NASA TECHNICAL NOTE H NASA TN D-7420

FILECOPY

FLIGHT INVESTIGATION OF

A STRUCTURAL MODE CONTROL SYSTEM

FOR THE XB-70 AIRCRAFT

by

Wilton P. Lock, Eldon E. Kordes, and James M. McKay

Flight Research Center

Edwards, Calif. 93523

and

John H. Wykes

North American Rockwell Corporation

Los Angeles, Calif. 90009

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C • OCTOBER 1973

1 Report No

NASA TN D-74202 Government Accession No 3 Recipient's Catalog No

4 Title and Subtitle

FLIGHT INVESTIGATION OF A STRUCTURAL MODE CONTROLSYSTEM FOR THE XB-70 AIRCRAFT

5 Report DateOctober 1973

6 Performing Organization Code

7 Author(s)

Wilton P. Lock, EldonE. Kordes, James M. McKay (Flight ResearchCenter), and John H. Wykes (North American Rockwell Corporation)

8 Performing Organization Report No

H-732

9 Performing Organization Name and Address

NASA Flight Research CenterP.O. Box 273Edwards, California 93523

10 Work Unit No

743-36-22-0011 Contract or Grant No

12 Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

13 Type of Report and Period Covered

Technical Note

14 Sponsoring Agency Code

15 Supplementary Notes

16 Abstract

A flight investigation of a structural mode control system termed identical loca-tion of accelerometer and force (ILAF) was conducted on the XB-70-1 airplane. Dur-ing the first flight tests, the ILAF system encountered localized structural vibrationproblems requiring a revision of the compensating network. After modification,successful structural mode control that did not adversely affect the rigid body dynam-ics was demonstrated.

The ILAF system was generally more effective in supersonic than subsonicflight, because the conditions for which the system was designed were more nearlysatisfied at supersonic speeds. The results of a turbulence encounter at a Machnumber of 1. 20 and an altitude of 9754 meters (32, 000 feet) indicated that the ILAFsystem was effective in reducing the vehicle's response at this flight condition.

An analytical study showed that the addition of a small canard to the modalsuppression system would greatly improve the automatic control of the higher fre-quency symmetric modes.

17 Key Words (Suggested by Author(s))

Structural mode controlElastic mode controlControl systems

18 Distribution Statement

Unclassified - Unlimited

19 Security Qassif (of this report)

Unclassified

20 Security Classif (of this page)

Unclassified

21 No of Pages

82

22 Price*Domestic, $3 75Foreign, $6 25

" For sale by the National Technical Information Service, Springfield, Virginia 22151

FLIGHT INVESTIGATION OF A STRUCTURALMODE CONTROL SYSTEM FOR THE XB-70 AIRCRAFT

Wilton P. Lock, Eldon E. Kordes, and James M. McKayFlight Research Center

and

John H. WykesNorth American Rockwell Corporation

INTRODUCTION

Structural flexibility must be considered in the design of large, high performance air-craft. Aeroelastic deformation affects not only basic flight characteristics such as performance,controllability, handling, and ride qualities, it also increases structural loads and fatigue. Theproblems associated with flexible aircraft are not new, however, the technology required tocontrol structural dynamics behavior was first developed for the inherently aerodynamicallyunstable launch vehicles. The success of the launch vehicle systems prompted the develop-ment of similar systems for aircraft, including systems for the control of structural moderesponse (ref. 1).

Two flight-test programs sponsored by the United States Government were initiated toachieve elastic mode control in large, flexible aircraft. The first program, which was conduct-ed by the Boeing Company and Honeywell, Inc., was devoted to the development of a loadalleviation and mode stabilization (LAMS) system for the B-52 airplane. Extensive analyticaland simulator studies were used to define the details of the system and also to demonstratethe system's potential (ref. 2).

The second program, which used the XB-70 airplane, was undertaken to develop anelastic mode control system called identical location of accelerometer and force (ILAF). Theconcept on which it is based was first developed in the analytical study reported in refer-ence 3. The design of the ILAF control system is described in reference 4. Reference 5discusses the analytical design and briefly evaluates the first flight-test results.

The ILAF system flight-test program was conducted to investigate the ILAF system con-cept rather than to develop an optimum operational system. To flight test the ILAF systemunder well-controlled conditions, the aerodynamic shaker system described in reference 6 wasused. The shaker system was capable of exciting the first four symmetric structural modes.

This paper describes the integration of the ILAF system with the XB-70's control systemand presents test results obtained in flight at subsonic and supersonic airspeeds. This reportalso includes the performance calculated for a number of altitudes and Mach • numbers thatare representative of the flight condition

SYMBOLS

Physical quantities in this report are given in the International System of Units (SI) andparenthetically in U S Customary Units. The measurements were taken in Customary UnitsFactors relating the two systems are presented in reference 7

/ forcing frequency

g acceleration due to gravity

gt structural mode damping (structural plus aerodynamic)

hp pressure altitude

A, = 0.11

Kf ) control system gain associated with subscripted parameter

/ distance from flight augmentation control system accelerometer to vehicle center ofgravity

M Mach number

n. normal load factor

q pitch rate about Y-axis

s Laplace operator

A increment

5 control surface deflection

5, wingtip deflection

6 rate of control surface deflection

8SV shaker-vane deflection, positive deflection produces positive lift force

r?( ) generabzed coordinate, subscript indicates mode

a root mean square value

at = 57 3-7—, real part of root of the characteristic equation as in s + a, ± a>,

<E> power spectral density

0 phase angle

0, ith normalized mode shape

0/ slope of i^h normalized mode

cj forcing frequency

co, natural frequency of itn mode

Subscripts

c command

e all elevens except inboard

P pilot's station

wg gust velocity

6 pitch rate

1 inboard eleven

2,3,4,5,6 individual eleven panels outboard of inboard panel

A dot over a quantity denotes the time derivative of that quantity

ABBREVIATIONS

BP butt plane

CADS central air data system

FACS flight augmentation control system

FS fuselage station

HS canard horizontal station

ILAF identical location of accelerometer and force

LAMS load alleviation and mode stabilization

TEST APPARATUS

Airplane

The XB-70 airplane (fig. 1) is a large, delta-winged, multiengme jet airplane designed byNorth American Rockwell Corporation for supersonic cruise at a Mach number of 3 0 andaltitudes above 21,336 meters (70,000 feet). Two airplanes, designated XB-70-1 and XB-70-2,were built This investigation was conducted using the XB-70-1. The general configurationand overall dimensions of the XB-70-1 are shown in figure 2 The basic design incorporatesa thin, low-aspect-ratio wing with a leading edge swept back 65.57°, folding tips, twin verticalstabilizers, and movable canards with traihng-edge flaps The XB-70-1 was manufactured withthe wings mounted at a dihedral angle of zero.

Stability Augmentation System

The flight-test data reported herein were obtained with the stability augmentation system,called the flight augmentation control system (FACS), engaged A brief description of thissystem is given below. A more detailed description is given m reference 4 along with thesystem's frequency response characteristics.

The FACS is a conventional command augmentation system designed to improve handbngqualities by operating simultaneously with the pilot's manual control system. A block dia-gram of the pitch augmentation system is shown as the unshaded blocks in figure 3. Pilotcommands for pitch control are processed through two paths to the eleven actuation system.The first path is purely mechanical. Pilot commands are transmitted to the master cylinder,which outputs through linkage to the eleven actuators to produce the desired control surfacemotion without force feedback to the controls. In the second path, pilot commands actuatea transducer that provides electrical signals that are in turn electrically summed with signalsfrom two aircraft response sensors (a gyro and an accelerometer). The combined signal isfiltered to reduce the transmission of high frequency signals to the servo. The signal is thengam scheduled according to altitude and Mach number information provided from the centralair data system (CADS). Finally, the signal positions the pitch servo, which sums mechani-cally with the pilot commands from the first path to drive the inboard eleven panel. Themotion of the inboard panel commands the motion of all the outboard panels (fig. 2) asdescribed below.

Redundancy was accomplished by dualizing the electronics from the sensors to the servo.Other safety provisions were incorporated to provide self-monitoring and to control servocentering rates upon disengagement.

The FACS controls only the elevens in the pitch mode. The elevens for each wing aredivided into six segments to prevent control surface binding under aerodynamic loading. Withthe wingtips in the 0° position, all five outboard panels are slaved to the inboard panel. Inthe 25° and 65° wingtip positions, the two outermost panels are disengaged and centered, andthe three remaining outboard panels are slaved to the inboard panel.

ILAF Structural Mode Control System

The ILAF system was first developed under a U.S. Air Force study contract using earlyXB-70 design information to develop an analytical model (ref. 3). The objective of the studywas to design a simple, stable system to maintain system stability and to provide damping tothe structural modes The system was to operate over a wide range of altitude and Machnumber conditions and vehicle weights. The principle on which the ILAF system is based isexplained in reference 3, and a more detailed description of the design and a performanceanalysis are given in reference 4. Briefly stated, the design synthesis locates the structuralmotion sensor (accelerometer) as close as possible to the force generator (eleven).

Design limitations.— Several constraints were imposed on the ILAF system. It was tocontrol only the first four airplane symmetric structural modes (those less than 8 hertz) Itwas to utilize the existing pitch FACS, but no modification was to be made to the FACSthat would affect the basic handling qualities of the airplane. A further requirement was thatthe pitch FACS was to be in operation before the ILAF system could be engaged in flight,and, to preclude instability problems (ref. 4), it was to be impossible to engage the ILAF sys-tem on the ground.

The ILAF system was mechanized as a dual channel system to make it compatible withthe existing pitch FACS and to make it possible to utilize the existing failure protectioncircuitry.

Description.— The shaded blocks in figure 3 show the ILAF system components incorpo-rated into the pitch augmentation system. A primary structural motion sensing accelerometerwas located in each wing near the number 2 elevon panel hinge line, and a secondary acceler-ometer was located near the airplane's center of gravity. Together, the three accelerometersprovided the ILAF system input signals (ref. 4) The signals from the two wing accelerome-ters were halved and then summed to eliminate whole-vehicle roll motion and antisymmetricmodes. The signal from the accelerometer near the center of gravity was subtracted fromthis signal to eliminate vehicle rigid body plunge motion

The combined signal from the accelerometers was passed through a notch filter networkto a manually adjusted gain control knob in the cockpit. The notch filter was designed toattenuate the signal associated with elevon natural frequencies at approximately 20 hertz.From the gain control, the signal passed through the compensation network into the pitchaugmentation system electronics. The ILAF system compensation shaping, shown in figure 4,was a lead-lag network designed to improve the damping for the third mode (ref. 4).

