Investigation Into the Impact of Agility in Conceptual Fighter Design

NASA Contractor Report 195079

lit', ", ;

dm_ 7-7

Investigation Into the Impact ofConceptual Fighter Design

Agility in

R. M. Engelbeck

Boeing Defense & Space Group, Seattle, Washington

(NASA-CR-195079) INVESTIGATIONINTO THE IMPACT OF AGILITY ON

CONCEPTUAL FIGHTER DESIGN (Boeing

Defense and Space Group) 19i p

N96-10737

Unclas

G3/05 0063497

Contract NAS1-18762

June 1995

National Aeronautics and

,Space Administration

Langley Research Center

Hampton, Virginia 23681-0001

PREFACE

The work reported here represents the final report for NASA Langley contract NAS1-18762 Spacecraft &Aircraft Guidance and Control Task 22, Agility Design Study.

The NASA Project Engineer was M. J. Logan, and the Boeing Principal Investigator was R. M. Engelbeck.

ACKNOWLEDGMENTS

The authors would liketo recognize the contributionsmade by all the people at Boeing who made thiswork possible:

B. Ray Stability and Control Engineer

Helped break the conceptual design barrier

P. GotliebG. Letsinger

-- Stability and Control Engineerm Stability and Control Engineer

Agility implementation issues

S. Bockmeyer Aerodynamics Tech Aide

Graphic art and document publication

D. Bock Aerodynamics Engineer

Sizing and Performance

W. Mannick -- Weights Engineer

Weights and Technology Assessment

W. MooreDJRuzicka

Configuration EngineerConfiguration Engineer

ii

3.0

4.0

TABLE OF CONTENTS

5.0

Introduction and Summary ................................................................................................. 1

Study Requirements and Guidelines ................................................................................ .'62.1 Design Mission Profiles ......................................................................................... 6

Air Interdiction Mission Description ........................................................................ 6Air Superiority Mission Description ........................................................................ 6Multi-Role Mission Description .............................................................................. 6

2.2 Maneuver Performance Requirements ................................................................. 10Air Interdiction Maneuver Requirements .............................................................. 10Air Superiority and Multi-Role Maneuver Requirements ...................................... 10

2.3 Agility Requirements ............................................................................................ 102.4 Observables Requirements ................................................................................ 102.5 Carrier Suitability Requirements .......................................................................... 10

Technology Risk Assessment ......................................................................................... 153.1 Technology Risk Assessment Approach ............................................................ 153.2 Technologies Used in Agility Study Configurations ............................................ 373.3 Weight and Cost Impact of Advanced Technologies .......................................... 37

Configuration Development ............................................................................................. 474.1 Assumptions and Groundrules ............................................................................ 47

Single Crew ......................................................................................................... 47Twin Engine ......................................................................................................... 47F-22 Core Avionics Suite .................................................................................... 47Observables Features ........................................................................................ 49Control Effector Selection .................................................................................... 50

4.2 Carrier Suitability Impact on Aircraft Designs ...................................................... 51Geometric Limitations ........................................................................................... 51Weight Limitations ................................................................................................ 51Landing Gear Design ........................................................................................... 51Wing and Fuselage Structural Reinforcement ...................................................... 51

4.3 Designing for Agility ............................................................................................. 554.3.1 Agility Metrics ........................................................................................... 554.3.2 Preliminary Design Guidelines ................................................................. 594.3.3 Method Development .............................................................................. 62

4.3.3.1 Time to Bank and Capture - 90 Degree ................................... 624.3.3.2 Longitudinal Control Requirements ........................................... 64

4.3.3.2.1 Minimum Nose Down Pitch Acceleration ................. 664.3.3.2.2 Maximum TrimmableAngle-of-Attack ...................... 67

4.3.3.3 Maximum Lateral Side Force ..................................................... 674.3.4 Configuration Evaluation ......................................................................... 73

4.3.4.1 Agility Impact on Low Observable Configurations .................. 734.3.4.2 Observables Impact of High Agility Design ............................. 73

4.3.5 Summary

Configuration Synthesis Results ................................................................................ 825.1 Global Impact of Requirements on Design Space .............................................. 825.2 Aircraft Synthesis Results ................................................................................ 86

5.2.1 Air-to-Ground Configurations ................................................................... 865.2.2 Air-to-Air Configurations ........................................................................... 895.2.3 Multi-Role Configurations ........................................................................ 895.2.4 Joint Service Customer ........................................................................... 97

..°

III

6.0

7.0

8.0

Aircraft Design Data and Performance ........................................................................... 1026.1 Air Interdiction Concepts .............................................................................. 1026.2 Air Superiority Concepts .............................................................................. 1146.3 Multi-Role Concepts .............................................................................. 139

Critical Assessment o°o .... oooo° .......... o°o°°oo ............. oo°°oooo ............ ooooo°o°o°°oo°°o163

Flight Research Needs Assessment ............................................................................. 175

iv

1.11.21.31.42.12.22.32.4

"2.52.62.73.1-13.1-23.23.33.43.53.63.73.83.93.103.113.123.133.143.153.163.173.183.194.14.24.34.44.54.64.74.84.94.104.114.124.134.144.15 °4.164.174.184.194.205.1

5.2

LIST OF FIGURE_;

Fighter/Attack Aircraft Group - Metric Selection ResultsAgility Design Study - Scope and ObjectivesAgility Design Study ProcessAgility Design Study ConfigurationsNASA (NG) Design Mission - Battlefield Air InterdictionNASA (NA) Design Mission - Defensive Counter AirMulti-Role Strike Mission

Air-to-Ground Energy/Maneuverability RequirementsAir-to-Air Energy/Maneuverability RequirementsAgility Design GoalsCarder Suitability RequirementsGuidelines for Probability of FailureGuidelines for Consequence of FailureTechnology Risk Assessment - Control EffectorsTechnology Risk Assessment - AerodynamicsTechnology Risk Assessment - PropulsionTechnology Risk Assessment - Structures & MaterialsTechnology Risk Assessment - AvionicsTechnology Risk Assessment - VMS TechnologiesTechnology Risk Assessment - Crew SystemsTechnology Risk Assessment - WeaponsSample Technology Risk Assessment CriteriaRisk Assessment

Technologies Selected - Control EffectorsTechnologies Selected - Aerodynamics and PropulsionTechnologies Selected - Structures and MaterialsTechnologies Selected - Avionics and Vehicle Management SystemsTechnologies Selected - Crew Systems and WeaponsAdvanced Technology Applications - Weight and Cost EffectsAustere Avionics Suite for an Air-to-Air Agile FighterAustere Avionics Suite for a Multi-Role Agile FighterConfiguration EvolutionHistorical Weight TrendsAir-to-Air Agility Design GoalsAir-to-Ground Agility Design GoalsWings Level Turning OptionsControl Surface Deflections and Actuation RatesNominal Control Surface Size and Actuation Rates

Roll Agility Bank and Capture 90 degreesTypical X-ChartNormal Force Referenced to Plan View Gross AreaCenter of Pressure Referenced to Plan View Gross AreaSide Force Control Effector Selection

Side Force Control Effector Effectiveness with Angle-of-AttackDelta Wing Vortex Lattice ModelDelta Wing Aerodynamic CharacteristicsDelta Side Force Control Concept988-115 Vortex Lattice Model988-115 Aerodynamic Characteristics988-115 Control Effector Selection

Airplane CharacteristicsGlobal Design Space; Twin Engine Air-to-Ground Deltoid;Takeoff Gross Weight (16)Generic A/A Configuration; NASA Maneuver Requirements; T/W Required

v

LIST OF FIGURES - continued

5.35.45.55.65.75.8

5.95.10

" 5.115.125.135.14

5.15

5.166.16.26.36.46.56.66.76.86.96.106.116.126.136.146.156.166.176;186.196.206.216.226.236.246.256.266.276.286.296.306.316.326.336.346.356.366.376.38

Aircraft Synthesis ApproachAir-to-Ground Configuration Sizing ChartsImportant Design ParametersDesign Sensitivities About the Design PointAerodynamic Effect of Leading Edge DevicesDesign Sensitivity to Fuel Load for Max Neg Ps Requirement

Air-to-Air, High Agility Configuration Sizing ChartAir-to-Air, Medium Agility Configuration Sizing ChartMulti-Role, High Agility Configuration Sizing ChartMulti-Role, High Agility Configuration Sizing ChartMulti-Role, Medium Agility Configuration Sizing ChartGroup Weight Breakdowns for USAF and Joint Service Aircraft -Full Mission PerformanceGroup Weight Breakdowns for USAF and Reduced Mission PerformanceJoint Service Aircraft

Carrier Suitability Assessment of Joint Service DesignsGeneral Arrangement Drawing - 988-122Inboard Profile - 988-122

Group Weight Statement - 988-122C. G. Envelope - 988-122Inertia Data at Combat Weight - 988-122Aircraft Geometry - 988-122Design Mission Segment Performance Breakdown - 988-122Summary of Point Maneuver Performance - 988-122General Arrangement Drawing - 988-123Inboard Profile - 988-123

Group Weight S{atement - 988-123C. G. Envelope - 988-123Inertia Data at Combat Weight - 988-123Aircraft Geometry - 988-123Design Mission Segment Performance Breakdown - 988-123Summary of Point Maneuver Performance - 988-123

General Arrangement Drawing - 988-119Inboard Profile - 988-119



Group Weight Statement - 988-115C. G. Envelope - 988-115Inertia Data at Combat Weight - 988-115Aircraft Geometry - 988-115

vi

LIST OF FIGURES - continued

6.396.406.416.426.436.446.45

6.46" 6.47

6.48

Design Mission Segment Performance Breakdown - 988-115Summary of Point Maneuver Performance - 988-115General Arrangement Drawing - 988-114Inboard Profile - 988-114

Group Weight Statement - 988-114C. G. Envelope - 988-114Inertia Data at Combat Weight - 988-114Aircraft Geometry - 988-114Design Mission Segment Performance Breakdown - 988-114Summary of Point Maneuver Performance - 988-114

vii

1.0 Introduction and Summary

The work contained in this report was accomplished as part of the NASA Langley Research Center(LaRC) Agility Design Study Activity. The purpose of the NASA Agility Design Study is to assess theimpact of spedfic agility requirements on the aircraft design decisions.

Previous work leading up to this phase of the study provided a set of agility metrics to be used tocategorize aircraft agilityand the methodologyto assess these metrics. These metrics are identified infigure 1.1.

The purpose of the current phase of study is to conduct configuration design studies to determinethe impact of varying levels of agilityrequirements on a wide spectrum of potential aircraft andmissions. Lockheed has investigated the impact of agility requirements on an existing airframe in thefulfillmentof a multirolefighter mission. McDonnell-Douglas has investigated new designs inthefulfillmentof the same multi-role fighter mission. This contract report addresss the effects of customerrequirements (NAVY Vs Air Force) and aircraft mission role (Air Superiority, Multi-Role, and AirInterdiction)on agility design decisions. The study process is presented in figure 1.3.

The requirements for the aircraft designs are presented in section 2.0. The concepts presented hereare intended to be representative of high end, next-generation replacements to the A-6 AirInterdictionand F-15/F-14 Air Superiority aircraft. The Multi-Role concepts represent a compromisedesign between the dedicated Air-Superiority designs and the dedicated Air-Interdiction designs. Inaddition to mission role, the impact of customer requirements (primarilycarrier suitability)andobservably levels were used to develop the matrix of configurations studied and presented infigure 1.4.

A technology risk assessment was accomplished using a list of suggested technologies supplied byNASA as a point of departure. The results of the risk assessment presented in section 3.0 were thenused as the basis of selecting subsystems and technologies available for use in the development ofthe individual configurations studied.

Several of the technologies on the NASA supplied listwere in reality a configuration conceptdependent list of control eftectors. As part of the configuration design trade studies presented insection 4.0, a selected subset of control effectors identified for use on each of four basicconfiguration types. Control sizing studies were conducted to determine the most effectivecombination of control effectors required to meet all the agility design requirements. Themethodology used and results are presented in detail for use as design guidelines in selectingindividual control effe'ctors,or combinations of control effectors, necessary to achieve an agility levelfor a given application.

Twelve configurations were studied under this contract, six Air Force aircraft and their six derivativejoint service counterparts. Trade studies documented in section 5.0 were conducted to identify theimportant design parameters and driving design constraints. These constraints were then used in theselection of the design points.

Once each individualdesign point was selected, three-view drawings and interior layouts werefinalized. Group Weight statements, Center-of-gravity envelopes, Inertia estimates, drag polars,maneuver point performance and mission breakdowns were also finalized and presented insection 6.0

The results of a criticalassessment are presented in section 7.0.

Section 8.0 contains recommendations for flight research.

Fighter/Attack Aircraft GroupMetric Selection Results

• Working group consensus• Government inputs: NASA, AF

• Industry inputs: BoeingjEidetics, General Dynamics,McDonnell-Douglas

• Metrics Selected:

PO Metric

1. Maximum negative Ps

2. Time-to-bank 90°

3. Minimum nose-down pitchacceleration

4. Maximum achievable, trimmedangle-of-attack

5. Maximum lateral acceleration

Conditions

0.6M @ 15,000 ft., max inst.450 kts @ sea level, max inst.

0.6M @ 15,000 ft, max inst. Nz450 kts @ sea level, 5g

Condition for Cm*

Subsonic

Max inst. Nz (air-to-air)lg wings level (air-to-ground)

Figure 1.1M $299-0363 -2 01/26/94

Agility Design StudyScope and Objectives

Purpose• Investigate impact of agility on design decisions• Identify NASA research needs• Develop agility design guidelines

...................... ROEIAYO

GO

Objectives:• Design 12 configurations to address the issues of:

- USAF vs Joint Service customers- Aircraft Mission Role- LO vs Agility

Mission

Air Superiority

Multi-Role

Air Interdiction

MediumAgility

Air Force Only

LowObservables

LowObservables

LowObservables

i HighAgility

ModerateObservables

ModerateObservables

LowObservables

Joint Service

MediumAgility

LowObservables

LowObservables

LowObservables

HighAgility

ModerateObservables

ModerateObservables

LowObservables

Figure 1.2M_ -401/26/94

I NASA DesignRequirements [_1

I NASA st[_

Technology Li I

Technology Risk IAssessment

I Section 3.0

Study Assumptionsand Groundrules

_, SecUon 2.0

Initial Configuration 0_._

Concepts

Section 4.

Control Effe_orSelection

I S_ion 4.0

ConfigurationSynthesis

_ Section 5.0

Aircraft DesignData and Performance

_ Section 6.0

CriticalAssessment

_ Section 7.0

Flight ResearchNeeds Assessment

_ Section 8.0

Conclusionsand IRecommendations

Section 9.0

Figure 1.3. Agility Design Study Process

spb-7-7-re-Ad8

4

01

Air-to-ground Multi-role Air-to-air

Low Moderate Low Moderate Low Moderateobservables observables observables observables observables observables

988-122

_,l_

988-123

988-118 988-114

988-119

Figure 1.4. Agility Design Study Configurations

m

988-115

USAF Customer

Joint Sen/Ice

spb-7-7-re-Adl0

2.0 Study Reauirements and Guidelines

2.1 Design Mission Profiles

Air Interdiction Mission Description

The Battlefield Air Interdiction Mission defined by NASA is a 1000 Nm High-Lo-Lo-High profilepresented in figure 2.1. The design payload consists of four GBU-27 laser guided bombs and twoAIM-9 missilesfor self defence. The intent of the mission is to describe a reasonable interdictionrange/payload. NASA contindeds that a significant air power deficiency was discovered during DesertStorm in that missionranges were limitedto about 600 Nm total radius, virtuallyall of whichwas flown atmoderate altitudes. Those missionswhich necessitated low-altitude attacks (eg. Tornado airfieldattacks, F-16 Interdiction) required numerous aircraft since fuel tanks were the majorityof the storeIoadings. The Navy A-6E is currently capable of 450/300 Nm leg distances using the J52 engine withan external payload similarto that called out here. To accomplish thisthe A-6 does carry a centerline300 gallon fuel tank. NASA expects an F404 engined A-6 would probably be able to accomplish themission described.

Takeoff fuel allowance is modeled by 20 minutes at idle power and 2 minutes at maximum augmentedpower. Both the inbound and outbound high altitude cruise legs are at optimum Cruise Mach numberand altitude, with a radius of 600 Nm. The ingress leg is 500 KTAS at 200 feet altitude for 400 Nm.The combat leg over the target consists of four sustained turns at Mach 0.8 at Military Power setting.All four GBU-27s are expended along with 500 rounds of 30 mm ammo. The egress from the targetarea is accomplished at 550 KTAS at 200 feet altitude for the same 400 Nm radius as the ingress.

Air Superiority Mission Description

The design mission for the dedicated air superiority concepts to replace the F-14 and F-15 is theDefensive Counter Air (DCA) mission presented in figure 2.2. This mission has a total radius of 450Nm with a payload of four AIM-120 missiles, two AIM-9 missiles and 500 rounds of amino. Takeoff fuelallowance is modeled by 20 minutes at sea level and idle power followed by 2 minutes at maximumafterburner. The outbound leg to the aircraft combat station consists of a 350Nm cruise legaccomplished at best cruise altitude and Mach number followed by a 90 minute loiter on station atMach 0.8 at 40000 feet. The stationkeeping is followed by a 1.5 Mach dash (dry power) to interceptinbound adversaries. Combat is modeled by four sustained turns at 40000 feet, Mach 0.9 at maximumaugmented thrust with the expenditure of four AIM-120 missiles and 50% of the ammunition. Afterthe combat segment, a military power climb is executed for the 450 Nm inbound cruise at optimumMach number and altitude. The aircraft lands with reserves of 5% mission fuel at its point of originaftera 20 minute loiter at sea level and optimum Mach number.

Multi-Role Mission Description

The multi-role mission presented in figure 2.3 is a compromise between the rigorous radiusrequirements of the air interdictiondesign mission radius and payload. The design weapons load istwo 2000 IbJDAM laser guided bombs. This payload was reduced from that of the air interdictionmission discussed because the two advanced laser guided weapons would not suffer unacceptably inPk relative to the four GBU-27s carded in the air interdictionmission. Combat maneuver performanceof the air superiority design is maintained. The takeoff allowance was reduced to ten minutes at idlepower, 15 seconds in intermediate power, and 15 seconds in maximum afterburner. After takeoff amilitarypower climb is initiated until the aircraft reaches its optimum cruise altitude. The outboundcruise leg is 650 Nm at optimum cruise Mach number. The aircraft then drops to a penetration altitudeof 20,000 feet to ingressto the target at 540 KTAS for 50 Nm. The 700 Nm total missionradius of themulti-role strike mission still exceeds the 600 Nm mile limitationpresented in the air interdictionmissiondiscussion, without the use of external tanks. Over the target, the aircraft drops its air-to-ground weapons load of two JDAM laser guided bombs. Combat over the target is modeled by a 180degree sustained turn at 540 KTAS at 20000 feet using dry power. The aircraft 50 Nm egress fromthe target area is accomplished at 540KTAS and 20000 feet. At this pointthe aircraft enters an airengagement modeled by a 360 degree turn at 540 KTAS and 20000 feet using maximum dry power.

6

NASA (A/G) Design Mission - Battlefield Air Interdiction(4) GBU-27 + (2) AIM-9 + 1000rd 30mm

Reserves:

• 20 ein. Loller at SL/Opt g• 5% of Mission Fuel

Optimum Mach/AIt Mil Pwr.Climb

Mil Pwr.Climb

WUTO: 20 Min. Idle + 2 Min. Max. A/B +

T.O. & Accel to Climb Speed in Max. Pwr.(No Distance Credit)

Combat: 0.8M/Mil Pwr.

• (4) Sust Turns• Drop (4) GBU-27s• Expend 500 RDS 30mm Ammo

Ingress: 500 KTAS/200 Ft.Egress: 550 KTAS/200 Ft.

600 NMi 400 NMi "-

Figure 2.1 _ spb.7-7.re-McD2

NASA (A/A) Design Mission - Defensive Counter Air(2) AIM-9. (4) AIM-120

Reserves:

• 20 Min. Loiter at SL/Opt M• 5% Mission Fuel

GO

Climb-Cruise BCA/BCM

Climb-Cruise BCA/BCM

1.5 DASH (Dry), 40,000 Ft.

Mil Pwr.Climb

90 Min. Loiter at 40K Ft., M= 0.8Accelerate to M=1.5

Mil Pwr.Climb

WUTO: 20 Min. Idle + 2 Min. Max. A/B +

T.O. & Accel to Climb Speed in Max. Pwr.

(No Distance Credit)

Combat: 40K Ft./M0.9/Max. Pwr.

• (4) Sustained Turns• Launch (4) AIM-120s & 50% Ammo• Accel. from M=0.9 to M=1.5

• Launch (2) AIM-9s

350 NMi

spb-7-7-re-McD1Figure 2.2

BoeingDefense &Space Group Multi-Role Strike Mission

Reserves: • 20 Min. Loiter at S.L./Opt. M• 5% of Mission Fuel

Optimum Mach/AIt__ Mil Pwr.Climb

Weapons: (2) JDAM - 2,000 Ib(2).AIM - 120400 RDS Ammo 20 mm

(D

Mil Pwr.Climb

Combat:• 360 ° Sust. Turn

54O KTAS

Max. Dry Pwr

WUTO: 10 Min. Idle +15 Sec Int. +15 Sec Max. AB

650 nm

Ingress: 540 KTAS, 20,000 ft.Egress: 540 KTAS, 20,000 ft.

_ .

Combat:• Drop (2) JDAM° 180 ° Sust. Turn

540 KTAS

Max. Dry Pwr

50 nm =I

Figure 2. 3.

The aircraft escapes the engagement an executes a militarypower climb to optimum cruise altitude.The aircraft then returnsto its base 650 Nm away. Reserves are specified as 20 minutes loiterat sealevel at optimum Mach number plus 5% of mission fuel.

2.2 Maneuver Performance Requirements

Air Interdiction Maneuver Requirements

The maneuver requirements for the air interdictionand multi-role designs is presented in figure 2.4.

Air Superiority and Multi-Role Maneuver Requirements

The NASA defined Air-Superiority Maneuver Requirements are intended to be approximately 10%better than the F-14 and F-15 maneuver capabilities. There are 25 maneuver conditionscalled out infigure 2.5. The multi-role designs meet the same maneuver requirements as the air-superioritydesigns.

2.3 Agility Requirements

There five agilitydesign metrics presented in figure 2.6 along with goals for aircraft designed for theAir Interdiction(AG) and Air Superiority (AA). The Multirole concepts are designed to the same agilityrequirements as the Air Superiority concepts.

2.4 Observables Requirements

The purpose of this study is not to develop low observables technology, but rather to assess agilityrequirements impact on aircraft with varying degrees of stealth characteristics. This purpose and thesensitive nature of observables technology lead to the establishment of the observablesrequirements used in this study. For the purposes of this study, low observables is defined as a levelof observables consistent with the B-2, and moderate observables is defined as a level of observablesconsistent with the F-22. No actual observables assessment will be conducted on the designs orreported. Observables are addressed purely as a qualitive measure and implemented by thedesigners to be consistent with the requirements and their experience.

2.5 Carrier Suitability Requirements

The joint service concepts must meet the carder suitability requirements presented in figure 2.7 inaddition to all the requirements met by their counterpart Air Force concepts. The catapult wind-over-deck required with the aircraft at its design gross weight is zero knots on a C13-1 catapult. The singleengine rate-of-climb after launch on a tropical day is 200 feet/minute. There is no specified arrestedwind-over-deck requirement, but the single engine rate-of-climb after an aborted approach is 500feet/minute. The desired carder deck spotting factor is 1.0 relative to the F-18, not to exceed 1.31.

10

BoeingDefense &Space Group

Air-to-Ground Energy / ManeuverabilityRequirements

..¢

Requirement

Combat Ceiling ...............................................................................................................................................................40,000 ft

Accelerate From 300 Kts to 550 Kts (Sea Level) .....................60 sec

Sustained Load Factor (Sea Level, Mach = 0.8) ........................6.5 g's*

Instantaneous Load Factor .....................................................................................................................9.0 g's

Unrefueled Ferry Range .....................................................................................................................3,000 nmi

'_ With Stores and 60% Fuel

Figure 2.4 K120 - 25 February 1994


Air-to-Air Energy / ManeuverabilityRequ=rements

PO

Mach Altitude Sustained Instantaneous Excess Power-(Kft) g's g's Ps (fps)

0.6 0 900......................... ....................................... I ............................................................. t ............................................................... I................................................................ I...............................................................

0.9 0 1,300

0.6 10 6.0 8.0 650................................................................ I ............................................................... f ............................................................... I.................................... : ........................... I ................................................................

..........................o.9.......................................................,Io...................................................., g.O,......................................................................................................................................................., , 1,ooo1.2 10 600

............................................................... I ................................................................ I ............................................ .................... I ............................................................... ! ...............................................................

0.6 20 4.4 450............................................................... I ............................................................... I ............................................................... I ............................................................... l ...............................................................

0.g 20 7.0 g.0 * 800................................................................ i ............................................................... f, ............................................................................................................................... i ................................................................

1.2 20 6.8 650

1.4 20 600............................................................... I ................................................................ I ................................................................ I ............................................................... 4 ................................................................

