that an LRU has failed under door 3R. On an F-15C/D, the JTIDS light on the BCP will illuminate. Additionally, there will be a failure message displayed on the MPCD, which will alert the operator of system problems. When the FDL system fails, these failures are recorded to the DTM.

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225. F-15E avionics interface unit The avionics interface unit set consists of two LRUs: AIU 1 and AIU 2. The AIUs set controls, processes, and routes interfacing signals between multiple aircraft systems. In this lesson, we will discuss exactly what makes up the AIU system and how the AIUs accomplish the routing of interfacing signals.

System and related components The avionics interface unit set is made up of two components: AIU 1 and AIU 2.

Avionics interface unit number 1 AIU 1 is rack-mounted and located in back of door 6L (fig. 3–19), and possesses two different processing channels (AIU1A and AIU1B) within the AIU itself.

Figure 3–19. Avionics interface unit no. 1.

Avionics interface unit number 2 AIU 2 is rack-mounted and located in back of door 3R (fig. 3–20). It possesses a single processing channel and serves as a back-up for AIU 1 for processing mission critical system data or those systems that directly impact the safety of flight.

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Figure 3–20. Avionics interface unit no. 2.

Principles of operation The following paragraphs describe the AIU principles of operation.

Communication with advanced core display processor One of the primary LRUs that the AIU set communicates with is the ADCP. The AIU and ADCP communicate by way of the 1553 MUX bus 5A/5B. Data transferred between the AIU set and the ADCP includes BIT data, up-front display and control data, aircraft systems discrete, mode, control, and status data, cautions, warnings, and advisories, UHF and identification friend-or-foe (IFF) initialization, as well as memory inspect data.

Communication with up-front control The UFC serial data bus provides communication between the UFC and the AIU set, as well as between AIU 1 and AIU 2. The AIU1A channel within AIU 1 functions as the UFC’s primary bus controller; however AIU 2 has the task of backup UFC bus controller, in case the AIU1A channel fails.

The UFC sends switch keyboard, panel controls data, and status to the AIU set on the UFC serial data bus. The AIU set sends menu and display data to the UFC. The AIU 1 and AIU 2 communicate data and status with each other by way of the UFC serial data bus.

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Aircraft systems serial data bus The AIU set interfaces with other aircraft systems by way of serial data. The aircraft systems, interfacing AIU, and signal function are listed as follows:

Ultra high frequency receiver-transmitter no. 1 UHF RT no. 1 (UHF radio 1) controls data from the UFC sent to the AIU1A channel. This data is made up of frequency and mode control. The radio 1 data is sent to the UHF COMM RT number 1 through the serial interface in the AIU1A channel, as well as to the ADCP by way of avionics 1553 MUX bus. UHF radio 1 frequency and mode data are stored in the AIU1A channel memory in case of power loss or failure.

Ultra high frequency receiver-transmitter no. 2 Serial data transfer for UHF RT no. 2 (UHF radio 2) data is the same as explained above, except that AIU 2 controls the data transfer.

Automatic direction finder control amplifier The automatic direction finder (ADF) sends serial ADF bearing data to the AIU1A channel. The AIU1A channel formats this data and outputs it to the ADCP on the avionics 1553 MUX bus. AIU1A sends the ADF a continuous clock and data initiate signal when AIU1A requires serial ADF bearing data.

Tactical air navigation receiver-transmitter TACAN control data from the UFC is sent to the AIU1A channel. This data is made up of channel and mode control. The AIU1A channel routes the TACAN data to the TACAN RT. The TACAN RT outputs serial control word, range and bearing data to the AIU1A channel. The TACAN data is sent to the ADCP by the AIU1A channel on the avionics 1553 MUX bus. The AIU1A channel sends the TACAN a continuous clock and data initiate signal when AIU1A requires serial TACAN data.

Combined radar altimeter receiver-transmitter The combined radar altimeter (CARA) RT routes serial radar altitude and altimeter status to the AIU 1 and AIU 2. The AIU 1 is the primary processor for this data with AIU 2 serving as a backup. CARA data is handled the same way by the AIU 1 or AIU 2. The AIU 1 processes the CARA serial data and then places the information on the avionics 1553 MUX bus for use by the ADCP.

Engine monitor display The AIU 2 receives serial data from the engine monitor display (EMD). This serial data is reformatted by AIU 2 and sent to the ADCP by way of avionics 1553 MUX bus. Serial data from the EMD is made up of EMD shop replaceable unit (SRU) status. Serial data from the EMD is sent to AIU 2 after the EMD receives a BIT initiate discrete signal from AIU 1.

Discrete inputs/outputs The AIU interfaces many aircraft systems using discrete inputs and outputs. These are usually two state signals (open or ground, open or 28 VDC, and 28 VDC or ground). Discrete inputs are processed by the AIU set and sent in serial format to the ADCP by way of the avionics 1553 MUX bus. Discrete outputs are enabled by the AIU set as a result of one of the following inputs:

• Avionics 1553 MUX bus data from the ADCP. • Serial data from the UFC and rear UFC. • Input from aircraft systems.

Analog interface The AIU1B channel and AIU 2 receive analog inputs from several aircraft systems. The AIU1B channel receives analog inputs from the ILS receiver, cockpit throttle grips, and left-hand controller. Synchro inputs are received from the left and right engine fuel flow transmitters. The AIU 2 receives analog inputs from the right cockpit throttle grip, right-hand controller, and SENSOR control panel.

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Analog data is processed by the AIU set and sent to the ADCP by way of the avionics 1553 MUX bus.

Avionics interface unit built-in test The AIU set has three types of BIT: power-up, periodic, and initiated.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

224. Fighter data link 1. What is FDL?

2. Which F-15C/D FDL component contains the controls used for modes of operation and what are the modes?

3. Through which FDL component do all TACAN signals pass?

4. Which FDL component provides operating power for the FDL RT and TACAN RT?

5. How many batteries does the FDL RT contain and what do they provide?

6. Which switch on the control stick grip is used for FDL control?

7. What is the purpose of the TDC and where is it located?

8. The F-15C/D PSDP receives commands from which component as to the positioning of FDL symbology?

9. Nearly all of the processing related to the FDL is contained within which component for the F-15C/D and F-15E?

10. What is the purpose of the AIUs in the FDL system on the F-15E?

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11. Which component(s) has a CRYPTO switch which may be used to zeroize FDL crypto codes on the F-15E?

12. On the F-15C/D JTIDS mode control panel, what is the function of the MODE selector in the HOLD position?

13. On the F-15E sensor control panel, what is the function of the MODE selector in the NORM position?

14. Which menu display is the main screen used with FDL?

15. Which menu display allows the aircrew to set up or modify the DTM data for their own aircraft and is normally loaded during initialization?

16. What ENTER NET menu display indication means that time synchronization is complete and the FDL can both send and receive messages?

17. How does the FDL system keep the transmission secure and jam resistant?

18. What is NTR?

19. List the five things needed for the pilot to achieve FINE synchronization and fully communicate along the Link-16 network.

225. F-15E avionics interface unit 1. How many processing channels are in the AIU 1?

2. Which component functions as the UFC’s primary bus controller?

3. What type of data does the AIU use to interface with other aircraft systems?

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4. Where is control data from the UHF radio 1 sent?

5. Where does the AIU1A route TACAN data?

6. Which LRU routes serial radar altitude and altimeter status to AIU 1 and AIU 2?

7. To which AIU does the EMD provide serial data?

8. Which AIU 1 channel receives analog inputs from the ILS receiver, cockpit throttle grips, and left hand-controller?

3–2. Electronic Warfare Systems The lessons in this section have one thing in common. They all need RF for them to operate. Whether it is landing an aircraft in inclement weather, performing navigation with radio signals, determining if another aircraft is the enemy, or interrogating other aircraft; they all use RF. In this section you will learn how the F-15 aircraft accomplishes all of the above mentioned tasks. We will begin with the ILS.

226. Air-to-air interrogator The air-to-air interrogator (AAI) system provides the pilot with the ability to interrogate another aircraft’s IFF system. The AAI system is similar to the IFF ground interrogator with only a few differences. Although the IFF system can reply in modes 1, 2, 3/A, C, and 4, the F-15’s AAI system can only interrogate modes 1, 2, 3/A, and 4. Mode C provides the ground operator with the aircraft’s altitude. In contrast, the AAI system establishes whether the target aircraft is a friend (with high or low confidence) or a possible foe. The altitude function is computed by the on-board radar system.

The AAI system integrates with the aircraft radar. AAI displays are shown on the multiple indicator control panel (MICP) for F-15C/D or MPD/MPCD for F-15E and correlated with the radar targets. The AAI system holds various programs for automatic operation which may be manually selected to meet changing mission requirements. The system transmits challenges and receives replies through 10 dipole antennas mounted to the radar antenna. Bit circuits detect and isolate system failures. The AAI system uses the IFF transponder system in a loop test (self testing your own aircraft using AAI to interrogate its IFF system) for mode 4.

System components Now we discuss the system components to the AAI system.

IFF interrogator receiver/transmitter (APX-114) Located in door 3R, the IFF interrogator receiver/transmitter (AAI RT) provides interrogation generation, reply evaluation, and radar interface functions. The AAI RT monitors failures and provides system status and LRU status signals to the AIUs. The APX-114 is a 2.5 kilowatt transmitter and receiver. The receiver is made up of two separate receiver sections, each operating on 1090 megahertz (MHz); one for the reception of sum antenna inputs and one for reception of difference antenna inputs. Monitoring circuitry samples all critical operating parameters. Aircraft cooling system

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forced air cools the AAI RT and exhausts through ports on the unit face. Two interrogation program switches S1 and S2 are mounted in the face of the unit. There are two LED (light emitting diodes) indicators on the face of the unit, the top LED is a power indicator and the bottom LED indicates an RT fault.