To protect the pitch servo from large amplitude high frequency commands generatedwith the ILAF system, an electronic limiter was added to the pitch electronics that reducedthe servo authority from ±7.5° to ±7.0°.

'The ILAF system electronics were blended into the FACS electronics just in front of thepitch augmentation gain, Kf, (fig. 3), which is an automatic function of the CADS.

To prevent possible instability of the ILAF system on the ground, a switch was installedin the landing gear system that prevented the ILAF system from being operated unless thelanding gear was fully retracted.

Vibration Excitation System

Movable aerodynamic vanes trapezoidal in planform and 0.19 square meter (2 square feet)in area were mounted on each side of the forward fuselage in front of the pilot's station(ref 6). The location of the shaker-vane system relative to the cockpit is shown in figure 5.The vanes constituted an excitation system capable of producing a controlled, oscillatory mo-tion in the XB-70 over the frequency range from 1 4 to 8.0 hertz. The system was installedto determine the dynamic response of the airplane in flight. The shaker-vane system, opera-tional procedure, and safety features are described in detail in reference 6.

INSTRUMENTATION AND RECORDING

The performance of the ILAF system as a structural mode damper was evaluated pri-marily with sensors already installed in the airplane, as shown in figure 6 The only sensorsadded to the airplane were the wing and fuselage accelerometers required for the operationof the ILAF system

The sensors used for the evaluation of the ILAF system are listed in table 1 along withtheir ranges and locations on the aircraft The natural frequencies of these sensors were suchthat their responses were essentially flat up to frequencies commensurate with those beingmeasured

The flight data used to evaluate system performance were recorded on magnetic tape byeither analog or digital techniques, depending upon the parameter and its frequency responserequirements The magnetic tape generated within the airborne system was processed by aground station computer to convert the data to engineering units (ref 6).

A data acquisition system previously installed on the XB-70 (ref. 8) made it possible toevaluate the ILAF system during turbulence encounters This data acquisition package con-sisted of a gust boom installed at the nose of the airplane and the sensors associated with it.The data obtained were used to correct for airplane motion at the nose. The procedureused to reduce the turbulence data is discussed in reference 8.

The analog and digital data acquisition systems are described in references 9 and 10

FLIGHT CONDITIONS

The flight conditions for the ILAF system evaluation tests were generally representativeof the XB-70's flight envelope (fig. 7) The airplane weights, the center of gravity locations,and other flight-test conditions are given in table 2.

TEST PROCEDURE

The tests of the ILAF system were performed in wings-level, Ig, trimmed flight at sev-eral airspeeds. The system was not operated during deliberate maneuvers, and during theopen-loop ILAF system evaluation the system was not operated in turbulence.

Open-Loop Tests

The first tests were performed by increasing the I LAP system gain in the open-loopmode without shaker-vane input and observing airplane response. The performance of theILAF system was then evaluated by setting shaker-vane input at a constant amplitude andmeasuring aircraft response and control system motion with and without the ILAF systemoperating. Shaker-vane amplitude was set, and the frequency was slowly increased to ap-proach and excite each symmetrical mode up to 8 hertz. The system was mechanized sothat a relay was intentionally opened between the output of the ILAF system electronics andthe input to the pitch FACS electronics, allowing the ILAF system from the accelerometersthrough the shaping and compensation network to be monitored during flight without disturb-ing the airplane. The frequency sweeps were repeated at higher levels of shaker-vane ampli-tude.

Closed-Loop Tests

Closed-loop tests were performed in the same manner, but with the interconnecting relayclosed. The tests were started with a low value of ILAF system gain and zero shaker-vaneinput and continued by increasing the gain while overall system performance and aircraft re-sponse were observed. Then the shaker-vane system was used to excite the structural modeswith and without the ILAF system operating.

The ILAF system was further tested during encounters with turbulence; the response ofthe airplane was measured with and without the ILAF system engaged.

Structural damping information was obtained for each mode with and without the ILAFsystem engaged. Because of fuel consumption and the associated change in airplane mass, itwas necessary to reestablish the mode being investigated. Once excited, the modal frequencywas allowed to stabilize. The shaker-vane system was then shut down abruptly, and datawere recorded until telemetry indicated that the responses were completely damped out.

During all the ILAF system evaluation tests, modal frequencies and amplitudes were mon-itored by a test engineer on the ground, who used telemetered data on airplane response andsystem operation. The signal from the nose ramp accelerometer at fuselage station 7.43meters (292.5 inches) was used to observe airplane structural response. The parameters tele-metered for ground monitoring of the ILAF system and aircraft structural integrity were mon-itored on strip charts.

DATA ANALYSIS

Detailed response calculations were made for the XB-70 airplane and reported in refer-ences 3 to 5 However, the conditions analyzed were specific design conditions not readilyobtainable in flight. To more readily compare analytical results with the response of the air-plane during these tests, calculations were made for the weight, altitude, and Mach numberconditions of the actual tests. These conditions are shown in table 2 for Mach numbers of0.87, 0.86, and 1.59. The analysis used the updated mass and stiffness data given in refer-ence 6 for the basic airplane and for the airplane with FACS engaged. The analysis of air-plane response with the ILAF system engaged is described in appendix A.

RESULTS AND DISCUSSION

System Stability

Original I LAP system.— Initially it was planned to evaluate the ILAF system stabilityduring climbout at a high subsonic speed and in a heavyweight condition. During this ini-tial stability evaluation, a limit cycle occurred at a low ILAF system gain. The first indi-cation of a limit cycle was a large amplitude oscillation observed in the pitch servo. Thetelemetry data indicated that the oscillation occurred at 26 hertz. Additional flight testsfor system stability in a supersonic, mediumweight flight condition and a high subsonic,lightweight flight condition confirmed the presence of a limit cycle. During one of the threeILAF system stability evaluation flights, a failed wing accelerometer allowed the system loopgain to approximately double before the limit cycle appeared. However, at the later twotest conditions the dominant frequency in the pitch servo occurred at 12 hertz

Figure 8 illustrates the various frequencies and amplitudes measured during the ILAFsystem stability tests with and without the ILAF system operating. Data are presented forfuselage acceleration, wing acceleration, inboard elevon deflection, and pitch servo deflection.The plots are composed of data from frequency analysis and time histories from the threeclosed-loop stability tests. The frequency analysis identified the peak responses, and the timehistory data aided in determining the magnitude of the response.

The data in figure 8 show that the predominant frequency on the pitch servo is 12hertz with a smaller peak at 26 hertz with the ILAF system engaged. On the other hand,the inboard elevons responded to the 12-hertz signal, but there were no noticeable elevon de-flections at 26 hertz. The fuselage accelerometer, which was located in the nosewheel well,indicated only a 12-hertz response, whereas the wing accelerometer did not respond appre-ciably to frequencies below 20 hertz.

A series of ground vibration tests was conducted to verify the source of the 12- and26-hertz vibrations and to determine any other significant mode of vibration. One test con-sisted of oscillating the pitch servo at various frequencies and amplitudes, forcing the elevonsurfaces to move and excite the vehicle structure. The fuselage accelerometer revealed a num-ber of significant modes above 10 hertz. The wing accelerometer also sensed large amplituderesponse from 10 hertz to over 32 hertz. Both accelerometers reflected an elevon mode con-tribution near 20 hertz which had been noted in previous ground tests.

The mounting arrangement of the ILAF system wing accelerometers was also checked foreffects on frequency response. Figure 9 shows the results of the tests. The ice and watermentioned in the figure were used to provide adequate cooling for the accelerometers for theduration of each flight. These test results showed that the mounting arrangement made nocontribution to the 26-hertz oscillation, but they did indicate poorly damped modes near 29hertz and 38 hertz for the accelerometer with mount.

The limit cycle instability measured from flight tests is shown schematically in figure 10.The two dominant frequencies are 12 hertz and 26 hertz for the fuselage and wing acceler-ometers, respectively. The 26-hertz oscillation was measured on the pitch servo; however, as-sociated elevon motion was not noticeable. A hydromechanical coupling must, therefore, haveexisted between the servo and accelerometers, causing a sustained oscillation between them asshown in the figure. Airplane response decreased with increasing frequency near 26 hertz,but at 26 hertz the gain of the accelerometer and ILAF system shaping network increased.

8

(See fig. 4.) The amplitude of the servo is also decreased at this frequency because of servorate limiting This combination of factors allowed a high gam closed-loop signal transmissionwhich caused the limit cycle near 26 hertz. The attenuation of the overall system, however,was great enough to eliminate any limit cycles above 26 hertz.

The 12-hertz oscillation is shown schematically in figure 10 as a mechanical couplingmode between the elevens and the ILAF system fuselage accelerometer. The ground test in-dicated that the relative response of the fuselage accelerometer was increasing with increasingfrequency in the 10- to 16-hertz range, and that the relative gain through the ILAF systemshaping network (fig 5) was decreasing with increasing frequency for the same range. How-ever, the gain magnitude was still greater than unity at these frequencies. The pitch servowas capable of driving the elevens at 12 hertz, reinforcing the signal until the limit cycle ex-isted. Relocating the ILAF system fuselage accelerometer to a less sensitive fuselage locationcould have alleviated this problem, however, this was not possible within the scope of theprogram. Instead, the compensating shaping network was revised.

Revised system — To prevent further limit cycle response and to obtain benefit fromusing the ILAF system, the shaping network was revised, even to the extent that the per-formance of the ILAF system was impaired. Acceptable system performance in the thirdmode, along with eliminating the limit cycle problems beyond 8 hertz, would have requiredthe mechanization of the desired shaping network shown m figure 11 to provide phase anglelead near 5 hertz. Several nonlinear filtering techniques were considered, but, again, this taskwas not within the scope of the program. The revised shaping network shown in figure 11was a compromise between potential performance, simplicity, reliability, and modificationtime. Phase characteristics were discarded in favor of satisfactory amplitude ratio, partly be-cause electrical noise became a problem with increased phase lead. In addition, as indicatedin figure 11, although attenuating the ILAF system signal in frequencies of 12 and 26 hertzwould have eliminated the limit cycle problem, the phase lag associated with a 5-hertz leadmight have adversely affected system performance at the third mode frequency.

After the modification of the shaping network was completed, the ILAF system wasflight tested for stability margins at each of the three flight conditions previously investi-gated ILAF system gain was increased to a value of 6 (0.043 radian per g) before anyhigh frequency oscillations were detected for the heavyweight condition at M = 0.87. Amaximum gain of 10 (0 185 radian per g) was attained for the meduimweight condition atM = 1.59, and a maximum gain of 10 (0 143 radian per g) was also attained for the light-weight condition at M = 0 86.