0.g 30 5.0 9.0" 550................................................................ I ............................................................... t ............................................................... I ................................................................ I ...............................................................

1.2 30 5.0 9.0 _' 500................................................................ I ....................................... :....................... t ............................................................... I ................................................................ I ...............................................................

1.4 30 600

• Acceleration Time from Mach = 0.9 to 1.5:<60 sec. at 40,000 ft. ,_Structural Limit• Combat Ceiling: >55,000 ft.• Unrefueled Ferry Range: >3,000 nmi with AIM-9 / 20mm Stores Retained

Kl19 - 25 February 1994

Figure 2.5

BoeingDefense &

Space Group

Fighter / Attack Aircraft GroupAgility Design Goals

Co

Metric

1. MaximumNegative Ps

A-A:

Conditions

Mach = 0.6, 15 Kft (Nz = 5.5g)

Low

-800ft/sec

Medium

-450ft/sec

2. Time-to-Bankand Capture 90 °

3. MinimumNose-DownPitchAcceleration

4. MaximumAchievableDeparture-FreeAngle-of-Attack

. MaximumLateralAcceleration

A-G: 450 Kts Sea Level (Nz = 7.5g)

A-A: Mach = 0.6, 15 Kft,Maximum Instantaneous Nz = 9.0

A-G: 450 Kts Sea Level, 5g

A-A: Condition for Cm*Use Mach = 0.6, 15 Kftfor Consistency

A-G: Same

With Air-to-Air Stores, Subsonic

A-A: Mach = 0.6, 15 Kft,Maximum Instantaneous Nz = 9.0

A-G: 450 Kts Sea Level, Wings Level

(Same)

3.0 sec

2.0 sec

-0.05rad/sec2

(Same)

25 deg

0.25 g

0.6 g

(Same)

2.5 sec

1.5 sec

-0.15rad/sec2

(Same)

40 deg

0.4 g

1.2 g

High

-100ft/sec

(Same)

1.5 sec

1.0 sec

-0.35rad/sec2

(Same)

70 deg

1.0 g

2.0 g

Figure 2. 6

spb-8-7-re-Ad6


Ca rrier .Su ita b iIityRequirements

Catapult(c13-1)

Requirement

• WOD Requirement at Design Gross Weight ........................................0 Kts• Single Engine Out Rate-of-Climb .............................................................200 ft/min

Arrest(Mk.7 Mod 2)

• WOD Requirement at Design Landing Weight .................................None• Single Engine Out Rate-of-Climb .............................................................500 ft/min

SpottingFactor

• Desired ............................................................................................................................................................................1.00• Required ......................................................................................................................................................................1.31

Figure 2. 7 K121 o 25 February 1994

3.0 Technoloav Risk Assessment

The objective of the technology assessment task was to identifythe technologies that provide thegreatest benefit for the twelve candidate Agility Design Study (ADS) concepts and also to help NASAidentify meaningful research needs which, if accomplished, will improve future aircraft design,manufacturing and performance. '

3.1 Technology Risk Assessment Approach

A "technology matrix"was developed by the Boeing Military Airplanes (BMA) technology staff usingthe technology list provided by NASA with additions and combinations as deemed necessary to bestidentify the technologies that might be configuration drivers or, required to satisfy the ADS missionperformance criteria. The basic ground rules used by the technical assessment experts were that theIOC date would be 2005 and development testing (materials, systems, aerodynamics, etc.) would beaccomplished. The technology assessors were also required to:

(1) Provide a brief description of the individual technology.(2) Provide a rationale for determining whether the technology should or should not be

selected for incorporation into ADS configurations.(3) Provide the expected impact, either beneficial or detrimental, the technology would have

on the configurations if incorporated into the design.(4) Provide a subjective assessment of the probability and consequence of failure as

determined by the ground rules shown in Tables 3.1-- 1 and 3.1-2 and described inSection 3.1.1.

(5) Provide a suggestion of research needed to bring the technology to maturity andvalidation.

The resulting "technology matrix" is shown in figure 33.1-3through figure 3.1-16.

Probability and Consequence of Failure Determination

Each technology was rated in terms of Probability of Failure (POF) and Consequence of Failure (COF)as outlined by the guidelines specified in figure 3.1- 1 and 3.1-2. The technology assessment usedPOF as the probability that the identified technology will or will not be available for aircraft application atthe IOC date specified. Likewise, COF is the consequence to the aircraft if the identified technologyis not available for application. Using the Probability of Failure guidelines, each proposed technologyhas been considered with respect to its maturity, complexity and level of support base. In assessingthe Consequence of Failure, each technology has been considered with respect to aircraftperformance, cost and schedule impacts. The POF and COF values shown in the tables were only tobe considered as guidelines and not absolutes. All technology assessors subjectively determinedPOF and COF risk levels for each proposed technology implication based on the imposed guidelines.These guidelines are a combination of Boeing and Defense Systems Management College (DSMC)criteda for determining risk.

The standard risk plot of POF vs COF which was used by the technical experts to assess the risk levelof each proposed technology is shown in figure 3.1- 17. On the plot are lines that represent whatBoeing Military Airplanes (BMA) believes are acceptable limits of POF and COF going into aDemonstration/Validation (DEM/VAL) phase and a full Scale Development (FSD) phase of an aircraftdevelopment cycle. Acceptable values of POF and COF for entering the DEM/VAL phase ar less thanor equal to 0.5. Acceptable values for entering the FSD phase are less than or equal to 0.3.

Each proposed technology's POF and COF were plotted and are shown in figures 3.1- 18 through3.1- 22. All the technologies are identified by a number on the plots for quick reference. Thetechnologies which were selected to be used in the evaluations of the ADS "point design"configurations are shown as shaded areas in the tabulations on the left of each figure. Examining therisk plots and considering the acceptable values as defined for DAM/VAL and FSD, the technologiesthat require the most attention can be identified and earmarked for future meaningful researchactivities.

15

Complexity of Support baseValue Maturity of hardware/software hardware/software

0.1 Existing equipment; in production

0.3

0.5

0.7

0.9

Minor redesign, prototype/engineering model flight tested;extensive lab demonstrations

Major change feasible,preliminary brassboard

Proof of concept in lab environment,complex hardware design, newsoftware similar to existing

Concept formulation, some research,never done before

Simple

Somewhat complex

Fairly complex

Very complex

Extremely complex

Multiple programsand services

Multiple programs

Several parallelprograms

At least one otherprogram

No additionalprograms

Figure 3.1-1. Guidelines for Probability of Failure

Value

0.1

0.3

0.5

0.7

0.9

Fall back solutions

Several acceptableaitematives

A few knownalternatives

A single acceptablealternative

Some possiblealternatives

No acceptablealternative

Cost factor

Highly confidentwill reduce LCC

Fairly confident willreduce LCC

LCC will not changemuch

Fairly confidentwill increase LCC

Highly confidentwill increase LCC

Schedule factor

90-100% confidentwill meet IOC





Downtime factor

Highly confident willreduce downtimesignificantly

Fairly confident willreduce downtimesignificantly

Highly confident willreduce downtimesomewhat

Fairly confident willreduce downtimesomewhat

Downtime may notbe reduced much

Figure 3.1-2. Guidefines for Consequence of Failure

spb-7-7-re-Ad 1

16

¢

Control Effectors

Technologydescription

ConventionalAilerons

Tiperons

Trailing EdgeManeuver Flaps

Leading EdgeFlaps or Slats

Blown ControlDevices...,t.

,,,j

Porous Leading/Trailing EdgeDevices

Leading EdgeSuction/Blowing

TangentialWing Blowing

Drag rudders

Spoilers/Speedbrakes

Selection rationale

Roll performance at highAOA better than spoilers

Proven

Increased high liftcapability

Use to reduce controlsurf size or to increasecontrol power

May provide increased highlift with low RCS & reducedcomplexity compared toslats/slotted flaps

Increased high lift

Provides high lift with lesscomplexity

Provide yaw control withno vertical fins

Wing spoilers very effectiveahead of high lift flaps

Configuration Impact

Benefit I

Effective tohigh AOA

Increasedmaneuver-ability

Low takeoff& approachspeeds

Significantincrease ineffectiveness

Low RCS,reducedmechanicalcomplexity

Lower T.O.& appr. speed

Simplersystem thanslot blowing

Low RCS

Low AOAeffectiveness

F/gure

Penalty

Heavy attackpivot req'd.

Weight

Increased IUTandcomplexity

Ineffective athigh speed

Unprovenconcept

Complexity

Ineffective athigh speed

Reducedeffectiveness

Weight, poorhigh AOAeffectiveness

_.1-3.

Proba-bilityof failure

.10

.60

.01

.10

.40

.70

.50

.70

.40

.10

Consmquenceoffallure

.I0

.30

.01

.I0

.6O

.7O

.30

.7O

.3O

.5O

Recommended research

Wind tunnel test databaseneeded to quantify benefits

Wind tunnel databaseneeded for flexible controlconcepts

Share LE slateffectiveness

More wind tunnel test dataneeded to prove concept

More wind tunnel test dataneeded to prove concept

Wind tunnel test databaseneeded to quantify benefitsvs. blowing requirements

Wind tunnel databaseneeded as a function ofdeflection and wing planform

spb-2/g4-re-Ad 1

Control Effectors (continued)

-.J.

CO


Horizontal TailWith Elevator

All-MovingHorizontal Tail

Variable IncidenceWing

Vertical TailWith Rudder

All-Moving Canard

Other MovingFin(s) or YawVanes

DoubleHinged/SplitControl Devices

ArticulatingForebody Strakes

ArticulatingChine

Selection rationale

This type of control is onlyappropriate for subsonica_rplanes

High speed, high agilityairplanes need high pitchcontrol power

This is an option if the highspeed design of the air-plane results in unaccept-able over-the-nose visibility

Standard low risk approachto yaw control/directionalstability

Provides both pitch andyaw control if positionedproperly

Better control at super-sonic speed than fin withrudder

Provide more control powerthan single hinge. Mayresult in reduced fin size.

May provide high AOA yawcontrol to supplementrudders


Configuration Impact --_

Benefit

May be some-what lighterthan all-movinghoriz, tail

Goodeffectivenessthroughoutspeed range

Providesgood over-the-nosevisibility

Proveneffective

Proveneffective

Good highAOAeffectiveness

Increased yawcontrol

High AOAyaw control

High AOAyaw control

Penalty

Pooreffectivenessat supersonicspeed

Requires highhorsepowerhydraulicsystem

Weight

Weight

RCS, poorpilot visibility

Weight

Weight

IncreasedRCS

Weight,complexity


.10

.10

.30

.10

.20

.10

.10

.70

.65

Conse-quenceof failure

.30

.30

.70

.30 •

.60

.30

.10

.60

.65


Develop wind tunnel database of effectiveness onchined forebodies

Wind tunnel research neededto quantify effectiver_ess forvarious forebody shapes

Figure 3.1-4. spb-2/94-re.P, d2

Control Effectors (concluded)

.-¢

_O


Forebody JetBlowing

Forebody SlotBlowing

Forebody Suction

Articulating NoseStrakes

Body Flaps

Fluidic ThrustVectoring

Pitch AxisMechanical ThrustVectoring

Multi-AxisMechanicalThrust Vectoring

Selection rationale





Pitch control due to bodyflap may allow smallerhorizontal tail

Provide increased yawcontrol

Low risk approach toincreased pitch controlpower

Low risk appraoch toincreased combined pitchand yaw control power


Benefit

High AOAyaw control

High AOAyaw control

Hig h AOAyaw control

High AOAyaw control

High AOApitch control

Low RCS

Increasedmanevuerability

Increasedmaneuver-ability

I Penalty

Weight,complexity

Weight,complexity

Weight,complexity

Weight,complexity

Weight

Complex

Weight

Weight

Proba-bility;of failure

.70

.70

.70

.60

.50

.70

.30

.40


.60

.60

.60

.60

.30

.70

.20

.30


Wind tunnel research neededto quantify effectiveness forvarious forebody shapes


Wind tunnel research needed

to quantify effectiveness forvarious forebody shapes


Continue to develop to attainincreased vectoringcapability

Figure 3.1-5.

$pb-2/g4-re-Ad3

Aerodynamics

POC_


Vortex Flap

Variable Camber/Mission AdaptiveWing

Natural Flow Wing

Porous LiftingSurfaceTechnology

Natural LaminarFlow/SupercriticalWing

Hybrid LaminarFlow System

Forward SweepWing Technology

Selection rationale

Improved cruise &manevuer IJD

Improved L/D overentire operating envelope

Improved cruise L/D

Reduce shock strength

Reduce cruise drag

Reduce cruise drag, t L/D

Improved stallcharacteristics & high-c_aero performance


Benefit I Penalty

10-20% L/Dimprovement(better onhigh-A wings)

10-40% L/Dimprovement,depending onmission profile

10% L/Dincrease insome cases

10% increasein Machcapability

10-20% dragreduction

10-40% dragreduction

• Highersustainableangle-of-attack

• Improvedhigh-e¢maneuver

• Higher wingweight

• Signaturepenalty

Wing weight& complexityincreased

Potentialcurvature/manufacturingproblems

• Potential drag

• penaltySurfacecomplexity

• Maintainability

Potential toincreasemanufacturingcost

Increasedweight &complexity

Increasedstructuralweight

Proba-bilityof faUure

.3O

• .30 forLE/TE

• .70 forfullchord

.30

.60

.2O

.40

.4O

Conse-quenceof fallure

.20

.30

.30

.40

.3O

.3O

.50


Wind tunnel testing forsha.rp & semi-sharpleading edges, with variousALE

• CFD & wind tunnel testsof variable geometry- LE/TE vs. full chord

• Structural concepts for fullchord system

CFD & wind tunnel tests ofnew technology appliedto realistic configurations

Flight test samplesDetailed CFD & wind

tunnel shock strengthvs. drag trade

Improve CFD capabilityfor transition prediction

Improve CFD transition

i predictionWind tunnel & flighttesting of options

• CFD & wind tunnel &flight tests on configurationsof interest in addition toX-29

• Aero-structural optimization

Figure 3.1-6. spb-2/94-re-Ad4

b

Propulsion

Po

Technologydescrlptlon

IHPTET Gen 5Engine Technologies

IHPTET Gen 6Engine Technologies

FADEC/PSCTechnologies

Variable CycleEngine Technologies

F100/F110 DerivativeEngine Technology

F119/F120 DerivativeEngine Technology

Selection ratlonale

• Higher performance• Lower weight• Standard performancelevel for year 1997

• Improved LO signature• Higher performance• Lower weight• Standard performancelevel for year 2008

•Controls with increasedcomputing capability

•Greater reliability• Reduced weight &volume

• One solution to highthrust yet longendurance missions

• Reduced fuel load

• Lower cost• Available now

• Lower cost• Available near term


Benefit

+30% T/W-20% TSFC

+60% T/W-30% TSFC

Optimizedengineoperation

Smaller vehicledue to reducedfuel load

Reliability baseexists• +20% FN• +25% T/W• -3% SFC

• +20% FN• 2 x turbine life• -5% SFC

I Penalty

Higher cost

Morecomplexsystem

•Valving h/wis heavy

• Reliability• Number ofmoving parts

Lower thrust/weight ratio

Lower thrust/weight ratio


.4O

.60

.50

.20

.20

.30


.I0

.50

.20

.30

.30

.30


• High strength, low weightmaterials

• New aerodynamic designof compressors & turbines

• Efficient cooling techniques

• Advanced cooling• Endothermic fuels• Engine controls• Materials• Ceramics

Variable/engine controlintegration

• Bypass vs. coreperformance

• Matching missionparameters

• Mission tailored enginecycle

• Derivative feasibility study• Cost vs. schedule vs.performance

• Mission tailored enginecycle

• Derivative feasibility study• Cost vs. schedule vs.

performance

Figure 3.1-7.

spb-2/9,4-m-Ad5

.Structures & Materials

PO


Advanced Aluminum-Lithium Alloys

Advanced TitaniumAlloys

Powder Metallurgy

(2 types)Current Materials

• Metal MatrixComposites

Intermetal Ceramic

Rare Earth Alloys -Sapphire

Graphite BasedComposites

Boron BasedComposites

Selection ratlonale

Reduce weight ofaluminum parts

Reduce weight oftitanium parts

Confusingl -

Only cost savings

Weight benefit:silicon titanium, etc.

• Save weight• Very smooth complex

surfaces

• Save weight• Very stiff


Benefit Penalty

10% weightreduction to30% ofstructure

10% weightreduction to30% ofstructure

$ only

30% weightsavings on50% of struct.

Weight savingsin HOT areas.20% of 1%.

10% weightsavings to40% of struct.

10% weightsavings to40% of struct.

20% costpenalty

20% costpenalty

Needsdevelopment

$ costincrease

$

• Very expensive• Hard to work

Proba-bllityof failure

.50

.50

.50

.50

.70

0.7

.50

0.3

Conse-quence!of fallure

.50

.50

.50

.50

.50

0.6

.10

0.7


Caution: this has been triedbefore and FAILS due topoor ductility

Caution: previous failuresdue to lack of weldability& crack growth

Probably not worth effort

Putting fibers in metalsmetal matrix composites)s potentially major benefit

Not used much inairframes - more applicationto engines

Not used in airframe

Improved materials are nice,but breakthrough will be newjoining and manfacturingmethods

Competes with graphitecomposites but moreexpensive

Figure 3.1-8.

spb-2/94-re-A¢110

Structures & Materials (continued)

POGO


Kevlar BasedComposites

Fiberglass BasedComposites

Advanced Resins

ThermoplasticMaterials (Arimid K.series developed byDuPont with HighGlass TransitionPolyemide Systems)

ThermosetMaterials

AdvancedManufacturing• Superplastic

Forming

T. WeldingComposite Welding

• Z Pinning

Selection rationale

Kevlar is very tough& impact resistant

Graphite stiffer?

Could save weight viaimproved toughness

• Very tough resin• Saves weight• Potential formanufacturingbreakthroughs

Could save weight viaimproved toughness

Could save costand weight

Could save costand weight


Benefit

May save 20%weight on 10%of structure

May save $

10% weightsavings on 40%of structure

20% savingson 40% ofstructure




I Penalty

$, plus onlyhelps impactsensitive parts

May costweight

May increasecost

$ for develop.but can save $in production

May increasecost




.50

.30

.30

.30 •

.30

.30

.30

Cons_quenceoffallure

.50

.30

.30

.I0

.30

.30

.I0


Past Kevlar use on 767withdrawn due to serviceproblems - watercontamination

Fiberglass widely used forlightly loaded parts, notnew technology

Manfacturing, etc. is criticalto success

Again, real breakthroughwill be innovative manu-facturing, etc. (welding,co-curing)

Manfacturing, etc. is criticalto success

Past efforts at SPF couldnot achieve minimumthicknesses required

All have major potentialfor future fighters

Figure 3.1-9.

spb-2/94.re.Ad11

Structures & Materials (concluded)

Po


Advanced StructuralTechniques• Welded Joints• Issogrid• Column Core• Z Pinning

Active FlutterSuppression

Selection rationale

Potential majorweight & cost savings

AeroelasticTailoring

Smart Structures

NEW - ControlSurface AdvancedAero (Blown Surface,etc.)

Potential major weightand drag savings

Saves weight

• Saves weight• Improves sensor vs.

tiny radome, etc.

Improvesmaneuverability

Benefit


Penalty

Save weight& cost inproduction,20% of structure

Save 10% ofstructure

10% wt reductionon aircraftstructure

10% of structureweight ifcleverly done

Developmenttakes time & $

High risk

Requires $& schedule time

Could addweight ifpoorly done

CAUTION -Adding weightto surface haslarge "hidden"penalty in flutterrequired hydraulicsystem changes


.50

.70

.30

.70

Cons_quenceoffallure

.30

.90

.30

.70


More work is needed

Needs development onunmanned drone

Figure 3.1-10.

spb.2/94-re-/_ 12

¢

Avionics

O1


Benefit I Penalty

Increased weightrelative to PavePace integratedavionics


JIAWG/Pave PillarClass IntegratedAvionics

Advanced TargetingFLIR, IntegratedNav FLIR/IRST/MLD

Tiled Array Radar

Off Board DataManagement

Common RFModules

Selection rationale

Off the shelfadvanced systemavionics

Multi-missionsupport

Reduced weight

Reduced weight

Reduced weight

Reduceddevelopment $

• PGM support• Night low levelflight

• Situationawareness

Potential for50% weightreduction inradar

Potential for50% weightreduction inavionics

• Reducedweight

• Lower LCC

Developmentcost

Developmentcost

Developmentcost

Developmentcost


.20

0.5

0.5

0.4

0.8


.I0

.30

.3O

.3O

.4O


Define growth path of RF& digital processingupgrades

• Combined multispectralapertures

• Staring focal plane array

• Advanced multilayerwafer IC on ceramicsubstrate

• Planar slotted radiators• MMIC• Packaging (component

& substrate integration)

• Reduced RCS comm.apertures & receiversensitivity

• Data fusion

Integrated Sensor Systems(ISS)

Figure 3.1-11.

spb.2/94-m-Ad7

VMS Technologies

roo_

Technologydescrlptlon

Photonics• Cables & Connectors

• I/O Interfaces

• Sensors

High Speed PhotonicDatabuses

High TemperatureElectronics

Smart Sensors/Smart Actuators

Improved Processing• Fault Tolerant

Processors• 32 Bit Processors

Modular Rack MountedElectronics

Rapid PrototypingHardware & Software

Selection rationale

Reduced system weight



• Reduced weight• Increased BW



• Increased system

• performance & reliabilityReduced maintenance

Reduced maintenance(LCC)

Reduced developmentcycle time

--,--- Configuration Impact

Benefit I Penalty

50% weightreduction

IncreasedBW

10% weightreduction

50% weightreduction

50% weightreduction

25% weightreduction

Reduced LCC(20%)

Reduced LCC(15%)

Reduceddevelopmentcost (35%)

Increasedinterface wt.

Increasedcomplexity

Increasedcomplexity

Increasedcomplexity

Increasedcoolingsys.weight

Increasedcomplexity

Increasedcomplexity


.20

.50

.50

.30

.20

.10

.10

.10

.20


.60

.60

.60

.60

.40

.40

.30

.20

.30


• Low loss connectors• Life testing• Field repair

• High temperature• High power sources• High sensitivity receivers

• Low loss sensors• Life testing

• Flight critical applications• Redundant bussynchronization

• High density/temperatureelectronics packaging

• Life testing

• Advanced actuatorpackaging

• Redundancy analysis

Redundancyconsiderations

Advanced packaging

Development tools

Figure 3.1-12.

spb-2J94-re-Ad 13

ro..,j


Reusable Software

Integrated Tool EnvironmentReliability & Performance

• Requirements & Specs

Subsystem Utilities Integration

Technology (SUIT) •Integrated Closed ECS

• Integrated Power Unit• Thermal & Energy

Management Module

Improved Hydraulic SystemConcepts

• Variable Pressure HydraulicSystems

• Variable Area Actuators• Power/Control by Light

More Electric Airplane Concepts• Electromechanical Actuators• Electrohydrostatic Actuators• Integrated Actuator Packages

Integrated Flight & PropulsionControl

• Surface Reconfiguration• Thrust Vectoring• STOVL• Optical Air Data• Flush Port Air Data

VMS Technologies (concluded

--,-- Configuration Impact _ Proba- Conse-blllty quence

Benefit I Penalty of failure of failure

Reduced Reduced .30 .30development developmentcost cost (25%)

Reduced Reduced .50 ,30development developmer_tcycle time & cost cost (25%)

50% weightreduction .30 .50

maintenancecost

Selection rationale

• Reduced weight• Increased energy

utilization• Reduced

Increasedvehicleperformance

Reducedmaintenancecost

Improvedperformance

5% increasedperformance

10% reducedmaintenancecost

10% increasedperformance

Increasedcomplexity

Increasedcomplexity

Increasedcomplexity

.60 .60

.50

.02

.50

.40


Modular softwaredevelopment tools

Abstract representationof system functionalityand requirements

Physical & functionalintegrationSuitability of differentfluidsEnergy utilizationAdvanced packaging

Energy optimizationHi._h powered, highrehability opticalsources

Reliability & lift testing

• Flight control surfaceredundancy

• Vehicle performance• Advanced control

laws

Figure 3.1-13.