Interrogator computer (KIV-119) The interrogator computer, commonly called the transponder computer, produces mode 4 interrogations and decodes mode 4 replies. The interrogator computer is located under door 3R. There is a single connector on the rear of the unit which mates with the AAI RT. The interrogator computer mounts external on the AAI RT. A FILL connector, through which the interrogation/reply code programs are inserted and a ZEROIZE switch, are on the front of the unit.

Antenna group The AAI antenna group consists of 10 dipole antennas arranged on the radar’s planner array (antenna), two power dividers, and a hybrid coupler. The dipoles are divided into two sets of five antennas, one set for the sum channel and one set for the difference channel. The sum and difference antennas are positioned on the planar array in a deliberate way to enable the AAI system to detect the direction of arrival of the reply signal. These antennas are delicate and you must take care when they’re exposed not to damage or move them.

Related system components Now we will discuss the related system components of the AAI system.

Right throttle grip The right throttle grip is located on the cockpit left console; the MULTIFUNCTION switch on the right throttle grip controls manual AAI interrogation initiation.

Interference blanker system The APX-114 supplies a suppression signal, which brackets the IFF interrogation to the blanker. It also receives a suppression signal from the blanker during IFF transponder transmissions. The APX-114 blocks suppression signals to and from blanker during mode 4 loop test (M/4 BIT).

AAI control panel F-15C/D The AAI control panel is located in the cockpit left console and it contains the switches used to control the AAI system. They include the MASTER switch and the MODE and CODE thumbwheels.

Multiple indicator control panel F-15C/D The AAI displays are shown on the MICP and correlated with the radar displays.

Up-front controller F-15E The UFC and rear UFC include 10 multifunction switches for selecting AAI submenu, AAI master modes and interrogation modes. The keyboard provides code selection capability. Either UFC can control the AAI system.

Multipurpose display system F-15E The multipurpose display system includes the seven MPD/MPCDs located in the cockpit and rear cockpit. The MPD/MPCD shows all the AAI system displays.

IFF transponder system The AAI system uses the aircraft IFF transponder system in a loop test configuration for mode 4 BIT.

Radar set/indicator group The AAI displays are shown on the MPD/MPCD and correlated with the radar displays. The IFF INTERROGATE switch and AAI BIT initiate signals are supplied to the RDP. The data processor supplies an interrogate command to the programmable signal processor (PSP) when the radar mode is compatible.

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System operation The following paragraphs describe the AAI system operation. The AAI system receives power from the left bus circuit breaker panel.

AAI master modes F-15C/D The MASTER switch on the AAI control panel is a four-position switch that allows selection of these options:

1. OFF – Removes power from the AAI system. 2. AUTO – When in the AUTO position, programming switches on the front of the

identification friend-or-foe reply evaluator (IRE) control the challenge mode and the interrogation time-out period.

3. NORM – The controller for the NORM mode is the MODE SELECTOR switch on the AAI control panel. The NORM mode will interrogate only the specific mode selected on the AAI control panel.

4. CC – The controllers for the CC (correct code) mode are the MODE SELECT switch and the four code selectors. The APX-114 will decode only replies that match the code selected on the AAI control panel. Thus, if the selective identification feature (SIF) code isn’t present or is incorrect, the AAI system won’t process or display the IFF target.

AAI master modes F-15E The selections on the UFC’s submenu (fig. 3–21) determine what combinations of modes, codes, master modes, and bars that are interrogated when an AAI interrogation is commanded. Any combination of codes can be selected on both UFC’s AAI submenu.

Code selection Codes for modes 1, 2, 3 are entered in the UFC’s scratchpad as indicated below:

• Mode 1: two digits from 00 to 73. • Mode 2 and 3: four digits from 0000 to 7777.

Figure 3–21. F-15E UFC AAI submenu.

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Interrogation initiation Interrogations are initiated either manually using the MULTIFUNCTION switch on the throttle quadrant or automatically by the radar data processor when radar switches to tracking mode.

Interrogation processing Pressing the MULTIFUNCTION switch on the right throttle grip to the IFF interrogate position initiates the AAI interrogation. The interrogate command signal from the MULTIFUNCTION switch enters interrogate control circuitry within the radar data processor (081 or 082). The interrogate command is sent as part of a data word that is transmitted between the radar PSP (042 or 044) and the APX-114. Data to the APX -114 is made up of information required for interrogation and reply processing. The APX-114 tests the data word for parity (validity).

AAI displays IFF replies are translated into target symbology by the radar system and displayed on the MICP or MPD/MPCD. There are many types of symbols used to indicate a detected aircraft and give the pilot important information about that aircraft. The letter I or A is displayed on the MICP or MPD/MPCD when the radar modes and displays are compatible with the AAI modes. An ID OFF message is displayed on the MICP when the APX-114 processes too many targets and becomes overloaded. An unidentified target will appear as a rectangle on the display. When a target is challenged (interrogated) by AAI, the symbol may change to indicate high or low confidence target. Any IFF mode can process as a low confidence target but only mode 4 replies will display as a high confidence target. A low confidence target is displayed as a diamond (non-tracked) or an igloo (tracked) symbol and a high confidence target is displayed as a circle (fig. 3–22).

Figure 3–22. AAI display.

M/4 light F-15C/D The M/4 light, mounted on the BIT control panel, indicates KIV-119 status and activity. It will light during normal operation to indicate a KIV failure, bad/no code loaded into the KIV or power removed from the KIV. During BIT or following an interrogation, the M/4 light remains on steady to indicate that the APX-114 has not received time decoded video from the KIV after receiving a valid mode 4 reply.

BIT operation The AAI system uses three types of BIT: automatic BIT, AAI IBIT, and mode 4 (M4) IBIT.

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Fault indications When the crypto fails, ASP fault indicator 35 will latch on the ASP panel as well as the fault indicator on the APX-114. When the AAI system fails, the ASP fault indicator 69 will latch.

227. F-15C AN/ALQ-135 internal countermeasures set components The AN/ALQ-135 internal countermeasures system (ICMS) is made up of three independent countermeasures sets. Two sets (band 1 and 2) are made up of an oscillator, RF amplifier, and one antenna each. The band 3 set is made up of an oscillator, two RF amplifiers (fwd and aft), and three antennas. The forward amplifier feeds continuous wave (CW) RF and pulse RF signals to a magic tee; the combined signal is transmitted out the forward up and forward down antennas. The aft amplifier feeds separate pulse RF signals and CW RF signals to separate elements in the aft antenna. The band 1 operates in the E through G bands, band 2 operates in the G and H bands, and band 3 operates in the H, I, and J bands. The F-15C ICMS is configured so that set 1 is band 3, set 2 is band 2, and set 3 is band 1.

Cockpit controls The following paragraphs describe the ICMS units. Each unit, where required, is cooled by the aircraft’s ECS, which provides the airflow necessary to dissipate heat. When any ICMS component is not installed in bay 5, ballast must be installed for flight in its place.

The cockpit controls for the ICMS consist of the tactical electronic warfare system (TEWS) immediate action (IA) control panel, TEWS control panel, TEWS display unit, and the MPCD.

TEWS IA control panel The TEWS IA control panel (fig. 3–23) is used to select the combat/training mode, place all the ICMS bands in standby or automatic mode, and enable individual band operation.

TEWS IA Control Panel Switch Position Function

RWR/ICMS (Radar Warning Receiver/ ICMS Conflict Cue)

COMBAT ICMS in the combat mode. The ICMS will operate using the combat preflight message (PFM) data.

TRAINING ICMS in the training mode. The ICMS will operate using the training PFM data.

ICS (ICMS Conflict Cue)

STBY ICMS in standby mode. AUTO ICMS in automatic mode.

MAN Pilot may manually select each set’s mode of operation at TEWS control panel.

TEWS control panel The TEWS control panel (fig. 3–23) is used to select individual band ICMS operating modes (AUTO/MAN) and turn on system power.

TEWS Control Panel Switch Position Function

ICS ON Applies power to the ICMS. OFF Removes power from the ICMS.

SET 1 AUTO Places band 3 in automatic mode. MAN Places band 3 in standalone mode.

SET 2 AUTO Places band 2 in automatic mode. MAN Places band 2 in manual mode.

SET 3 AUTO Places band 1 in automatic mode. MAN Places band 1 in manual mode.

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The TEWS control panel also contains set fail lights that indicate when a specific set (band) has failed.

Set Light Indication SET 1 FAIL Band 3 failure. SET 2 FAIL Band 2 failure. SET 3 FAIL Band 1 failure.

TEWS display unit During radar warning receiver (RWR) IBIT, the TEWS display unit displays ICMS program information. When band 3 jamming is present, the jammed threat is identified as being jammed.

Multipurpose color display The MPCD (fig. 3–23) provides ICMS status on the tactical situation display (TSD) and programmable armament control set (PACS) air-to-air (A/A) display. ICMS BIT log and detail BIT display selection is made on the TEWS BIT page.

Figure 3–23. ICMS cockpit components.

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Equipment bay 5 ICMS components The following components are located in equipment bay 5 on the left side. Refer to figure 3–24 for their location in relation to the other bay 5 components.

Summing network The summing network receives RF samples from only band 1 and band 2 RF amplifiers and applies a composite output of these samples to the RWR high band receiver for comparison to the initial threat data.

Figure 3–24. Bay 5 ICMS components.

Band 1 oscillator (set 3) The band 1 oscillator receives jamming parameter commands from the RWR and uses tuning units and function generators to produce low-power amplitude and/or frequency modulated RF signal. In manual (emergency) mode, the jamming commands and parameters are received from a self-contained read-only memory (ROM) contained in the programmer. The low-power RF signal is then sent to the band 1 RF amplifier. The RWR routinely turns off the tuning units to look at the covered frequency range to determine if more jamming is required.