Although the original ILAF system shaping network was developed from an early analyt-ical model of the XB-70 airplane, this shaping network revealed problems that can be en-countered during flight test that would be difficult to predict, even with the best design in-formation.

Performance With Shaker-Vane Excitation

To evaluate the performance of the ILAF system, a baseline was established for com-parison purposes The airplane, which was normally operated with the FACS on, is re-ferred to as the basic vehicle, or just FACS, and the vehicle with the ILAF system opera-ting is referred to as FACS + ILAF

M = 0 86, lightweight condition — Figure 12 presents the airplane vertical accelerationresponse measured at the pilot's station with and without the ILAF system engaged for alightweight flight condition at M = 0.86 The data presented are for a shaker-vane ampli-tude of ±4°, but are normalized to 1 unit of shaker-vane input. The second and third modefrequencies were very close at this flight condition and appear as one peak near 6 hertz inthe figure. It should be noted that the peak response of the vehicle to shaker-vane inputshifted to slightly higher frequencies and was greater with the ILAF system engaged thanwith FACS only The greater response of the second-third mode with the ILAF system en-gaged was not unexpected because of the change in phase angle in the ILAF system shapingnetwork However, the ILAF system was expected to improve performance for the firstmode The reason it did not is not known, but it is believed to be related to the non-linear characteristics of the system at this flight condition

The effect of the ILAF system upon eleven surface activities was also investigated Ele-von motion due to ILAF system operation for the lightweight condition at M = 0 86 isshown in figure 13. The figure shows the inboard and the outboard eleven positions perdegree of shaker-vane input for the first and second-third structural modes. The differencein amplitude between the inboard and outboard panels is attributed to the mechanization ofthe control system, in which the three active outboard panels follow the motion of the in-board eleven. The eleven deflections measured during these tests were relatively small, how-ever, the elevens do respond to the demands of the ILAF system at the mode frequencies,indicating that the primary reason for the lack of effectiveness in damping the modes isphase lag and not the system's nonlinear characteristics It was anticipated that the dampingof the third structural mode would be compromised by these phase lags in the revised shap-ing network. Further, a node line of the third mode crossed the elevons, rendering themrelatively less efficient for generating generalized forces for third mode control (ref. 4)

Additional shaker-vane frequency sweeps were conducted for the first structural mode forseveral combinations of shaker-vane amplitudes and ILAF system gains, and these results areshown in figure 14 A frequency sweep with FACS only was performed again for a shaker-vane amplitude of ±4° to establish baseline data A similar frequency sweep with the ILAFsystem engaged was made for an ILAF system gam of 6 (0.086 radian per g) The peakresponse measured at the pilot's station with the ILAF system engaged was somewhat higherthan the peak measured with FACS only Frequency sweeps were repeated at an ILAF sys-tem gam of 4 for shaker-vane amplitudes of ±8° and ±12° The data in figure 14 show thatthese larger shaker-vane inputs produced peak accelerations less than those measured withFACS only, indicating the nonlinear characteristics of the eleven.

As was expected, the elevon deflections commanded by the ILAF system increased withlarger shaker-vane inputs An elevon deflection of ±1.1° was measured during the peak re-sponse for a vane input of ±12° and an ILAF system gam setting of 4 (0.061 radian per g).The phase lag for vertical acceleration at the pilot's station with respect to the inboard ele-von position was found to be similar to that encountered with the smaller shaker-vane inputs,indicating that the poor response with the ILAF system for the first mode cannot all beattributed to the nonlinear characteristics of the system.

M = 1 59, mediumweight condition — The acceleration response at the pilot's station fora mediumweight condition at M = 1.59 is shown in figure 15 for an ILAF system gamsetting of 4 (0.072 radian per g) The first and second-third modes are well defined withand without the ILAF system engaged. The first mode peak response was reduced by theoperation of the ILAF system, however, the system caused vehicle response at the second-third mode frequency to increase, which was expected A comparison of figures 12 and 15

10

shows that the frequency at which the structural modes occurred was higher for the lightervehicle weight The data also indicate that after the airplane response had approached a par-ticular mode, only a small change in frequency was necessary to cause the mode to peak.The abruptness of the peak indicates that the total damping of these modes was low at bothsubsonic and supersonic flight conditions Even though the shaker-vane frequency sweepswere begun at frequencies higher than the basic airplane short period dynamics, successfulstructural mode control was demonstrated without adversely affecting the rigid body dynamics.

The elevon motion measured during the operation of the ILAF system at the medium-weight condition at M = 1.59 is presented in figure 16. The data show that the inboardelevon amplitudes were greater at the second-third mode frequency than at the first modefrequency, even though the second-third airplane mode was not damped effectively. This loss?f effectiveness at the higher frequency was due primarily to the phase lag in the shapingletwork and to the position of the node lines, as discussed previously

A comparison of figures 13 and 16 along with figures 12 and 15 indicates that althoughhe inboard elevon has approximately the same amplitude for the first mode, the ILAF sys-em was more effective in damping the first structural mode at the M = 1.59 flight condi-lon. The phase lag measured at the first structural mode between the inboard elevon and.he normal acceleration at the pilot's station for the lightweight condition at M = 0.86;fig. 13) was approximately -60° for an ILAF system gain of 6 (0.086 radian per g), ascompared with approximately -45° for an ILAF system gain of 4 (0.072 radian per g) forthe medmmweight condition at M = 1 59 (fig 16). Because of the FACS gain changewith altitude, the overall ILAF system gain was approximately 0.014 radian per g higher atthe lightweight flight condition at M = 0 86

The phase relationship between the inboard elevens and the normal acceleration at thepilot's station was also compared for the second-third mode frequency. Phase lag was foundto be -205° for the medmmweight condition at M = 1.59 <and -155° for the lightweightcondition at M = '0.86 Although the inboard elevens responded to the ILAF system com-mands, phase lag was such that the ILAF system signal reinforced the shaker-vane input andcaused higher acceleration response throughout the vehicle

M - 2 38, mediumweight condition.- The acceleration response at the pilot's station fora mediumweight flight condition at M = 2.38 is shown in figure 17 for the first mode.The effect on the aircraft's response of varying the ILAF system gain is shown in the figure.The peak vehicle response with FACS only was established by extrapolating the test resultsat the previous flight conditions The FACS-only peak response occurred at 2 34 hertzPeak response was also established with the ILAF system engaged, for gain settings of 4(0.099 radian per g), 6 (0 149 radian per g), 8 (0.199 radian per g), and 10 (0.248 radianper g). Although the resonance frequency shifted with ILAF system gain, so that the 2.34-hertz data did not correspond to the peak response, the data show that there was a reduc-tion in response with increased system gain.

The performance of the ILAF system was expected to be better at the supersonic thanat the subsonic flight conditions; the aerodynamic forces generated by control surface deflec-tion in supersonic flight are concentrated at the control surfaces, so the conditions for whichthe ILAF system was designed are more nearly satisfied (ref. 5).

M = 0.87, heavyweight condition— It was believed that the elevon deflection was sosmall (±0 6°) that the nonlinear characteristics added sufficient phase lag to the system tocause the elevens to produce vehicle accelerations instead of structural damping To

11

evaluate the ILAF system under conditions of larger elevon deflection, either an increase insystem gam or a larger shaker-vane input was required. The decision made was to keep thegam constant and increase the shaker-vane amplitude for the next series of tests. The firststructural mode was the only mode evaluated during these tests, and the results of severalfrequency sweeps with and without the ILAF system are shown in figure 18 for a heavy-weight condition at M = 0 87. A reference sweep with FACS only was first obtained for ashaker-vane input of ±4°, followed by a frequency sweep with the ILAF system engaged forthe same shaker-vane input The other curve shown is for a shaker-vane amplitude of ±6°,also with the ILAF system engaged The figure shows that a small reduction in peak ampli-tude occurred when the larger shaker-vane amplitude was used, however, the data with theILAF system engaged indicated peaks somewhat higher than the FACS-only data. The phaseangles between the inboard elevon and the vertical acceleration at the pilot's station for thetwo frequency sweeps with the ILAF system engaged are not appreciably different

Performance With High Surface Rates and ILAF System Gains

Elevon surface rate limiting was investigated, but it was not considered to present aproblem for operation at the first mode because higher surface rates had been measured foroperation at the second-third mode The peak elevon displacements corresponding to peakvehicle response for the first and second-third modes for all test conditions are summarized infigure 19 as a function of elevon surface rates The figure indicates that all measurementsobtained during the first mode tests fall along a straight line with a slope of 0.048 degreeper degree per second As noted in the figure, the solid symbols represent the test condi-tions where the ILAF system reduced the vehicle response The data obtained from thesupersonic test conditions indicated that successful mode damping was obtained with theILAF system with elevon surface displacements greater than ±0 52°, whereas the data at sub-sonic test conditions for the first mode indicated that a minimum elevon deflection of±0.66° was necessary for the ILAF system to improve the structural response.

The maximum inboard elevon deflection measured at peak vehicle response is shown infigure 20 as a function of ILAF system gain The data were obtained for the first modeonly, and with the exception of two data points the data were for high subsonic flight Forthe subsonic flight condition, the data appear to follow a pattern according to shaker-vaneamplitude. The subsonic data indicate that positive damping was not necessarily achieved byincreasing the ILAF system gain alone, but rather that it also depended upon vane excitationamplitude The data show that for vane amplitudes of ±4° and for several ILAF system gainvalues between 0.04 and 0 095 radian per g, positive damping did not result However, at avane amplitude of ±8°, positive damping could be achieved with an ILAF system gam settingof 0 06 radian per g.

The percentage of change in vehicle acceleration at the pilot's station is shown in figure21 as a function of ILAF system gain for the first structural mode The data presented arefor both the subsonic and supersonic test conditions. The subsonic data for a shaker-vaneinput of ±4° indicate that increasing the ILAF system gam only was not enough to provideadequate damping for the first structural mode.

Performance With Turbulence Excitation

The results of a turbulence encounter with the shaker system off at M = 1 20 andhp = 9754 meters (32,000 feet) are shown in figure 22 The ILAF system was engaged for

12

approximately 15 seconds. The data are presented in power spectral density format showingthe vehicle's response at the pilot's station with and without the ILAF system operating.Turbulence intensity was measured by means of a gust boom installed on the airplane (ref. 8)and was found to be different for the time periods when the ILAF system was and was notoperating, therefore, the data were normalized to a root mean square gust input of 0.30meter per second (1 foot per second). These data show that the ILAF system was effec-tive in reducing the vehicle's response. The pilots also reported a noticeable reduction inthe airplane's response with the ILAF system operating. The peak responses associated withthe aircraft structural modes do not show in these data because of the filter bandwidth thathad to be used to give good statistical accuracy with the short sample time.