- spb-2/g4-re-Ad 14

Crew Systems

POCO


Helmet-MountedDisplay -Monochrome

Helmet-Mounted

Selection rationale

• Reduce control/display

•suite wt. (replaces HUD)Increase situationawareness

• Reduce workload

Display -Color

Laser-HardeningTechnologies

Night VisionSystems

PanoramicDisplay

3-D Audio

Flat PanelDisplay

; Technology

• Reduce control/display

suite wt. (replaces HUD)Increase situationawareness

• Reduce workload

• Increased pilot

.survivabilityMission effectiveness

• Improve low light

• operations perf.Mission effectiveness

• Reduced C&D suite

weightReduced no. of units

Increased situation/spatial awareness

• Reduced display weight• Power & cooling needs


Benefit Penalty

• 10-15% reductionin C&D suite weight

• Reduced restrictionon fore canopyshape

• 10-15% reductionin C&D suite weight

• Reduced restrictionon fore canopyshape

• Increased pilot

•survivabilityMission effectiveness

• Improve low light

• operations perf.Mission effectiveness

>25% reduction indisplay weight

Increased situation/spatial awareness

°>25% reduction in

display wei_thtPower, cooling needs

• Less behind-paneldepth required

Canopy mayneed to besli_lhtly widerornigner(-1o%A)

Canopy mayneed to be

sli_lhtly widerornigner(~10% A)

~10% increasein canopyweight orcockpitsystems weight

<5% weightincrease

Front panelshape will bemorerectangular

Minor weightincrease

_-- Proba-blUtyof failure

.45

.75

.70

.30

.90

.50

.30


.3O

.70

.60

.30

.50

.75

.25


• Optical design/fov/weightreduction

• Position trackingaccuracy & throughput

•S_mbology• Pilot performance A

• Color mini-CRT•Above topics-Color-coding

• Multiple wavelengthsensitivity

• Response time to first

•pulseAircraft vs. pilot-mounted

•Compatible cockpitlightin_

oSys. s=ze & wt. reduction

• Large color flat paneldevelopment

oSymbology design

• Determine task perf.improvement

• Position tracking system

improvementPCB enhancement

• Increase display perf.(brightness, resolution,color)

• Manufacturing methods

Based on F-16 baseline and IOC of 2005 Figure 3.1-14. " ,pb.2,_.,o_15

Weapons

PocO


Internal WeaponsCarriage

External/PylonMounted Carriage

Conformal Carriage

Gravity Weapons

Laser GuidedWeapons

AutonomousGuidance Weapons

Selectlon ratlonale

Reduce signatureand drag

• Reduce aircraftweight

• Simpler loading

Reduce aircraftweight and size

Cheap & availablein large quantity

Requirement forprecision delivery

• Standoff requirement• Eliminate man-in-the-loop


Benefit I Penalty

!. Signaturereduction of30-40%

,. Drag reductionof 10-20%

• Smaller aircraft • Drag increase• Lighter weight of 10-30%

• Not LO highsignature

• Smaller aircraft• Lighter weight• Reduced

signature fromexternalcarriage

Asset or liabilitydepending oncarriage modeselected

Asset or liabilitydepending oncarriagemode selected

Improvedsurvivability

Weight increaseof 5-15%

• Higher drag thaninternal carriage

• Lower signaturethan externalcarriage

Asset or liability

Asset or liability

More complexweapons &avionicsintegration


.30

.20

.5O

.10

.10

.30

Conse-quenceoffallure;

.30

.I0

.30

.10

.10

.30


• Weapons separation• Aeroacoustics

• Suspension & releaseequipment

• Weapons separation• Suspension & release

equipment

• Conformal weapons• Conformal suspension

& release• Aircraft design

• Weapons separation• Suspension & release

equipment

• Avonics integration• Suspension & releaseequipment

Fiber opticsi OperationsSensor fusion

• Stores managementsystem

• Pave Pillar architecture

Figure 3.1-15.

spb-2/94-re-Ad8

,Weapons (concluded)

COO


"All Envelope"Air-to-Air Weapons

Selection rationale

Air-to-air and self-defense requirement


Benefit I Penalty

• Survivability• Offensivecapability

• Doors, launchers,pylons, etc.

Ballistic Weapons"Guns"

YePervelocityapons

HARM or OtherSEAD Weapons

Cruise Missileor UAV Carriage

• Close-in millrequirement

• Simplicity

Hardened targetmill standoffrequirement

Self-defensecapability

Standoff or RECCErequirements

None

Can besubstantialdepending oncarriage mode

None

None

• Weight• Avionicsintegration

• Weight• LO integration• Space foreffective guns

• Weight• Rocket motorblast

Substantialimpact to config.is carried ininternal baysor conformally

• Substantial ifcarried internallyboth in weight,bay volume, andaircraft size

• Conformalcarriage maynot be possiblefor UAV dueto size


.50

.20

.20

.30

.30

iConse-!quenceof failure

.40

.10

.20

.30

.30


• Sensor fusion• Helmet mounted sight• Weapons separation• Advanced suspension& release equipment

• Improved guns• Body/wing integration

• Weapons geometry• Weapons separation• Suspension & releaseequipment

• Compact/conformalweapons

• Suspension & releaseequipment

• Suspension & releaseeq.ulpment

Wing designFuselage design &weapons integration

Figure 3.1-16.

spb-2/g4.re-Ad9

1.0

0.9

0.8

0.7

ALI.Oa. 0.6

"" 05o

.Q 0.42

n

0.3

0.2

0.1

0

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///////_

DEMNAL limit//11/I//1/111/11/I/1111/ "1111111/f

¢FSD limitY//////_ "///////, "/, !

,'_-- W

l °

0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

Consequence of failure (COF)

Figure 3.1-17. Sample Technology Risk Assessment Cnteria

0.9 1.0

spi_7-7-re-Ad2

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3.2 Technologies used in Agility Study Configurations

The technology elements selected to be used for each point design configuration are shown onTables 3.2-1 and 3.2- 5. The majority of the chosen technologies are common to all configurationswith the only exceptions being in the "Control Effectors" selections. Also, most of the applied 'technologies risk levels are within the pre-established DEM/VAL limits. The exceptions being the all-moving canard, power/control-by-light and IHPTET Gen 6 engine technology.

Principal impact on configuration development resulted from incorporation of projected technologybenefits in five major functional areas.

• Main Engines Use of IHPTET "Gen 6" engines resulted in significant weight andsize reductions in the overall propulsion system (inlet, diffuser,engine bay and exhaust duct). Engine mass location within theairplane was less of a driving issue to achieve air vehicle balance.

• Avionics Principal benefits to airplane configuration resulted from reductionsin weight and volume for both the modules or units and theinterconnection system. Cascading benefit to the environmentalcontrol system for reduced cooling loads results in further volumereduction.

• Subsystems Expanded technology development in flight controls actuation,secondary power generation and control, ECS, andmanagement/integration of functional components are consideredas contributions to obtaining sufficient or expanded capability withinavailable or reduced airframe envelopes. The resultant anticipated isimproved installation density or volume utilization.

• Structural Materials Application of next generation composites, such as Titanium MatrixComposites (TMC), permits the implementation of unique designfeatures not feasible with conventional materials because offabrication complexity, environment limits, or weight impact onvehicle performance.

3.3 Weight and Cost Impact of Advanced Technologies

The Boeing developed parametric/statical Level 1 weight prediction methods used to estimate thegroup weights of the ADS "point design" configurations contain weight considerations for some ofthe technology items selected for incorporation into the designs. These items are not considered"advanced technology" and include items such as conventional ailerons, leading edge flaps or slats,all-moving horizontal tail, supercritical wing, electromechanical actuators, etc. Weight increments forincorporation of these devices are not specifically called out as special features. Tables 3.3-1and 3.3-2 show the advanced technology application weight effects. These features required specialconsideration, outside the standard method, when estimating their weights.

Projected weights for IOC 2005 avionics suites for the air-to-air and multi-role missions are shown onfigures 3.3-3 and 3.3-4. The air-to-ground avionicssuite was considered to be identical to the multi-role. Advanced technology assumptions used to generate these weights are presented on thetables. F-22 avionics weights were used as the base points and the advanced technology weighteffects were applied on a system-by-system basis.

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Figure 3.2-3

Page 1

TechnologyItem

Note:Shadedareadesignates"used on" technology

JIAWG/Pave Pillar Class Integrated Avionics

Advanced Targeting FLIR

Integrated Nay FLIR/IRST/MLD

Tiled Array RadarOff Board Data ManagementCommon RF Modules

Avionics and Vehicle Management Systems

:f

::)

z

E1

E2

E3

E4

E5

E6

Air Force Joint Services

A/A M/R A/G A/A M/R A/G

988 988 988 988 988 988 988 988 988 988 988 988

-114 -115 -118 -119 -122 -123 -116 -117 -120 -121 °124 -125

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Reusable Software

Subsystem Utilities Integration Tech. (SUIT)

Integrated Closed ECS

Integrated Power Unit

Thermal & Energy Management Module

Variable Pressure Hydraulic SystemsVarlable Area Actuators

Power/Control by LightElectromechanlcal Actuators

Electrohydrostatic Actuators

Integrated Actuator Packages

Surface Reconfiguration

Optical Air DataFlush Port Air Data

Photonlcs - Cables & Connectors

Photonics. I/O Interfaces

Photonics - Sensors

High Speed Photonic Databuses

High Temperature ElectronicsSmart Sensors/Smart Actuators

Fault Tolerant Processors

32 Bit Processors

Modular Rack Mounted Electronics

Rapid Prototyping Hardware and SoftwareIntegrated Tool Environment

High Pressure Hydraulics

V1

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Figure 3.2-4

Page 1

Crew Systems and Weapons

Po

Technology Item

Note: Shaded area designates "used on" technology

Helmet-Mounted Display - Monochrome

Helmet-Mounted Display - Color

Laser-Hardenlng Technologies

Night Vision Systems

Panoramic Display3-D Audio

Flat Panel Display Technology

Internal Weapons Carriage

External/Pylon Mounted Carriage

Conformal Carriage

Gravity WeaponsAutonomous Guidance Weapons

"All Envelope" Air-to-Air Weapons

Ballistic Weapons "Guns"Hyperveloclty Weapons

HARM or Other SEAD Weapons

Cruise Missile or UAV Carriage

Laser Guided Weapons

n-

m

:S

z

P1

P2

P3

P4

P5

P6

P7

Wl

W2

W3

W4

W5

W6

W7

W8

W9

Wl0

Wll

Air Force Joint Services

A/A M/R A/G A/A M/R A/G

988 988 988 988 988 988 988 988 988 988 988 988

-114 -115 -118 -119 -122 -123 -116 -117 -120 -121 -124 -125

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Page 1

03

Technology Description

Yaw Vanes -Advanced Composite

Spilt Control Surfaces

PitchAxis Thrust Vectoring -Aft Body Flaps in Exhaust

Internal Weapons Carriage

2005 IOC Integrated Avionics

JIAWG Integrated AvionicsAdvanced Targeting FLIR• Integrated NavigationFLIR/IRST/MLD

• Tiled Array Radar• Off-Board Data Management• Modular Rack Mounted• Flush Air Data Port• Reusable Software• Helmet-Mounted Display-Monochrome

Night Vision SystemsFlat Panel Displays

Integrated Actuator Packages

SelectionRationale

Extendable lowriskyawcontrol surfaces

Increasedyaw control

• High AOA pitchcontrol

• Increasedmaneuverability

Signature anddrag reduction

Reduced weight

GroupApplication

Yaw vanes

Controlsurface

Body

Body

Avionics

---,,--- Weight Impact ----,-.-

Weight Effects

6.32 Ibs/sqft of surfacearea (includingcontrols)

31% weight penalty

10 Ibs/sq ft of flap area(includingcontrols)

18 to 23% body weightpenalty depending oncutout size

1,000 to 1,200 Ibssavingsover present dayintegrated avionicsinstallations

EMD

$12,700/ft 2

(+) $10.5M

$8,045/ft 2

(+) $80-$104M

(-) $65M

Cost Impact

Average UnitProduction

$218/ft 2

(+) $0.18M

$138/ft 2

(+)$1.0-I .4M

(-) $1.7M

B=,

250NG BuyProduction

(+) $45M

(+)$250-350M

(-) $425M

Note: see tables forMission Avionicsweightsbuildups

Reducedmaintenancecost

Weaponmulti-modelaunchers

50 Ib penalty to eachlauncher, reduced functionsforthe main aircrafthydraulicsystem savesweight depending on thenumber of weapons carried

(+) $1.8M/launcher

(+) 0.05M/launcher

Figure 3.3-1. Advanced Technology Applications - Weight and Cost Effects

sp_8-7-re-Ad I

Cost Impact

Combined effects of:• Thermoplastic Materials• Thermoset Materials

Graphite Based CompositesFiberglass Based Comp.• Advanced Manufacturing

- Titanium Welding

.,,-.-- WeightImpact

Selection Group Average Unit 250 NG BuyTechnology Description Rationale Application Weight Effects Production Production

(-) $1.31 M (-) $327M

- Z Pinning• Advanced StructuralTechniques

- Welded Joints- Z Pinning

Combined effects of:• Titanium Matrix Composite• Powder Metallurgy

Superplastic FormingAdvanced Carbon-CarbonComposites

IHPTET Gen 6 AdvancedEngines Technologies(including FADEC/PSC)

High Pressure Hydraulics

Power and Control-by-LightFlight Controls

Weight savings: Cost savings• Improved

° toughnessPotential formanufacturingbreakrhroughs

: Weight savingsi Use at exhausttemperatures

• High strength

• Higher performanceLighter weightReduced SFC

Lighter weight

Cable/wire weightsavings

Low riskapproachto yaw controlpower

• Wing structuralbox

• Wing controlsurfaces

• Wing secondarystructure

• Horizontal and"vertical tails

• Body structure• Air inlet

Exhaust nozzles

Engine

Hydraulic system

Surface controls

Exhaust system

-17%

-20%

-22%

-25%

-12%-15%

Note: weightsavings are relativeto an all metal a/cNote: assumesapproximately 55%of the airplanestructure weight isadvanced GR/EPmaterials

35%

50 to 60% T/Wincrease overexistingdry gasturbines

-12%

-22%

EMD

(-) $92.3M

(-) $27.8M

(+) $1.2B

(-) $3.7M

(-) $17.4M

(+) $33.3-41.3MYaw Axis Vectored Thrust+ 45 Degrees

(-) $0.9M

Use CER* 1.0816

r

(-)$0.05M

(-) $0.3M

42 to 52% increaseover a nonvectoringdry or NB nozzle

(+) $1.1-1.4M

(-) $225M

(-) $12.5M

(-) $75M

(+) $275-350M

Figure 3.3-2. Advanced Technology Applications - Weight and Cost Effects

spb.8-7-re-Ad2


Austere Avionics Suite for anA r-to-Air Agile Fighter

4_cn

Weic

Subsystem Uninstld Instln

CNI

EW

SMS

rll

I Sub Total

VMS

Misc.

318 81

Total

Total Capabilities Comments

..- ..........-.......:

::::::;;::::

iiiiiiii:82i!

399

105 58

54 15 69

1332 469 _/ 34.3

UHF (Have Quick), VHF, IFF Int/Trans.,

Band 2 DF, ESM, JTIDS, Landing Aids,GPS, IRS

RWR (4_), Forward PDF, ESM,Countermeasures

Monitoring/Control AA & AG Weapons,Gun, CM, Doors, Spoilers & Launchers

Utility Mngmt Comp., Flight ControlSensors. Air Data

Stick, Throttle, Pedals & Misc. Instrument_

21.9

F-22 TechnologyAdditional functions to consider:

;ATCOM, IFDL, TACTS

MLD/Laser Warn. Provided by EO

tiiiii!ii!iiiiii!iii!i.iii!!i_IAdvanced TechnologyFigure 3.3-3.

BoeingDefense &

Space Group

Austere Avionics Suite for aMulti-Role Agile Fighter

O3

Wei(

Subsystem Uninstld Instln Total

-...-.-.........

CNI 327 84 411

114 360EW 246

i_iiiiiiiiiii!SMS 91

JSub Total 1287 _@_

VMS 105 58 163

Misc. 54 15 69

Total 1446 486 _

148 239

........ Technology,_:_:_:_:_:_:_:z:_Advanced-=.:!:i:?:i:!:i:!:i:!:!:i:i:?:i:!:i:!:i:!:i:::::::::::::::::::::::::::::::::::::::::::

30.8 24.7

Capabilities

UHF (Have Quick), VHF (SINCGARS,ATHS), JTIDS, IFF Int/Trans., Band 2DF, ESM, RAIt., Landing Aids, GPS, IRS

RWR (4_), Forward PDF, ESM,Countermeasures

Monitoring/Control AA & AG Weapons,Gun, CM, Doors, Spoilers & Launchers

Utility Mngmt Comp., Flight ControlSensors t Air Data

SUck, Throttle, Pedals & Misc. Instruments

Figure 3.3-4.

Comments

F-22 TechnologyAdditional functions to consider:

;ATCOM, IFDL, TACTS

MLD/Laser Warn. Provided by EO

4,0 Configuration Development

The process used to develop the concepts is presented in figure 4.1. The initial configuration matrixconfigurations and desirable features were developed in round table discussions by the DesignTeam. The selected assumptions, ground rules, number of engines, crew size and observablesguidelines are all a product of team decision-making. In parallel to the Design Team, a technology riskassessment was undertaken by the Boeing Military Airplanes (BMA) Technology Staff. The results ofthis technology risk assessment guided the subsystems and technologies selected for incorporationinto the design concepts.

4.1 Assumptions and Ground Rules

Single Crew

A single crewman concept was selected as the basiss of all the configurations in this study.Improvements in avionics and crew systems technologies will allow a single pilot to manage theworkload now being accomplished by a pilot and a weapons officer. Reducing the number ofpersonnel to enemy fire and reduced overall operating costs are added benefits of a single man crewover a two man crew concept.

A single pilot/crew station is incorporated in each air vehicle concept. Mission and flightsubsystemspostulated for usage in these vehicles will permit operation and control throughout all flight phases byone person.

Benefits accrue, from the single person crew, in reduced airframe and subsystems volume, weightand cost while satisfying misison performance requirements.

Survivability in threat environments or intense workload mission segments (terrain following, targetarea, and air combat), where extra eyes have proven valuable, will now require systems technologytoprovide situation awareness, threat positiondata, and target acquisition/trackingfor single personoperation at flightcriticalreliability levels.

Twin Engine

The use of twin enginesfor all the concepts was a ground rule established early as a result of anumber of observations. The Navy has a strongbias for twin engine designs because of the fail safeengine lossover water issues. All of the aircraft these designs are to replace; the A-6 and F-15/F-14aircraft have twin engines. Early sizing studies indicated that the aircraft would be very large and wouldrequire two engines to keep the engines within the airflow ranges seen for these classes of aircraft.Selection of a common engine arrangement for all concepts would eliminate the confusion of dealingwith a mixture of single and twin engine designs in comparison of other design issues.

Airframe integration for Joint Service usage is achieved more efficientlyin a twin engine configurationby use of a centerline structural keel to directly carry both launch and arrested landing loads.

Survivability and general safety of flightdata show an advantage for the redundancy in both primaryand secondary poer sources integrated in a twin engine configuration.

F-22 Core Avionics Suite

The NASA provided technology list had a large number of technologies already utilized in the F-22avionics suite. Any differences in avionics suite requirements to handle different mission roles will behandled as additions or deletions to the baseline hardware or software of the existing F-22 avionicssuite. Improvements to the avionics systems have also been considered.

47

configuration Aircraftsizing Final configurationsconceptualization and performance

O: -9

Initialconfiguration matrix

--_ _ - Assumptions and ground rules

.._ ! Number of engines

= Number of crew_Technology assessmentObservables guidelinesO • Subsystem selections

GW = 67,350 Ib ,_\

T/W = 0.7 /-[_!i\ _w/s 45psf,--I =AR = 3.0 --

GW = 67,350 Ib_ _ T/W = 0.7

_: W/S 45psfAR = 3.0

T/W = 0.9 _%_o E W/S 60psf ._,,..- %AR =3.0 :_

GW = 57,750 Ib 4'_,%

_ T/W = 0.9 _h= W/S = 55 psf _ :,_:_,AR= 2.7 v-._/_.

GW = 47,700 Ib ]_/,',!;o T/W = 1.1 _._,,1._W/S = 55 psfAR = 3.0

GW = 47,700 Ib j__'._T/W = 1.1W/S -- 70 psfAR = 3.5

• Control effectorselection/sizing

• Conceptual layout/size

_ "'"

i II

II

_,. I

;i

II I

• Planform selection

• Engine sizing

. , , , , , ,

_,6,,,qCq,

I

!iI

i

• . . .,...# = ° °

u_ i

I'_i.

. . = = =,.=,. ....

• Weight and balancecheck

• Fuel volume check

988-122GW = 73,145 Ib _"..T/W = 0.34 _'_-W/S = 50 psf ::_AR = 3.5

988-123GW = 73,145 Ib _"

W/S = 50 psf ' "-AR = 3.5

GW = 50,899 IbT/W = 1.1W/S = 46 psfAR = 3.52

988-119 i,,'._.GW = 48,801 Ib _,.,,,T/W = 1.1W/S = 59 psfAR = 3.97

,8.,,T/W = 1.13W/S = 46 psfAR = 3.64

988-115 j_GW = 59,549 IbT/W = 1.13W/S = 58 psfAR = 3.8

Figure 4.1 Configuration Evolution

$pb-7-7-re_Ad7

Observable Features

Moderate levels of observability,as a general classificationregarding both RF and IR signaturecharacteristics, is taken to describe vehicles as similarto the YF-22/YF-23 airplanes, or better, incertain frequency bands.

Lowobservable levels for both RF and IR signature characteristics are considered to place a vehicle inthe region approaching B-2 levels.

In order to achieve the general levels directlythree primary configuration items have been establishedfor integration within each air vehicle type.

Internal Weapons Carriage - mission loads are carded within the vehicle basic moldlineindedicated weapons bays. Stores are either ejection released or rail launched from thesebays. No conformal or external carriage is considered for the primary/sizing missionspecified.

Tail Surfaces - directional control traditionallyobtained by use of either vertical or cantedfin/rudder, or all moving surfaces, have been eliminated from consideration because oftheir inherent penalty to signature reduction. Additionally, in the high alpha combat flightregimes, directional control effectiveness becomes degraded rapidly.

Inthis study each air vehicle type incorporates a thrust vectoring rotating nozzle toprovide yaw control power by direct control of engine exhaust. Supplementary directionalcontrol is obtained by use of Yaw Vane panel pairs integrated into the forward bodysurfaces fairing into each nozzle.

Additionally, Yaw Vane pairs are provided on the lower aft vehicle surface for use duringthose inflight phases requiring increased directional control or side force momentgeneration.

The combination of a thrust vectoring rotating nozzle with co-located Yaw Vane panelsresults in a unique method of generating sufficient directional control power throughoutthe flight envelopes and maneuver range of these vehicles at greatly reduced signaturelevels.

Vehicle Shaping/Arrangement - Moderate observables levels are to be obtained bydeveloping local body maximum half breadth slopes at or near to forty (40) degreesrelataive to the horizontal reference plane. Wing body integration will be blended to avoidcorner reflector conditions. Where wing and tail, or wing and canard combinations areemployed for agility the approach taken will be to minimize platform edge mis-alignment orbreaks and dissimilar sweep angles. Where these conditons exist, observvability levelswill degrade as a direct result of obtaining the required agility metric.

The approach to obtaining low observables in a configuration type will employ alignededges with minimum breaks or dissimilar angles. However, in each air vehicle type, agilityperformance metrics will be the dominate consideration.

In the case of Air Interdiction type, where the prescribed mission requires a longdistancepenetration segment, the configuration will be based on an all flying wing design conceptemploying long straight edges to the maximum extent possible with the objective ofachieving lower observability at the lower frequency threat levels.

49

ControlEffectors Selection

Selection of control effector devices for each air vehicle type was based on the following listing.These devices are combined/integrated with a particularconfiguration concept to generate therequired control forces. Most of these devices are well known and used widely in actual application;

Yaw axis thrust vectoring is included here as a primary control effectorwhich operates synergeticallywith the Yaw Vane panels to produce directional/side-force moments, or alone as speed brakes.

• Yaw Axx ThrustVectoring

• Yaw Vanes

AoolicationlUsage

•Directionalcontrolwith._+45degreesdeflection

range

•Sideforcemoment generator

• Pop-up surfaces integratedwith Yaw axisrotatingnozzle

•ProvidesupplementaryYaw axiscontrolpower,sideforcemoments, oractas speed brakeswhendeployedas fullpairs

• Canard-Lifting • All-moving surface deflected symmetrically forpitch and asymmetrically for roll cotnroll

• Horizontal Tail • All-moving surface deflected symmetrically forpitch and asymmetrically for roll control

• Elevons • Single panel used for lateral/itch control

• Split panels used for lateral/pitch andasymmetrically for side force or Yaw momentgeneration.

• Leading Edge Slats • Increased lift for maneuver conditions

• Trailing Edge Flaps •Increased liftfor maneuver/field performance

50

4.2 Carder Suitability Impact of AircraftDesigns

Carder suitabilityis clearly the overdding requirement of any aircraft design operating from an aircraftcarrier. Operations from Navy Aircraftcarders at sea impose a broad range of geometry constraints,and performance requirements on aircraft designs. The issues of carder suitability involve all designdisciplinesincluding support functions such as ILS, maintainability, and supportability. Cardersuitabilityhas many interwoven effects such as launch/recovery/basing geometry constraints,maintainability access, weapons loading, and landing gear geometry for efficient structureand gooddeck handling. Control effector sizing designed to trim the high liftsystem while maintaining adequatedynamic margins is also an important design issue.

Geometric Umitations

The catapult launch imposes hard limitson the overall length of the aircraft and the minimum heightabove the ground for the fuselage and any of its externally carded stores such as centerline tanks andweapons.

The tight quarters of the flight and hanger decks, the large number of operating aircraft, personnel,and support equipment contribute to a maze of Navy unique design requirements.