Band 1 RF amplifier (set 3) The band 1 RF amplifier receives the low-power RF signals for the band 1 oscillator. The band 1 amplifier also receives CW and high-level modulation commands from the oscillator for automatic leveling and AM modulation. The band 1 amplifier uses two traveling wave tube (TWT) circuits to produce leveled high power RF jamming signals. A sample of band 1 RF is sent to summing network

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before final amplification. The final amplified band 1 RF signal is sent to the band 1 antenna over a coaxial transmission line. The amplifier has a unit fault indicator and a voltage standing wave ratio fault indicator.

Band 2 oscillator (set 2) Band 2 works identical to band 1 only the frequencies in which band 2 transmits are in the G and H bands while band 1 operates in the E/F/G bands.

Band 2 RF amplifier (set 2) The band 2 RF amplifier receives the low power RF signals from the band 2 oscillator. The band 2 RF amplifier also receives CW and high level modulation commands from the band 2 oscillator for automatic leveling and AM modulation. The band 2 amplifier uses two TWT circuits to produce leveled high power RF jamming signals. A sample of band 2 RF is sent to the summing network before final amplification. The final amplified band 2 RF signal is transmitted through pressurized waveguides to the band 2 antenna. When the band 2 amplifier is not installed, the waveguides are secured to the aircraft structure in a stowed position; pressurization integrity is preserved by quick-disconnect end caps. The amplifier has a unit fault indicator and a voltage standing wave ratio fault indicator.

Band 3 oscillator (set 1) The band 3 oscillator receives RF signals from the RWR high band antennas through the preamplifier. The RF signals are processed, analyzed, and compared with threat signal tables located in the oscillator itself. Threats are prioritized and jammed based on the threat analysis. Tuning commands, amplitude and/or frequency jamming output signals, and frequency information is sent to the forward and aft band 3 RF amplifiers.

Band 3 forward RF amplifier (set 1) Final TWT amplification of band 3 oscillator signals takes place in the band 3 RF amplifier. The band 3 forward RF amplifier also provides operating voltages to the band 3 oscillator. The pulse and CW RF signals from the band 3 RF forward amplifier are applied through waveguides to the magic tee and then to the band 3 ICMS forward up and down antennas.

Band 3 aft RF amplifier The band 3 aft RF amplifier provides final TWT amplification of band 3 oscillator. The band 3 aft RF amplifier also provides operating voltages to the preamplifier. The pulse and CW RF signals from the band 3 aft RF amplifier are applied to the coaxial adapters. RF signals from the coaxial adapters are sent to the band 3 aft antenna. The aft amplifier is pressurized when the pressurization line is connected to it. The amplifier has a unit fault indicator and a voltage standing wave ratio fault indicator.

Magic tee The magic tee is used to combine band 3 forward amplifier CW RF and pulse RF signals into a single signal, split the combined signal, and send the split signal to the forward up and forward down band 3 ICMS antennas.

Programmer The programmer is a small matrix ROM device, which provides band 1 and band 2 oscillators with predetermined program inputs. The predetermined program inputs are used by the ICMS when operating in manual mode. A separate programmer is provided for and installed on each oscillator (not band 3) installed in the aircraft.

Antennas The ICMS antennas, shown in figure 3–25, are transmit only antennas; they do not receive!

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Band 1 antenna (set 3) The band 1 antenna is located on the bottom centerline of the aircraft just behind the nose wheelwell.The band 1 antenna is a coaxial cable fed, omnidirectional, broad band blade-type antenna.

Band 2 antenna (set 2) The band 2 antenna is a waveguide fed, broad band omnidirectional blade-type antenna.

Figure 3–25. F-15C forward ICMS antennas.

Band 3 forward up antenna (set 1) The band 3 forward up antenna is a coaxial cable fed horn, pulse/CW antenna. Together with the band 3 forward down antenna, it provides a cone of coverage forward of the aircraft.

Band 3 forward down antenna (set 1) The band 3 forward down antenna is a waveguide fed horn pulse/CW antenna. Together with the band 3 forward up antenna, it provides a cone of coverage forward of the aircraft.

Band 3 aft antenna The band 3 aft antenna is a dual element antenna. The upper element is a coaxial cable fed horn CW RF antenna. The lower element is a coaxial cable fed horn pulse RF antenna. Each provides a cone of coverage directly aft of the aircraft.

ICMS integration with other LRUs The ICMS integrates with several other components on the aircraft such as LRUs like the central computer, bit control panel, and avionics panel. Minor parts hardware like the waveguide seals and coax adapters are important to the system as is the operational flight program (OFP) and preflight messages (PFM), all of which are covered in the next paragraphs.

Preamplifier The preamplifier receives four RF inputs (originating at the high-band antennas) from the RWR high band receiver and one agile tuning unit 1 (ATU 1)/BIT RF signal input from band 3 aft RF amplifier. The band 3 ATU 1/BIT RF signal is used with an automatic leveling circuit (ALC) to provide a calibrated amplitude source for the BIT mode of operation. The preamplifier provides initial system

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amplification of the received RF threat signals from the RWR high band antennas. The antenna input signals from the RWR are then routed to the band 3 oscillator.

Coax adapters The coax adapters are used to couple RF signals from the band 3 aft RF amplifier waveguide ports to cable assemblies.

Waveguide seals Waveguide seals are not a major component, they are consumable items. Wherever there is a mating surface between waveguide sections, between waveguides and antennas, and between waveguides and amplifiers, there is a waveguide seal. These seals prevent pressure loss, arcing, and contamination.

Left main landing gear weight-on-wheels relay When an aircraft is flying (weight is off the left main landing gear (MLG)), the left MLG WOW relay is energized and passes a ground to the ICMS oscillators. When this ground is absent (aircraft on the ground), the oscillators send an inhibit command to the amplifiers.This prevents accidental high power RF signal transmission when the aircraft is on the ground.

Central computer The CC sends MUX data to the band 3 oscillator. This data consists of radar operating frequencies, aircraft altitude, and aircraft speed. A reply word is sent to inform the CC that the band 3 has received valid CC data.

Band 3 OFP and PFM While not physical components, the OFP and PFM are necessary for ICMS operation. As C-shoppers, you’ll be required to load the OFP and PFM into the band 3 oscillator on a regular basis. Changes in threat environments (new or improved radar systems and missiles) require changes in the ICMS operating programs. The OFP and PFM are loaded in the band 3 oscillator over the 1553 avionics MUX bus. A reprogramming receptacle is assessable when the ground communications panel is open.

BIT control panel The BCP is used to start and monitor the BIT function of the ICMS system.

Avionics status panel Remember, an ASP fail indication is represented by an orange indicator. ASP 9 means a failure in the aft band 3 amplifier and ASP 59 means a failure in a bay 5 ICMS component.

228. F-15C internal countermeasures set operation Band 1 and band 2 are “transmit only” systems under the control of the RWR during automatic mode (AUTO) or the programmers during manual mode. Band 3 is a self-contained TEWS package that can receive and transmit independently of RWR. Band 3 requires the RWR high band antennas for threat RF signal detection.

Modes of operation The three modes of operation for the ICMS are STBY (standby), AUTO (automatic), and MAN (manual).

Standby mode When STBY is selected, the ICMS receives power to warm up the system, but the ability to transmit is inhibited.

AUTO mode AUTO mode is selected by placing the ICS switch on the TEWS IA control panel to AUTO.

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Band 1 and band 2 auto mode The RWR controls the band 1 and band 2 oscillators with ICMS control words sent over the RWR/ICMS Manchester data bus. The RWR selects jamming mode, establishes jamming priorities, and determines the jamming frequencies. Sample jamming RF from the band 1 and band 2 amplifiers is sent from the summing network to the high band receiver for comparison to the commanded jamming parameters.

Band 3 auto mode The RWR provides threat information for comparison with threat information received by the band 3 oscillator. The band 3 oscillator uses the comparison of this information to select jamming mode, establishes jamming priorities, and determines jamming frequency or frequencies to be covered.

Manual mode Manual mode is selected for each set of the ICMS by setting the ICS switch on the TEWS IA control panel to MAN, and setting the respective SET–1, SET–2, or SET–3 switch on the TEWS control panel to MAN.

Band 1 and band 2 manual mode The programmer self-contained ROM provides the data words that control jamming parameters and commands for the band 1 and band 2 manual modes. Manual mode, an emergency mode, is used during RWR or ICMS failure.

Band 3 manual mode The band 3 manual mode is a stand-alone mode. The band 3 oscillator ignores all communication with the RWR and develops jamming signals with no comparison to RWR developed threat data.

Band 1 and band 2 operation Both band 1 and band 2 are controlled in the AUTO mode by the RWR and by their programmers during manual mode. Refer to figure 3–26 for the following discussion.

Figure 3–26. Band 1 and 2 block diagram.

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When the RWR detects a threat, it will select a jamming mode, establish jamming priorities, and determine jamming frequency or frequencies to be covered. The selection of jamming mode, priority, and frequency is dependent on COMBAT/TRAINING mode selection and missionized PFM selection.

The low band receiver processor will send ICMS control words sent over the RWR/ICMS Manchester data bus to the band 1 and band 2 oscillators. The oscillators use the control words to enable the tuning units and function generators to produce low-power amplitude and/or frequency modulated RF signal. In the manual mode, the control words come from the ROM of the programmers.

The low power RF signal from the oscillator is sent to the amplifier. The amplifier also receives CW and high level modulation commands from the oscillator for automatic leveling and AM modulation. The amplifier uses two TWT circuits to produce leveled high power RF signals. A sample of the RF is sent to summing network before final amplification.