In figure 23 the average number of zero crossings is shown as a function of the incre-mental vertical acceleration at the pilot's station for the turbulence encounter. The datashow that accelerations larger than 0.03g were reduced with the ILAF system engaged.

Improving the Performance of the ILAF System

Because of the degraded performance of the ILAF system after the revision of the shap-ing network, a study was made of methods for improving the system's capability. Since theshaker-vane exciter proved to be an effective means of forcing the aircraft modal response,the shaker-vane system was evaluated as a mode damper in conjunction with the ILAF sys-tem (appendix B). The calculated results show that with a simple modification the shakervane would be effective in damping the higher modes and would aid the ILAF system indamping the first mode The flight program on the XB-70 was completed without installingthis system for evaluation j I

CONCLUDING REMARKS ' '

A flight investigation of a structural mode control system termed identical location ofaccelerometer and force (ILAF) was conducted on the XB-70 airplane The ILAF systemencountered localized structural vibration problems requiring a revision of the compensatingshaping network. However, successful structural mode control was obtained without adverselyaffecting the rigid body dynamics. Although the ILAF system was developed with informa-tion from an early analytical model of the XB-70 airplane, flight tests of the modified shap-ing network and associated filter revealed problems that can be encountered during flight thatwould be difficult to predict, even with the best design information.

In general, the ILAF system was more effective at supersonic than subsonic flight condi-tions because the aerodynamic forces generated by control surface deflections in supersonicflight are concentrated at the control surfaces; thus the conditions for which the ILAF sys-tem was designed were more nearly satisfied. The ILAF system reduced the response of thefirst symmetric mode when eleven deflections were greater than ±0.66° in subsonic flight andgreater than ±0.52° in supersonic flight.

The results of a turbulence encounter at a Mach number of 1.20 and an altitude of9754 meters (32,000 feet) indicated, that the ILAF system reduced vehicle response at thisflight condition.

13

The results of an analytical study showed that the addition of a small canard to themodal suppression system would greatly improve the automatic control of the high frequencysymmetric modes

Flight Research Center,

National Aeronautics and Space Administration,Edwards, Calif., August 1,1973.

14

APPENDIX A

THEORETICAL ANALYSIS OF THE XB-70 ILAF SYSTEM

The design of the ILAF system was based on early estimates of the XB-70 airplane'smass and its structural and aerodynamic characteristics (refs 3 to 5). Therefore discrepanciesappeared when vehicle response characteristics obtained in flight (ref 5) were compared withanalytical responses based on these early estimates. Because of these discrepancies, a studyusing the XB-70 airplane was initiated to determine just how well the response characteristicsof a flexible airplane could be predicted The results of the new analysis, with a detailedaccounting of the updated mass, structural, and aerodynamic data, are given in reference 6.The results show the response of the vehicle with and without FACS operating. The datapresented herein attempt to reconcile the analytical and the flight-measured ILAF system per-formance.

Flight Conditions

Three specific flight conditions were selected for which analyses using updated data weremade. The conditions, which were representative of the flights during which the ILAF systemwas operated, are the heavyweight, M = 087, lightweight, M = 086, and mediumweight,M = 1 59 conditions shown in table 2

ILAF System Nonlinear Characteristics

During the flight tests of the ILAF system on the XB-70, it became apparent that thesystem's nonlinear characteristics were largely responsible for the lack of agreement betweenthe analytical and flight-measured system performance. Figures 24 and 25, both for the firstmode, illustrate this problem. They show the same basic trend, that is, the ILAF system didnot improve performance at the subsonic flight condition for the low shaker-vane amplitudesat which most of the flight-test data were obtained (6SV = ±4°). However, some improvementis shown at M = 1.59 (fig. 26) The data show that had larger shaker-vane inputs beenused to obtain responses with the ILAF system operating, the predicted performance improve-ments at subsonic Mach numbers for the first mode would have been obtained

Analytical Description of Elevon System

The analytical model for the FACS was described in reference 6 using two sets of trans-fer functions. The first set described the motion of the inboard elevon, the second set themotion of the remaining three elevon segments.

The analytical model for the ILAF system is described herein in a similar way using servotables. The development is general, but the specific numerical data used in the examples arefor the mediumweight, M = 1.59 case Table 3 shows the numerical data for this case.Tables 4 and 5 show similar data for the heavyweight, M = 0 87 and lightweight, M = 0 86

15

cases, respectively.

The inboard eleven deflection is given by the expression:

/ 6j \ / A/ V" A& \5t = (ILAF system gam adjust) K,,p t — - j [ KILAF —s q - > KILAP — s2 77, J

\ "ZILAF/ \ g ^^ g I

With ^ILAF = 1-0 and Kf, = 0.62, and expanding as a function of independent variables,the expression becomes

61 = - — [ (57-1) A/? + (57-2) A0,r?1 + (57-2) A02r?2

+ (57-2) A03773 + (57-2) A04774 + (57-

where

Kf = 0.11 radian per g

A/, A0, differences in data at the ILAF system wing accelerometers and theaccelerometer at the center of gravity (table 6)

57-1,57-2 servo tables in table 3

The expression for outboard eleven deflection with the addition of the lag between theinboard elevon and the outboard elevens can be written.

( Si \/62-4\/ A/ \^v M, \A« - )l T~ ) 1^ILAF ~ s q ~ / ; A:ILAF — *2 ^ J

2^4 = - — [ (57-3) A&? + (57-4) A01r?1 + (57-o

+ (57-4) A03773 + (57-4) A04T?4 + (57-

where

57-3, 57-4 servo tables in table 3

The .KILAF term used in the numerical example above should not be confused with thevalue of the pilot control panel ILAF system gam select (ref. 4). The value A^ILAF = 1-0as used herein is a computer control system gain input that, when combined with Kf, , yieldsan overall ILAF system gain, KIt equal to 0.11 radian per g.

Elevon Response Analyses8 i 52_4

It remains to explain and qualify the elevon frequency response data - - and — — .A"*ILAF 5l

16

Because of the mechanics of obtaining frequency response with the ILAF system engagedusing a digital program, there was no automatic way to coordinate the magnitude of the elevenresponse with the amplitude characteristics of the nonlinear system dynamics. Because of this,an iterative scheme was used. If the eleven amplitudes were not in agreement with those as-sumed for the system dynamics, a new estimate was used and the eleven response was recalcu-lated. This procedure was used to obtain the results herein However, completely convergedsolutions were not obtained in all cases because of the computer time required.

Figure 27(a) presents the frequency response from the ILAF system's blended accelerome-ter signal through the inboard eleven deflection. The solid curves are constructed from datapresented in reference 4. The solid-line curves are not curves of constant 5j, as they are inreference 4, but rather are curves passed through various magnitudes of 5t at the indicatedfrequencies The magnitudes of 5] used in constructing these curves are taken from flight-test results and were used to start the previously mentioned iterative process The dashed curvesare the calculated results that best represent what was measured in flight Only amplitude datacould be obtained from flight records with accuracy The curves showing the phase characteris-tics (fig 27(b)) were based on the estimates made in the iterative procedure that produced thebest agreement between the measured and computed amplitude characteristics

Table 7 compares flight-test data and analytical data at several points in the control sys-tem as well as at the pilot's station. These flight-test data reflect the dashed-bne data offigure 27. Since the dashed-lme data produce the best analytical agreement with flight-testdata, and the solid-line data are based on ground vibration test measurements, it can beinferred that flight aerodynamic loads or other unidentified influences had changed the sys-tem's frequency response characteristics

Vertical Acceleration Responses

Vertical acceleration responses were calculated for various locations on the airplane. Flight-test measurements were made at these locations, and data measured with the FACS only oper-ating are compared with the predicted response in reference 6, where the better agreement ob-tained from the refined analyses is shown Hence, only results using the refined analysis areused herein, and the airplane with FACS on is used as the basic vehicle since the aircraftnormally operated with the FACS engaged. The ILAF system evaluations were made fromthis base configuration, and the calculated results are shown for the pilot's station in figure 28for the heavyweight condition, M = 0 87, in figure 29 for the lightweight condition, M = 0 86,and in figure 30 for the mediumweight condition, M = 1 59

Control Surface Responses

The calculated frequency responses of eleven action due to ILAF system operation perunit of shaker-vane input are shown in figures 31 to 33 for the three flight cases studied.As explained, the FACS servo drives the inboard eleven and the motion of the inboard elevonactivates the remaining elevon panels (panels 2 to 4 for the flight cases analyzed with thewingtips deflected). Because of this arrangement, it was desirable to determine separately theinboard and outboard elevon motions in the analyses with the ILAF system engaged. Thiswas easy to do analytically, but it could not be done with data from the actual airplane.

The elevon deflection data presented can be used to obtain elevon rate information by

using the relationship 8 = cj5 for sinusoidal oscillations.

17

Structural Mode Damping

For a lightly damped dynamic system, the calculated damping (structural plus aerody-namic) can be obtained from the dynamic system characteristic determinant in the form ofphase angle as a function of forcing frequency (ref. 6). Using this technique requiresknowledge of the mode natural frequency, co,, and the phase angle slope with frequency,d<i>

, at that natural frequency. The damping calculated for the vehicle with and withoutacothe ILAF system operating is presented in table 8.

18

APPENDIX B

DESIGN OF A STRUCTURAL MODE CONTROL SYSTEM FORTHE XB-70 AIRPLANE USING A SMALL AUXILIARY CONTROL SURFACE

The results of this study and studies reported in reference 4 show that the XB-70 ele-von ILAF system has significant potential as a means of damping the first and second struc-tural modes. However, it has less potential as a means of damping the third structural' mode.This is because a third mode node line (the locus of zero displacement) runs between theelevens, rendering the elevens relatively less efficient for generating generalized forces for thirdmode control. Further, flight-test results show that the eleven ILAF system is ineffective indamping even the first structural mode when elevon amplitudes are less than ±0.66° Severalother factors associated with the use of the elevon surfaces are discussed in reference 4.

When ways to improve the ILAF system's performance were examined, it became ap-parent that a more effective structural mode control system might be implemented with arelatively small modification to the XB-70 system Specifically, it appeared that the shaker-vane system, which was utilized to excite the XB-70 airplane during the elevon ILAF systemevaluation, could be converted to perform the structural mode control function Previousstudies have shown that a small aerodynamic control surface located at the nose of a flexiblevehicle is effective in damping the lower frequency modes An inspection of the lower fre-quency mode shapes showed that the existing shaker-vane location was well placed to adddamping to the third mode (which could not be controlled adequately with only the elevonILAF system) as well as to augment the elevon ILAF system in damping the first and secondmodes

Shaker-Vane Characteristics

The shaker-vane system, which is described in reference 6, is capable of continuousoperation in the frequency range from 1.4 to 8.0 hertz The vane amplitudes are variablefrom 0° to 12° on either side of a preselected vane trim position (no load condition).