The elevator clearances require that hinges for folding wing aircraft be employed with power actuation.

The hanger deck imposes a height limit to the vertical tail and wings in the folded position to17 feet.

Weight Limitations

The aircraft takeoff weight, fully loaded, is limited the 90,000 Ib capability of the C-13-1 catapult.However, to efficiently conduct flight operations, the elevators must support two mission readyaircraft, one tractor and the associated personnel. The fueled aircraft without stores must thereforenot exceed 54,500 pounds, using the new TA-12 tractor.

The landingweight, with reserve fuel and retained weapons, is limitedto the 65000 Ib limit of the Mk7-MOD3 arresting gear.

Landing Gear Design

The landing gear strength and stroke length are driven by the impact loads of arrested landings. Theweight penalty applied to the main gear to adjust the Air Force version of the configuration to the jointservice configurations amounts to 37.8% to the main gear weight.

A stored energy nose gear is assumed during this study. The stored energy nose gear uses thevertical reaction of the nose gear with the deck during the deck run on the catapult power stroke toimpart both an optimum pitch rate and attitude to minimize launch flyaway airspeed. The nose gearmust be fully casterable for roll back after arrestment. The dual tire nose gear must also have built-intow and holdback fittings for catapulting. The resulting weight penalty used to adjust from Air Forcelanding loads to joint service landing loads results in an increase in nose gear structural weight of 63percent over its Air Force counterpart. See figure 4.2.

Wing and Fuselage Structural Re-lnforcement

Structural adjustments to the wing structure to accommodate landing gear punch loads and foldingmechanism adds 17.5 percent to the wing structural weight. The fuselage structure is increased 5percent to handle the loads of the tail hook and nose gear during landing.

Engine Installation

Engine air intakes must be placed to avoid ingestionof steam on the catapult stroke.

51

L

0

0

0

ro

¢o

o

_S

Landing Gear Weight (Klb)(Including Arresting Gear

e=o

I

Z

I-"(I)

O

"1"iOm

¢2.CD

"O

¢Dm

==dl.

Wing Fold Mechanism andStructural Penalty (KIb)

!

.............................................. : ...................... i

0_

.............................................. _.................... ..................................... _..o.= ........................... ; ...................... <

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¢D e,n¢_

8,=e-

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Engines must be located to allow complete removal and replacement without cranes, while wings arefolded, in the hanger deck area.

Weapons Loading

Weapons must be loaded while wings are folded, withoutthe use of cranes or ladders.Suspended weapons must be high enough above the ground to clear the catapult shuttle and toavoid deck impact in a wing-low arrested landing.

High Uft Devices

The wing designs used have historicallyused highlift flap systems and other devices to allow safe lowspeed flight after catapult release, fully loaded.

The wing must also provide low lift for safe go around without touchdown on an aborted landing.

Support Equipment

Steps or ladders for entry of the crew must be built-into minimize deck clutter safety hazards on theflight deck and the hanger deck.

Fueling and routine servicing or rearming must not require platformsor external hoists,only dollies.

Environment

High sea-states and low-visibility/nightoperations demand an aircraft with superior stability and controlcharacteristics to accomplish the required high recision flight path control necessary to routinelyaccomplish recovery safely.

Landing Recovery

Carrier approach speed, approach angle-of-attack, stall margin, vision angle, pop-up maneuver,longitudinalacceleration, thrust response, single engine rate-of-climb analysis are all inherent analysiscapabilitywithin the Fighter Aircraft Sizing Tool (FAST) aircraft sizing and performance code. Thecarder suitabilityanalysis modules in FAST parallel the conceptual level methodology of the NAVAIRCAT and APR codes. In addition, FAST is capable of determining a rough order estimate of carrierspotting factor.

The main driver in carrier recovery is the requirement for significantlylower airspeeds during approachand arrestment. This drives the designer to maximize the use of high lift recovery devices. Use ofsuch devices frequently conflict with the need to use thinner, cleaner airfoils optimized for high-speedup and away flight. Safe recovery of Navy aircraft force the design to emphasize low speed stabilityand Control regions driving the size of the horizontal surface up. Naval aircraft become a balancebetween the uncompromising need for safe flying qualities at the low speed end of the flightenvelope while minimize_ng maneuvering and performance penalties at the high speed end.

Catapult Launch

Catapult launch analysis determines the minimum safe launch airspeeds while maintaining acceptableflight characteristics in this low altitude, high angle-of-attack regime. Approach and landing requiresthe slowest possible approach airspeeds while retaining the performance and handling qualities needfor precision glide slope control. Keeping approach airspeeds low results in reduced ship's operatingspeed and thus enhances the operational flexibility of the aircraft carder.

Catapult launch presents the danger of operating too close to the aircraft minimum control airspeed.Since catapult end-speed is constrained by catapult performance, the requirement for a 10% stallmargin at the end of the deck-run and an angle-of-attack margin 20% below stall drives the designer tomaximize Clmax in the takeoff configuration. The requirement of a 500 foot/minute minimum rate of

53

climb inthe event of an engine failure and a fly-away longitudinalacceleration greater than 0.065 gsimply the need to maximize L/D beyond that required for an equivalent Air Force Aircraft.

The naval aviator has to be able to see the carder during approach at relatively high angle-of-attack.The location of the pilots eye and shape of the aircraft nose must accommodate this approach angle-of-attack. A 3.5 degree glideslope mandates an 18. degree over-the-nose vision angle for carrierapproach.

Wave-off and Bolter

Wave-off and bolter present further constraints on the propulsion and drag-brake systems, which inturn directly affect stabilityand control through rapidly occurring, transient changes accompanyingtypically large thrust commands. The major challenge is obtaining quick engine response, coupledwith an adequate amount of pitch control.

Combat Maneuvering

Up-and-away maneuvering requirements have traditionally been more stringent for the Navy becauseof its insistence of utilizing as much of the flight envelope originally designed into the aircraft. TheNavy expects their pilots to fly to the edge of the envelope and consequently drives the designer toprovide Level 1 flying qualities to the maximum limits of the operational envelope. This has a numberof implications to departure resistance, angle-of-attack limiters, and maneuver devices.

The Navy requires high departure resistance at highangle-of-attack sufficient to prevent Ioss-o controlwhile maneuvering close to and possibly through the flight envelope where aerodynamic controltraditionallybeginsto diminish. The Air Force will typically accept limitersto avoid approaching CLmaxboundaries throughout the maneuvering envelope. The Air Force F-16 employes an angle-of-attacklimitingschedule which shrinks the left boundary of the energy maneuverability envelope significantlybeyond comer speed. Unique maneuver devices normally found on naval aircraft to ensure maximummaneuvering performance over a full flight envelope. These devices usuallytake advantage of analready unique low speed, high lift system such as the maneuver flap or slat.

54

4.3 Designing for Agility

This section discusses studies conducted to relate the agility metrics to design considerations (figure4.3-1). Before the design studies could be carded out, a framework of design guidelines wasestablished. Aerodynamic characteristic needs were derived from the metrics and the designguidelines. Techniques were formulated to bridge the gap between metrics/guidelines and effect'orsizing. Finally, this approach was used to size effectors on three different airplane configurations: onemedium agility and ten high-agility concepts.

Agility metrics are defined in figure 4.3-1 below. The agility performance of conceptual configurationsis discussed in terms of aerodynamic forces and moments required to meet these performance goals.

4.3.1 Agility Metrics

Maximum Negative Specific Excess Power

Maximum negative specific excess power (Ps) is a metric that was created to describe the energy lossof an aircraft while executing an unsteady turn. This metric attempts to quantify an aircraft's potentialfor losing energy by measuring the minimum (or maximum negative) PS (rate of change in specificenergy) achieved during a maneuver. Maximum negative specific excess power corresponds to anaircraft's maximum instantaneous turn rate capability

Energy exchange during combat is a combination of speed loss (kinetic energy) and/or altitude loss(potential energy) and depends on the controls applied by the pilot or flight control system and theaircraft's aerodynamic characteristics. The classical approach to combat management is to minimizeenergy loss during combat.

Maneuver employed to attain the maximum instantaneous turn rate consists of using the elevator toincrease the aircraft angle-of-attack and, in some cases, the application of aileron, rudder, speedbrakes, and maneuyer flaps. Although a reduction of thrust would result in a reduction of the net axialforce on the aircraft (and thus a reduction of specific excess power) this technique is not normallyused. Engine response time is of the same order of magnitude as the time needed to achieve thedesired conditions. Furthermore, the capability to gain speed following the turn would be seriouslycompromised.

Computation of the maximum negative specific excess power is identical to specific excess powerperformance. This is addressed in section 5.0 along with the maneuver performance requirements..

Time-to-Bank and Capture 90-Degrees

In air combat, the offensive pilot attempts to achieve target acquisition. To achieve his objective ofdestroying the enemy, the pilot must successfully deploy his weapon, which requires aiming orlocking-on. To lock-on or aim a weapon the pilot must precisely control his aircraft. During this phase,the defensive pilot tries to evade the offensive pilot's attempt by jinking of-of-plane and changing thebattle geometry. The offensive pilot has to reacquire the target and track sufficiently to deploy hisweapon. The cycle of acquire, jink, reacquire, jink, etc., is characterized by the offensive pilot'sbanking with the intent of capturing a specific bank angle as determined by the jinking maneuver ofthe defensive participant. Time-to-bank to and capture 90 -degrees was chosen as an agility metricbecause it quantifies an aircraft's ability to offensively reacquire an evading target.

Airplane roll performance is measured with respect to a single-degree-of-freedom system. While thepilot may use the rudder peddles to slip the airplane and increase roll acceleration, the designer is notpermitted to take advantage of this maneuver. Indeed, for a class IV airplane, automatic turncoordination is already required, insuring that the airplane behaves as a single-degree-of-freedomsystem in roll. Therefore, the performance of the roll control system can, to a great extent, bedescribed by two terms: maximum roll acceleration and the roll time constant. Maximum rollacceleration is proportional to the roll control moment available. Roll time constant is related to theairplane roll damping. Roll damping can be influenced by roll rate feedback if required. Much research

55

# Flight AgilityCondition axis

1 M = 0.6

Hp= 15,000 Ft pitch

qbar = 301 psf maneuveragility

2 _t Pitch

angular

Agility

3 "_ Pitch

envelope

agility4 _t Roll

Agility

5 _t Lateral

agility

_t

AIR TO AIR

Agility Metric

The airplane will have a specified value ofdeceleration at the maximum instantaneous turn

rate Deceleration is given in terms of specific

power. Load factor to be greater than 5.5 g'sMinimum nose down angular acceleration

at the design critical alpha.

(taken as the alpha 'pinch point')Maximum departure free alpha

Time to roll and capture .

Start at _ = -45 Deg. Then roll thru 90 deg and

capture _ -- 45 Deg.

(Adequate yaw control power to roll around the

velocity vector is required.)Maximum lateral acceleration with the wings level.Max maximum load factor

Evaluation is shown in the performance section

Figure 4.3-1. Agility Design Goals (sheet 1 of 2)

Medium AgilityDesign Goal

Ps = -450

fps

-.15 Rad/Sec 2

40 Deg

2.5 Sec

Ny = 0.4 g's

High Agility

Desi_In Goal

Ps = -800

fps

-.35 Rad/Sec 2

70 Deg

1.5 Sec

Ny = 1.0 g's

O1

# Flight AgilityCondition axis

V= 450 KEASM = 0.68 Pitch

Hp= SLS maneuver

qbar = 686 psf agility2 _ Pitch

angular

agility

3 _t Pitch

envelope

agility

AIR TO GROUND

Agility Metric

The airplane will have a specified value ofdeceleration at the maximum instantaneous turn

rate Deceleration is gi_/en in terms of specific

power. Load factor to be greater than 7.5 g's

Minimum nose down angular acceleration

at the design critical alpha.

(taken as the alpha 'pinch point')

Maximum departure free alpha

4 _t Roll Time to roll and capture .

agility

5 _t Lateral

agility_t

Start at _ = --45 Deg. Then roll thru 90 deg and

capture (_ -- 45 Deg.

(Adequate yaw control power to roll around the

velocity vector is required.)Maximum lateral acceleration with the wings level.

At load factor = 1.0 g's.

Evaluation is shown in the performance section

Figure 4.3-1. Agility Design Goals (sheet 2 of 2)

Medium Agility

Design Goal

Ps = -450

fps

-.15 Rad/Sec 2(not

applicable)

na

(Alphalimiter at

0.9 CL max)

1.5 Sec

Ny = 1.2 g's

High Agility

Desi_ln Goal

Ps = -800

fps

-.35 Rad/Sec 2(not

applicable)

na

(Alphalimiter at

0.9 CL max))

1.0 Sec

Ny = 2.0 g's

has been done to determine optimum values for roll acceleration requirements and time constants.Specifying a minimum rollacceleration capability and time constant, along with a control rate input,results in a unique rollangle time history. Frequently, specificationsare expressed as the timerequiredto roll througha certain rollangle. For a class IV airplane at combat flightconditions, this isusually 90 degrees in 1 second. It is the task of the preliminary design engineer to ensure enough rollcontrol to meet this specification. Adequate roll control must be designed into the airplane during'preliminary design. The designer has some control over the time constant through roll rate feedback.

Maximum Nose-Down Pitch Acceleration

Many times in air combat the roles of the offensive and defensive pilotsare reversed. When anoffensive pilot is faced with role reversal his objective changes from that of destroying the enemy tonot being destroyed. A frequently successful defensive tactic is to disengage, break off the battle,and retum to safe air space. As the defensive pilotattempts this action, the offensive pilotwillcontinue his pursuit. The success of the defensive pilot depends on his ability to transition from anengagement mode characterized by high load factors and highturn rates to an escape modecharacterized by high longitudinalaccelerations to maximize the separation distance. This maneuverrequires the pilot to unload his airplane as quickly as possible and achieve a minimum drag flightangle-of-attack. Maximum nose-down pitch acceleration was chosen as an agility metric to quantify theaircraft's transition from a highly loaded air combat flight condition to an escape or maximumlongitudinal acceleration condition.

Maximum Achievable Trimmed Angle-of-Attack

Modem air combat research has shown that high angle-of-attack or post-stall flight may provide atactical advantage on both offensive and defensive aerial engagements. In an offensive mode thepilot's abilityto turn at higher turn rates with smaller turn radii provides him with the optionto morequickly achieve shot opportunity by out-maneuvering his opponent. In a defensive mode high-angle-of-attack capability can be utilized by a pilot to bleed energy more quickly, thus forcing the offensivepilot to overshoot and providingrole reversal. In either case high-angle-of-attack capability will beutilized by a pilot ohly if the airplane remains controllable and has good handling qualities. MaximumAchievable (Departure-Free) Trimmed Angle-of-Attack was chosen as an agility metricto quanti_j anaircraft's abilityto utilize the post-stall flight regime.

Maximum Lateral Acceleration

It has been proposedthat an aircraft's abilityto laterally translate its position may be of significanttactical advantage. In a real engagement this ability may provide useful defensively as a jinkingmaneuver. However, this characteristic may be of even greater importance in a ground attack mode.Typically, highvalue ground targets are attacked in a manner requiringa single pass or flyby for eachtarget. An airplane with substantial lateral displacement capabilitymay be able to attack a target,laterallydisplace its position, acquire and attack a second target on the same pass. Maximum lateralacceleration was chosen as an agilitymetric to quantify an airplane's abilityto attack multiple groundtargets on a single pass.

Before discussing the scope analysis a few words must be said about how the agility is used and itsimportance. Tactics using flat turns were flight tested by the USAF in 1983 on the AFTI/F-16. Therecommendations from that testing (more than 15 unique flight modes were tactically tested) singledout flat turns as important for new airplanes.

The maneuver was best for a/g and not as good for a/a. It was best for strafing runs and deliveringdumb bombs. Delivery of smart bombs may not be an agility issue. The same is true for a/a. The flatturn would be best for a/a gunnery and not guided a/a missiles.

Flat turns made the airplane more lethal and at the same time more survivable. The use of flat tums iscomplex. For example, the optimum dumb bombing technique combined classical rolland pitch forgross heading changes with flat turn for small changes. The pilot used rollstick to quicklyget thepipper in the vicinityof the target. Remaining directional errors were removed with flat run rudder

58

Themaneuverwasbestfora/gandnotasgoodfora/a. Itwasbestforstrafingrunsanddeliveringdumbbombs.Deliveryofsmartbombs may notbe an agilityissue. The same is true for a/a. The flattum would be best for a/a gunnery and notguided a/a missiles.

Flat tums made the airplane more lethal and at the same time more survivable. The use of flat turns iscomplex. For example, the optimum dumb bombingtechnique combined classical roll and pitch forgross heading changes with flat turn for small changes• The pilotused roll stick to quickly get thepipper in the vicinityof the target. Remaining directional errorswere removed with flat run rudderpedals. This combined technique reduced exposure time to hostile fire and increased bombingaccuracy. Savings of 0.90 second on a 3 to 4 second final dive were routinelydemonstrated. Strafingwas markedly improved. On a single pass there was time to strafe more than one target•

It was importantto note that a finding was that flat tums were used only in the case of small (5 degree)heading changes. Beyond about 5 degrees, it was best to roll.

The conclusion is that flat turns are an importantflight mode as long as guns and dumb bombs are animportant part of the inventory.

4.3.2 Preliminary Design Guidelines

Design guidelines were established along with assumptions necessary to provide a realzst=cpreliminary design framework for the study. The following issues are individually discussed insubsequent paragraphs.

a. Departure Free Flight Operations.b. Airplane Weight and Balance.c. Finless S&C Design Criteriond. High Alpha Aerodynamicse. Thrust Vector Consideration.f. Moment of Inertia Consideration.g. Engine I_ailure Consideration.h. Axis System Consideration.i. Multi-Axis Simultaneous Control Consideration.

Departure Free Flight Operations

No studies to define ingredients to make an airplane departure free were made. It is felt that none ofthe currently available evaluation criteria has proven to be necessary and sufficient to guaranteedeparture free flight operations. Consequently, it is assumed that a smart and fast digital FCSNMScombined with active thrust vectoring for pitch roll and yaw control would make the airplane departurefree. It is believed that departure free flight operations will result from effectively used thrust vectoringcontrol power.

Airplane Weight and Balance

The designs shown in this report have been balanced. The balance of each configurations is basedon huiristics that establish location of the aft limit of cg. Once the aft limit is established then theweight of engines fuel and subsystem equipment are adjusted. Often the wing planform must beadjusted to get a satisfactory cg location. These huiristcs have evolved from comprehensive studiessuch as ATF, MRF AX. and ASTOVL. The assumptions used are listed below.

All the design rules are based on the location of the aerodynamic center. This location is predictedfrom simple and rapid vortex lattice analysis. This process is routine in Boeing preliminary design.

Type Airplane Aft cg Limit LocationFlying wing On the ac

Aft Tail 5% mac aft of the acCanard At the 'canard off' ac

59

These rules are based on recovery from any alpha with only aerodynamic control effectors.Consequently, the airplane is not dependent on pitch thrust vectoring for safety of flight.

Finless Airplane S & C Design Criterion.

The S & C design criterionforfinless airplanes is as follows:. The airplane shall be recoverable from abeta upset with the use of only aerodynamic control effectors. This means that the airplane then canbe safely flown in spite of a defective thrust vectoring system.

A finless airplane must have certain special characteristics. These characteristics are listed below.

a. Large yaw vectoring range (20 to 45 deg) with gas angle rates of from 80 to 100 deg/secb. Fast differential thrust magnitude that produces significant levels of yaw control.(even at

low power settings).Co An alternate source of yaw control that is independent of the engines. (Yaw vanes and

B-2 type split flaps)d. A means of controlling the thrust magnitude for flight conditionswhen the airplane at trim

requires low throttle settings. (Aero speed brakes and/or in-flight thrust reversing can beused .)

Items a. and b. are for normal flight operationswhen the airplane is stealthy. Items c. and d. are forabnormal conditions when flight safety, not stealth, is the main consideration.

High Alpha Aerodynamics

Methods to predicting forces and moments for flight conditions at high angle of attack are not reliable.This short coming was overcome by predicted high alpha data based on empirical data or based ondata extrapolations from wind tunnel test of similar configurations.

Thrust Vectoring Considerations

Thrust vectoring philosophy emphasizing yaw vectoring was adapted early in the study. This allowedtwo unusual features to be considered during development of the configurations:

a. The configurations could be fin-less.b. The configurations could have widely separated engines.

The thrust vectoring mechanization selected for this study is unique and innovative. The thrustvectoring has 45 degrees capability. The vectoring nozzle when exhausting over flap can producepitch. A two-engine arrangement could produce moments for pitch roll and yaw control. Thisrepresents a different philosophy from current designs. 'Now' airplanes emphasize pitch vectoring of20 to 25 degrees with no yaw vectoring or multiaxis axisymetric nozzles with limitedauthority (10 to 12degs).

Thrust vectoring is a nozzle term. It is the gross thrust that is being vectored and not the net thrust asused in the performance calculations. The gross thrust is often quite different from the static thrustand can be larger or smaller that the static thrust. The breakdown of net thrust into gross thrust andram drag is tabulated. The data is for a unity engine at power setting 1.0. Engines are scaled from thedata below..

60

Case # 1 2 3

Flight task - Base A/G A/AMach - 0 0.68 0.60Altitude (feet) 0 0 15,000Power Setting 1.0 1.0 1.0Gross Thrust (pounds) 20r966 28,899 16,398Ram Drag (pounds) 0 8,574 4,266Net Thrust (pounds) 20,966 20,325 12,132Fgross/Fgross sis - 1.0 1.38 0.78

Moment of Inertia Considerations

Moments of inertia have been estimated using empirical data. These moments of inertia are defined inthe body axis. These data are predicted for each airplane.

Pitch, roll,yaw and product of inertia values were estimated using historical data on actual airplaneswhich have significantparameters very similarto the ADS "design point"configurations. Values ofradii of gyration in percentages of wing span, body lengthor an average of the two were determinedfrom existing aircraft which have similar wing-span-to-body-length ratios, engine number and enginelocations. The percentages were then applied to the ADS airplane(s) dimensions and the inertia datagenerated at the combat weight conditions. In some cases the statisticalvalues were amended toaccount for specific peculiarities of the design and, therefore, improve the validity of the estimates.

Engine Failure Considerations

Powerful yaw vectoring allows the engines to be far apart. This design degree of freedom is notusually available. In case of one engine out the operating engine can be vectored so that the nozzleforce acts through the cg. This means that the mission can be terminated and the airplane can safelyreturn to the base.

Axis System Considerations

Forces and moments in both dimensional and non-dimensional form are given in the stabilityaxissystem. Analysis in the stabilityaxis system is the standard at this divisionof The Boeing Company.Conversion of inertias to the stabilityaxis system is routinelydone. Analyses shown in this report isdone in the stabilityaxis system.

Multi Axis Simultaneous Control Considerations

Agility metrics are defined for single axis. There is no intent to design the airplane for simultaneousapplication of 100% of control power to meet all the metrics at once. The control power definitions arefor a single axis based on a 1-DOF analysis.

Obvious trim and/or cross axis coupling is considered. Simultaneous control activity in several axis atonce is normal for a maneuvering airplane. For example, rollaround the velocityvector at high alpharequires adequate moments to null the inertia coupling and aerodynamic coupling to both pitch andyaw axes. Hence, there would be control activityin three axis.

The airplanes have been reviewed in a cursory fashion to ensure that there is adequate control powerfor realistic levels for simultaneouscontrol. For the rollexample; If flaperons are used for three axis(roll, pitch, and yaw) then there would be a separate allocationof span for simultaneous roll, pitch,and yaw; if the full available span is used to meet the rollmetric, the airplane would have a fatal fatalflaw.

61

4.3.3 MethodDevelopments

4.3.3.1 "rime to Bank and Capture 90 Deg

Design of the roll control system should be approached as a single degree-of-freedom roll about thevelocity vector. Military specificationsdo not allow the designer to take any credit for rolldue tosideslip. Coordinated flight must be maintained during roll maneuvers. Also an important part ofdesigning the airplane consists of ensuring that the vertical tail and ruder are adequate to hold zerosideslip (coordinated flight) during the roll maneuver. The easiest way to do this is to predict the timehistoryof a single degree-of-freedom rollmaneuver and then predict the maximum yawing momentsthat occurred. The rudder must have adequate control power to balance that yawing moment. Theyaw control power requiredto balance the yawing moment due to roll is a strong function of angle ofattack.

Total aerodynamic yawing moment duringthe coordinated roll maneuver is

n= IxzP (1)

where

n = aerodynamic yawing momentI xz = product of inertia about the x-z stabilityaxes

= rollacceleration

Aerodynamic yawing moment consists of contributionsfrom roll rate, the roll control system, and therudder. The design problem is to

a. Size the ailerons, spoilers,etc., so that adequate roll performance is attained. Aileron eeffects can probably be predicted using linear aerodynamics. Aeroelastic effects andspoiler characteristics are ignored.

b. Design the vertical tail rudder so that the yawing moment due to roll is balanced out.Notice that directional stability requirements might be more critical than turn coordinationwith regard to vertical tail size. Also, however, keep in mind that the tail has to accomplishdirectional stability and turn coordinationconcurrently and this has important implicationswhen artificialdirectional stability is used. If, for example, the airplane is artificiallystabilized by feeding sideslip to the rudder, the turn coordination signal cannot bepermitted to bottom out the rudder.