The final amplified band 1 RF signal is sent to the band 1 antenna over a coaxial transmission line. The final amplified band 2 RF signal is transmitted through pressurized waveguides to the band 2 antenna. The summing network receives RF samples from the band 1 and band 2 RF amplifiers and applies a composite output of these samples to the RWR high band receiver for comparison to the initial threat data.

Band 3 operation The band 3 ICMS was designed as a stand-alone system. Refer to figure 3–27 for the following discussion.

Figure 3–27. F-15C band 3 block diagram.

Threat RF signals are detected by any combination of the RWR high band antennas. Inside the high band receiver, the threat RF signals are split and sent to the preamplifier. The preamplifier provides initial system amplification of the received RF threat signals from the RWR high band antennas. The antenna input signals from the RWR are then routed to the band 3 oscillator.

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The receiver in the band 3 oscillator processes, analyzes, and compares the threat RF signals with threat tables located inside the band 3 oscillator itself. This analyzed information results in prioritization and jamming method selection. Tuning commands, amplitude and/or frequency jamming output signals, and frequency information are sent to the forward and aft band 3 RF amplifiers.

The forward amplifier sends the pulse RF and CW RF signals to the magic tee where they are combined and the composite pulse/CW jamming signal is sent to the forward up and forward down antennas. The forward antennas provide jamming coverage for the front of the aircraft.

The aft amplifier pulse RF signal goes to a waveguide-to-coaxial adapter and is sent to the pulse element of the aft antenna. The aft amplifier CW RF signal goes to a waveguide-to-coaxial adapter and is sent to the CW element of the aft antenna.The aft antenna elements provide jamming coverage for the rear of the aircraft.

When band 3 is jamming, the oscillator commands the RWR to display an open X (fig. 3–28) around the jammed threat on the TEWS display. If no RWR/ICMS jamming correlation exists, the open X will be displayed at the center of the TEWS display.

Figure 3–28. Band 3 jamming indications.

Band 1 and band 2 BIT Built-in test of bands 1 and 2 are separated into three categories: continuous BIT, intermittent BIT, and manual initiated BIT.

ICMS band 3 BIT The band 3 BIT categories are initialization, continuous, intermittent, and manual initiated BIT.

Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

226. Air-to-air interrogator 1. What does the AAI RT provide to the air-to-air interrogation system?

2. What makes up the AAI antenna group?

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3. How are the dipoles divided?

4. Why are the sum and difference antennas positioned on the planar array in a deliberate way?

5. What is the purpose of the MULTIFUNCTION switch on the right throttle grip in the AAI system?

6. What is the function of the F-15C/D AAI control panel?

7. Which F-15E component includes 10 multifunction switches for selecting AAI submenu, AAI master modes, and interrogation modes?

8. When the AAI MASTER switch on the AAI control panel is in the AUTO position on the F-15C/D, what controls the challenge mode and the interrogation time-out period?

9. Which F-15E component determines what combinations of modes, codes, master modes, and bars that are interrogated when an AAI interrogation is commanded?

10. How are interrogations initiated?

11. Explain how interrogation processing works.

12. What happens to IFF replies that are displayed?

13. During BIT or following an interrogation, what does the M/4 light provide?

227. F-15C AN/ALQ-135 internal countermeasures set components 1. Which countermeasures sets make up the AN/ALQ-135 ICMS?

2. In which frequency range does band 1 operate?

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3. In which frequency range does band 2 operate?

4. In which frequency range does band 3 operate?

5. What do the set fail lights indicate?

6. What ICMS information does the TEWS display provide?

7. What ICMS information does the MPCD provide?

8. What is the function of the summing network?

9. What does the band 3 oscillator provide the forward and aft band 3 RF amplifiers?

10. What does the band 3 forward RF amplifier provide?

11. What is the function of the magic tee?

12. What is a programmer?

13. What is the function of the preamplifier?

14. What is the function of waveguide seals?

15. What does the left MLG WOW relay provide to the ICMS, and why is it important?

16. What information does the CC provide to the band 3 oscillator?

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228. F-15C internal countermeasures set operation 1. What controls band 1 and band 2 during AUTO mode and what type of system are they?

2. What controls band 1 and band 2 during MAN mode?

3. When should MAN mode be used for band 1 or band 2 operation?

4. What signal does the oscillator send the RF amplifier?

5. Where do the band 1 and band 2 amplifiers send an RF sample?

6. How is final amplified band 1 RF signal sent to the band 1 antenna?

7. How is final amplified band 2 RF signal sent to the band 2 antenna?

8. What is the RWR commanded to display when the band 3 ICMS is jamming?

3–3. PODS Since its birth, the F-15E aircraft has been the superior fighter aircraft on the planet. In order to stay at the top of the food chain the F-15E has been and is still going through many modifications. Two important upgrades of the F-15E are the navigation and targeting capabilities of the aircraft. They are happening so fast that by the time you study this material on these pods they may be extinct. In this section we will discuss the LANTIRN NAV pod. We will conclude by diving into the operation of the data link pod. Let’s begin with the LANTIRN NAV pod.

229. F-15E LANTIRN navigation pod The LANTIRN NAV pod is made up of two systems: the terrain following radar (TFR) and the fixed imaging navigation sensing (FINS), also known as the navigation forward look infrared (NAV FLIR) system. The TFR subsystem uses radar emissions and returns to maintain flight at a discrete preselected clearance (100 to 1000 ft) above ground level. This provides low altitude operation and under-the-weather flight. The FINS subsystem uses variations in forward looking infrared emissions to develop a video image. This image is projected to the wide-field-of view head-up display (WFOV HUD). This provides day or night operations. Snap look and look-into-turn (LIT) modes add to normal operation.

Component description The following paragraphs describe the components to the F-15E LANTIRN NAV pod system.

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LANTIRN NAV POD adapter The LANTIRN NAV POD adapter (fig. 3–29) mounts under the right engine inlet and provides the mating surface between the F-15E and the NAV POD.

Figure 3–29. LANTIRN NAV pod adapter and pod.

LANTIRN NAV POD The NAV POD contains five functional subsystems, which contain 11 physical subassemblies. Opening the access covers or removing the radome cover provides access to the subassemblies in the pod. The 11 subassemblies are listed below and shown in figures 3–30, 3–31, and 3–32.

1. NAV pod power supply. 2. Advanced pod control computer. 3. NAV FLIR assembly. 4. Environmental control unit (ECU). 5. ECU controller/maintenance panel. 6. Radar interface unit. 7. Power supply. 8. Electronic equipment pressurization set. 9. Radar transmitter. 10. Radar receiver. 11. TFR antenna/gimbal.

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Figure 3–30. LANTIRN NAV POD components (1 of 3).

Figure 3–31. LANTIRN NAV POD components (2 of 3).

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Figure 3–32. LANTIRN NAV POD components (3 of 3).

Related systems description We briefly describe the interface between the LANTIRN system and other aircraft systems.

Head-up display unit The display image shown on the HUD is made up of a video representation of a FLIR image provided by the FINS and vertical steering commands and other symbology derived from the TFR.

Horizontal warning lights assembly The forward crew station horizontal warning lights assembly is located in the 30 degree visual cone on the main instrument panel (fig. 3–33). The horizontal warning lights assembly provides discrete warnings, cautions and advisory light information. The LOW ALT, OBST, and TF FAIL are unique to the LANTIRN system and provide the following information to the aircrew:

• LOW ALT comes on if the aircraft radar altitude descends below 75 percent of the terrain following (TF) set clearance value plus a predictive variable based on negative vertical velocity. The LOW ALT light and a low altitude voice warning message is also provided and set off by the altitude selected on the UFC.

• OBST comes on to warn the aircrew of an obstacle appearing in the aircraft flight path as detected by the terrain following radar and requiring more than 2 gravity (G) to clear. A TF flyup will be produced when this condition occurs. A voice warning message (OBSTACLE AHEAD) is heard as long as the obstacle is detected.

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• TF FAIL comes on if a TF system fail has been detected by the NAV POD or flight control computer (FCC). A TF flyup will be produced when this occurs. A voice warning message (TF FAIL) is also heard for TF FAIL.

Figure 3–33. Horizontal warning lights assembly.

Up-front control/rear up-front control The UFC and rear UFC are the major interface units for control of the NAV FLIR and other avionic systems. Each UFC has a processor with paths to the other unit to provide a redundancy when a UFC or processor fail occurs. The UFC provides access to system functions, status, and control through data displays, menus and submenus.

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Multipurpose display system The multipurpose display (MPD) system provides visual representation of the FLIR imagery and terrain following radar. The MPD system is made up of the ADCP and the seven MPDs/MPCDs. Display interface is provided by the ADCP.

The MPD/MPCD and the HUD displays system data, sensor video, and weapon information in monochromatic or multicolor format. The primary format used with the LANTIRN system is the TF format. The TF format provides vertical steering information, pod raster video, moding control and status, and TF cautions/warnings. The TF display is composed of video from the NAV POD and the MPD/ADCP.

MASTER CAUTION light/reset switch The MASTER CAUTION lights (fig. 3–33) on the upper instrument panel in both cockpits come on when any MPD/MPCD caution legend comes on. The MASTER CAUTION lights do not come on when the AV BIT or FLY UP ARM caution lights come on. However, pressing either MASTER CAUTION light resets the circuit and causes the AV BIT light to go out.

FLY UP ENABLE switch The FLY UP ENABLE switch (fig. 3–33) is located outboard of the forward throttle quadrant, below the left canopy sill. The toggle switch is guarded in the ON position. Lifting the guard and selecting OFF disables the automatic flyup mode during manual TF.

Control augmentation system control panel The control augmentation system (CAS) control panel, located on the cockpit left console, contains controls for the automatic flight control system (AFCS) and auto TF engagement.

Sensor control panel The sensor control panel provides the operator interface for the radar altimeter, TFR system, and NAV FLIR sensor.