Laboratory tests were conducted to define the shaker-vane steady-state inputs to excitethe symmetric structural modes of the airplane The specific actuation transfer function wasnot available from laboratory tests, however, and flight-test data were used to make estimatesof the actuation system's dynamics. The transfer function for the linear range of the actu-ator was estimated to be 60/(s + 60), and this estimate was used in all the design analyses.

Structural Mode Control System Design

The conversion of the shaker-vane system to a structural mode control system requiredinstalling an accelerometer in the vicinity of the shaker vane (as required by the ILAF systemtechnique), subtracting the existing FACS accelerometer signal from the signal of the acceler-ometer, shaping the net signal, and then feeding it back through the-actuation system. The

19

primary objective of the design was the augmentation of the elevon ILAF control system,however, the shaker-vane ILAF system and the elevon ILAF systems were designed to oper-ate independently of one another within their own limitations

To minimize the effort of installing the shaker-vane ILAF system, existing componentsin the FACS and the original shaker-vane system were used in the design study Figure 34shows a block diagram of the shaker-vane ILAF system that was designed The accelerometerto be installed near the shaker vane was the required ILAF system primary sensor, whilemeasurements from a second sensor (a FACS accelerometer was used to maximize use ofexisting equipment) were used to cancel all the rigid body plunge and some of the rigidbody pitching acceleration signals The reason for canceling the rigid body signals in theshaker-vane ILAF system signals was to avoid amplitude saturation, which is a possibility withsuch a small control surface. It is possible that rigid body signals alone could command allthe available surface authority of the small shaker vane, leaving nothing for the structuralmode signals to command For example, at a gam of 0.25 radian per g, a rigid body signalamplitude of 0.5g would command all of the 12° available in the shaker-vane system Toavoid such amplitude saturation the ILAF system and FACS accelerometer signals were com-bined and shaped before commanding the shaker-vane actuator The shaping networks werelimited to use of the spare components already existing within the FACS equipment to mini-mize the need for wiring, cooling, packaging, and so forth Manual engagement and selectionof input frequency and amplitude, automatic disengagement, and fail-safety features of theoriginal shaker-vane system were retained and incorporated in the shaker-vane ILAF system.The primary ILAF system accelerometer and the gain selector were the only new components

Two shaping networks, shown in figure 35 and designated shaping 1 and 2, were evalua-ted for the shaker-vane ILAF system Shaping 1 was a simple first order lag or 5/(s + 5)and was selected to provide a lag of approximately 90° at the frequencies of the structuralmodes to be controlled However, a first order lag was unsatisfactory for the proposed shaker-vane system because the large attenuation at the mode frequencies required a very high gainThe FACS equipment initially restricted the maximum gain (occurring at / = 0 hertz) to bewithin 0 3 radian per g This restriction, together with the attenuation characteristics of afirst order lag, meant that gams available at the third mode frequencies would be limited towithin 0.05 radian per g, which is too low to be effective at some flight conditions Shap-ing 2 was selected as 1600/(s + 40)2 and alleviated the 0.05-radian-per-g gain limitation byincreasing the gain at mode frequencies to 0 2 radian per g while providing approximatelythe same phase lags at the third mode frequencies However, shaping 2 phase lags could betoo small at the first mode frequencies, too high at fourth-fifth mode frequencies, or both.Thus the shaker-vane ILAF system could adversely affect the pilot station acceleration at thesefrequencies.

Performance and Stability Characteristics

A typical estimate of the performance of the shaker-vane ILAF system is shown in fig-ure 36 for supersonic and subsonic flight conditions as a function of shaker-vane forcing fre-quency.

Most of the performance estimates were restricted to a maximum gain of 0 3 radian per g,although a few higher gains were investigated to determine whether any modification of the ex-isting equipment was necessary. Estimated performance with shaping 1 (fig 36) was unsatisfactory

20

because of shaping attenuation at the third mode frequencies The estimated performancewith shaping 2 was promising enough to be considered for installation in the XB-70 airplanefor flight-test evaluation However, the two lag time constants were designed to allow modi-fications to be made prior to each flight Investigation of the air vehicle and shaping phasevariations with frequency indicate that the estimated performance with shaping 1 and 2would be significantly better if it were not for some adverse phase effects at frequenciesslightly away from the peak response frequencies which apparently shift these peaks Becauseof these shifts m response peaks (usually in the direction of higher frequency), the vehicleresponse at some specific frequencies with the ILAF system engaged are worse than withoutthe ILAF system

The stability analyses corresponding to the performance data in figure 36 are shown infigure 37 The stability analysis technique, whjch is not a conventional one, is described indetail in reference 4 In effect, a continuously decreasing phase angle with increasing frequen-cy is considered to denote a stable system, and a mode that shows a reverse trend is un-stable As these figures demonstrate, the systems investigated were stable

As indicated in figure 38, the shaker vane as being used here has a dual function—toexcite and also to provide damping to the structural modes Figure 38 also shows requiredshaker-vane ILAF system deflections as functions of ILAF system gain and pilot-selectedshaker-vane excitation input Because shaker-vane excitation input and shaker-vane ILAF sys-tem feedback signals are subtracted before actually commanding a net shaker-vane motion,required deflections and rates decrease with increasing ILAF system gam. Therefore, nosaturation problems are expected to occur if phase estimates are correct When possible,previous flight-test data were examined for phase characteristics and compared with estimatesof the analytical model used to represent the air vehicle in an at tempt to anticipate phasingproblems, and none were uncovered

Most of the estimated performance data were obtained by using the shaker vane as theinput because of on-demand availability and repeatability advantages over gust inputs for datafor comparison with analytical data However, some analytical data with gust as the inputwere obtained to estimate potential saturation problems Figures 39 and 40 show typicalperformance and required shaker-vane rates and deflections, respectively, for random gust in-puts These estimates indicate that good performance can be expected up to a root meansquare gust magnitude of 1 22 meters per second (4 feet per second) Possible adverse ef-fects can occur when root mean square gust magnitudes exceed 1 52 meters per second(5 feet per second) because of shaker-vane rate saturation This area was considered worthyof further investigation prior to flight-test evaluation

Although this study showed the shaker-vane ILAF system to be effective in reducing themodal response, the XB-70 airplane was taken off flight status before the system could beinstalled and tested

21

REFERENCES

1 Garrick, 1 E Perspectives in Aeroelasticity Israel J Technology, vol 10, no 1—2,1972, pp 1-22.

2 Burns, P M , and Bender, M A Aircraft Load Alleviation and Mode Stabilization(LAMS) AFFDL-TR-68-158, Air Force Flight Dynamics Lab , Wright-Patterson AFB,Apr 1969

3 Wykes, John H , and Mori, Alva S An Analysis of Flexible Aircraft Structural ModeControl AFFDL-TR-65-190, Part I, Air Force Flight Dynamics Lab , Wright-PattersonAFB, June 1966

4 Wykes, John H , Nardi, Louis U , and Mori, Alva S XB-70 Structural Mode ControlSystem Design and Performance Analyses NASA CR-1557, 1970

5 Wykes, John H , and Kordes, Eldon E Analytical Design and Flight Tests of a ModalSuppression System on the XB-70 Airplane Part 1 Design Analysis Part 2. FlightTests Aeroelastic Effects From a Flight Mechanics Standpoint, AGARD Conf. ProcNo 46, Mar 1970, pp 23-1 - 23-18

6 McKay, James M , Kordes, Eldon E., and Wykes, John H Flight Investigation of XB-70Structural Response to Oscillatory Aerodynamic Shaker Excitation and Correlation WithAnalytical Results NASA TN D-7227, 1973

7 Mechtly, E A The International System of Units-Physical Constants and ConversionFactors NASA SP-7012, 1969

8 Wilson, Ronald J , Love, Betty J , and Larson, Richard R Evaluation of Effects ofHigh-Altitude Turbulence Encounters on the XB-70 Airplane NASA TN D-6457, 1971

9 Edwards, EL A Data Processing Facility for the XB-70 Flight Test Program FlightTest Instrumentation, AGARD Conf Proc No 32, 1967, pp 243-258

10 Ince, D B Application Experience With the B-70 Flight Test Data System AerospaceInstrumentation Vol 4—Proceedings of the Fourth International Aerospace Symposium,College of Aeronautics, Cranfield, Eng , March 21-24, 1966, M A Perry, ed , PergamonPress, Ltd , 1967, pp 195-208

22

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25

TABLE 3 - COMPUTER PRINTOUT OF SERVO TABLE DATA FOR THE ILAF SYSTEM

[ Mediumweight, M = 1 59, h = 11 , 918m (39,100ft), 5, -65°, rad/g

(a) Inboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

ST-]

0.199S9999E OCi

0.5999999tE OC0.100COCOCE Cl0 . 2 C O O O C O O E 01o .ecocococE 010.10000000E 020.2000000CE 020.3000000CE 02O . A O O C O C O C E 0 20 . 5 0 0 C O C O O F 020 . 6 0 0 C O O O C E 020 . 7 0 0 C O O O C E C 20.8000000CE 020.9000000CE 020.100COOOCE 02

0.0C.O0.0c.73^}<;<;<;(;5E cc0 .35999994F ClC.86999<;<; f iF Clc.iosgs«;<;9E C2C . 4 K C C C C C C F C l

-0 .3 l99< ;cceE Cl- C . 7 5 0 C C C C C P C l- C . 6 5 9 9 9 9 S 4 E ci-c . ^35999< ;7E ci

C.135C < ;<;<;7E ClC . 2 5 1 9 9 9 9 5 E Cl0.20<59<;cc<,E ci

C.1S999999E 00C.5<;999996E 00C . 1 C O O O O O O E 01C.1E599997E 01C.A7999992F 01

C.2399P9S6E 01- C . 6 C C O O O O O E 01-C.11700000E 02-C.IC799999E 0?- C . 6 C C O O O O O F 01

C.17SQ9992T 01

C.46399996E 01C.47399998E 01C.26099997E 01C.S9q09996E 00

ST-2

0.19999999E CC0.59999996E OC0.10000000E 010.200COOOCE 010.6000000CE 010.100COCOCE 020.2000000CE C20.300COOOCE 020.4000000CE 020.500COOOCE 020.600COOOCE 020.700COOOCE 020.80000COCE 020.9000000CE 020.10000000E 03