Roll performance and tail size requirements must be analyzed at several flight conditions. The tailrudder size design point is very likely not at the same flight condition at which the rollcontrol surfacesare critical. For example, the rollcontrol system will be designed to provide a minimum level of rollperformance at some point inthe combat flight envelope. Rollperformance will be higher every placeelse in the combat flight envelope. Vertical tail and rudder design requirements will be determined bysome combination of high angle of attack (high Ixz) and high rollacceleration, not necessarily the rollperformance design point.

Figure 4.3.3-1 was developed from the time-to-bank and capture algorithm developed in reference 1._This chart predicts rollcontrol power required to meet any specified time-to-bank and capture 90degrees agility metric goal. This figure assumes that the rudder is sized so that sufficientyaw controlpower is available to balance out the yawing moment due to roll about the aircraft's velocity vector.The figure has dimensional rolldamping and rolltime as the independent variable and initial angularacceleration as the dependent variable. The rolling moment coefficient required can then becomputed from the equation below.

62

ooI I I I

E

¢'L_I

8_

E

I I I

0

co_o

AClrequired=Sb

where.-

ACIrequired

8 - Roll acceleration rad/sec 2

0 lyYstab - Roll moment of inertia about the stability slug-ft2axis

- Dynamic pressure Ibs/ft2

S - Wing reference area ft2

b - Wing span ft

Lp - Roll damping inthe stabilityaxis l/tad

Figure 4.3.3.1-1 illustrates the exponentially increasing roll control power requirements necessary torealize time to bank and capture 90 degrees in less than 1 second.

4.3.3.2 Longitudinal Control Requirements

Any of the agility requirements relating to longitudinalcharacteristics need s to be considered at thesame time as control surface sizing, c.g. envelope requirements, and optimum landing gearplacement. These three issues must be accomplished simultaneously. Regardless of what controldevices are selected to accomplish the extreme angle of attack, or what devices are used to meet thepitch acceleration agility goals, the center-of-gravity location is of critical importance. The traditional"X-Plot" shown in figure 4.3.3.2 with the addition of the longitudinal agility requirements is therecommended approach.

The X-chart is a plot of horizontaltail arae, SH, versus fuselage station, F.S. forward and aft c.g. limitsare then plotted. These lines hopefully cross, forming an X. Thus the name: X-chart. For a flying wingdesign, fap-to-wing-chord ratio might be plotted in place of SH. A sample X-chart is shown in figure4.3.3.2. There is usually a best order in which to place the lines on the X-chart. The first stop is topredict aerodynamic center versus tail area. Methods used will depend on the configuration, windtunnel available, etc. Aerodynamic center will depend on Mach number and dynamic pressure(aeroelasticity effects). During the initial design phase, aeroelastic effects are seldom available.Judgement is needed in order to choose what flight condition the ac curve is predicted for. As theproject continues and more and more is learned, ac curves for more flight conditions will appear on theX-chart. In figure 4.3.3.2, two ac curves are shown: one curve represents low speed flight and theother represents a high-speed flight condition. At this point an important decision must be made:What stability level will the airplane be designed to?

The table below lists suggested points of departure for conceptual design location of the aft center ofgravity relative to the aerodynamic center. As more information becomes known about theconfiguration, this information sould be updated.

Type AirplaneFlying wing

Aft TailCanard

Aft c_l Limit LocationOn the ac

5% mac aft of the acAt the 'canard off' ac

64

0301

(0

oN

"C0

,1-Taft areawith high(xcapability _

Tail areawithouthigh¢zcapability

(_ = -.35 rad/sec2h = 15,000 ft

M=.6 --_

\$

\$

c.g. envelope withhigh _ capability

n = 6.0h = 20,000 ft

M=1.8 -,_

c.g. envelope withouthighc¢capability

\\\

--IIt

\$

a.c. low speedAft c.g. limit \ _ a.c.

5% static margin -_ "__105._O00ft=./- Forward c.g. /-- Forward c.g. _ _ "%,

/ limit nose / limit if airplane " _, ,p wheel liftoff P has capability i i /

to trim at e I /

_=70o I / /h = 15,000 ft e /

M = .6 ///i -"/

• ------ Tip over limit

• i

/

/=

/

Fuselage station of the center of gravity l

Main gear

Main _lear locationwith htgh ¢xcapability

Figure 4,3.3.2.

spb-8.7-m-Ad3

In the sample X-chart, a conventional stability level of 5% MAC has been chosen. The aft cg limit canbe drawn a distance of 5% MAC ahead of the critical ac. Notice that at low tail areas, the low speed ac iscritical and at the larger tail areas, the high speed ac is critical. This is not a "probable" result, merely anillustration of one of the things that can happen. Once the aft cg limit is established as a function of tailarea, the landing gear location can be put on the chart. Optimum landing gear location will also be afunction of tail area. The configurator will determine how far the main gear must be behind the cg toprevent tip-over. We don't want the gear to be much farther aft than this because that aggravatesnose wheel lift-off problems. The main landing gear location can now be drawn on the X-chart. It isdrawn at the minimum tip-over distance between the aft cg limit As the design progresses it is usuallydifficult to maintain the optimum gear location and it will end up a little bit aft of the gear location curveshown on the PD X-charts. This may cost a small increase in tail area depending on how criticalnosewheel lift requirements are. The next step is to start putting forward cg limit lines on the chart.

Forward cg limits can results from a number of different requirements. Nose wheel liftoff is a commonlimitingfactor, especially jet airplanes with slab tails. Sincethe cg is always ahead of the main landinggear, it is harder for the tail to rotate the airplane around the gear than the cg. Horizontal CLmax mustbe determined, or assumed; the cg location is found where the airplane balances on the main gear(nose gear reaction is zero) at the required rotation speed. This is done for a variety of tail areas so cglocation can be plotted on the X-chart. The resultantcurve is the forward cg limitwith regard to nosewheel lift-off. Maneuver requirements can also determine the forward cg limit, especially if the airplanehas a supersonic capability. Design requirements might call for certain maneuver capabilities atvarious points in the flight envelope. They will all have to be analyzed eventually but a littlejudgementcan usuallyyield the criticalones for PD purposes. As an example, the airplane may be required to pull6 g's at 20,000 ft and mach = 1.8. This condition is represented on the X-chart by predicting the cglocation with various tail areas with the airplane at the specified flight condition. The tail is, of course,loadedto its maximum CL in each case. Notice in the example X-chart that this condition did not turnout to be as critical as nose wheel lift-off. So far we have not addressed any of the "special" agilityrequirements. They belong, however, on the X-chart.

There are some additional X-chart features that should be discussed. The cg envelop must be fittedin between the forward and aft cg limits. Notice that when you do this you don1 get to choose thelocation. If the actual cg envelope is someplace else, the design does not balance and must be re-configured. In the case of conventional airplanes, this is usually easy. The wing just "slides" forwardor aft and analysis begins again. In the case of a flying wing, there may not be enough material tomove around. Sometimes a flying wing plan form must be abandoned because it cannot be made tobalance.

Canard configurations are another special case. Canard area replaces tail area on the vertical axis. Asthe canard area grows,the ac moves forward instead of aft. All the lines ,forward and aft cg limits, leanto the left. There is no guaranteed solution. The cg envelop may have a negative length at anycanard size. Our design approach to the canard configurationis to put the cg at the canard off ac. Allthe aft cg limitsar the vertical lines on the X-chart. This, however, results in extremely unstableairplanes with canards of any significantsize.

4.3.3.2.1 Minimum Nose Down Pitch Acceleration

Firstwe address the problem of minimum nose-down pitch acceleration using mach = 0.6 at 15,000 ftas a sample flight condition. There may or may not be some special devices to help meet thisrequirement. In any case the tail should be used to help so pitch-down acceleration will be a functionof tail size. Even if the tail is not used as a controller, it will affect the problem through its stabilitycontribution. Assume, for this example, that thrust vectoring is used to aid in pitching down. Aconstant nose-down pitching moment might be assumed from the thrust vectoring plus an additionalincrement proportionalto horizontal tail area. A horizontal tail CL max must be determined or assumedfor this flight condition. A thrust level must also be assumed. In the sample X-chart, this requirement isnot criticaland has no affect on the cg limits.

66

4.3.3.2.2 Maximum Trimmable Angle of Attack

Determining the forward cg limit for trim at the highangle of attack, ((_= 70 degrees for example), the

case requires a knowledge of the nonlinear aerodynamics not generally known during the PD phase.

Assumptions for variations in the aerodynamic center location and the magnitude of the normal forceenable the evaluation of control requirements for trim at high angles of attack. Figures 4.3.3.2.2-1 and-2 present the nonlinear behavior of normal force coefficient and the center of pressure for the F-16,F-18 and a flying wing configuration. When the normal torce is normalized with total projectedplanform (includingthe canard and tail) the data collapses along a singletrim line. This high-alphatrend can then be faired into the linear low alpha data computed using simple vortex lattice methods.

At angles of attack near 90 degrees, the normal force is equivalent to the drag of a flat plate and has itscenter of pressure at the centroid of the area of the projected planform.

Predictionof the pitch moment to trim at any alpha is then based on the equation:

ACmtdm = CNgross (Xcg : Xcp)Sref

where:CN - Normal force coefficient

to total aircraft projected platformas a function ofangle of attack

Xcg - Longitudinal position of the center of gravity

Xcp - Longitudinal position of the center of pressure

If thrust vectoring is used, effects of angle of attack on inlet characteristics must also be known. In anycase, the tail is probably a factor and the cg location to balance the airplane with all the control efforts atmaximum capability will be a function of tail size. An example of how this function might look is shownon the sample X-chart. The curve is shown as a "painful" result. This is done not because of anyoption regarding trim requirements at high angle of attack, but to illustrate what might happen whenunusual requirements are imposed on a design. The X-chart in the sample case shows us that thehigh angle of attack trim requirement is very expensive in terms of tail size and, therefore, airplaneweight and cost. All the other forward cg limit lines are grouped together. If there were no nosewheel lift-off requirement, the tail could be made smaller, but not much smaller. The trim at 6 g's or thepitch acceleration forward cg limit lines are encountered at only slightly smaller horizontal tail areas. Tailsize required to meet the high angle of attack requirement, however, is much larger than that requiredto meet any of the other criteria. In this case, the X-chart is telling us we have a detective design. Onesolution might be to use some other or additional pitch control devices to accomplish the high angle ofattack trim. In any case some re-evaluation is indicated.

4.3.3.3 Maximum Lateral Sideforce

There ar two basic approaches to generating the sideforce necessary for a wing's level turn. The firstwould be a control effector that would develop a sideforce without any sideslip. These devices couldbe vanes with skewed hinge lines, bomb bay doors with skewed hinge lines, ventral fins, folding wingtips, and landing gear deployment. These devices would have to be located at or near the center ofgravity or they would generate a sideslip that would have to be balanced out by some other controldevice to achieve zero sideslip. Stealth requirements would require the devices be retracted untildeployed. Deployable devices that operate at high dynamic pressures (690 Ib/ft2 is the point ofinterest for the air-to-ground designs) and have substantial structure. Large and structurally stronglanding gear have structural placards at 200 to 250 KEAS. The maximum lateral sideforce at zerosideslip approach was therefore abandoned.

67

3.2!

2.8

2.4

2.0

CNgross

1.6

Normal force

CNgr°ss = q'x S gross

Wingstall

F-16 WBH

Sgros s_<

F-18 WBH

Delta wing

Used in analysis

Delta wing - <>F-16- 0F-18 - El

I I

10. 20. 30. 40. 50. 60. 70. 80.

Alpha

Figure 4.3.3.2.2-1. Normal Force Referenced to Plan View Gross Area

68

XCp

.8

.7

XCp = [Xcp "XLEMAC t

Wingstall

F-16 WBH

Used in analysis

F-18 WBH

Delta wing

.1-

Deka wingF.18F.16"_ I

I I0 I I I I I10. 20. 30. 40. 50. 60. 70. 80.

Alpha

Figure 4.3,3.2.2-2. Center of Pressure Referenced to Plan View Gross Area

69

The second approach would be to allow 10 degrees of sideslip while maintaining wings level. This islarger than the +5° effective wing level sideslip angle findings on the AFTI F-16 discussed in section4.3.1.

The sideforce agility design goal is expressed by the equation:

_' = ny _/ ft2q" GOAL q

And the sideforce generating capability of a control device is given by:

= Cy S ft2q DEVICE DEVICE

Assuming a combat gross weight of 50,000 Ibs, the sideforce requirements can be computed andpresented in figure 4.3.3.3-1.

MissionCombat

Gw=50T000 Lbs

A/A

qbarPSF

301A/A 301

NGNG

686686

AltitudeFT

15,000151000

Agile level

MediumHigh

MediumHigh

alphadegrees

3030

nyg's

0.41.0

1.22.0

Y/qbarSq Ft

67167

87145

Figure 4.3.3.3-1

Figure 4.3.3.3-1 shows that the air-to-air requirements for maximum lateral sideforce are the mostdemanding because of the lower dynamic pressure of the requirements flight conditions, andbecause of the loss of controller effectiveness at high angles of attack (figures 4.3.3.3-2 and4.3.3.3-3).

It is clear from the analysis that the side force agility goals can only be reached by using severalaerodynamic devices in combination. It is also clear that yaw thrust vectoring is the most effectivedevice.

B-2 type splitflaps are a powerful means of producing yawing moment. There is a small loss of lift androlling moment to consider for trim of these flaps. The resultant increase in drag is large.

The early aJg designs showed swept-forward trailing edges. This was changed to scalloped-trailingedges so that B-2 type split flaps could be used to trim yawing moment developed by the side-force-producing devices.

7O

Wings left turn

Control eflectors provide the force

Vo

13=0

Cm=Cn=CI=0

Combat GW = 50,000 Ibs

200

Ft2

I00

A/G

o_.5o- =686psf

11y"-2.0high

145 Aft finfwd finS=240

TVY20o.2 + nosefin

"NBlowing

nose Jl

strakesJJNose _. II

str_es U U

TVY 30°50 ftfwd fin

+ strake

20O

I lly = 1.0high

100

I

67

0

A/A

(z=30°_= 300 psf

TVY 20°50 ft fwd fin

pibt

Blowingnose

strakes Aft finfwd finS=240

Airto ground Airto air

TVY 30°50 tt fwd fin

+ strake

Figure 4.3.3.3-2. Side Force Control Effector Selection

71

1.0

I CYtest limitcos (cant). K carryover

.5

Basedontrue area

1

1

0 I I I0 10 20 30

o_~ deg

Rudders

Pivoting fins

High lift mddem

Nose yaw strake hinged device

Nose yaw by slow-blowing

Thrust vectoring5 = 20°Unbalanced

Figure 4.3.3.3-3. Sideforce Control Effector Effectiveness With Angle-of-Attack

72

4.3.4 Configuration Evaluation

4.3.4.1 Agility Impact on Low Observable Configurations

A vortex lattice model of the air-to-ground flying wing concept is shown in figure 4.3.4.1-1. Theaerodynamic characteristics shown in figure 4.3.4.1-2 are the results from the vortex lattice method. Tl_eflight condition shown is a combat gross weight of 58,270 Ib and an airspeed of 450 KEAS.

Both the high agilityand low agilityversions of this aircraft have the control power necessary to trim at the9g limit load factor. This limitload occurs at an angle-of-attack of 10 degrees because of the lowwingloading of the flying wing concept. The trim at this 9g condition requires less than a 10-degree trailingedge up deflection from the inboard flaps. An alpha limiterwill be required to prevent inadvertentexcursions outside the aircraft structural envelope. Limited pitch thrust vectoring in combination withtrailing edge flaps yield a responsive capability in load factor while retaining powerful control power foralpha limiting.

The time to bank and capture 90 degrees was accomplished using the method outlined in section4.3.3.1. The results are summarized in figure 4.3.4.1-3.

The time to bank and capture 90-degree agilityrequirements are well within this configurations ability toachieve. The yaw control pwoer to balance the rolluses only 10% of the total available yaw control power.

A combination of four control effectors were used to meet the maximum lateral side force agilityrequirements. The control effectors and their contributionto the lateral side force are shown in figure4.3.4.1-2.

The engine thrust is the dominate control effector, contributing 57% of the control power for the medium-agilityaircraft and 77% of the control power for the high-agilityaircraft. The aircraft T/W requiredto meetthe high-agility level is 1.6, well outside what could be reasonably expected to be available on a fighter.

A smart digital flight-control system is required for the effective integration of the control effectors shownin figure 4.3.4.1-4. The rollcoupling from the B-2 type splitflaps were found to be small. Yaw controlduringthe side force maneuver can be achieved by differentiallyvarying split flap deflections or yaw thrustvectoring.

Xaw thrust vectoring is the most effective side force producingcontrol effector. Oversizing the enginewould translate into gains into maneuver performance at the expense of aircraft weight and range.Aerodynamic control effectors to achieve the side force requirements would increase the weight of theaircraft without any additional synergistic improvements anywhere except for the maximum lateral sideforce.

4.3.4.2 Observables Impact of High Agility Designs

The vortex lattice model presented in figure 4.3.4.2-1 is a high agility moderate observable air-to-air fighterconcept. The aerodynamics resulting from the vortex lattice analysis is presented in figure 4.3.4.2-2.Figure 4.3.4.2-3 presents the agility levels achieved by the concept aircraft broken down by controleffector.

The pitch control power to trim the aircraft at high angles of attack is much greater than that required tomeet the nose down pitch acceleration agilityrequirements. Pitchthrust vectoring is again the mosteffective control effector. The ability to trim the aircraft at 70 degrees angle-of-attack will require thecombined use of pitch vectoring and over-rotating the horizontal tail.

73

• 1,196 panels• 27 trapezoids

Delta Wing Vortex Lattice Mode/

Figure 4.3.4.1-I

74 sl_8-7-m-C1

.05

Cm 0

-.05

-.10

BaseI

1.0

Moment reference

Xref = FS 247.2"Sref = 1,496 ft 2

Bref = 68.33 ft

Ore f = 335.18"

CIl¢z

0.2

0

-0.2

-0.4

10

BaseI

2O

CL

CL(z

I IE

.06 _

.04

.02

I00 10

G.002

Cm(z 0

-.002

-.004

.001

10(z

.o.--o---o

-.002

I

Basei

20

BaseI

2O

BaseI

20

BaseI

2O

.001

0

-.001

-.002

-.003

-.004

-.005

.0002

0

-.0002

-.0004

-.0006

-.0008

-.0010

I I101

20

Base Cmctll 0" 2D

" 30" 4ZX

4 321

I I10 20(z

Base CIIctl 0" 2D

- 3<>_ " 4A

Figure m.3.4.1-2

75

spb-8-7-re-Ad4

Time-to-bank and capture 90 °

Roll damping (stability axis)

Roll acceleration

Rolling moment coefficient required

Control surface deflection

- sec t

- 1/rad Lp

- rad/sec

ACIreq

- deg

Mediumagility

1.5

3.8

5.5

.0146

10

Highagility

1.0

3.8

9.8

.0256

30

Figure 4.3.4.1-3. Time-to-Bank and Capture 90 Degree

spb-8-7-re-Ad9

76

pg042794

AG 2501Side Force Agility

Vanes

CD= = l.O

Flap

(_ =50to70Deg

C D= = 1.0 _'_

Delta Side Force Control Concept

77 Figure 4.3.4.1-z_

988-115 Vortex Lattice Model

Figure 4.3.4.2-I

spb-8-7-te-C2

78

.05

Cm 0

-.05

-.10

1.5

0.5 1.0CL

Moment reference

Xref = FS 392"Sref -- 1,062 ft2

Bref = 48.8 ft

Cre! = 192" 0.2

CIIcc 0

-0.2

-0.4

I I

10 20

-0

CL1.0 --

0.5

0O- 10 20

(z.06 -

CL(z

.04

.02

0

.002

-<i)

I I0 10 20

Cma 0

-.002

-.004

I I10 20o[

Clll3 o

-.0005

-.001

-.0015

I I

.001

0

-.001

-.002

-.003

-.004

-.005

.0004

0

-.0004

-.0008

-.0012

-.0016

-.0020

I10(z

o.-.o---o-

[] Inb. TEFI

20

-J_ Out. TEF

--fO H. tail

B-DERA_PGMOD Cmctll C)" 21-1

" 30

(z10 20I I

,,I-I Inb. TEF0 H. tail

__.(_.,.,=_.,,.,,.._ Out. TEF

B-DERA_PGMOD CIIctl O" 2D

" 30

988-115 Aerodynamic Characteristics

Figure 4.3.4.2-2 spb-8-7-re-Ad5

79

0.2

Trim(x= 60°

aft limit = .33C I

PitchACm

0.1

0

.08

Roll .04ACI

(7.=35 °

.35r/s2andmaneuver

Full spanailerons.057

m

-- .013o¢= 20 °R&S 90°1.5 sec

|.08 --

Over rotatetail for controland trim

17

_H = -30° II

II

_51_V_p = -30°

Horizontaltail roll

r7

Yaw .04ACn

0

Trim roll.009

|

Yawvectoring

m

I

2.0

o

Side force

ny, 1.0~gs

0

OTVy = 45 °

max pwr.707 g's

= 15°.122

[-1

Enginesize

Doors _-I

290 ft 2

M988-115 Control Effector Selection

Figure 4.3.,_. 2-3

B Goal

D Available

Vo = 634 fps

Hp = 15kM = 0.6

sl:_.8-7-re-Ad8

80

Roll control power is sufficientto meet the time-to-bank and capture 90 degrees agility requirement.Sufficient yaw control power is available to balance the yawing moment generated by the rollabout thevelocity vector.

Side force is dependent on yaw thrust vectoring as were the low observable designs presented in the ,previous section.

One of the key design traits of low observable designs is the emphasis on keeping the number of controlsurfaces down to a minimum. The most obvious impact is the lack of control surfaces available to addressany handling quality or agility requirement. Side sector signature is completely counter to the availability ofefficient lateral control devices to meet the maximum lateral side force agility requirements. Any significant

_"side sector signature requirement drives the aircraft to a highly coupled V-tail configuration and eventuallyto eliminating the tails altogether. Analysis has shown that yaw thrust vectoring is the most effectivecontrol effector in achieving the maximum la_teralside force agility requirements.

The signature impact on longitudinal agility requirements are not as extreme as that of the lateral-directional agility just discussed. This observation is primarily due to the horizontal orientation of the mosteffective pitch control effectors is favorable to signature requirements. Even with the availability ofnumerous options for pitch control effectors, pitch thrust vectoring is the most effective control effector.

Time to bank and capture 90 degrees does not seem to be affected by the signature issue. This isbecause the most effective roll control devices have the favorable horizontal orientation and wing aileronsseem to have the control power necessary to meet the agility requirements

81

5.0 Conflauratlon Synthesis Results

There are primarily two approaches to aircraft synthesis studies, numerically optimization and thetraitionai trade study approach. The numerical optimization approach provides a highly refinedoptimized solution subject to all the constraints supplied and limitationsof the parametric sizingmodels. The tradition trade study approach is a long cumbersome series of trade studies thateventually reveal an optimal solution. The traditional trade study approach was selected because itprovides the visibility into what the important design parameters are, where the design constraintboundaries are relative to each other, and what the sensitivities are about the design point. The"blackbox" nature of numerical optimization does not lend itself to visualizing these global issues.

USAFCustomer

Six of the twelve study configurations were designed for an USAF only customer. Traditionally themost important design parameters determining the size and cost of a concept is engine size (T/W),wing size (W/S), and wing shape (AR). Generally the aircraft thrust-to-weight (T/W) ratio was driven bythe maneuver requirements, wing loading (W/S) was driven by the instantaneous turn requirements,and aspect ratio (AR) was varied to minimize the empty weight/cost of the designs.

Agility requirements of maximum achievable angle-of-attack, minimum nose down pitch acceleration,and time-to-bank and capture 90° are primarily determined by the control power and inertialcharacteristics of the basic concept. Traditionally these issues are ignored in the configurationscreening stages untilwind tunnel data becomes available to address these and many other handlingqualities issues. In this study, control effector sizing for agilitywas built into the overall concept usingthe process discussed in Section 4.3. Control effector volume coefficients were held constant duringthe synthesis studies with the assumption that scaling control effectors size using constant volumecoefficients wouldyield similar handling characteristics. There is no data to support this assumption.The agility requirements for maximum negative specific excess power only drove the air-to-groundconfigurations until the maneuvering flap was added to the concepts. The maneuver requirementsfor the A/A and A/G configurations were demanding on aircraft T/W requirements.

The most demanding agilityrequirements for these tailless configurations is the maximum lateralsideforce requirements. Yaw vectoring is the single most effective means of achieving the sideforceagility requirements for the high T/W A/A configurations. The T/W level required to meet the maximumnegative specific excess power agility, on the A/G configurations with the leading edge device, wastoo low to have sufficientyaw control power from yaw vectoring alone. Deployable yaw vanes and splitailerons were added to increase the yaw control power to meet the maximum lateral sidefore agilityrequirement.