GND PWR control panel The GND PWR control panel contains toggle switches for activating various aircraft systems with external power. Positioning switch 1 to B ON, switches 2, 4, and ADCP/AIU1 to ON, applies external power to all LANTIRN related systems.

Remote intercommunication control panel The RICP is on the left console in the front cockpit. The communication system provides voice warning ability. The voice warning is activated when a failure condition exists causing the following warning lights to come ON: OBST and TF FAIL. The control related with the LANTIRN NAV POD is the voice warning silence pushbutton. Pressing this pushbutton silences the voice warnings and the audio tones in progress for one minute. If the warning still exists after one minute, the warning will be heard again.

Intercommunication set control panel The ICSCP is on the left console in the rear cockpit. The voice warning silence pushbutton provides the same function as that on the RICP.

Throttle grips The left cockpit throttle grip (fig. 3–33) provides the control and interface to enable the FLIR line of sight (LOS) alterations. The snap-look and look-into-turn alternatives are enabled by holding the right MULTIFUNCTION switch in the DOWN position. The TDC function switch enables electrical slewing to align the NAV FLIR image with the real world as viewed through the HUD.

Control stick grip The control stick grip (fig. 3–33) provides the display control/snap-look switch to command the display preference and LANTIRN snap-look direction.

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Automatic flight control system The AFCS provides pitch, roll, and yaw CAS and autopilot modes of pitch/roll attitude and altitude hold. The autopilot can be coupled with TF subsystem to provide AUTO-TF operation when all three CAS axes are engaged.

Avionic interface unit 1/AIU 2 AIU 1, in parallel with AIU 2, receives analog and discrete signals from the sensor control panel. The AIU set receives two discrete signals from the LANTIRN NAV pod: NAV POD PRESENT and NAV POD READY.

Advanced display core processor The ADCP is the bus controller and display processor for the whole aircraft. The ADCP produces symbology for the HUD, MPD, and MPCD. It also does video processing for the MPD and MPCD. It initiates and controls data transfer with the HUD, MPD, and MPCD; and communicates with NAV pod and other systems on the 1553 MUX bus.

Inertial navigation set The inertial navigation set (INS) supplies the primary attitude reference for the aircraft and provides continuous present position monitoring. The INS provides aircraft attitude, heading, velocity, and acceleration information to the radar, ADCP, LANTIRN and AFCS. INS data is used in determining TF operating envelope, flyup cues, and vertical steering commands.

Mode description The NAV FLIR and TF RADAR switches on the sensor control panel control primary power application to the NAV POD. They provide independent control of the two functional subsystems of the NAV POD: NAV FLIR and TFR.

NAV FLIR mode The UFC menu 1 (fig. 3–34) displays NAV FLIR mode status (N/R, STBY, NORM or BRST). Both cockpits can select submode and alternate changes, using the UFC. A change started in either cockpit is displayed on both UFCs and is overridden by the next sequential change in either cockpit. There are two NAV FLIR submodes: NORM and BRST.

Figure 3–34. UFC menu 1 display.

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In NORM submode, the HUD image depicts an area within the pilot’s LOS. There are two selections available to change the field of view: snap look and LIT.

In BRST submode (fig. 3–35), two types of boresight inputs are available for display control: electrical and mechanical. Electrical boresight is selected to slew the NAV FLIR video to provide correlation with the real world image. Mechanical boresight displays the yaw, pitch and the roll correction values. The correction coefficients consider manufacturing tolerance errors of the NAV POD/adapter hard point mounting pads. These unique numbers are stored in the ADCP.

Figure 3–35. NAV FLIR submenu display (BRST selected).

TFR mode In addition to normal (NORM) operation of the TFR, weather (WX 1 or WX 2), low probability of intercept (LPI), electronic counter-countermeasures (ECCM), and very low clearance (VLC), are available for specific operating conditions. Radar functions include set clearance control, ride control, and frequency selection. The TFR submodes and radar functions (fig. 3–36) are selected on the TF display (MPD/MPCD). When the TFR is initialized the last selected submode will be active and boxed on the TF display.

• The NORM submode provides the best combination of system operational parameters. Normal submode provides the highest range and angle measurement accuracy. This enables precise flight profiles at selectable altitude clearances with limited auto ECCM.

• The WX submode allows the radar to operate in adverse weather conditions with rainfall rates up to 10 millimeters/hour (mm/hr).

• The LPI submode uses power management techniques and special/time control of transmitter emissions to minimize the probability of detection. Minimizing the radiated microwave energy and radar scan patterns results in a low electronic profile.

• The ECCM submode provides maximum immunity from electronic countermeasure threats.

• The VLC submode provides for performance capability down to 100 feet. VLC automatically reverts to 200 feet set clearance when operating at 100 feet set clearance over rough terrain.

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• Terrain following can be done in manual or automatic submode. Manual TF allows the aircraft to be flown at a selected terrain clearance with normal handling qualities of the flight control system.

Figure 3–36. MPD/MPCD TF display.

NAV FLIR operation Infrared (IR) emissions are received by the FINS in the NAV FLIR assembly. The IR scene is received through the prisms and focal lens and reflected by a scan mirror to the IR imager. From the IR imager, the scene is supplied to the pod’s detector sensors. A focus wedge provides focus control of the IR imager. The data on each of the 180 channels of the detector are supplied to the digital scan converter (DSC) for formatting before being sent to the aircraft for display.

NAV FLIR submode and alternative selection Submode or display alternatives are selected in the STBY or OPERATE modes. The active submode is displayed next to S1 on the NAV FLIR submenu on the UFC.

The NORM submode allows selection of NAV FLIR display alternatives listed below: • GRAY SCALE. • MAN (AUTO) - GAIN/LEVEL. • WHOT/BHOT. • BRST.

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The UFC processes the selected switches into a format compatible with the AIU. AIU 1 and AIU 2 receive switch status from the UFC and digitize the output for transmission to the ADCP.

The ADCP receives the GRAY SCALE command and then commands the NAV POD to the GRAY SCALE mode. The FINS overlays the GRAY SCALE on the FLIR video.

The MAN (AUTO)-GAIN/LEVEL provides manual or automatic control of the gain or level of the infrared image. The ADCP reads the NAV FLIR status based on the information from the AIU. The output quality of the NAV FLIR video is controlled by two gain/level pots on the sensor control panel. AIU 2 receives and digitizes the analog signals from the sensor control panel. The ADCP encodes the FLIR manual level and gain settings and sends the commands to the FLIR. The FLIR then adjusts its sensitivity and video output accordingly. When AUTO-GAIN/LEVEL is selected, the FLIR automatically adjusts level and gain using pod determined criteria.

Pressing and releasing of the BHOT (black-hot) key or WHOT (white-hot) key changes the video polarity of the FLIR video. White-hot corresponds to a standard video format in which hot areas are displayed as green on the HUD. Black-hot corresponds to a reverse video format. In this format, hot areas are displayed as black.

When the BRST submode is selected, the UFC display changes to indicate that it has been selected. Two types of boresight inputs for display control are electrical and mechanical. The only function that is exclusively enabled in BRST submode is electrical slewing to align the NAV FLIR.

TFR operation The LANTIRN TFR subsystem uses radar emissions to the ground and their return signals to produce terrain following commands. These commands are used as steering cues displayed on the HUD, electronic attitude director indicator (EADI) and TF display for MANUAL and AUTO-TF modes of operation.

The antenna/gimbal RF section receives the transmitted signal from the receiver and radiates it in a beam from the antenna. The antenna captures the radar return signal from the terrain and delivers it to the receiver. The return data passed to the receiver is converted to the first of several intermediate frequencies (IF). The IF signal is then processed into video data and provided to the radar interface unit. The radar interface unit stores the consecutive radar returns and correlates the current radar video with the radar returns stored from previous transmissions to declare targets. The radar interface unit uses the nearest declared target to compute the range for each radar transmission. Video data enhanced with display data is processed into a format compatible with the processing done by the ADCP.

Primary control and interface between subsystem functions and between each subsystem and the aircraft is provided by the APCC. Other TFR functions include obstacle warning, jam detection, TF blanking and stick force feedback.

TFR submode and function select ion The TFR subsystem operates in one of five modes:

1. Normal (NORM). 2. ECCM. 3. Weather (WX1 or WX2). 4. LPI. 5. VLC.

Radar functions include set clearance control, ride control and frequency selection. Pressing and releasing the required switch on the TF display (MPCD/MPD) sends the switch status to the ADCP. The ADCP provides submode command to the NAV POD. A box appears around the selected submode.

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Set clearance Set clearance is entered using the UFC or the TF display. If the VLC submode is selected, the ADCP allows the 100 foot set clearance to be displayed for selection on the TF display or through the UFC. Selection of any submode (other than VLC) commands the removal of the 100 foot set clearance on the TF display.

Ride control The ride control function provides a ride quality control over terrain contours. There are two available settings: HARD and SOFT. Ride control is sort of like having shocks in the seat of your car; the tension is either hard or soft so it will dictate the comfort of your ride as you go over bumps but in the pilot’s case it affects the comfort of the flight as the aircraft follows terrain.

Frequency selection The ADCP allows manual frequency selection using the TF display. If the selected submode is either ECCM or LPI, the ADCP does not display/allow manual frequency selection. In ECCM or LPI frequency selection ability is removed from TF display.

Built-in test The LANTIRN system has two types of BIT: periodic and initiated. PBIT is non-interruptive and automatic. PBIT is enabled when power is applied to the NAV POD and operates continuously.

IBIT provides a comprehensive system interruptive test. NAV FLIR and TF RDR IBIT are started from any MPD/MPCD BIT or MAINT BIT menu.

230. F-15E data link pod Now you have a firm understanding of the LANTIRN NAV and targeting pod system along with a glimpse into the reconnaissance (RECCE) pod; let’s now dive into the data link pod. This lesson contains description and principles of operations for the data link pod(s).