0.3999S999E-C10.359999S5E CCC. IOOOCCCOE Cl0.37199<;<;3E ClC . 2 8 7 9 9 9 P 8 E C20 . 2 A O O O O C C F C2

- C . 1 2 0 C C C C O E C3-C .3510CCCCE 03- O . A 3 2 0 C C C C E C3- C . 3 0 C C O C C C E C3

C . 1 0 8 C C C C C E C3C.327000COF C3C . 3 A O O C C C C E C3C . 2 3 6 C C C C C E C3C . 6 C C C O C C C F C2

C.CC.CC.C

-C.1A799995E 01-C.21599991E 02-C.87COOOOOE 02-C.21200000E 03-C.13500000F 03

C.12800000E 03C.37500COOE 03C.39600000E 03C.23^00000E 03

-C.11COOOOOE 03-C.22800COOE 03-C.21COOOOOF. 03

26

TABLE 3 - Concluded

(b) Outboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

ST-3

0.19999999E OC0 .59999996E oc0 . 1 0 0 C O C C C E Cl0.20000000E 010.6000000GE 01O . I O O O O O O C E C 20 . 2 0 0 0 0 C O C E 020 . 3 0 0 C O C O C E C 20 . 4 0 G C O C O C E C20 .5000000CF C20 .600COOOOF 020 . 7 0 0 C O C O C F 020 . 8 0 0 C O C O C E 020 . 9 0 0 C O C O C E C 20.1000000CF C3

0.00.00.00.10199995E ClC .47999992E Clc . a o c c c c c c F ciC . 3 0 0 0 0 C C C E Cl

-0 .32999992E Cl- C . 6 0 C C O C C O F Cl- I C . 4 C C C C C C C E Cl '- ' C . 3 5 9 9 9 9 C 5 E c c

C . 1 5 4 C O C C C E C lI0 .45999994F ClC . 1 0 C O O C C C F ClC. 199999995 CC

C.19999999E 00C.59999996E 00C.1CCCOOOOE 01C.17399998E 01C.31199999E 01

-C.1199999flE 01- c .ecooooooE 01-C.65999994E 01-C.1599999^F 01

C.25COOCOOE 01C.3COOOOOOE 01C.16799994E 01

-C.31999999E 00- C . I C O O O O O O E 01-C.79999995E 00

ST-4

0. 19999999E OC0.59999996E OC0 . 1 C O C O C O C E C l0 . 2 0 0 C O C C C E Cl0 .60000COCE ClO . l O O C O O O C f c C20.20000COCE C20 .30000COCE 020 . 4 C O C O C O C E 020 . 5 0 0 C O C O C E C20.6000000CE C20.70000000E 020.80000000E C20.90000000F C20.100COOCCE C3

C . 3 9 9 9 c c c c E - c iC . 3 5 9 9 9 C 9 5 F C CC.10000CCCF ClC . 3 4 7 9 9 9 9 5 E ClC .18599991E C2

-C .1200CCCCE C2- C . 1 6 C C C C C C F C3- C . 1 9 P C C C C C F C 3-C .6400CCCOE C2

C . 1 2 5 0 0 C C C E C3C . 1 8 C O O C C C E C 7

C . 1 1 8 C C C C C E Cl-C .25599991E C2- C . 9 0 C C C C C C E C2-0 .8000CCCCE C2

C.CC.CC.C

-C.2C400000E 01-C.28799989E 02- C . 8 C O O O O O O E 02-C.6COOOOOOE 02-C.99000000E 02

C.24000000E 03C . 2 C C O O O O O E 03C.21599991F 02

-C.1C800000E 03-C.12800000E 03-C.9COOOOOOE 02-C .2CCOOOOOE 02

27

TABLE 4. - COMPUTER PRINTOUT OF SERVO TABLE DATA FOR THE ILAF SYSTEM

[ Heavyweight, M = 0 87, hp = 7620 m (25,000 ft), 5, - 25°, K, = 0 07 rad/g ]

(a) Inboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

ST-\

O. lQQqaqgQC DO

0. 599Q9996F 000.1000000TF 010 .20000HOOF °»1o.60ooonoot 01O . I O O O O O O O F o?o.2onnoooo r o?0. ^ P O O O O O O F 0?O . I O O O O O O O F n?O . I O O O O O O O F 070.60000000F 02

0.70000000^ 0?O . I O O O O O O O F 02o.cioooooooc o?o; loooonocp m

o.o0.0o.oC.TIoqoqq^r 00

0.359 c 'q994F 01( - ^Q^CQOC^CIp p^

O. lO^^OqT^F °^0.£ 5000000^ 01

- 0.^ l^oqqoRf- Q1

- 0 .7^000000^ 01- o. oQqcq/vp 01- O.i • '^Qqqqvr oi

0.] ??9^ f ;97F Cln.PSl0^?^ 010.?09Qoqo4F 01

C. 1 qqoooaq p noC.cqgqqQq,cc QO

r. 10000000^ ciC. l«sqoQQ7F 01C.47qqnqQ? F 01n.p^^qqqo^F 01

-C. ^ 0000 000 c HI-C. 1 1 70000 r>r o">-C. 1 07Q°aoc>c O*>- C . 6 C O O O O O O F 01

0. 1 790990,7 P oi

O .A^aaoqo^F 01O.A?.39qqqaF 01C.26CT3Q07F 01p.Sqqoooq^F OO

ST-2

0.19999999F 000.ei999«99^F 000.10000000E 010.2000000TF 01

O . f O O O O O O C F 01

0.10000000E 020.20000000F 0?0.3000000CE 0?O.AOnonoorr r>20.500COOOOF 02O . f t O O O O O O O F 0?

0.70000000F 0?0.80000000F 0?0.90000000F 0?O. IOOOOOOOF. c?

0.39999999F-010.3 t5c>9Qqq5F QC

0.1000CCOOE 010.^71Q9991F 010.2f l7Q9qRgp o?

C,240nooOQF o^-C.12000000F 03-C .3^1 OOCOOb 0?-0 . ^3?OOOOOF C3-C.30000000F o^

0.10«OOOOOE O"1

o.^27ooccn': ciC . ^ ^ O O O C C O F 03n.?360000QF 030.600000COF 0?

0.00.0C. 0

-C. 14799Q9 t ;F 01

-o, 21*5^909^ o?-C. 87OnooOOF 02-C.21200000F 0^5-0. 1 3 500000 c C^

C. 12ROOOOOF 03

0.37500000E 03r .30^>OOOOOF 0^

0.2^40000^F 0^-C.ll OOOOOOF 0^>

-C.??80nooOF 03-0.21 OOOOOOF. 0?

28

TABLE 4 - Concluded

(b) Outboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

57-3

O.iqoqqqqqp QC

o.^q^gqqqtF ooO . I O O O O O O O F 010.20000000F 010. *>0000000P 0 1O . I O O O O O O O F 020.20000000F 020.30000000F n?0.4.0000000F 020.5000000PF 020.60000000F 020.7000000CF 0?O . B O O O O O O O F 0?0.9000000CF C2O . I O O O O O O O F 0^

0.00.00.0o. io i99995F 010.47 09a992F CIC . R O O O O O O O F 01C.?0000000P 01

-0 .3299 c 9q2F 01- O . ^ O C O O O O O F 01- r . 4ooooccc c n- 0.3 5°9Qq9cF CC

0.1 5400000F 0]

Cl 1 oqqqOAF 01C . I O O O O O O O P c iC. iqoqqqqQP 00

G. lCqqqqoqF 00

0. sqqqqqoAF 00O . I O O O O O O O F 01C. 1 73qoqqR H 01C. 31 1 qqaccp 01

-C. 1 1 OQOOOR F 01

- O . B O O O O O O O F 01-0.6S ^qooq^F 01

-0. 1 5 T O O Q O A F 01

0. '5000000F 01

C.30000000F 01C. 1 f,7oqQO£ F 01

-C, "*1 qqooqo F O1"1

-C. 1 O r!OO°OC F 01-0. 799qqocn; r 00

57-4

o.i9qqqqqqc oc0. «5qonqqqf,F PC

0. 10000000F 010.20000000F 01

0.60HOOOOOF 010 . lOOnooonr C ">0.2000000f 02

O . A O O O O O O n r 02O.BOOOOOOOF n?O . B O O O O O O O F 020.70000000F 02O . R O O O O O O O F 02O . q O O O O O O O F 020.1000000CF C"1

O.^qcgqqqqc-oi

Q.3SOqoqqt ,F 00

^.1 O O O O C O O F 010 . ^ ^ 7 9 9 9 0 ^ ^ 01C.I S^encq i F pp

-c . ip ioor r rp C2

'r^oqo™0^ °lp ^~/4.pooppn^ ^^o ipsnppppt P"^

C. 1 R O O O C O C r C^0.11 q O C O O O c C3

- C .2 ^5Qcqn iF 02- O . C O O O O O O O F 02- O . R O O O O O r C F 02

C.OC.O

C.C-0 .204HOOOOP 01— 0. 2 R7qqr>p q p n"?

~C. f O O O O O O ^ F 02- 0 . 6 C O O O O P O F 02

"c i^OOOOOOF 0^C.20000000F 0^0.21 S Q O O f / l F 02

- C . 1 C R O O O O O F 0^- C . l ' R O O O O O F 03- O . C O O O O O O C F o?-C.20000nooF 0?