Joint Service Customer

The remaining six of the twelve configurations are derivatives of their Air Force counterparts.Generally, a 15 to 17% increase in empty weight over their Air Force counterparts to do the samemission and meet the same maneuver requirements. This increase in empty weight is due toincreased structure to accommodate higher design sink speeds for landing gear design, tail hook,nose wheel shuttle, and wing folding mechanism.

5.1 The Global Design Space

The Air-to-Ground Maneuver requirements were examined in a Global Design Space Study presentedin figure 5.1. This figure shows the variation of aircraft thrust-to-weight required to meet the air-to-ground maneuver and agility requirements with the aircraft geometry varying in a historically relevanttrend. The 6.5g maneuver requirement was the dominate requirement sizing the engine except forthe Maximum Negative Specific Excess Power Agility requirement. The interpretation of the MaximumNegative Specific Excess Power Agility requirement at the time this data was generated was that theflight condition occurred at CLmax. The conclusion drawn from this chart was that configuration swithpoor high lift capabilities had an advantage over more maneuver able designs because they could not

82

reachthesamehighliftconditions.ThereforetheMaximumNegativeSpecificExcessPowerAgilityrequirementwasmodifiedtooccurataconstant load factor to negate influence of obtainable CLmax.

Similar historically relevant trend data were used to examine the global design space of the air-to-airmaneuver requirements. The results presented in figure 5.2 show that the medium agility level ofMaximum Negative Specific Excess Power requirement does not drive the size of the enginerequired. The high agility levels match the maneuver requirements. Concern about the transonic,acceleration requirements are only relevant at the low wing loading, high aspect ratio portions of thedesign space. Expected aircraft thrust-to-weight ratios in configuration sizing trade will be from 1.1to 1.3.

83

.o_

¢.-

¢=

0

2

0

I--

0.9

0.8

0.7

0.6

0.5

0.4

0.360

89,000

87,000

000

I

i73,000

0

6.sg

Figure 5.1.

70

!

69,000

000

80 90

Takeoff wing loading (Ib/ft2)

100

Global Design Space; Twin Engine Air-to-GroundDeltoid; Takeoff Gross Weight (Ib)

110

spb-7-7-re-Ad3

84

Max neg P== -100 ft/seo

Max neg P= ,_.q,= 450 ft/sec

i I I I

40 45 50 55 60 65 70 75

Takeoff wing loading (Ib/It 2)

Generic AJA Configuration; A/A Max Neg Ps Requirement; T/W Required

80

OR,G,,_,_ F;:.CE IS

OF POOR "QUALITY

_:=-7.7-r_

GOO"1

I

=o

GObI

O

Wing aspect ratio

•_ .1=.b LnI I

Accel M = 0.8 -_ 1.5@ 40,000 It ..-=

Fuel

bI

I

I

I

5.2 Aircraft Synthesis Results

The aircraft synthesis approach consists of four steps, as illustratedon figure 5.3, and discussedbelow.

=_p,,12..l_"PreliminaryLayout and Sizing"

This consists of preliminary layout and sizing of an aircraft that will be used as a starting point for theparametric analysis. This is typically a one ortwo day effort to (1) identify specifictechnologies, (2) sizefixed equipment and develop general arrangement of crew accommodations, instruments, avionics,gun and provisions, ammunitions weapon bay...etc., (3) develop overall shape to best meet systemrequirements and (4) estimate fuel requirements and size and layout of the aircraft.

"Parametric Analysis"

This effort requires the rapid analysis of a large number of aircraft designs that meet all systemrequirements; it is computation intensive and has been mechanized. The "Fighter Aircraft SizingTool" (FAST) of reference (1) is employed. Specific tasks are:

(i) Determine aerodynamics characteristics, fuel requirements and maneuvercapability of a specific configuration on a specific mission.

(2) Package fixed equipment and fuel perform loads, stress and mass propertyanalyses and size aircraft (iterative process required with (1) above).

(3) Conduct configuration trade studies to identify the minimum weightconfiguration that will perform the specified missionwithin the imposedsystem constraints. Typically, wing loading, thrust-to-weight ratio, andwing aspect ratio, leading edge sweep and thickness-to-chord ratio arevaried. Considerable interaction between final layout and sizing (Step 3)ex}sts during the selection of a configuration.

"Final Layout and Sizing"

Using the parametric sizing results as a guide, apply sound engineering sizing, packaging and massproperties analyses to develop a final aircraft design.

"Final Performance"

Determine the performance capability of the final configuration using the FAST program.

5.2.1 Air-to-Ground Configurations

The preliminary layouts of both the high agility and medium agility designs consisted of a deltaconfiguration with a saw-tooth trailing edge to satisfy the low observable requirement. Theconfiguration consisted of (1) a 1500 ft2 wing with aspect ratio of three, (2) no leading edge deviceand (3) a thrust level of 22,840 lb.

The parametric sizing results are provided on figure 5.4. The analysis consisted of an investigation of(1) wing leading edge sweep, (2) then wing aspect ratio, (3) then thrust-to-weight ratio and finally (4)the addition of a leading edge device to meet the maximum negative specific power requirement atsignificantly reduced gross weight.

The selected configuration was a compromise between (1) minimum weight, (2) the ability to balance adelta configuration, and (3) the ability to maintain a 6.5g sustained maneuver at sea level and mach =0.8 lb. It has the following characteristics.

86

Technology

Step I

l re" ina requirements layout andsB.=ng

Step 2

Parametric sizing

Identify minimumweightconfigurationthat meetsall system requirements

Fighter Aircraft SizingTool (FAST)

Step 3

Finallayout ands_.mg

Step 4

_..I Finalperformance

Figure 5.3. Aircraft Synthesis Approach

spb-5/94-re-Ad2

87

A

Q¢:>v

._m

(/}¢/)

2C9

88

86

84,<

,_ Max neg Ps=

// .... _-s95

_; S = 1,500 ft 2• AR -- 3.0• No LED• T = 22,840 Ib

EVA I I I40 45 50

Wing leading edge sweep (deg)

55

A

.o00o

¢,.-._m

¢/}

2(.9

85

84-

-700

_Max neg Ps =

- "__._-550• S = 1,500 ft2• ALE = 52.5 deg• No LED• T = 22,840 Ib

I I I3.0 3.5 4.0

Wing aspect ratio

4.5

+300 I I I I I

+200

+100

v

oY_ 0

X

-100

-500

• S = 1,462 it 2• AR = 3.5• ALE = 48.75 deg

High. agilityrequ,rement

80,000 Ib

Thrust-to-weight ratio0.5 0.6 0.7

85,0001b

I I I= 75,000 Ib

Selected configurationLeading edge deviceMinimum T/W • GW = 73,820 Ibto maintain • Max neg Ps = -58 90,000 Ib6.5 g'ssustainedload factor 85,000 Ib

GW = 80,000 Ib

90,000 Ib

I I I I I I I

I

100,000 Ib

95,0001b

Figure 5.4. Air-to-Ground Configuration Sizing Charts

spb-5/94-re-Ad 1

88

• Wing area = 1462 ft2• Aspect ratio = 3.5• Wing leading sweep = 48.75 degs• Thrust-to-weight ratio 0.336• Gross weight = 73,820 Ib

The same configuration was selected for both the high agility and medium agility design because itrepresents the minimum weight design. As shown on figure 5-4, it is possible to satisfy the 450 fpsmaximum negative specific power requirement of the medium agilitydesign, but the weight is greater.

The important design parameters are tabulated in the order of significance in table 5.5. The leadingedge device provides a significantweight reduction due to its aerodynamic effect, as shown on figure5.7. It provides a significant improvement in left coefficient at low angles of attack and a slightlyimproved maximum liff-to-dragratio.

The design sensitivities about the design point are provided on table 5.6. Another interestingsensitivity, although it is not about the design point, is provided on figure 5.8. Presently, therequirement is to calculate the maximum negative specific power at the point in the mission where60% of the fuel remains on board. If this requirement were changed to 50% fuel remaining on board,the aircraft gross weight could be reduced 6,700 Ibs.

5.2.2 Air-to-Air Configurations

High Aailitv

The a variation of aircraft gorss weight with the significant design parameters and design constraintboundaries is illustrated on the configuration sizing chart on figure 5.9. The chart was constructe(;Iwith a thrust-to-weight ratio of 1.125 which allows a small design space betweenthe sustained loadconstraint of 4.4 g's (H = 20,000 It at M = 0.6) and the specific power constraint of 550 fps (M = 30,000ft at M = 0.9). The other twenty-four constraints are all satisfied. Note that the minimum weight designoccurs at an aspect ratio of approximately 5.5. It is anticipated that this design would be subject to asevere weight penalty due to flutter. Without a detailed analysis, we have selected a design with alower aspect ratio to avoid flutter. The selected configuration lies on the sustained load designconstraint at an aspect ratio of 3.75. It has a gross weight of 59,835 lb. Other characteristics aretabulated on figure 5.9.

Medium Agility

The variation of aircraft gross weight with the significant design parameters and design constraintboundaries is illustrated on the configuration sizing chart on figure 5.10. The chart was constructedwith a thrust-to-weight ratio of 1.13 which allows a small design space between the instantaneous loadconstraint of 9 g's (M = 30,000 It at M = 0.9) and the specific power constraint of 550 fps (H = 30,000 ftat M = 0.9). Again, the minimum weight configuration occurs at a higher aspect ratio, but aconfigurationwith an aspect ratio of 4.0 was selected to avoid flutter.

5.2.3 Multi-Role Configurations

Hiah Aailitv

The variation of aircraft gross weight with the significant design parameters and design constraintboundaries is illustrated on the configuration sizing chart on figure 5.11. The chart was constructedwith a thrust-=to-weight ratio of 1.1 which allows a small design space between the sustained loadconstraint of 4.4 g's (H = 20,000 ft at M = 0.6) and the specific power constraint of 550 fps (H = 30,000ft at M = 0.9). The other twenty-four constraints are all satisfied. Although a slight weight saving isindicated at higher aspect ratios, a configuration with aspect ratio of 4.38 and wing loading of 67.5 psfwas selected to avoid flutter.

89

Table 5.5. Important Design Parameters

Parameter

1. Wing leading edge device

2. Thrust-to-weight ratio

3. Wing aspect ratio

Significance

Results in a 25,000 Ib weightreductionat maximum negativespecific power of -100 fps

Maximum negative specific powerand 9ross weight are extremelysensitive to these parameters

Table 5.6.

Q)r_

r'-

Design Sensitivities About the Design Point

Partial of:

Gross

weight (Ib)

Maximumnegative specificpower (fps)

Thrust-to-weight +35,600 +765

Leading edge sweep (deg) -240 +2.3

Aspect ratio -1,430 +125

spb-7-7-re-Ad4

9O

._o:I=(l)

8.m.-J

._o

8

,.J

1.0

0.8

0.2

0

With leading / ss SS

edge device -._ / sSSs S

SSSS S

/ "(s SSSSS No leading

I s S" edge device

ISS S

I I I I2 4 6 8

Angle-of-attack (deg)

10

With leadingedge device(L/D)max= 17.4

No leadingedge device(L/D)max= 16.5

I I0 0.1 0.2 0.3

Drag coefficient

Figure 5. 7. Aerodynamic Effect of Leading Edge Device

spv.SF34.re-Ad3

91

=£

C

X¢;

0 8,,oo0//Max neg Ps

@ 50% fuel _82,000#/J" 4

T/W = 0.543 _/" #GW = 78,450 Ib -_ 80.000,._ ,-,,,, /

¢u,uuuj GW = ,/

/ 78,000 Ib_,s"

Max neg Ps_@ 60% fuel

-50

-100

-150

-200

High agilityrequirement

-250 I I I I I I I0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8

Thrust-to-weight ratio

Figure 5.8. Design Sensitivity to Fuel Load for Max Neg Ps Requirement

spv.5/94-re-Ad4

92

70,000

¢.OCO

A

v

1=._m

u)u)2(.9

68,000

66,000

64,000

62,000

60,000

58,000

56,000

/_ Selected cOnfolguratlon988-115• T/W = 1.125 I / , w/s = 58psf

-- 40o_ /I I'_ AR = 3.75/ J [ _ .GW=59,8351b

.,7E I I _ • Sref = 1,032 ft2__ ,/ • I J _ "b=62.2ft

-1.9oX/_ _,, j___ .ILOAD3=4.38g's"_ Y _ / "_40 ° • Max neg Ps = -60 fps

E'-- Selected_ \ configuration

\ _'/_._ /- Specific power ="-'_ \ ,-(,._-. / 550 fps

• _ \ _'/"_,.Z'.Z._Z . . / • H = 30,000 ft"_'_ _"/_/--Z..E,Z_/.(./../__)// ... • i = 0.9 _n

_ - _ _.,.r../..(,/_

/ -- " _ _ ___ . /- Minimum weight design_ / - / _ / Design space /

Sustained load --

4;4 ig'SH= 20,000 ft° M = 0.6 70

<_I I I I I I I I I I --I I I I I

M 3.6 3.8 4.0 4.2 4.4 4.6 4.8 5.0 5.2 5.4 5.6 5.8 6.0 6.2 6.4 }.6

Aspect ratio

Figure 5.9. Air-to-Air, High Agility Configuration Sizing Chart

spb-5/g4-re-Ad8

67,000

¢.OJ:=

O©

A.Q

¢0¢/3

2

66,000

65,000

64,000

63,000

62,000

61,000

A Selected configuration 988-114• w/s = 45.05 psf• AR = 4.0• GW = 64,030 Ib

Sref -- 1,420 ft2b = 75.4 ft• ILOAD3 = 8.99• Max neg Ps = 318 fps

w/s = 44.5

45.0

• T/W = 1.13

_. Mlnstantaneousload =9 g's

• H = 30,000 ft

= 0.9

F Specific power =

550 fps• H = 30,000 ft.M=0.9

V

V

Selectedconfiguration988-114

45.5

46.0

Leading edgesweep = 50 deg

3.3 3.4

Figure 5.10.

I I I I I3.5 3.6 3.7 3.8 3.9

Aspect ratio

Leading edgesweep = 45 deg

4.0 4.1 4.2

Air-to-Air, Medium Agility Configuration Sizing Chart

spb-5/94-re*Ad5

58,000

• T/W= 1.1

COO1

A

v

E-.g_

2(.9

56,000 -- _ _¢,, w/s = 52 psf

(v'_(, Specific power =

_"_ • H = 30,000 ft54,000 -- 550 fps

.M=0.9

52,000

50,000

48,000

Sustained load =4.4 g's

• H = 20,000 ft• M =0.6 Design space 56

6O

Selected

configurationw/s = 67.5 psf

AR = 4.38

>

___ I i I3.0 3.5

_--- Max neg Ps= -100 fps.H = 15,000 ft

• M =0.6• Fuel fraction = 0.6

i I4.0

Aspect ratio

L_ Selectedconfiguration

I I

64

68

4.5

72

Figure 5.11. Multi-Role, High Agility Configuration Sizing Chart

5.0

spb-5/94-re-Ad10

COOb

A.O

(tJf/)2r5

58,000

56,000

• T/W = 1.05 //_.Note: no design space

1) Psl0 >-550-- 2) Max neg Ps ->-100 fps

w/s = 52=,_======w

54.000 " """ "" _ _ .. ,--- Specific power = . .....

"_.... / 530 fps* H = 30.000 ft @ M = 0.9

_ ._ _ 60

52,000 - . _"_ Max neg Ps_ -_00 fps'''''--..'"""

50,000 -- __. ....

, ,ooo- N1" 'p°wer=--/ -'"P

I , i I i , i , I ,"-_ 3.0 3.5 4.0 4.5 5.0

Aspect ratio

Figure 5.12. Multi-Role, High Agility Configuration Sizing Chart

spb-6/94-re-Ad8

To illustrate how the design space disappears at lower thrust levels, another sizing chart wasconstructed using a thrust-to-weight ratio of 1.05. Note how specific power constraint of 550 fpsmoves to lower gross weight designs and no design space remains.

Medium Agility

The variation of aircraft gross weight with the significant design parameters and design constraintboundaries is illustrated on the configuration sizing chart on figure 5.13. The chart was constructedwith a thrust-to-weight ratio of 1.1 which allows a small design space between the instantaneous loadconstraint of 9 g's (M = 30,000 tt at M = 0.9) and the specificpower constraint of 550 fps (H = 30,000 ftat M = 0.9). The other twenty-four constraints are all satisfied. The selected configuration lies on theinstantaneous load constraint of 9 g's at an aspect ratio of 3.6. A slightweight reduction is indicated athigher aspect ratios but there is a concern for flutter. Characteristicsof this design are:

Wing loading = 51.5 psfAspectratio = 3.6Span = 62.6 ft.Leading edge sweep = 38 deg.Gross weight = 56,060 lb.

5.2.4 Joint Service Customer

The original intention for showing the impact of customer on the aircraft designs was a two stageapproach. The firststage was to fix the aircraft missionand maneuver performance capability and thengrow the aircraft structurally until it met the structural requirements of a joint service customer. Thesecond stage would then address the impact of carder suitability, such as Launch and Recovery wind-over deck on aircraft size.

The result of the first stage structural growth is shown in Figure 5.14. This figure is a complete side byside comparison of the Air Force and Joint Service Design Weight Breakdowns enforcing thecondition that both aircraft have the same mission and maneuver performance. In general thestructural penalties associated with carrier suitability increased the aircraft empty weights 14 to 17percent and the design takeoff gross weights 11 to 15 percent.

The span of wing panels outboard of the wing fold range from 17 to 25 feet making them difficult tohandle below deck. All the designs except the multi-role designs exceed the 54,500 lb. zeropayload/maximum fuel weight corresponding to the elevator limit required for efficient flightoperations.

The Joint Service Air Interdiction Design already exceeds the 80,000 lb. launch weight of the A-3,largest aircraft to ever operate from an aircraftcarrier.

Increasing the aircraft size further in response to launch, recovery, and single-engine rate-of-climbrequirements is not a feasible approach. Instead, the 54,500 lb. elevator limit was used to define themaximum launch weight of the Joint Service Designs. The basis of comparing the Joint Servicedesigns with their Air Force counterpartswill be missionradius and maneuver performance.

97

6pweJ._R_e

I

OL

,ueqo 6u!z!s uopeJnB!tuoo Xi!I!6V mn!pei/v 'elol:t-!llnl/V "_'1."9 eJn6!_,-!

o!leJloedsv

O't_ 9"g 9"g t,'g g'g O'g

I I I I I I9"0 = IN •

P,000'01. = H •S,69

= peol peu!elsns9"0 = IN •

11000'Og =liB"S= peol peu!e|sns

$9

O00'OS

O00'gg

Q

o)o)

(I)

000'1_gCT

000'9S

000'8g

CO0")

Shee.

DeSkln MissionObseravbles Level

Agility LevelModel NumberService

Takeoff Gross WeightWing Reference Area

Air Interdiction Multi-Role Multi-Role Air-To-Air Air-to-Air

Low Low Moderate Moderate Low Low Moderate IModerat( Low Low

Moderate Moderate Hio_h High Moderate Moderate HicJh High Moderate Moderate988-122 958-122N 988-119 988-1191_ 988-116 958-118N 988-115 1988-115t_ 988-114 988-114_

USAF Joint USAF Joint USAF Joint USAF Joint USAF Joint

UnitsIbs 73145 80910 48801 54704 50899 56947 59549 67397 65230 75312 _

sqft 1463 1618 830 931 1112 1244 1032 1167 1421 1641tt 72 75 57 61 63 65 53 67 72 77

fl 27 - 27 27 27 27Wing SpanFolded Wing Span

Takneff TNV 0.34 0.34 1.10 1.10 1.10 1.10 1.13 1.13 1.13 1.13Takeoff W/S 50 50 59 59 46 46 59 58 45 46

Wing Aspect Ratio 3.50 3.50 3.97 3.97 3.52 3.52 3.80 3.80 3.64 3.64

structuresGroupWingForeplaneHmizim_l Tail

BodyMain Gear

Nose GearAir Induction

Engine SectionYaw Vanes

Ibs 8454 10287 5423 7920 6931 9530 6293 7835 9275 11779Ibs 406 457 634 718

Ibs - 621 7O3Ibs 8988 9437 6488 6512 7618 7999 8303 8718 9057 9510

Ibs 1633 2489 1249 1929 1288 1956 1317 2054 1382 2199Ibs 320 948 306 925 330 989 301 912 313 968

Ibs 320 320 581 581 1145 1145 1088 1088 787 787

Ibs 110 110 275 275 293 293 316 316 335 335Ibs 675 747 294 330 294 329 294 333 374 432

Total Structure Ibs 20500 24338 16026 19229 17899 21270 19167 22676 21523 26010

Propulsion GroupEngines Ibs 2044 2261 3100 3475 3352 3750 3660 4142 3934 4542AMADS Ibs 200 200 200 200 200 200 200 200 200 200

Engine Controls Ibs 40 40 40 40 40 40 40 40 40 40Starl_ System Ibs 80 80 80 80 80 80 80 80 80 80

Fuel System Ibs 1063 1176 1004 1125 1020 1141 1065 1205 1103 1273Vectoring Noz:des Ibs 548 606 1529 1714 1412 1580 1532 1734 1631 1883

Total Propulsion Ibs 3975 4363 5953 6634 6104 6791 6577 7402 6988 5019

Fixed Equipmentmight_APU Ibs

Ine_rumenls Ilos

Hydraulics IbsElectrical Ibs

Avionics IbsArmament Ibs

Furnishings and Equipment Ibs

Air Conditioning IbsAnt_=e U3s

Load and Handling Ibs

Total Fixed Equipment Ibs

1563 1729 1267 1420 1380 1544 1434 1623 1347 1555

210 210 210 210 210 210 210 210 .210 210270 270 270 270 270 270 270 270 270 270518 593 588 682 506 586 485 568 457 546

527 627 618 618 618 618 688 688 690 6901700 1700 1700 1700 1700 1700 1569 1569 1569 1569

85 85 204 204 204 204 242 242 242 242371 386 371 386 371 386 371 386 371 386

640 640 659 659 658 658 713 713 712 71210 10 10 10 10 10 10 10 10 1010 10 10 10 10 10 10 10 10 10

6004 6260 5907 6169 5937 6195 6002 6289 5888 6200

We_t r=lq_ Ibs 30479 34961 27886 32033 29940 34257 31746 36367 34399 40229

Rxed Useful Load

Crew Ibs 215 215 215 215 215 216 215 215 215 215

Craw Equipment Ibs 40 40 40 40 40 40 40 40 40 40

Oil & Trappad Off Ibs 100 111 100 112 100 112 100 113 100 115Trapped Fuel Ibs 456 504 213 239 214 239 360 407 405 468Gun Installat_ Ibs 243 243 252 252 252 252 262 252 252 252

Launchem/E)ectm_ Ib¢,, 980 980 700 700 700 700 760 760 760 760Ammo Cases Ibs 450 450 90 50 50 90 113 113 113 113

Non-Expendable Useful Load Ibs 2484 2543 1610 1648 1611 1648 1840 1901 1685 1963

Operating Wekjnt Ibs 32963 37504 29496 33681 31551 35906 33586 38268 36284 42162

Missies Ibs 9100 9100 4990 4990 4990 4990 1800 1800 1800 1800

Ammo Expandable Ibs 710 710 110 110 110 110 137 137 137 137Fuel Ibs 30372 33596 14205 16023 14246 15941 24026 27192 27009 31164

Design Takeoff Gross Wei_lhl Ibs 73145 80910 48801 54704 50899 56947 59549 67397 95230 75312

Zero Payload, Max Fuel Weight Ibs 71810 - 49714 51957 65597 73512

Figure 5.14

99

Sheet1

Design Mission IObseravbies Level I_t Wt.eval

NumberService

Units

Takeo_ Gross Weight IbsWing Reference Area SClR

Wm9 Span flFolded Wing Span ft

TakeoffTNVTakeoff W/S

w_ _ Ratio

Stxuctures Groupw_ r_Foreplane insHoriziontal Tail Ibs

BoW ,_Main Gear IbsNose Gear IbsAir Induction Ibs

En_line SecOon IbsYaw Vanes Ibs

Total Stmotum Ibs

Propulsion GroupF_.ng_es _osAMADS Ibs

Engine Controls Ibssta_n9system msFualSystem insVectoring Nozzles Ibs

Total Propulsion Ibs

Fixed EquipmentFr_htContro_ IbsAPU I_.Instruments Ibs

Hydraulics IbsElsctdcal insAvionics insArmament Ibs

Fumlshlngs and Equipment IbsAir Conditionin_l Ib_Anti-loe Ibs

Load and Handfin_l IbsTotal Fixed Equipment Ibs

We_ht_

Air Interdict_m Multi-Role MulU.Role Air-To-Air Air.to-AirLow Low Moderate Moderate Low Low _moderate Moderall Low Low

Modend_ Moderatu High HkJh Moderate Wlodera_ High High ,Moderate Moderate988-122 988-122h 988-119 988-1191_ 988-118 988-1181_ 988-115 988-1151_ 988-114 988-114N

Joint I._ Joint _ Joint U_ Joint U_ ,Joint

73145 63600 48801 59490 50899 59490 59549 56300 65230 563001463 1272 830 1012 1112 1300 1032 975 1421 1227