System description The AXQ-14 data link pod (DLP) (fig. 3–37) or an improved data link pod (IDLP) is carried on the centerline pylon, station 5, of the F-15E. The electrical interface of pod to pylon and pylon to aircraft is done through the electrical cables in the nose of the centerline pylon. The pod can also be carried on one aircraft to control a weapon launched from another aircraft. The pod is used to communicate with the GBU-15 guided bomb or AGM-130 guided missile (referred to as the weapons). The weapons must have a data link RT installed. The weapons can be controlled and locked on to a target after launch by the launching aircraft or another aircraft equipped with a data link pod.

Figure 3–37. Data link pod.

System operation The DLP receives power from the aircraft right circuit breaker panel no. 3. The video recorder is installed in the front compartment of the AXQ-14 pod and provides recording and playback functions

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when the pod is powered. Access to the recorder is through the front access door. The system allows pre-recorded tapes to be loaded for a mission and allows tapes which were recorded during a mission to be used in debrief.

Operating modes There are three DLP operating modes: AUTO, MAN (manual), and FWD (forward), which are selected and boxed with the MPD/MPCD switch (S2). The mode selection on the MPD/MPCD is sent to the ADCP on the MUX bus and is returned to the PACS. The PACS receives the data indicating the selected operating mode and sets specific pod operation command bits to their applicable levels, logic 0 or logic 1. The operation command data is then sent to the DLP on the PACS/DLP MUX bus. The operating modes are described below:

AUTO mode Automatic mode allows the phase scanned array (PSA) antenna to automatically scan the left, right or both quadrants for the strongest RF signal.

MAN mode Manual mode allows weapon control commands to be transmitted through the rear PSA antenna. The antenna can be manually moved to find the strongest signal as indicated by the weapon video. The antenna is controlled by movement of the cursor functions/sequence point select switch on the hand controllers. Two positions (forward and rear) are used to move the antenna at a fixed rate until the switch is released. The forward position moves the antenna toward the nose and aft moves it toward the tail. Reference marks for the antenna position are displayed on the displayed weapon video for forward and aft. The antenna caret indicates left or right. The PACS receives the PSA antenna control signals from the ADCP and converts them into an analog voltage that represents the antenna position command from the hand controllers. The PSA antenna control voltage is sent hardwired to the DLP. MAN antenna control is only functional in the XMIT (transmit) mode.

FWD mode In forward mode, all commands are transmitted through the forward horn antenna on the DLP.

Controls and displays DLP functions available for selection on the MPD/MPCD are listed below:

STBY (standby) In this function, power is applied to the DLP. The selected function (standby or operate) is sent by the ADCP to the PACS. The PACS receives the data indicating the DLP is in either standby or operate and sets the operate command to the applicable level. It is recommended that the pod be on for a minimum of five minutes before being used. This is to make sure stable operations are available. The pod can receive video and provide it for display in the STBY function, but it does not transmit commands.

TEST When this function is selected, weapon control commands are sent at low power through the rear PSA antenna. This function allows the aircraft and weapon data link to be tested before weapon release. The weapon cannot be launched in TEST function.

Terminal FUSE ARM After weapon release, fuse arming is irreversible. If in terminal command, fusing circuits are automatically armed if not already selected. FUSE is selected at S5 on the MPD/MPCD. Fuse enable command is received by the PACS from the ADCP. The PACS sends the fuse enable data to the DLP on the PACS/DLP MUX bus. Manual selection of fuse arm may be required to prevent a weapon dud if weapon release is very close to the target and terminal command is required immediately. The weapon must be in terminal function a minimum of 5 seconds before impact to detonate.

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Transition command Transition command (TRSN) is available for selection at S6 when in the XMIT function. Transition is selected to change the weapon flight phase from midcourse.

Terminal command When terminal command (TERM) is selected and TERM is displayed, the weapon can be steered from the cockpit in both pitch and yaw.

PSA LFT/BOTH/RT PSA command provides for control of the search vector of the DLP PSA antenna from the MPD/MPCD. The options are PSA LFT, which allows for search on the left side of the aircraft. PSA RT allows for search on the right side of the aircraft. PSA BOTH allows for search on both sides of the aircraft.

PSA FWD and PSA AFT If PSA MAN is selected, the PSA antenna direction can be controlled by the MPD/MPCD. Selecting PSA FWD moves the PSA antenna towards the aircraft nose, or selecting PSA AFT moves it toward the tail. The weapon video displays a triangular caret, indicating antenna position.

EDG BLK and EXP WHT The MPD/MPCD is used to toggle the track function between EDG BLK and EXP WHT. In EDG BLK function, the weapon seeker tracks contrasting vertical edges of a target. In EXP WHT function, the weapon seeker tracks two vertical edges to center the tracking gate over the target.

MODE This function allows the pod to talk to the correct weapon. There are eight operating modes available A, B, N1, N2, J1, J2, C, and D. Selecting the correct mode will normally be required when the pod is used to control a weapon launched from another aircraft. Modes A and B are used for training only.

CHAN Channel selects the operating frequency and power level. Switch S17 on the MPD/MPCD increases the frequency with each press from 1 to channel 8; then the cycle starts over. Switch S18 is used to decrease the frequency. Frequency 5 thru 8 are the same as 1thru 4, but at a higher power output.

LOCK Switch S4 on the MPD/MPCD is used to control the DLP lock enable. Lock must be selected to unlock the PSA antenna on the DLP to allow control of the weapon seeker by the vertical and horizontal slew commands.

WPN1/2 When a DLP/IDLP is selected, the indirect weapon video display VID5, will display selected GBU-15/AGM-130 weapon functions. Switch S16 can display the legend WPN1 for station 2 or WPN2 for station 8. This provides a one-touch button that changes pod mode and frequency to a different weapon.

MCG/R-ALT Mid-course guidance (MCG) selection is used for aircraft launch zones and control/display functions, and selected to match what is programmed in the weapon.

System checkout You can use a tester to simulate the pod and troubleshoot the data link pod system. The 224 aircraft pod interface test set (APITS) (fig. 3–38) is used to verify/fault isolate the AXQ-14 DLP and the ZSW-1 IDLP systems at the umbilical, pylon, and fuselage disconnects. This is accomplished using interconnect cables from the APITS to the appropriate disconnects on the umbilical, pylon, or fuselage. A cable connects the APITS to the aircraft utility power receptacle for test set power.

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The test routines reside in electrically erasable programmable read-only-memories (EEPROM) in the APITS and can be executed in automatic or manual mode.

The remote control panel (RCU) (fig. 3–38) is a portable, handheld interface used to guide an operator through various diagnostic tests. Eight pushbutton switches mounted around the perimeter of the display are used to allow the user to select various test options.

Figure 3–38. 224 aircraft pod interface test set and remote control panel.

The AUTO/MANUAL pushbutton defaults to AUTO and NEXT mode of operation and AUTO is displayed on the RCU display. When AUTO pushbutton is pressed, AUTO is removed and MANUAL is displayed as the mode of operation. In MANUAL mode, MANUAL and STEP are displayed instead of AUTO for each test.

The NEXT pushbutton provides a mode to advance the test set to the next screen or test sequence. The BACK pushbutton provides a mode to go back to the previous screen within a test routine. The REPEAT pushbutton provides a mode that will run the previous test.

The RESTART pushbutton provides various break points for the aircraft test routines. RESTART goes back to the start of the test routine. This allows the operator to exit from a test routine and return to Power-up BIT PASSED and continue with other tests.

The REPEAT pushbutton provides retesting of the failures by repeating only the section of the test routine that is of concern.

If, after completion of the test, the operator needs to know the value of any failed measurements that the test set made, the operator presses and releases the RECALL pushbutton and the measured value is displayed. If more than one test parameter was measured, press the RECALL pushbutton again and the second parameter is displayed. Pressing RECALL switch again will display the next failure. Repeat this procedure until all recorded failures are displayed.

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Self-Test Questions After you complete these questions, you may check your answers at the end of the unit.

229. F-15E LANTIRN navigation pod 1. Which warning light comes on if the aircraft radar altitude descends below 75 percent of the TF

set clearance value plus a predictive variable based on negative vertical velocity?

2. What warning light comes on to warn the aircrew of an obstacle appearing in the aircraft flight path as detected by the TF radar and requiring more than 2 G to clear?

3. Which component provides the operator interface for the radar altimeter, TFR system, and NAV FLIR sensor?

4. What throttle grip switch provides the control and interface to enable the FLIR snap-look and look-into-turn alternatives?

5. What are the two functional subsystems of the NAV POD?

6. List the NAV FLIR submodes.

7. When in TF radar mode, what submode provides the best combination of system operational parameters?

8. Where is the NAV FLIR active submode displayed?

9. How does the LANTIRN TFR subsystem produce TF commands and where are they displayed for MAN and AUTO-TF modes of operation?

10. List the modes of the TFR subsystem.

11. How many types of BIT does the LANTIRN system have and what are they?

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230. F-15E data link pod 1. With what weapons is the DLP used to communicate?

2. What does the video recorder provide?

3. How many operating modes are there for the DLP and what are they?

4. How does the operation command data get from the aircraft to the DLP?

5. Which DLP function allows the aircraft and weapon data link to be tested before weapon release?

6. Which DLP function allows the pod to talk to the correct weapon and has eight operating modes available labeled A, B, N1, N2, J1, J2, C, and D?

7. Which function must be selected to unlock the PSA antenna on the DLP to allow control of the weapon seeker by the vertical and horizontal slew commands?

8. What would you use to check out the DLP system or to troubleshoot it?

Answers to Self-Test Questions 224 1. A data link system that transfers information at high rates. The data is encrypted to provide security, and is

jam resistant to yield high reliability in a hostile RF environment. It also provides the ability to interconnect many widespread sources and users of information.