29

TABLE 5 - COMPUTER PRINTOUT OF SERVO TABLE DATA FOR THE ILAF SYSTEM

[ Lightweight, M = 0 86, h = 7620 m (25,000 ft), 5, - 25°, K, = 0 07 rad/g ]

(a) Inboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

ST-\

0.1<59<;<;<;9CE CC0.599<,9<;9eE CCo . i c c c r c c c E c iC . 2 C C C O C C C F 010 . 6 C O C O C O C E C i0 . 1 C O C O C O C E C ?C . 2 C O C O O C C F C 2U . J C O C O C U C E C 20 . 4 C C C O C C C F C ?o . ^ c c c c c ^ c t c?O . f t O C C O C C L E L2C . 7 0 0 0 C C G C E C2O . e C O C O C C C F 02C . < 5 C O C O C O C E C.2C . 1 C O C O C O C E C ^

C.OC.OC.Or .73<599<;9 t)F CCC . ' ^ S ^ g S ^ ^ A E 01C . ^ f e Q g q ^ c j S E Cic.i c^gQc j^qp C2C . < f 5 C C C C C C F C I

-C.3 1 9 9 9 C C 8 E 01- C . 7 5 C O O O O O F CI-C.t 5<5099S^F Gl- C . 3 3 5 9 9 9 9 7 F 01

C . 1 3 5 9 9 C 9 7 E n.C .2519999SE 01C . 2 C n 9 9 9 9 ^ E CI

C.19999999E 00C.59999996t 00C.10000000E 01O. IR599997E 01U.A799999^b 010.23999996E 01

-C.60000000E 01-U. 11 700000 f: 0?-C.1C799999E 02-C.6COOOOOOE 01

0.17999992E 01O.Af -899996E 01O.A239999«F 010.26099997E Oi0.59999996E 00

ST-2

o. 19999S99F cc0 . 5 S S 9 9 9 9 6 E CC0 . 1 0 0 C O O O C F 010 . 2 C C C O C O C F C IC . f c C C C C C O C E 01n . l C O C O C C C E C 20 . 2 C O O O C O C F C 20 . 2 C O C O C O C E C 2o . A C r c o c c c r c?o . ^ c o c c c o c r C 2C . e C C C O C O C F C 20 . 7 C C C C C O C F C2O . e C C C G C C C E C 2C . C O O C C C O C E 0 20.1000000CT C3

C.399999C9E-C1C.35999995E OCC . 1 C O O C O C O F CIC.3719SS53F CIC.287999R8E 02C . 2 4 C C C O O O E 02

-C .12000CCOF. 03- C . 3 5 1 C O O O O E 03- C . 4 3 2 C C O C C E C3- C . 3 0 C O O C C C E C3

C . l C f i O C C C C f - C3C .327COOCCE 03C . ^ ^ O O O C C C E C3r . 2 3 6 0 C C C C F C3r . 6 C C O O C C O F C2

0.00.0O.C

-C.14799995E 01-0.21599991E 02- C . 8 7 C O O O O O E 02-C.21200000E 03-0.13500000F 03

C.12800000E 03C.37500000E 030.39600000F 030.23AOOOOOF 03

-C.11000000E 03-C.22800000E 03-C.21000000E 03

30

TABLE 5. - Concluded

(b) Outboard eleven

Frequency,rad/sec

Real,rad/sec

Imaginary,rad/sec

ST-3

0.1999999CF CCc.5999999tF cc0 . 1 C O C O C C C F C lC.PCOCOCOCE 01C . 6 C O C O C O C E C l0 . 1 C O C O C C C F C 20 . 2 C O C O C O C E C 20 . 3 C C C O C O C F C 20 . 4 C C O O O C C F C ?0 . 5 0 0 0 C O C C E C ?0 . 6 C C C O C O C E 02C . 7 C C C C C C C E C ?O . P O O C O C C C F C ?0.900CUGUU C?c . i c c c o n o c E C3

C.Oc.oC.O0.10199995E ClC.47999992E 01C . f l O C O O C C O E 01C . 3 C C O O C C C E C l

-C.32999992E Cl- C . 6 C C O O C O C E C l- C . 4 C C C C C C C E Cl-0.35999995F: CC

C . 1 5 A O O C C C E 01C.15999994F. 01C . 1 C C O O C C O E ClC.19Q99999E CC

0.19999999E 000.59999996E 000.10000000E 010.17399998E 01C.31199999E 01

-0.11999998E 01-0 .8COOOOOOE 01-0.65999994E 01-0.1599999AE 01

C.25000000E 010.30000000E 01C.1679999AE 01

-0.31999999E 00-C.ICOOOOOOE 01-0.79999995F 00

57-4

O . l S S s q q q c p C C

0 . 5 9 9 9 9 9 9 6 E cc0 . 1 C O C O O C C E ClQ . 2 0 0 0 0 C O C F Cl0 .6COCCCCCF. C!0 . 1 C C C C C C C E 020 . 2 C O C O C C C E C 20 . 2 C O C O C C C E C 20 . 4 C O C O C C C E C 2O . S C O C O C O C b 020 .6000000CE 020 . 7 0 C C C O O C E 0?0 . 8 C C C O C C C E T 20 . 9 C O C O C C C E C?C.1COCOOOCF C3

C . 3 9 Q 9 9 9 9 9 E - C 1C.35999995E CCC.I C C O C C C C F . ClC.14399996E Cl

- C . 1 0 P O C C C C E C2- C . 9 2 C O O C C C E C2- C . ? ? n O O C C C F 03- C . 1 0 C O O C C O E C3

C . 1 7 « > O C C C C F ClC . 3 2 ^ 0 0 C C C F C3C . 2 C 2 0 C C C C E C3r . i c c c c c c c F o?

-0 .153000COE C3-C.1800CCCCF 03-C .1COOCCOCC 03

0.0C.O0.0

-0.55199995E 01-0.473Q9994E 02-C.68COOOOOE 02

O. iOOOOOOOE 03C.29700000E 03C.3C400000E 030.750COOOOE 02

-0.15800000E 03-0.22600000E 03-C.1C200000E 030.0C.600nOOOOE 02

31

TABLE 6 - MODE SHAPE CHARACTERISTICS AT SELECTED LOCATIONS

[ Medium weight, M = 1 59, 5, = 65°]

Location

Fuselage nose,FS495 m (194 75 in)

Pilot station,FS 1 1 12m (438 in )

Nosewheel well,FS3261 m (1284 in )

Near center of gravity,FS 37 72m (1485 in.)

Wing accelerometer,FS 56.18m (2212 in.)

Center of gravity,FS 41. 99m (1653 in)

Mode

12345

12345

12345

12345

12345

12345

0,

' 2 1 200-0 1500

3 730000680

-0 8600

1 2500-0 0650

1 10000.2100

-0 1500

-0 420000250

-0 1 900-0 0045-0 1100

-03817000370212500068

-0.1529

06000-0 0600-0 9000-00550-0 4300

-0.2992-0.0162

0.424000112

-0.1190

0;

_

— —0000077

-0 02930-0 00072

0 00720

— — —

32

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(a) Amplitude.

50 r-

, deg -50

-100

-1502 4 6 8 10 20 30

f. Hz

(b) Phase angle.

Figure 4. XB-70 ILAF system initial flight-test shaping network.

38

FS 8.24 m(324.5m.)

Figure 5. XB-70 shaker-vane location

39

03

£p

eQ)

In

be

05

cO

03

03

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Z03

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o Analytical results

n Flight data

Nominal XB-70 flight envelope

-i 80 x 10,3

60

40 hp, ft

20

5 1.0 1 5M

2.0 2.5 30

Figure 7. ILAF system flight-test envelope.

41

FACS

FACS + ILAF

± 15

Fuselageacceleration,

g

Wingacceleration,

26

,- ,/ \ /

' J

+ .5

Inboardeleven

deflection,deg

Pitchservo

deflection,deg

Figure 8. Limit cycle oscillations due to ILAF system operation fromcomposite data for three flight conditions.

42

co

+1_CD

IwCJ

•rH

ECB

0•4-J

<DeO

oocflbDC

cCO

boc

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43

Electronics

I LAP system wingaccelerometers

Pitch servo

26 Hz

Actuators andelevens

12 Hz

Rigid body andstructural

modes

I LAP system fuselageaccelerometer

Figure 10. Deduced sources of limit cycle instabilities.

44

Amplitude ratio

<p, deg

Revisedshapingnetwork

(a) Amplitude.

Desiredshapingnetwork

-150 20 30

(b) Phase angle.

Figure 11. ILAF system shaping network.

45

FACS

FACS + I LAP, gam =6 (0.086 rad/g)

Figure 12. Flight-measured vertical acceleration response at the pilot's stationwith and without the ILAF system engaged. Lightweight, M = 0.86; h = 7620 m(25 ,000f t ) ; 6 = 25°. P

46

Inboard eleven

Outboard eleven

,16 i—

.14

.12

.10

sv .08

deg/deg

.06

.04

.02

4f, Hz

Figure 13. Eleven motion due to ILAF system operation. Lightweight, M = 0.86;h = 7620 m (25,000 ft); 5t = 25°; ILAF system gain = 6 (0.086 rad/g) .

47

04

03

.6sv '

g/deg

.01

FACS

FACS + ILAF

FACS + ILAF

FACS + ILAF

SVdeg

±4

±4

±8

±12

Gam (rad/g)

6 (0 086)

4 (0 061)

3f, Hz

Figure 14. Flight-measured vertical acceleration response at the pilot'sstation with and without the ILAF system engaged. Lightweight, M = 0.86;h = 7620 m (25,000 ft); « = 25°.

48

svg/deg

07

.06

05

04

03

.02

.01

FACS

FACS + ILAF, gam =4 (0.072 rad/g)

f. Hz

Figure 15. Flight-measured vertical acceleration response at the pilot'sstation with and without the ILAF system engaged. Mediumweight,M = 1 . 5 9 ; h = 11,918 m (39,100 ft); 5 = 65°.

P t

49

svdeg/deg

Inboard eleven

Outboard eleven

Figure 16. Eleven deflection due to ILAF system operation.Mediumweight, M = 1 . 5 9 ; h = 11,918 m (39,100 ft) ; 6t = 65°;

ILAF system gain = 4 (0.072 rad/g) .

50

07

.06

05

04

sv

03

02

01

Gam (rad/g)FACS

FACS + ILAF 4 (0 099)

FACS + ILAF 6 (0 149)

FACS + ILAF 8(0.199)

FACS + ILAF 10 (0 248)

f. Hz

Figure 17. Flight-measured vertical acceleration response at the pilot'sstation with and without the ILAF system engaged. Mediumweight, M = 2.38;h = 18,898 m (62,000 ft); 6 = 6 5 ° .

P *

51

.04

03

n

6sv'g/deg

.02

01

FACS

FACS + ILAF ±4 j

FACS + ILAF ±6 i

sv'deg Gam (rad/g)±4

4 (0.053)

2f. Hz

Figure 18. Flight-measured vertical acceleration response at the pilot'sstation with and without the ILAF system engaged. Heavyweight, M = 0.87;h = 6400 m (21,000ft) ; 5 = 25°.

P *

52

0>O)o.

+1o

oJS

CVJvO OO i—'+1 +1 +1 +1D O

T3<D

w

0)•*-»CO

CD(NJ

2 S1

o(D

O• r-(-*-»

O

CO

enccCDCO

•rH•4-'

OQJ

C!0)

CO

cp

03o.0

(1)

+1

CM

+t

oo-H -H -H

cvi

-H

CD

<o~

53

±1 4

+1 2

±1.0

deg

± 6

± 4

± 2

02

6svdeg Airspeed

+4

±6Subsonic

±8

±12

±4 Supersonic

Solid symbols indicate positiveI LAP system damping

04 06 08

Kiu\p rad/g

10 12

Figure 20. Inboard eleven motion measured at peak vehicle response as afunction of ILAF system gain. First mode only.

54

6 , deg Airspeed

Subsonic

20

Change inacceleration

response,percent

-20

-40

-6002 04 .06 08 .10 12

Figure 21. Percentage of change in vehicle acceleration with the ILAF systemengaged for the first structural mode as a function of ILAF system gain.