72 67 57 63 63 68 63 61 72 67

27 27 27 27 27

0.34 0.34 1.10 1.10 1.10 1.10 1.13 1.13 1.13 1.1350 50 59 59 46 46 58 58 46 46

3.50 3.50 3.97 3.97 3.52 3.52 3.80 3.80 3.64 3.64

8454 8086 6423 8613 6931 8911 6293 6545 9275 8806408 497 634 599

621 587

8988 9437 6488 6812 7618 7999 ,8303 8718 9057 95101633 1957 1249 2098 1288 2074 1317 1716 1382 1644320 745 308 1005 330 1033 301 762 313 723

320 320 581 581 1145 1145 1088 1088 787 787110 110 275 275 293 293 316 316 335 335675 587 294 358 294 344 294 278 374 323

20500 21242 16026 20241 17899 21799 19167 20609 21523 22128

2044 1777 3100 3779 3352 3918 3660 3460 3934 3395200 200 200 200 200 200 200 200 200 20040 40 40 40 40 40 40 40 40 4080 80 80 80 80 80 80 80 80 80

1063 740 1004 1328 1020 1251 1065 884 1103 773548 476 1529 1864 1412 1650 1532 1448 1631 1408

3975 3314 5.953 7291 6104 7139 6577 6113 6988 5896

1563 1359 1267 1545 1380 1613 1434 1356 1347 1163210 210 210 210 210 210 210 210 210 210270 270 270 270 270 270 270 270 270 270

518 466 588 742 506 612 485 475 457 408627 627 618 618 618 618 688 688 690 690

1700 1700 1700 1700 1700 1700 1569 1569 1569 156985 85 204 204 204 204 242 242 242 242371 386 371 386 371 386 371 386 371 386

640 640 659 659 658 658 713 713 712 71210 10 10 10 10 10 10 10 10 1010 10 10 10 10 10 10 10 10 10

6004 5763 5907 6353 5937 6291 6002 5928 5888 5670

Ibs 30479 30319 27886 33885 29940 35229 31740 32651 34399 33694

F'med Useful LoadCrew Ibs

Crew Equipmem insOil & Trapped Oil insTrapped Fuel IbsGun Installation Ibs

Launchers/Ejectors IbsAmino Cases lbs

Non-Expendable Useful Load Ibs

ope,_ng We_ht

215 215 215 215 215 215 215 215 215 21540 40 40 40 40 40 40 40 40 40

100 70 100 132 100 123 100 83 100 70456 318 213 282 214 263 360 299 405 284243 243 252 252 252 252 252 252 252 252

980 980 700 700 700 700 760 760 760 760450 450 90 90 90 90 113 113 113 113

2484 2315 1610 1711 1611 1682 1840 1762 1885 1734

Ibs 32963 32635 29496 35596 31551 36911 33586 34413 36284 35428

Missies Ibs

Amino Expendable Ibs_uel Ibs

Design Takeoff Gross Weight Ibs

Zero Payload, Max Fuel Weight Ibs

9100 9100 4990 4990 4990 4990 1800 1800 1800 1800710 710 110 110 110 110 137 137 137 137

30372 21155 14205 18794 14248 17479 24026 19950 27009 18935

73145 63600 48801 59490 50899 59490 59549 56300 65230 56300

54500 54500 54500 54500 54500

Figure 5.15

100

Fully Mission Capable

ConfiQuration UnitsAspect RatioWing Reference Area sq ft

LE SWeep degt/c@rootCl_max

Launch Weight Ibs

C-13-1 Endspeed kts

Operating Weight Ibs

Approach Weight Ibs

Powered Approach Stall Speed kts

Arresting Speed kts

Mk7Mod3 Engaging Speed kts

o bility--" Reduced Mission Capa

Carrier Suitability

988-122 988-119 988-118 988-115 988-114

3.5 4. 3.5

1618. 931. 1244.

49. 42. 38.

.08 .05 .05

1.1 1.53 1.72

80910. 54704. 56947.

138. 154. 152.

37504. 33681. 35906.

46504. 42681. 44906.

89. 96. 80.

116. 125. 104.

I 136 141. 138.iiii!!i!ii !i!i i iii i i!i i!!ii!i i iiiiiiiiiiiiiiiiiii}iii i iiiiiiiiiiiii

Confi_luration Units 988-122 988-119 988-11 8Aspect Ratio 3.5 4. 3.5

Wing Reference Area sq ft 1272. 1012. 1300.

LE SWeep deg 49. 42. 38.t/c@root .08 .05 .05CLmax 1.1 1.53 1.72

Launch Weight Ibs 63600. 59490. 59490.

C-13-1 Endspeed kts 148. 151. 151.

Operating Weight Ibs 32635. 35596. 36911.

Approach Weight Ibs 41635. 44596. 45911.

Powered Approach Stall Speed kts 95. 94. 79.

Arresting Speed kts 124. 122. 103.

Mk7Mod3 Engaging Speed kts 143. .............13:9;:.............. 137.=_._:i::;iii=:::.:._iiiii::ii:.._:=_i::_i_i::i!:._::::::!ii:i.:_:::::::iiiiii_iiiiiiii__i_i_iii!ili;iili!iiiil..........................................

Figure_16

Page 1

3.8

1167.

40.

.05

1.63

67397.

146.::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

38268.

47268.

87.

113.

135.

::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

988-115

3.8

975.40.

.051.63

56300.

153.

-..- ........... ..._..... ._.. ........ _.,

34413.

43413.

91.

119.

140.:::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::::

"!'!'!'!'!'!'!'!'!'!'!_'!'_; .! !:i:i!!!!!!!!!!ii_

3.6

1641.

48.

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75312.

142.

42192.

51192.

89.

115.129.

i!iiiiiiiiiiii ii ! i ii!iiiiii!iiiiiii!iiiiii

988-114

3.6

1227.

48.

.05

1.21

56300.

153.

35428.

44428.

96.

124.

139.

(},0 Confiouration and Performance Results

Joint Service Usage

Configuration Issues

The general arrangements of each USAF concept for Models 988-115, -118, and -122/-123 embodybasic features that permit incorporation of Joint Service unique items without voiding the design.These major unique items consist of carder landing gear, arresting hook, and wing fold. Performancepeculiar and mission resizing for zero fuel weight growth will impact size as a function of visibilityrequired, and mission fuel increases required to perform the mission.

USAF Service

Air Interdiction Concept

Conflauration Descdotion

The vehicle type is stipulated as a low observable configuration for both the medium and high agilityperformance conditions, Sizing iterations resulted in a decision to represent both vehicle types inone configuration arrangement with the only principal differences being engine thrust level andmission fuel required.

This vehicle, Model 988-122/-123, is a single-place, subsonic all flying wing design powered by twinlow bypass engines of 13,865 pounds dry thrust each. Externally, the vehicle, shown in generalarrangement drawing ASC988-122-1, is characterized by the moderately swept leading edge at 48.75degrees, lower surface inlet apertures, full span trailing edge elevons, and upper surface thrustvectoring exhaust nozzles. The wing leading edge incorporates large powered slats that are used toachieve critical maneuver conditions.

Control effectors include the yaw thrust vectoring exhaust nozzles with +45 degrees of deflection,upper and low Yaw Vane pairs integrated with the nozzle and lower surface, and four elevons persemi-span. Elevons are single panel at the most outboard and inboard position, with the two mid-spanpanels being split on the wing reference plane.

The interior layout, shown on inboard profile drawing ASC 988-122-2, provides sufficient room for allfunctional systems and features required. Principal features are the deep (approximately 15% t/c)center section for weapons bay, fuel tankage, crew station and equipment installations, include 30mmgun system installation. Basic thickness ratio decreases to approximately 8.5% at the main landinggear and then to 5% in the outboard panel.

The propulsion installation occupies a bay full chord length for each engine. The inlet is pitot typewith a slightly offset diffuser duct.

Exhaust system features for the non-augmented engine consist of the fully offset duct turningthrough a circular bearing/rotation plane to direct exhaust gas through the rotating nozzle exit plane.The exhaust nozzle is a fixed throat SERN type that utilizes the inboard upper surface as anexpansion surface.

Fuel tankage in the outer wing panel is integral. Center section fuel tankage above the weapons bayhas main tank volume allocated as bladder protected "get-home" fuel.

102

U

6 I

%,

A I Rl:q. Able

CENTEI_. |NE

ISTATIC GRO. L)NE

@)4).4 ft

2)24 --

O_GIP,:._L FP,_GE IS

OF POOR'QUALITY

6

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I_"s """'"" r-t-i]

103

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99

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FOL'I_OUT F_A -N_._,! t_'_s'_

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.+

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SYSTEM b& R SI_

INLETAPERTURF"

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SYSTEMBAY

"GEN 6" [NOINE

VI[W C /

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ROTAT I N( _ NOZZLEEXHAUST SYSTEM ¥1_ 0

A +',,,"

= •

Ju,..l_,l, o,,• im tree I_I =+ C_,_.a• i • i .v,

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104

4 I

141S0 IDe

A p_.. C _.. ..(._rZONT/U_ KEFEmENC[ L.I_

I"' r" I'" r"

Z _% 200 Y

L

GROUP WBGHT STATEMENT

MISSION: Air-to-Ground

MODEL: 988-123 HiAJLO

W_G

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENGINE SECTION

YAW VANES

TOTAL STRUCTURE

E]_31_ES

AMADS

ENGINE CONTROLS

STARTING SYSTEM

FUEL SYSTEM

VECTORING NOZZLES

TOTAL PROPULSION

FUGHT CONTROLS

APU

INSTRUMENTS

HYDRAULICS

ELECTRICAL

AVIONICS

ARMAME]_T

FURNISHINGS & EQUIPMENT

AIR CONDITIONING

ANTI-ICE

LOAD AND HANDLING

TOTAL FIXED EQUIPMENT

WE_d-fTBWPTY

CREW

CREW EQUIPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHERS/EJECTORS

AMMO CASES

NON-EXP USEFUL LOAD

WEIGHT

(LB)

8454

8988

1633

320

320

110

675

20500

2044

200

40

80

1063

548

3975

1563

210

270

518

627

1700

85

371

640

10

10

6004

30479

215

4O

100

456

243

980

450

2484

OPF_F_NGW/_3/-r/" 32963

M_SLES

AMMO EXPENDABLE

FUEL

9100

710

30372

GROSS _ 73145

NOSESTATION

_NGMAC

LEMAC

_DYLENG_

0IN

320 IN

159 IN

518 IN

BODY STATION

320

250

302

69

142

291

395

283

PERCENT MAC

291

218

195

248

274

383

293

371

380

105

296

184

168

120

180

172

172

286

239

276 36.4%

100

100

253

274

170

299

175

237

273 35.5%

306

175

274

275 36.1%

SpreadSheet: 988-123 W'l-. STATEMENT105 5/6/94

C)

8E:L _6/61/_0

9L'LG- O0"OOI- 00"0 00"0

8L'6L6- 00"0 00"0000_ OG'I

_g'I- 00"0 00"0000_ O_'I

_9"gg8- 00"0 00"0000_ O_'I

8_'E 00"0 00"0000_ 0_'I

O0"g_ 00"09 00"0 ES"O

8_'L911 00"00_ 00"0 00"0

00"6 00"6 00"0 00"0

09"9 0_'9 00"0 00"0

90"_ OE'I 00"0 00"0

i_n_ov a_!nb_ _ _pn_T_IV qo_H

00"0 89"0

00"0 08"0

00"0000_ 08"0

00"0 08"0

00"0000_ 08"0

00"0 _'0

00"0000_ 00"0

00"0 08"0

00"0 08"0

00"0 00"0

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00"I

00"I

00"I

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00 I

00 1

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< ....... leu!_ ....... >< ..... i_qIuI ...... >

00"0 00"000II 09"0

00"0 00"0gII 00"I

00"0 00"0gII 09"0

00"0 00"000II 00"I

00"0 O0"O00II 00"I

00"0 00"00011 09"0

00"0 O0"O001I 09"0

00"0 00"00011 09"0

00"0 O0"O00II 09"0

00"0 00"0 00"0

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II

GROUP WEIGHT STATEMENT

MISSION: Air-to-Ground

MODEL: 988-122 MedA/LO

WNG

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENGNE SECTION

YAW VANES

TOTAL STRUCTURE

ENGNES

AMADS

WEIGHT

(LB)

8454

8988

1633

320

320

110

675

20500

ENGINE CONTROLS

STARTING SYSTEM

FUEL SYSTEM

VECTORING NOZZLES

TOTAL PROPULSION

FLIGHT CONTROLS

APU

INSTRUMENTS

HYDRAULICS

ELECTRICAL

AVIONICS

ARMAMENT


AIR CONDITIONING

ANTI-ICE

LOAD AND HANDLING

2044

200

40

80

1063

548

3975

1563

210

270

518

627

1700

85

371

640

10

10

TOTALRXEDEQUIPMENT 6004

t4Fc/CV_EMPTY 30479

215

40

100

456

243

980

450

CFEW

CREW EQUIPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHERS/EJECTORS

AMMO CASES

NON-F_XP USEFUL LOAD 2484

OPERARNGHFc/GHT 32963

9100

710

30372

MLSSEES

AMMO EXPENDABLE

FUEL

__ 73145

NOSE STATION

WING MAC

LEMAC

BODY LENGTH

BODY STATION

320

250

302

69

142

291

395

283

291

218

195

248

274

383

293

371

38O

105

296

184

168

120

180

172

172

286

239

276

100

100

253

274

170

299

175

237

273

306

175

274

275

0 IN

32O IN

159 IN

518 IN

PERCENT MAC

36.4%

35.5%

36.1%

SpreadSheet: 988-122 WT. STATEMENT 1 1 1 5/1 1/94

n-On"

I--"l-

uJ

Nose @ BS 0LEMAC @ BS 159.37MAC Length --- 320.31 in.

AC @ 35.94% MAC

Model 988-122 MedAJLO

85000

80000

75000

70000

65000

60000m

55000

50000

45000

40000

350O0

30000

0.31

E;asic Air to-Groul;d Missk)n..._-,.,.

\\ t

'Wing FL el

i

E (,__GBU-27J3 _--"

- ,.o , . - - •Wlr g Fuel

'_ ""Aft Bo, ly Fuel e•,21 ,_

A:_ ".

:wd Body Fuel ....... -I...... j

0.315 0.32 0.325 0.33 0.335 0.34 0.345 0.35 0.355 0.36 0.365 0.37 0.375 0.38

C.G. LOCATION - %MAC

Chart: C.G. CHART 988-122 5/1 1/94

Inertia Data at Combat Weiqht

Parameter

Combat Weight

Longitudinal C.G. (Body Sta)

Vertical C.C. (from static ground line)

Ixx Roll Inertia

lyy Pitch InertiaIzz Yaw Inertia

ixz Product of Inertia

A/G Model

Units 988-122

MedA/LO

Ibs 58506

in. 274

in. 80

slug-ft^2 182307

slug-ft^2 92317

slug-ft^2 293526973slug-ft^2

113

6.2 Air Superiority Concepts

Model 988-115, High Agility - Moderate Observables

The high agility, moderate observables vehicle, Model 988-115, is a single place, three-surfacesupersonic design powered by two turbojet engines of 33,660 pounds augmented thrust e_ch.Externally the vehicle general arrangement, shown on drawing ASC 988-115-1, includes a liftingcanard or foreplane ahead of the main wing and a horizontaltail aft of main wing.

Each surface (wing/canard/tail), is of identical planform with forty (40) degrees leading edge sweep.The canard and tail are identical plan areas and the canard is set at +10 degrees dihedral, with the wingand tail set at -5 degrees relative to the horizontal reference plane.

Inlets are integrated/nested with the lower forebody, inboard of canard deflection path. Exhaustnozzles are located side-by-side on the upper aft fuselage and Yaw Vane pairs are integrated with thenozzles and on the lower aft body.

Control effectors include the yaw thrust vectoring exhaust nozzles, with _+40degrees of deflection,the Yaw Vane pairs above and below aft fuselage, and main wing trailing edge plain flaps, in additiontothe canard and horizontal tail.

Initial sizing optimizations for the high agility metric conditions resulted in main wing size and aspectratio which established overall span at a size that was considered impractical to achieve in a high agilityfighter. The approach taken was to extract the equivalent horizontal tail exposed area from thetheoretical main wing and incorporate a lifting canard/foreplane. This arrangement replicates thatcurrently in use on the F-15/SMTD research vehicle.

The interior layout, shown on inboard profile drawing ASC 988-115-2, accommodates the crew,subsystem, weapons and propulsion system volume allocations within a low profile body shape. Theforebody is conventional in arrangement and includes avionics, crew station, gun system, andavionics subsystem. Center body contents are main fuel tanks, inlet system, weapons bay, and mainlandinggear. The aft body provides engine and exhaust system accommodation.

Propulsion system installation features consist of the nested external compression fixed ramp inlet,long vertical offset inlet diffuser running over the weapons bay to engine face.

The exhaust system includes augmentor spray bars, fully offset duct to nozzle exit plane. The ductturns through a circular bearing/rotation plane to direct exhaust gas out the nozzle aperture. Asignificant and challenging risk issue is presented here in this concept of making the rotating nozzlesystem augmentor capable. A discussion of this issue is contained in Section 7.0, Areas of HighTechnical Risk.

Nozzle concept is that of a variable throat SERN type that utilizes the upper aft deck as the expansionsurface.

Fuel tankage in the main wing panel is integral and center section tankage contains fuel inconventional bladder cells.

114

6

AIR_JU_C['q TERL I

CAJ_A.qO/F OR (I:_. AN( i

AT .10 _ DIHE{_I,_I. VINO • HORIZO_IT_M. TAIL

HRL - vl. 200 _ /_ I I _

STMnC _0. LINE I_11 I

,I I'1 d

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115 ]

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t I

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5

1

GROUP WPlGHT STAllEMENT

MISSION: Air-to-Air

MODEI_ 988-115 HiA/ModLO

W_GHORIZONTAL TAIL

YAW VANES

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENGINE SECTION

FOREPLANE

WEIGHT

(LB)

TOTAL STRUCTURE

E_3NES

AMADS

ENGINEc rmOLSSTARTING SYSTEM

FUEL SYSTEM

VECTORING NOZZLES

TOTAL PROPULSION

FUGHT CONTROLS

6293

621

294

8303

1317

301

1088

316

634

19167

3660

200

40

8O

1065

1532

6577

APU

INSTRUMENTS

HYDRAUUCS

ELECTRICAL

AVIONICS

ARMAMENT


AIR CONDITIONING

ANTI-ICE

LOAD AND HANDLING

TOTAL F/XED EQU/Ph'EgVT

1434

210

270

485

688

1569

242

371

713

10

10

6002

WE/GF/TB_/PTY 31746

215

40

100

360

252

760

113

1840

CFE_N

CREW E(_IPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHERS/EJECTORS

CASES

NON-EXP USEFUL LOAD

OPERARNGWE/GHT 33587

1800

137

24026

MISSILES

AMMO EXPENDABLE

FUEL

G_/OSSV___F/T 59550

NOSE STATION

WING MAC

LEMAC

BODY LENGTH

BODY STATION

550

760

705

468

549

212

410

625

312

504

625

550

389

580

504

730

626

585

680

158

587

436

200

228

264

410

120

468

396

509

153

153

590

504

228

460

228

387

502

460

228

504

500

0 IN

242 IN

434 IN

806 IN

PERCENTMAC

30.8%

28.1%

27.4%

SpreadSheet: 988-115 Wt. Statement 1 17 5/6/94

CO

62000

57000

52000

47000

w 42000

37000

QO

q _ 32000

Nose @ BS 0LEMAC @ BS 434.17MAC Length = 241.76 In.AC @ 30.0% MAC

C.G. MOVEMENT

Model 988-115 HIA/Med LO

E'-1d<52

OIJ.

Wing + Bo,

At

t

I

I

I

|y FueJI

=

i

I

I

A,

Payload

/ Basic A

f

&,&- -,&

i

, Vling + BodyL

L

L

Jr-to-Air Mi

;uel

3slon

E

d2<

<

0.22 0.23 0.24 0.25 0.26 0.27 0.28

C.G. LOCATION - % MAC

0.29 0.3 0.31 0.32 0.33

Chart: C.G. CHART 5/6/94


A/A Model

Parameter Units 988-115

HiA/ModLO

Combat Weight Ibs 47540

in. 498Longitudinal C.G. (Body Sta)

Vertical C.C. (from static ground line)Ixx Roll Inertia

in.

slug-ft^2

87

84951

lyy Pitch Inertia slug-ft^2 240255

Izz Yaw Inertia slug-ft^2 329116

Ixz Product of Inertia slug-ft^2 2137

119

//gandalf/user/deb9848/agility_dir/aa_dir/988 - 115_weight s

Weight Statement Weight(LBS)

Airframe Structure

Fuselage ........................ 9117Wing ............................ 5144Canard .......................... 0Horizontal Tail ................. 618

Vertical Tail(s) ................ 205

Engine Mounts ................... 300Inlet(s) and Duct(s) ............ 1192

Exhaust Duct(s) ................. 0Pivots .......................... 0

Main Landing Gear ............... 1260Nose Gear ....................... 309

Total 18145

Propulsion System

Engine(s) and Nozzle(s) ......... 5731.Engine Start and Control ........ 120.Fuel Tanks ....................... 280.

Fuel Pumps ...................... 91.Fuel Distribution System ........ 531.

Air-Refueling System ............ 75.Fuel Inerting System ............ 75.Gear Box and Accessories ........ 200.

Total 7104.

Fixed EquipmentInstruments ..................... 270.

Surface Controls ................. 1433.

Crew Accomodations .............. 371.Armaments ....................... 1012.

Avionics ........................ 1569.

Electrical System ............... 622.Hydraulics and Pneumatics ....... 391.

Radar Absorpton Material ........ 0.

Auxiliary Power System .......... 210.Airconditioning and De-Icing .... 584.

Total 6461.

Empty Weight .................... 31710.

Operational Items"Crew ............................

Trapped Fuel and Oil ............Gun and Provisions ..............

Operational Empty Weight ........

PayloadAmmunition ......................Air-to-Air Missles ..............

Air-to-Ground Munitions .........

Total

Mission Fuel

Wing Fuel ...........

Body Fuel ...........External Fuel .......

Design Gross Weight .............

WeightFraction

0.1549

0 0874

0 00000 0105

0 00350 0051

0 0203

0 00000 0000

0 02140.00530.3083

0.09740.0020

0.00480.0016

0.0090

0.0013

0.00130.0034

0.1207

0 0046

0 0244

0 00630 0172

0 0267

0 01060 0066

0 0000

0 0036

0.00990.1098

0.5389

255. 0.0043

441. 0.0075

365. 0.0062

32771. 0.5569

Volume

(cuft)

91

510

62

3

827

047

15306.

4.

2.32.

2.

13.2.

2.

5.

62.

7 .

36.70.

34

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07

29

238

0

4.

9.

7.

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CG

(ft)

39 4044 89

0 000 00

0 00

62 7060 51

0 00

0 0040 56

25.3240.79

62.70

36.9838.27

38.27

38.2738.27

38.2762.70

58.33

28.12

38.27

22.43

38.2726 86

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35 180 00

0 00

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42 30

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0.00

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0. 0.0000 240. 38.271937. 0.0329 340. 0.00

12645. 0.2149 260.

11492. 0.1953 236.

0. 0.0000 0.

58845. 1.0000 2091.

0.000.00

38.27

07/16/94

Moment

(ft/ib)

359236

230929

00

018816

72155

00,

51113.

7829.740077.

359338

221910729

3494

203222870

2876.12540.

414388.

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8321.38729.

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R.Engelbeck

7/16/94. I

REVISED

DATE

Z ! JOH J(po8 I o_,°l

988-1 i5 High Agility, Moderote Observobles

Torget Cross Sectionol Areo Distribution

THE BOEINOCOMPANY PAGE

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6

0.0

6 6 6 c; 6 6 o 6 6 6 6 6 6

03

ilZ

ng., ,I'81 411 ,SEOOATEICHECK I

APPD. I

APPD.

"]3

988-115 High Agility, Moderate Observables

Target Cross Sectional Area Distribution

THE BOEINGCOMPANY ,,CE123

07/16/94 12:20 PM//gandalf/user/deb9848/agility_dir/aa_dir/988-115_mission

Design Mission Segment Performance Breakdown

Initial Final Fuel Time Range Mach Altitude CL CD

Weight Weight Burned (min.) (n.mi.) (feet)

Warm-up and taxi58845. 58150. 695.1 20.00 0.0 0.000 0 0.257 0.0168

Warm-up and taxi58150. 53698. 4452.3 2.00 0.0 0.300 0 0.257 0.0168

Acceleration from Mach 0.300 to Mach 0.900

53698 53023. 675.0 0.27 1.8 0.900 0 0.047 0.0114

Climb from 0.0 ft. to 44262.7 ft. at 100.8 ft/sec

53023 51821 1201.4 3.49 27.9 0.844 44263 0.257 0.0189

Cruise at Mach 0.845

51821 49543 2278.4 39.72 320.3 0.845

Loiter at 40000. ft and max L/D of 13.89

49543 44605 4937.6 90.00 687.7 0.800


44605 43788 817.2 0.88 9.6 1.500


43788 41638 2149.6 6.31 90.4 1.500

One Combat Turn at 8.2 deg/sec and 4.0 g's

42318 41105 533.3 0.73 6.3 0.900


41785 40579 526.0 0.72 6.2 0.900


41009 40060 518.9 0.71 6.1 0.900


40131 39548 511.8 0.70 6.0 0.900 40000 0.739 0.1000


39548 39310 238.1 1.64 12.9 0.829 49739 0.310 0.0222


39310 36869 2441.2 55.51 437.1 0.820 49815. 0.325 0.0230


36869. 35858 1011.5 20.00 67.8 0.307 0. 0.257 0.0168

Total Mission Fuel = 22987. ibs Reserve Fuel = 1149. ibs

44994 0.323 0.0226

40000 0.263 0.0190

40000 0.073 0.0284

40000 0.068 0.0276

40000 0.739 0.1000

40000 0.739 0.I000

40000 0.739 0.I000

Power Net Fuel Error

Setting Thrust Flow Code

0.040 1713. 2085.