2. The JTIDS mode control panel. The modes include FDL operating mode, battery support mode, mission channel selection, FDL crypto code zeroizing, and the TACAN/FDL master reset.

3. FDL RT. 4. FDL RPS. 5. It holds three 3.0VDC lithium batteries. The batteries provide enough power to store the crypto codes and

run the internal clock when system power is off. 6. Castle switch. 7. Located on the right throttle grip and controls the acquisition symbol on the FDL display. 8. The CC. 9. The CC for the F-15C/D and the ADCP for the F-15E.

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10. They provide an interface from the UFCs and HOTAS to the ADCP for control of aircraft systems. HOTAS controls for the FDL include the control stick grip castle switch and auto ACQ switch, throttle control TDC switch, and hand controller switches. AIU 1 also receives an FDL LRU fail discrete that drives asp fault indicator 3 and the AV BIT fail light.

11. The RICP and ICSCP. 12. Battery power allows retention of the initialization data and crypto variables. In this position, the crypto

variables cannot be zeroed using the CIPHER switch action. To load crypto variables, the MODE switch must be in HOLD or OFF.

13. NORM allows full participation in the Link-16 community and provides full transmit and receive ability. 14. The SIT. 15. OWN DATA. 16. FINE. 17. The information is spread out over a wide frequency spectrum. The FDL RT will hop around on the 51

possible FDL channels in a pattern controlled by the Crypto code key loaded in the FDL system. 18. It is the first terminal, which is transmitting on the network, and its clock establishes system time. All other

network participants enter the network via the NTR. 19. Properly loaded FDL TERM data, a common channel number (either mission or fighter channel), a host to

establish the NTR, time sync with the NTRs established time, and common crypto codes.

225 1. Two. 2. The AIU1A channel within AIU 1. 3. Serial. 4. AIU1A. 5. TACAN RT. 6. CARA RT. 7. AIU 2. 8. AIU1B.

226 1. It provides interrogation generation, reply evaluation, radar interface functions, monitors failures and

provides system status and LRU status signals to the avionic interface units. 2. It consists of 10 dipole antennas arranged on the radar’s planner array (antenna), two power dividers, and a

hybrid coupler. 3. Into two sets of five antennas, one set for the sum channel and one set for the difference channel. 4. To enable the AAI system to detect the direction of arrival of the reply signal. 5. It controls manual AAI interrogation initiation. 6. It contains the switches used to control the AAI system. 7. The UFC and rear UFC. 8. The programming switches on the front of the IRE. 9. The UFC’s submenu. 10. They are initiated either manually using the multifunction switch on the throttle quadrant or automatically

by the radar data processor when radar switches to tracking mode. 11. Pressing the multifunction switch on the right throttle grip to the IFF interrogate position initiates the AAI

interrogation. The interrogate command signal from the multifunction switch enters interrogate control circuitry within the radar data processor (081 or 082). The interrogate command is sent as part of a data word that is transmitted between the radar PSP (042 or 044) and the APX-114. Data to the APX-114 is made up of information required for interrogation and reply processing. The APX-114 tests the data word for parity (validity).

12. They are translated into target symbology by the radar system and displayed on the MICP or MPD/MPCD.

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13. The M/4 light remains on steady to indicate that the APX-114 has not received time decoded video from the KIV after receiving a valid mode 4 reply.

227 1. Band 1, band 2, and band 3. 2. The E and G frequency band. 3. The G and H frequency band. 4. The H, I, and J frequency band. 5. Indicate when a specific set (band) has failed. 6. During RWR IBIT, the TEWS display unit displays ICMS program information. When band 3 jamming is

present, the jammed threat is identified as being jammed. 7. The MPCD ICMS status on the tactical situation display and PACS A/A display and BIT information is

selected from the TEWS BIT page. 8. It receives band 1 and band 2 RF amplifier RF samples and applies a composite output to the RWR high

band receiver. 9. Tuning commands, amplitude and/or frequency jamming output signals, and frequency information. 10. Final TWT amplification of band 3 oscillator signals, operating voltages to the band 3 oscillator, and pulse

and CW RF to the magic tee. 11. It combines pulse and CW RF signals into a single signal and then splits the combined signal and sends the

signals to the band 3 forward up and down antennas. 12. A small matrix ROM device that provides band 1 and band 2 oscillators with predetermined program

inputs. 13. Provides initial system amplification of the received RF threat signals from the RWR high band antennas

and routes them to the band 3 oscillator. 14. They prevent pressure loss, arcing and contamination. 15. When the aircraft is flying, the left MLG WOW relay is energized and passes a ground to the ICMS

oscillators. When the aircraft is on the ground, the oscillators send an inhibit command to the amplifiers that prevents accidental high power RF signal transmission when the aircraft is on the ground.

16. Radar operating frequencies, aircraft altitude, and aircraft speed to the band 3 oscillator.

228 1. Band 1 and band 2 are transmit only systems under the control of the RWR during automatic mode. 2. The programmers. 3. During RWR or ICMS failure. 4. A low power RF signal, CW and high level modulation commands from the oscillator for automatic

leveling and AM modulation. 5. To the summing network. 6. It is sent to the band 1 antenna over a coaxial transmission line. 7. It is transmitted through pressurized waveguides to the band 2 antenna. 8. An open X around the jammed threat on the TEWS display, or if no RWR/ICMS jamming correlation

exists, the open X will be displayed at the center of the TEWS display.

229 1. LOW ALT. 2. OBST. 3. Sensor control panel. 4. Right multifunction switch. 5. NAV FLIR and TFR. 6. NORM and BRST. 7. NORM submode. 8. Next to S1 on the NAV FLIR submenu on the UFC.

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9. It uses radar emissions to the ground and their return signals to produce TF commands and these commands are used as steering cues displayed on the HUD, EADI, and TF display.

10. NORM, ECCM, WX1 or WX2, LPI, and VLC. 11. Two types of BIT: periodic and initiated.

230 1. With the GBU-15 guided bomb or AGM-130 guided missile. 2. The video recorder provides recording and playback functions when the pod is powered. It also allows pre-

recorded tapes to be loaded for a mission and allows tapes which were recorded during a mission to be used in debrief.

3. There are three DLP operating modes, AUTO, MAN, and FWD. 4. The operation command data is then sent to the DLP on the PACS/DLP MUX bus. 5. TEST. 6. MODE. 7. LOCK. 8. The 224 APITS.

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Unit Review Exercises Note to Student: Consider all choices carefully, select the best answer to each question, and circle the corresponding letter. When you have completed all unit review exercises, transfer your answers to the Field-Scoring Answer Sheet.

Do not return your answer sheet to the Air Force Career Development Academy (AFCDA).

80. (224) Which component would be used to zeroize the F-15C/D fighter data link (FDL) crypto code? a. Sensor control panel. b. FDL remote power panel. c. JTIDS mode control panel. d. Main communications control panel.

81. (224) On the F-15, what system has all signals passed through the fighter data link (FDL) receiver/transmitter? a. Instrument landing system (ILS). b. Identification friend or foe (IFF). c. Tactical air navigation (TACAN). d. Control augmentation system (CAS).

82. (224) Which F-15C/D component sends fighter data link (FDL) symbology positioning information to the programmable signal data processor (PSDP)? a. Central computer (CC). b. JTIDS mode control panel. c. FDL receiver/transmitter (R/T). d. Main communications control panel.

83. (224) Which component would be used to zeroize the F-15E fighter data link (FDL) crypto codes? a. Sensor control panel. b. Up-front controller (UFC). c. Intercommunications set control panel (ICSCP). d. Remote intercommunication control panel (RICP).

84. (225) What functions as the F-15E up-front controller’s (UFC) primary bus controller? a. Avionics interface unit (AIU) 2. b. AIU2A. c. AIU1A. d. AIU1B.

85. (225) On the F-15E, where does the engine monitor display (EMD) send its serial data? a. Avionics interface unit (AIU) 2. b. AIU2A. c. AIU1A. d. AIU1B.

86. (226) The F-15E radar data processor supplies an air-to-air interrogator (AAI) system interrogate command to what radar component when the radar mode is compatible? a. Oscillator. b. Transmitter. c. Antenna array. d. Programmable signal processor.

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87. (226) How many dipole elements does the F-15 antenna array have? a. 5. b. 10. c. 15. d. 20.

88. (226) On the F-15 air-to-air interrogator (AAI) display, what is a low confidence target displayed as? a. Circle. b. Triangle. c. Diamond. d. Rectangle.

89. (227) What action must be taken before flight when any internal countermeasures system (ICMS) component is not installed in bay 5? a. No action is required. b. Ballast must be installed. c. Ballast must be removed. d. F–15 aircraft cannot fly with any ICMS component not installed.

90. (227) How is the final amplified band 1 radio frequency (RF) signal transmitted to the band 1 antenna? a. Through the summing network. b. Over a coaxial transmission line. c. Through pressurized waveguides. d. Through unpressurized waveguides.

91. (227) What component performs the final traveling wave tube (TWT) amplification of band 3 oscillator signals? a. Magic tee. b. Preamplifier. c. Summing network. d. Band 3 RF amplifier.

92. (227) What component is used to combine band 3 forward amplifier continuous wave (CW) radio frequency radiation (RF) and pulse RF signals into a single signal, split the combined signal, and send the split signal to the forward up and forward down band 3 internal countermeasures system (ICMS) antennas? a. Magic tee. b. Preamplifier. c. Band 3 oscillator. d. Summing network.

93. (227) The preamplifier receives four radio frequency (RF) inputs and one agile tuning unit 1/built-in test (BIT) RF signal input from what components? a. Low band processor and band 3 aft RF amplifier. b. High band processor and band 3 aft RF amplifier. c. Low band receiver and band 3 aft RF amplifier. d. High band receiver and band 3 aft RF amplifier.