55

CO+-Io

« I I0) ej

c .o „-,CO 0

s qO 0)

« £f-i 0)d) •*-*

^5 CO

£ >>o «,CO &H

o^ <u3 -rt COCO > ^

« SQJ 73C c <=>T § «

2

LPiOO

•rH C^ c3 Tt

•6- CO

56

CD.Q l/>

IfC wion33 Ocn >-ro o

cfl ;3o o

•o

G §<u C

C g^"«H ~ coOS ^

to W> E05 .g rrCQ 5j J«bD 3 ^c -o °»

•S c "CO O £

° '-S £

o -2O W O^ CO CM

N "Q i-l

*X "^ NO Cl,

^ CD Scu ;;

*- 'Oco cphe<D 2 befciD 5? C

2 ft*- EQJ

•*-»- C J?

co-BCM 05

2 ^3 'oJ

3C

S ws <u

co

05

57

± 5

+ 4

± 3

± 2

± 1

—o—Calculated

—o—Flight data

sv' I LAP system gam,deg rad/g±12 _o^ 0

sv'deg±12 I LAP system gam,•o- ^. rad/g

^x .07

\\

2 2 2.4 2 6 2 8f, Hz

Figure 24. Effects of the nonlinear characteristics of the ILAF system onsystem performance. First mode; lightweight, M = 0.86; h = 7620 m(25 ,000f t ) ; 5 = 25°. ' P

58

±.3

± 2

± 1

—o—Calculated

--o—Flight data

6sv' I LAP system

deg gam, rad/g±8 0

sv'deg 1 LAP system±g gam, rad/g

^ .05

.05

2 0 2 2 2.4 2 6f, Hz

Figure 25. Effects of the nonlinear characteristics of the ILAF system onsystem performance. First mode; heavyweight, M = 0.87; h = 7620 m(25,000ft) ; 5 = 25°.

59

± 3

± 2

± l h -

0 20

sv' I LAP systemdeg gain, radfg

±4 ' 0xO^

/ \/ \

/ \

±4 .11

2.2 2.4 2 6f. Hz

Figure 26. Effects of the nonlinear characteristics of the ILAF system onsystem performance. First mode; flight-test data; mediumweight,M = 1 . 5 9 ; h = 11,918 m (39,100 ft); §t = 65°.

60

o Flight data

Data from reference 4

Calculated data used in this study

20

\— iviei

1n

rad/g

10

08

' 06

04 -

02

Mediumweight6j, cleg

± 50 Mediumweight,(VI = 1.59, h =11,918m (39,100 ft)

Heavyweight,M =0.87, h = 7620m(25.000ft) P

Heavy and lightweight —' \•Lightweight,

M = 0.86,h = 7620 m

(25,000ft)

10 20 30 40 50

w, rad/sec

(a) Amplitude.

Figure 27. Frequency response from the blended ILAF system accelerometerinput through the inboard elevon output.

61

o Flight data

Data from reference 4

Calculated data used in this study

<p, deg

-100 -

-120 —

-140 -

-160 —

-180

Heavyweight,M =0.87, h = 7620m (25,000ft)

Heavyweight,mediumweight,

Mediumweight,±.35 M=1.59, h = 11,918m

(39,100 ft) P

Lightweight,= 0.86,

h = 7620

(25,000ft)

10 20w. rad/sec

30 40 50

(b) Phase angle.

Figure 27. Concluded.

62

svg/deg

01

10 20 30w, rad/sec

FACS

FACS + ILAF

40 50

Figure 28. Calculated vertical acceleration response at the pilot's stationdue to shaker-vane input. Heavyweight, M = 0.87; h = 7620 m (25,000 ft) ;6 = 25°. P

63

.06

5svg/deg

05

04

03

02

01

FACS

FACS + ILAF

10 20 30w, rad/sec

40 50

Figure 29. Calculated vertical acceleration response at the pilot's stationdue to shaker-vane input. Lightweight, M = 0.86; h = 7620 m (25, 000 ft);5 = 25°. P

64

n

svg/deg

FACS

FACS + ILAF

10 20 30w. rad/sec

40 50

Figure 30. Calculated vertical acceleration response at the pilot's stationdue to shaker-vane input. Mediumweight, M - 1.59; h = 11,918 m(39,100ft) ; 5 t = 6 5 ° . P

65

svrad/deg

0008 i—

0004

10 20 30w. rad/sec

(a) Inboard elevon, FACS .

u2-4

rad/deg

0008

0004

10 20 30u), rad/sec

(b) Outboard elevens , FACS .

Figure 31. Calculated elevon deflection due to shaker-vane input with FACSor FACS + ILAF operating. Heavyweight, M = 0.87; h = 7620 m (25,000 ft);5 t = 2 5 ° . P

66

0012 r-

0008

1

sv

rad/deg0004

0008

2-4

sv .0004

rad/deg

10 20 30w, rad/sec

40

(c) Inboard elevon, FACS + ILAF.

10 20 30w, rad/sec

40

(d) Outboard elevens, FACS + ILAF.

Figure 31. Concluded.

50

50

67

Jsv

rad/deg

0008

0004

10 20 30u), rad/sec

40 50

J2-4

svrad/deg

0008

(a) Inboard elevon, FACS.

0004

10 20 30w, rad/sec

40 50

(b) Outboard elevens, FACS.

Figure 32. Calculated elevon deflection due to shaker-vane input with FACSor FACS + ILAF operating. Lightweight, M = 0.86; h = 7620 m (25, 000 ft) ;8 = 25°. P

68

0012 (—

0008

6S

rad/deg

0004 -

10 20 30w, rad/sec

40

(c) Inboard eleven, FACS + ILAF.

0008 r-

svrad/deg

0004-

10 20 30w, rad/sec

(d) Outboard elevens, FACS + ILAF.

40

Figure 32. Concluded.

69

0020

0016

0012

svrad/deg

0008

0004 -

10 20 30w, rad/sec

(a) Inboard eleven, FACS .

40

Figure 33. Calculated eleven deflection due to shaker-vane input with FACSor FACS + ILAF operating. Medium weight, M = 1.59; h = 11,918 m(39,100ft) ; 6 = 65°. P

70

00201-

0016

"2-46sv

rad/deg

.0012

10 20 30w, rad/sec

40 50

(b) Outboard elevens, FACS.

Figure 33. Continued.

71

0068 i—

rad/deq

.0004

10 20 30

w, rad/sec40 50

(c) Inboard eleven, FACS + ILAF.

Figure 33. Continued.

72

0028

0024 —

0020 -

0016 -

"2-46sv

radldeg0012

0008-

0004 —

0 10 20 30 40w, rad/sec

(d) Outboard elevens, FACS + ILAF.

Figure 33. Concluded.

50

73

Existing shaker-vanesystem

Amplitude,frequency

select

\ Remove input

Disengage

Actuator60

s +60

svAir vehicle

Gam select I Added component

Shaping usingexisting

components

Existing FACS iII

FACS accelerometer

Shaker-vane ILAFsystem accelerometer

Added component

Figure 34. Block diagram showing modification of existing shaker-vane andFACS equipment required to perform structural mode control function.

74

Amplitude ratio

1 08

.4

deg

1

0

-20

-40

-60

-80

-100

-120

-140

s + 40i

Shaping 2

J I

— Shaping 1

5 6 10

Shaping 2-

20 30 40 50 60

w, rad/sec

80 100

Figure 35. Shaker-vane ILAF system shaping networks. Analytical data,

75

08 r-

06

6sv

g/deg04

02

FACS

FACS + ILAFRevised system

FACS + ILAFRevised systemShaker vane

shaping 1

FACS + ILAFRevised systemShaker vane

shaping 2

10 20 30

w. rad/sec

KILAF'rad/g

-0 11

- 11

-.3

-.11

- 3

40 50

(a) Medium weight, M = 1 . 5 9 , h = 11, 918 m (39,100 ft) , 5 = 65°.

Figure 36. Performance of shaker-vane ILAF system. Shaker-vaneexcitation based on analytical data.

76

sv

g/deg

1.0

09

08

07

06

.05

04

.03

.02

01

FACS

FACS + ILAFRevised system

FACS + ILAFRevised systemShaker vane

shaping 2

KILAF'rad/g

-0.05

- 05

- 3

Ml

I10 20 30

w, rad/sec40 50

(b) Lightweight, M = 0.86, h = 7620 m (25,000 f t ) , = 25°.

Figure 36. Concluded.

77

K

FACS + I LAPRevised system

FACS + ILAFRevised systemShaker vane

shaping 1

FACS + ILAFRevised systemShaker vane

shaping 2

ILAF1

rad/g

-0 11

- 11

- 3

- 05

-.3

-90

tp, deg -180

-270

-36010 20 30

w, rad/sec40 50

(a) Mediumweight, M = 1.59, h = 11,918 m (39,100 f t ) , 5 = 65°.

Figure 37. Control system stability analysis, characteristic determinantphase angle . Analytical data.

78

K

FACS + I LAPRevised system

FACS + ILAFRevised systemShaker vane

shaping 2

ILAF'rad/g

-0 05

- 11

-.3

-90

<p, deg -180

-270

-36010 20 30

w, rad/sec40

(b) Lightweight, M = 0.86, h = 7620 m (25,000 ft) ,

50

= 25°.

Figure37. Concluded.

79

6 excitation input + 6 requiredsv sv Air vehicle

6 I LAP systemsv 3

feedback

Gam (shaping)

(a) Block diagram of shaker-vane system.

12

10

6 required,iv o

deg

10

ILAF system gam,rad/g

- 075

126 excitation input, deg

(b) Deflection requirements.

Figure 38. Shaker-vane deflection requirements operating both as anexcitation source and a structural mode control.

80

K

0005 r-

0004

0003 -

w,

2

rad/sec .0002

.0001 -

FACS

FACS + ILAF

FACS + ILAFRevised system

FACS + ILAF 'Revised systemShaker vane

shaping 2

10 20 30

w, rad/sec

40

|LAp

rad/g

-0 10

- 05

-.05

.3

0 0540

.0344

.0453

0253

Peakvalues

0 00157

00091

50

Figure 39. Performance of shaker-vane ILAF system . Gust excitation basedon analytical data. Heavyweight, M = 0.87; h = 7620 m (25,000 ft); 8 = 25°;o , = l ft/sec. P l

81

400

300

SVdeg/sec 200

100

14 i-

12

10

6 ,svdeq

o1—

ow , ft/secq

Required

Maximum

1 0 1.5ow , m/sec

Figure 40. Shaker-vane ILAF system deflection and rate requirementsas a function of gust intensity . Analytical data.

82 NASA-Langley, 1973 2 H-732

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