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94. (227) Into which component of the internal countermeasures system (ICMS) are the band 3 operational flight program (OFP) and preflight message (PFM) loaded? a. Preamplifier. b. Band 3 oscillator. c. Band 3 aft RF amplifier. d. Band 3 forward RF amplifier.

95. (228) What data bus does radar warning receiver (RWR) use to control the F-15C internal countermeasures set (ICMS) band 1 and band 2 oscillators? a. 1553. b. H009. c. Serial data bus. d. Manchester data bus.

96. (228) What do the band 1 and band 2 amplifiers use to produce leveled high power radio frequency (RF) signals? a. 1 traveling wave tube (TWT) circuit. b. 2 TWT circuits. c. 3 TWT circuits. d. 4 TWT circuits.

97. (228) Where is the aft amplifier pulse radio frequency (RF) signal sent during band 3 operation? a. To the preamplifier. b. Both elements of the aft antenna. c. To the pulse element of the aft antenna. d. To the continuous wave element of the aft antenna.

98. (229) On the F-15E LANTIRN navigation pod system, what is the primary display format used? a. Hot/cold mode. b. Fly up enable mode. c. Terrain following (TF). d. Forward looking infrared (FLIR) mode.

99. (229) On the F-15E, what component provides the operator interface for the radar altimeter, terrain following radar (TFR) system and NAV FLIR (navigation forward looking infrared) sensor? a. Control stick grip. b. Sensor control panel. c. Fly up enable control panel. d. Intercommunication set control panel.

100. (230) What is the purpose of the F-15E data link pod (DLP)? a. The pod is used to communicate with the GBU-15 guided bomb or AGM-130 guided missile. b. The pod uses radar emissions to the ground and their return signals to produce terrain following commands. c. The pod provides the ability to precisely geo-locate points of interest and conduct surveillance activities day or night, in adverse weather conditions. d. The pod permits day or night delivery of infrared (IR) guided AGM-65D/G maverick weapons and allows acquisition and tracking of targets for automatic handoff to the maverick missiles.

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Glossary

Abbreviations and Acronyms °C degrees Celsius

A/A air-to-air

AAI

AAI RT

air-to-air interrogation

IFF interrogator receiver/transmitter

AC alternating current

ACQ acquisition

A/D analog-to-digital

ADC air data computer

ADCP advanced display core processor

ADF automatic direction finding

ADI attitude director indicator

ADP air data processor

ADS autopilot disengage switch

AFCS automatic flight control system

AFMSS Air Force mission support system

A/G air-to-ground

AGR air-to-ground ranging

AHRS attitude heading reference system

AIC air inlet controller

AIU avionics interface unit

ALC automatic leveling circuit

ALT

AMI

altitude

airspeed Mach indicator

AOA

A/P

angle of attack

autopilot

APITS aircraft pod interface test set

ARI aileron rudder interconnect

ASA acceleration sensor assembly

ASP

AUTO

AV

AWACS

avionics status panel

automatic

avionics

Airborne Warning and Control System

G–2

BC bus controller

BCN beacon

BCP built-in test (BIT) control panel

BIT built-in test

BMM bulk memory module

BSM bulk storage module

BST boresight

CARA combined altitude radar altimeter

CAS control augmentation system

CASI control augmentation system interconnect

CAU

CB

counter accelerometer unit

circuit breaker

CC central computer or correct code

CCC central computer complex

CCDL cross channel data link

CDC career development course

CDIP continuously displayed impact point

CDU control display unit

CENC convergent exhaust nozzle control

CFI computerized fault isolation

CFRS computerized fault reporting system

CFT conformal fuel tanks

CM continuously monitored

CND could-not-duplicate

CRT cathode ray tube

CSBPC control stick boost and pitch compensator

CW

D/A

continuous wave

digital to analog

DC direct currrent DCL declutter DEEC digital engine electronic control

DFTA directional feel trim actuator

DLP

DMP

data link pod

digital map processor

DMS digital map system

G–3

DP digital processor OR data processor

DPM data processing modules

DPS

DSC

differential pressure sensor

digital scan converter

DSS differential stabilator servocylinders

DTED digital terrain elevation data

DTM data transfer module

DVRS digital video recording system

EADI electronic attitude director indicator

EAIC electronic air inlet controller

ECCM electronic counter-countermeasures

ECS environmental control system

ECU environmental control unit

EDM external data management

EDU engine diagnostic unit

EEC engine electronic control

EEPROM

EGI

electronically erasable programmable read-only memory

embedded GPS/INS

EHSI electronic horizontal situation indicator

EHV electrohydraulic valve

EMD engine monitor display

ETI elapsed-time indicator

EU

FBIT

electronics unit

flight BIT

FBS flap blowup switch

FCC flight control computer

FCSTS flight control system test set

F/D frequency-to-digital

FDL fighter data link

FF fuel flow

FINS fixed imaging navigation sensing

FLIR forward look infrared

FTIT fan turbine inlet temperature

FUS fuselage

G gravity

G–4

GBIT ground BIT

GMT ground moving target

GPP general purpose processor

GTWT gridded traveling wave tube

HDU helmet display unit

HEX hexadecimal

Hg mercury

HOTAS hands on throttle and stick

HPRF high pulse repetition frequency

HRM high resolution map

HSI horizontal situation indicator

HUD head-up display

Hz hertz

IA immediate action

IBIT initiated built-in test

ICCP integrated communications control panel

ICMS internal countermeasures system

ICSCP intercommunication set control panel

IDC input data converter

IDLP improved data link pod

IF intermediate frequency

IFF identification friend-or-foe

IG indicator group

ILS instrument landing system

INS inertial navigation system or inertial navigation set

INST

INTL

instrument

internal

INU

INV

inertial navigation unit

inverse video

I/O input/output

IOM input output modules

IOP input output processor

IP

IPF

IPM

identification point

Interference Protection Feature

image processor module

G–5

IR infrared

IRE identification friend-or-foe reply evaluator

ISM intelligent serial module

JHMCS

J-STARS

joint helmet mounted cueing system

joint surveillance and target attack radar system

JTIDS

LANC

joint tactical information distribution system

local application control bus

LANTIRN

lbs

LCD

LED

low altitude navigation and targeting infrared for night

pounds

liquid crystal display

light emitting diodes

LFTA longitudinal feel trim actuator

LIT

LO

look-into-turn

local oscillator

LOS line of sight

LPI low probability of intercept

LPRF

LRBST

low pulse repetition frequency

long range boresight

LRS long-range search

LRU line replaceable units

LTS left tail stabilator

LVDT

MAINT

linear variable differential transformers

maintenance

MC

MCG

mission cartridge

mid-course guidance

MCP mode control panel

MCV master control valve

MHz megahertz

MICP

MIL STD

multiple indicator control panel

Military Standard

MIT mass items

MLG main landing gear

MLV

mm

MN

memory loader verifier

millimeter

mission navigator

G–6

MOC maintenance operation center

MPCD multipurpose color display

MPD

MPDP

multipurpose display

multipurpose display processor

MPRF medium pulse repetition frequency

MTBMA mean time between maintenance actions

MTF manual terrain-following

MUX

NAV

multiplex

navigation

NAV FLIR navigation forward look infrared

NCl navigation control indicator

NMI nautical miles

NOWS overload warning system (OWS) not operational

NTR net time reference

ODC output data converter

OFP

OPS

ORT

operational flight program

operations

operational readiness test

OWS overload warning system

PACS programmable armament control set

PB push button

PBIT periodic built-in test

PFM preflight message

PLV

POST

programmable loader verifier

power-on self-test

PPH pounds per hour

PPI pixels per inch

PRC pitch rate controller

PRCA pitch and roll channel assembly

PRF

P/S

pulse repetition frequency

parallel to serial

PSA pressure sensor assembly or 3-phase scanned array

PSDP programmable signal data processor

psi pounds per square inch

psia pounds per square inch absolute

G–7

PSP programmable signal processor

PTC pitch trim controller

PVU precision velocity update

PYL pylons

Qc dynamic pressure

RBM

RCU

real beam map

remote control panel

RDP

RDR

RECCE

radar data processor

RADAR

reconnaissance

RF radio frequency

RFO

RGH

radio frequency (RF) oscillator

range dated high

RICP remote intercommunication control panel

ROM read-only memory

RPM

RPS

revolutions per minute

remote power supply

RSA rate sensor assembly

RT remote terminal

R/T receiver/transmitter

RTS right tail stabilator

RWR radar warning receiver

SAI standby airspeed indicator

SAR synthetic aperture radar

SDRS signal data recording system

SFS

SIF

SIT

stick force sensor

selective identification feature

situation display

SOV shutoff valve

SRS short-range search

SRU shop replaceable unit

SS supersearch

STT single target track

TACAN tactical air navigation

TC theater cartridge

G–8

TCN TACAN

TCTO time compliance technical order

TDC

TDM

target designator control

timing and discrete module

TDMA time division multiple access

TEWS tactical electronic warfare system

TF terrain following

TFR

TGT

terrain following radar

target

TO technical order

TOT takeoff-trim

TSD tactical situation display

TT total temp

TWS track-while-scan

TWT traveling wave tube

UFC up-front controller

UHF ultra high frequency

VAC volts alternating current

VDC volts direct current

VHF very high frequency

VHSIC very high-speed integrated circuit

VI indicated velocity

VLC

VME

very low clearance

VERSA module Eurocard

VSD vertical situation display

VSI vertical speed indicator

VTRS video tape recording system

VTS vertical scan

VVI vertical velocity indicator

WFOV wide-field-of view

WNG wings

WOW weight-on-wheels

WSO

ZM

weapons system officer

zoom

Student Notes

AFSC 2A374 2A374N 02 1310 Edit Code 01