2.000 70292. 133570.

2.000 83174. 172232.

2.000 11680. 11348.

0.387 3538. 3407.

0.305 3387. 3248.

2.000 36153. 70381.

0.961 17345. 20426.

2.000 22658. 43870.

2.000 22658

2.000 22658

2.000 22658

2.000 7554

0.378 2695

0.055 2370

43870.

43870.

43870.

7394.

2592.

3020.

//gandalf/user/deb9848/agility_dir/aa_dir/988-115_perf 07/16/94 12:24 PM

26 WEIGHT

xNegPs 48510.23CCEL 47698.18

axRC 48510.23

WFuel 55267.35

LOAD2 48510.23

LOAD3 48510.23

LOAD4 48510.23

LOAD1 48510.23

LOAD2 48510.23

LOAD3 48510.23

LOAD4 48510.23

LOAD5 48510.23

LOAD6 48510.23

LOAD7 48510.23

S1 48510.23

$2 48510.23

$3 48510.23

$4 48510.23

54 55267.35

$6 48510.23

$7 48510.23

$8 48510.23

$9 48510.23

SI0 48510.23

SII 48510.23

S12 48510.23

%Fuel Payload

0 60 1120.00

0 60 1120.00

0 60 1120.00

0 00 1120.00

0 60 1120.00

0 60 1120.00

0 60 1120.00

0 60 1120.00

0 60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

0.60 1120.00

DeltaF

0.00

0.00

0.00

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0 O0

0.00

0.00

0.00

0.00

0.00

0.00

0.00

< ...... Inital ..... >< ....... Final ....... >

Pset

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2.00

2 O0

2 O0

2 O02 O0

2 O02 O0

2 O0

Mach Altitude

0.60 15000.00

0.90 40000.00

0.60 50000.00

0.60 50000.00

0.90 20000.00

0.90 30000.00

1.20 30000.00

0.60 10000.00

0 90 i0000.00

0 60 20000.00

0 90 20000.00

1 20 20000.00

0 90 3OOO0.O0

1 20 30000.00

0 6O 0.00

0 90 0.00

0 60 i0000.00

0 90 10000.00

1 20 10000.00

0 60 20000.00

0 90 20000.00

1 20 20000.00

1 40 20000.00

0 90 3OO0O.00

1 20 30000.00

1 40 30000.00

Mach

0 00

1 50 40000.00

0 00 0.00

0 00 0.00

0 00 0.00

0 00 0.00

0 00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

Altitude Require Actual0.00 -i00.00 -53.41

60.00 51.60

500.00 6639.37

500.00 21080.20

9.00 9.00

9.00 9.00

9.00 9.00

6.00 6.40

9.00 9.00

4.40 4.40

7.00 8.32

6.80 7.90

5.00 5.49

5.00 5.83

900.00 968.22

1300.00 1442.85

650.00 716.27

1000.00 1143.51

600.00 998.04

450.00 484.50

800.00 840.16

650.00 781.47

600.00 811.69

550.00 554.71

500.00 628.82

600.00 752.68

..A

_OOn

Model 988-114, Medium Agility - Low Observables

The medium agility, low observables vehicle, Model 988-114, is a single place tail-less supersonicdesign powered by two afterbuming low-bypass turbofans of 36200 pounds (augmented) thrusteach. It is capable of Mach 1.5 on un-augmented engine thrust. Low observable characteristicsinclude low sideslope angles, long planform outline edges, edge alignment, lack of any vertical tailsurfaces, and inlets integrated into the wing-body junction.

The wing planform was chosen to allow some forward sweep on the trailing edge while maintaining thedesired (reference) aspect ratio of 4. The tip was "beveled" to alleviate undesirable aerodynamic,structural and RCS effects.

Inlets are canted F-22 type, with angles chosen to integrate with the leading edge while meeting sideslope and inlet ramp angle requirements. Placement at the wing-body junction results in the intakeduct passing alongside rather than over the weapons bay. Yaw vectoring exhaust nozzles are locatedon the upper aft fuselage; their fairing widths determine the thrust centerline spacing.

Control effectors include those on the trailing edge of the wing, the yaw vectoring nozzles, "yawvanes" forming the forward part of the nozzle fairings (and on the underside of the aft fuselage), andaft body flaps to provide thrust vectoring in pitch. There are large leading edge slats to enhancemaneuvering. The mid-outboard elevons are splitto act as drag rudders.

The interior layout, as shown by the Inboard Profile (ASC988-114-2) is conventional for tactical aircraft,with the exception of the internal weapons bay (side-by-side missiles) and exhaust nozzlearrangement. Most the fuel is contained in the large integral wing tanks, with a smaller, protected tankabove the weapons bay for balance.

The exhaust system includes full augmentation, and a rotating nozzle with variable throat and exitplane areas; an alternate aft body integration scheme is shown (versus 988-115 or -118). In the cruiseposition, the aft body flaps provide a "SERN" expansion surface; at any substantial vector angle, thenozzle must act as a 2D-CD nozzle. Achieving acceptable efficiencies and effective vectoring is asignificant technical risk (see Section 7.0).

126

4 T

FO

OF POOR

...... :',T._,I°_._..','*.=''_

L, s 4 I

127

2

ME;=301.6Oln

MID SPLIT E_

39, 28qf ?./s I Oe

INSDELEVON

MOOEL x

l_¢I too. [.alX Vm_RE L-------

?_ENERAL ARRANGEMENT

MODEL 9B8-114

_,[R SUPERIORITY CONFIG

i¢ic¢ S/de IPlfl l[ _'_ I

1

•

VIE_ A

VIEW D

_n_urT

tit

4 I 3

128

_ I0 2 I

ilal i i_t I_IOlI iII t lilt i

lINE II_(I_TII_I @DIITAIIMI@ NIIIIIIII III II_IIILTIi'V 1FI Till Illlqllll @ll_ll_

l_lll4OmlllYlt)ll I II II1111 C_sPI. I

o

.%.___.__.E."__ --.j __-.-...._.......

o 2 I

D

C

B:

GROUP WBGHT STATEMENT

MISSION: Air-to-Air


W_G

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENG_qE SECTK_N

YAW VANES

WEIGHT

(LB)

9275

9057

1382

313

787

335

374

TOTAL STRUCTURE

ENGNES

AMADS

ENGINE CONTROLS

STARTING SYSTEM

FUEL SYSTI_I

VECTORING NOZZLES

TOTAL PROPULSION

FUGHT CONTROLS

APU

INSTRUMENTS

HYDRAULICS

ELECTRICAL

AVIONICS

ARMAMENT


AIR CONDITIONING

ANTI-ICE

LOAD & HANDLING


_EMPTY

ClEW

CREW EQUIPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHERS/EJECTORS

AMIVlO CASES

NC_EXP USEFUL LOAD

21523

3934

2OO

4O

8O

1103

1.631

6988

1347

210

270

457

69O

1569

242

371

712

10

10

5888

34399

215

40

100

405

252

760

113

1885

OPERARNGWE/GHT 36284

BOMBS_IlSSlLES

AMMO EXPENDABLE

FUEL

GROSSV_GNT

1800

137

27009

65230

NOSE STATION

WING MAC

JEMAC

BODY I_B'qGTH

0 IN

303 IN

344 IN

725 IN

BODY STATION

49O

419

461

140

360

5O9

604

451

PERCENT MAC

509

434

320

464

444

626

522

543

580

135

485

369

152

210

219

300

90

419

332

445 33.3%

130

130

474

444

210

340

210

316

438 31.1%

337

210

444

437 3O.6%

SpreadSheet: 988-114 Wt. Statement 129 5/11/94

8

W

70000

65000

60000

55000

50000

45000

40000

35000

30000

23.0 %

Nose @ BS 0LEMAC @ BS 344MAC Length = 303 In.AC @ 30.4% MAC

Model 988-114 MedA/LO

dJ

d).o

i

Ba=ic Air-to-Aim

\

"iPayload

.-'" W;i-,§ F

"" _Ning F

,dy Fuel

•A

Jal

°m

.E

O

2<

<

24.0 % 25.0 % 26.0 % 27.0 % 28.0 % 29.0 %


30.0 % 31.0 % 32.0 % 33.0 %

Chart: C.G. CHART 5/1 1/94

Inertia Data at Combat Weiaht

A/A Model


Med A/LO

Combat Weight Ibs 51 730in. 429Longitudinal C.G. (Body Sta)

Vertical C.C. (from static ground line) in° 79

Ixx Roll Inertia slug-ft^2 123597

iyy Pitch Inertia slug-ft^2 195819

izz Yaw Inertia slug-ft^2 352451

Ixz Product of Inertia slug-ft^2 2002

131

//gandalf/user/deb9848/agility_dir/aa_dir/988-114_geom

Aircraft Geometry

Thrust-to-Weight =

Takeoff Gross Weight =

Wetted Area =

1.13 Wing-Loading =

63309.6 Reference Area =

3623.7 Swet/Sref =

44.3

1429 .i

2.54

07/16/94

Body Geometry

Fineness Ratio = 8.80

Length = 60.30 Width =

Wetted Area = 1347.0 Volume =

Wing Geometry

11.83

1222.1

Area = 1429.1

Aspect Ratio = 3.64

Span = 72.12

Mean t/c = 0.05

Sweep Angles

Leading Edge = 47.70

Quarter Chord = 39.50

Trailing Edge = 0.01

NOTE: ARPITCH = 6.73, ARWE=

Wetted Area = 1998.3

Taper Ratio = 0.00

Mean Aero Chord = 26.42

3.64 Wing STABLE in Pitch at High Angles-of-Attack

Vertical Tail Geometry (each)

Number of Vertical Tails = 2.

Area = 69.6

Aspect Ratio = 1.70

Span = 10.87

Mean t/c = 0.05

Sweep Angles



Trailing Edge = -7.04

Engine Geometry

Engine Scale = 0.9028

Engine Diameter = 33.37

Sea-Level Static Thrust = 35769.9

Engine Weight = 3096.6

wetted Area = 139.2

Taper Ratio = 0.I0


Capture Reference Area = 7.54

Nozzle Base Drag Reference Area = 60.76

132

//gandalf/user/deb9848/agility_dir/aa_dir/988-114_weights 07/16/94 II :19 AM

Weight Statement Weight Weight Volume(LBS) Fraction (cu ft)

Airframe Structure

Fuselage ........................ 8941•Wing ............................ 8595.Canard .......................... 0.

Horizontal Tail ................. 0.

Vertical Tail(s) ................ 375.

Engine Mounts ................... 324.Inlet (s) and Duct (s) ............ 1265.Exhaust Duct (s) ................. 0.Pivots .......................... 0.

Main Landing Gear ............... 1356.Nose Gear ....................... 333

Total 21189

Propulsion System

Engine(s) and Nozzle(s) ......... 6193Engine Start and Control ........ 120Fuel Tanks ...................... 291

Fuel Pumps ...................... 97

Fuel Distribution System ........ 547

Air-Refueling System ............ 75Fuel Inerting System ............ 77Gear Box and Accessories ........ 200

Total 7600

Fixed EquipmentInstruments ..................... 270Surface Controls ................ 1157

Crew Accomodations .............. 371Armaments ....................... 1012

Avionics ........................ 1569

Electrical System ............... 622Hydraulics and Pneumatics ....... 423

Radar Absorpton Material ........ 0

Auxiliary Power System .......... 210

Airconditioning and De-Icing .... 628Total 6261

Empty Weight .................... 35050

0.14120.1358

0.00000 0000

0 00590 0051

0 02000 0000

0 00000.0214

. 0•0053

• 0•3347

0.09780.0019

0.0046

0.00150.0086

0.0012

0.00120.0032

0•1200

0 0043

0 01830 0059

0 01600 0248

0 0098

0 00670 0000

0 0033

0 0099

0 0989

0 5536

Operational Items"Crew ............................ 255.

Trapped Fuel and Oil ............ 475.Gun and Provisions .............. 365.

0.00400•0075

0.0058

Operational Empty Weight ........ 36145• 0•5709

Payload"Ammunition ...................... 137.

Air-to-Air Missles .............. 1800.

Air-to-Ground Munitions ......... 0.

Total 1937.

Mission Fuel

Wing Fuel ........... 24138.Body Fuel ........... 1089.External Fuel ....... 0.

0.0022

0.0284

0.00000.0306

0.3813

0.01720.0000

Design Gross Weight ............. 63310• 1.0000

89.

86.0.0.

4.

385

80

53

16344

4

233

2

142

2

5

64

7 •

29.

70.34.

30.

16.ii.

0.

7.

31.234.

0.

4.

9.

7.

0 .

5 .

95.240.

340.

496.22.

0.

1793•

CG

(ft)

35.27

46.26

0.000.00

0.0055.68

53.41

0.000.00

36.4722.45

40.37

55.68

33.4734 37

34 37

34 3734 37

34 3755 68

52 02

25•80

35.6021.06

34.37

24.1220.55

31.96

0.00

0.0011.18

25.86

40.31

11•25

45.03

0.00

0.00

0.00

0.00

34.37

0.00

0.000.00

34.37

Moment

(ft/ib)

315304.

397608•0.

0.

0.18056.

67587.0.

0_

49450.7470.

855475•

344829

200810013

3323

188152578

264111137

395344

6966

411827813

3478337838

12779

135140

0

7020

161896

2869.21378•

0.

0 •

0 .

O.O.O.

0 .

O.

133

I CALC

CHECK J

APPD. I

APPD. I

NO

o-00.-0o

J#

/

/i

\ ,/_

\

_- tO tO u3

REVISED

/

S

d/

J

//

o01.tO

P-

.°

OOp,.

,o

00_0

0U_u_

0

o

0

0

o 5°-

o

.o_

0_0_3

0

D_

0_0C_

0

0.0

THE BOEINGCOMPANY P,0E134

988-114 Med. Agility, Low Observables


Z ! JO H 6U !M /[poll 101ol

\ .o

"OI

,,¢_C) 0 0 0 _ {D C) 0 0

CO II 0 I! C.) II C.) II . 0 II CJ II 0 II _.) II 0 II C..)

C%1

0

_b

0

CO

6

r_

0

0

4.-

0

(3o

0

.0

0

_0

6

.0

O

,O

OO

O

0164

CO

"5-D

CAhC R.Engelbeck

lO

988-114 Med Agility, Low ObservoblesCruise Drog Polor

THE BOEINGCOMPANY "'°_ls5

//gandalf/user/deb9848/agility_dir/aa_dir/988-114_mission


Initial Final Fuel Time Range

Weight Weight Burned (min.) (n.mi.)

Warm-up and taxi

63310. 62558. 751.2 20.00 0.0 0 000

Warm-up and taxi62558. 57747. 4811.4 2.00 0.0 0 300


57747. 57020. 727.3 0.27 1.7 0 900


57020. 55515 1504.2 4.63 38.9 0 885 49699


55515. 53359 2156.4 36.78 309.3 0 876 49698


53359. 48190 5168.6 90.00 687.7 0 800 40000


48190 47330 860.4 0.86 9.4 1 500 40000


47330 44899 2430.8 6.32 90.6 1 500 40000


45579 44434 465.0 0.59 5.1 0.900 40000


45114 43974 460.0 0.58 5.0 0.900 40000


44404 43519 455.0 0.58 5.0 0.900 40000


43590. 43069 450.1 0.57 4.9 0.900 40000

-_ Climb from 40000.0 ft. to 49981.7 ft. at 75.0 ft/sec

CO 43069. 42805 263.9 1.76 13.3 0.799 49982

O_ Cruise at Mach 0.780

42805. 40313. 2492.2 57.91 436.7 0.780 49689. 0.276 0.0177


40313. 39281. 1032.2 20.00 64.4 0.291 0. 0.222 0.0132

Total Mission Fuel = 24029. ibs Reserve Fuel = 1201. ibs

Mach Altitude CL CD

(feet)

0. 0.222 0.0132

0. 0.222 0.0132

0. 0.036 0.0091

0.224 0.0155

0.287 0.0187

0.202 0.0140

0.056 0.0225

0.052 0.0221

0.694 0.0768

0.694 0.0768

0.694 0.0768

0.694 0.0768

0.261 0.0171

Power Net

Setting Thrust

0.040 1851. 2254.

2.000 75961. 144342.

2.000 89882. 186123.

2.000 10507. 10422.

0.435 3543. 3457.

0.294 3527. 3410.

2.000 39069 76057.

1.003 19559 23025.

2.000 24486 47409.

2.000 24486 47409.

2.000 24486 47409.

2.000 24486 47409.

2.000 7858 7609.

0.356 2668. 2527.

0.051 2370. 3083.

Fuel Error

Flow Code

0

0

0

0

0

0

0

0

0

0

0

0

0

0

0

07/16/94 11:17 AM

//gandalf ./deb9848/agility_dir/aa_dir/988-114_perf 07/I 11:21 AM

26

MxNegPsACCEL

MaxRC

dWFuelILOAD2

ILOAD3ILOAD4

SLOADISLOAD2

SLOAD3

SLOAD4

SLOAD5SLOAD6

SLOAD7

PSIPS2

PS3

PS4

P54PS6

PS7

PS8PS9

PSI0

PSI1

PSI2

..K

Co-4

WEIGHT %Fuel Payload

52540.09

51691.4252540.09

55267.35

52540.0952540.09

52540.09

52540.0952540.09

52540.09

52540.0952540.09

52540.09

52540.0952540.09

52540.09

52540.0952540.09

55267.35

52540.09

52540.0952540.09

52540.09

52540.0952540.09

52540.09

0.600.60

0.60

0.00

0.600.600 60

0 60

0 60

0 600 60

0 600 60

0 6O

0 600.600 600 600 600 600 600 6O0 6O0 600 600 60

1120.00

1120.00

1120.001120.00

1120 001120 00

1120 00

1120 001120 00

1120 00

1120 001120 00

1120 00

1120 00

1120 001120 00

1120 00

1120 001120 00

1120 00

1120 001120 00

1120 00

1120 001120.00

1120.00

DeltaF

0.60

0.000.00

0.00

0.000.00

0.00

0 000 00

0 00

0 000 00

0 00

0 000 00

0 00

0 000 00

0 00

0 00

0 000 O0

0 000 00

0.00

0.00

< ...... Inital ..... ><

Pset Mach Altitude

2.00

2.002.00

2.00

2.002.00

2.002.00

2 00

2 002 00

2 00

2 002 00

2 00

2 002 00

2 00

2 00

2 002 00

2 002 00

2 00

2 00

2 00

0.60

0.90

0.600.60

0.90

0.90I .20

0.600.90

0.60

0.90

1.200 90

1 20

0 600 9O

0 60

0 901 20

0 6O

0 9O1 20

1 40

0 90

1 201 40

15000.00

40000.00

50000.0050000.00

20000.003O0OO.0O

30000.00

10000.00I0000.00

20000.00

20000 0020000 00

3O0OO 00

30000 000 00

0 00

I0000 0010000 00

I0000 00

20000 0020000 00

2O0O0 00

20000 0030000 00

30000 00

3O0OO 00

Final >Mach Altitude

0 00 0 00

1 50 4OOOO 000 00 0 00

0 00 0 00

0 00 0 000 00 0 00

0.00 0.000.00 0.00

0.00 0.00

0.00 0.000.00 0.00

0.00 0.000.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

0.00 0.000.00 0.00

0.00 0.00

0.00 0.000.00 0.00

0.00 0.000.00 0.00

0.00 0.00

0.00 0.00

0.00 0.00

Require

-450.00

60.00

500.00

5O0.OO9.009.00

9.00

6.00

9.004.40

7.OO6.80

5.00

5.00900.00

1300.00

650.00I000.00

600.00

450.00800.00

65O.O0

600.00550.00

50O.00

600.00

Actual

322.02

49.80

6605 78

23868 539 009 00

9 00

7 70

9.005.21

9.008.84

6.70

6.49964.28

1430.59

714.231135.82

998.04

484.25836.14

818.98798.68

553.80

653.25

746.35

1

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6.3 Multi-Role Concepts

Model 988-119, High Agility, Moderate Observables

The high agility, moderate observables vehicle, Model 988-119, is a single place tail-less transdnicdesign powered by two afterburning low-bypass turbofans of 28500 pounds (augmented) thrusteach. It is capable over Mach 1.5 on augmented engine thrust. Reduced signature characteristicsinclude moderate sideslope angles, edge alignment, lack of any vertical tail surfaces, and inletsintegrated into the wing-body junction.

The modified trapezoid planform was chosen to allow a higher aspect ratio without excessively narrowtip chords. Placement of the wing on the body for proper balance required the use of a canard insteadof a conventional horizontal tail.

Inlets are F-22 type, with angles chosen to align with the trailing edge while meeting side slope andinlet ramp angle requirements. Placement at the wing-body junctionresults in the intake duct passingalongside rather than over the weapons bay. Yaw vectoring exhaust nozzles are located in the aftfuselage.

Control effectors include those on the trailing edge of the wing, the canards, the yaw vectoringnozzles, "yaw vanes" forming the forward part of the nozzle fairings, and aft body flaps to providethrust vectoring in pitch. There are large leading edge slats to enhance maneuvering. The mid-outboard elevons are split to act as drag rudders.

The canards require high deflection capability to allow for effectiveness in high-Alpha maneuvers.They have 10 degrees of dihedral to reduce interference with the wing and inlets.

The interior layout, as shown by the Inboard Profile (ASC988-119-2) is conventionalfor tactical aircraft,with the exception of the internal weapons bay (side-by-side bombs/missiles) and exhaust nozzlearrangement. The fuel is contained in integral wing tanks, and a protected tank above the weaponsbay.

The exhaust system includes full augmentation, and dual rotating nozzles with variable throat and exitplane areas; this is an alternate nozzle arrangment from the single rotating nozzles shown on the otherconfigurations. It appears to offer reduced flow-turning losses and improved aft-body integration. Italso offers better pitch vectoring effectiveness (with the vectoring flap located between the nozzles),along with more flexibility for simultaneous yaw and pitch vectoring through differential pivoting of theupper and lower nozzles. There is not as much duct offset, but this is acceptable for a moderateobservables aircraft. Achieving acceptable efficiencies and effective vectoring is a significanttechnical risk (see Section 7.0).

139

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'" 3ENERAL ARRANGEHE'_", _,_ MODEL 986-119

MULTI -ROLE CONF !

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INBOARD PROFILE

MODEL 988-I19

MULTI-ROLE CONFIO

Icli.l I/in Jill ! I_. i

GROUP WEIGHT STA1--P...MENT

MISSION: Multi-role Mission

MODEL: 988-119 HiA/MedLO

WNG

HORIZONTAL TAIL

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENGINE SECTION

YAW VANES

TOTAL STRUCTURE

ENGNES

AMADS

ENGINE CONTROLS

STARTING SYSTEM

FUEL SYSTEM

VECTORING NOZZLES

TOTAL PROPULSION

FUGHT CONTROLS

APU

INSTRUMENTS

HYDRAUUCS

ELECTRICAL

AVIONICS

ARMAMENT


AIR CONDITIONING

ANTI-ICE

LOAD & HANDLING


HnEICY-fi'EMPTY

CREW

CREW EOJIPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHER_EJECTORS

AMMO CASES

NQN-EXP USEFUL LOAD

WEIGHT

(L.B)

6423

408

6488

1249

308

581

275

294

16026

3100

200

40

80

1.004

1529

5953

1267

210

270

588

618

1700

204

371

659

10

10

5906

27885

215

40

100

213

252

700

90

1610

OPERA_NGi4F=/G/-/T 29495

4990

110

14205

48800

BOMBS/MISSILES

AMMO EXPENDABLE

FUEL

GRCSS _

NOSE STATION

WING MAC

LB_AC

BODY LENGTH

BODY STATION

425

196

366

408

122

375

499

540

390

499

427

312

462

358

568

489

478

540

130

454

345

180

230

2O9

210

90

366

306

393

125

125

472

358

230

310

230

280

387

314

230

358

370

0IN

191 IN

334 IN

644 IN

PERCENT MAC

31.0%

27.7%

19.0%

SpreadSheet: 988-119 Wt. Statement 142 5/9/94

Nose @ BS 0LEMAC @ BS 334MAC Length = 191 In.AC @ 29.3% MAC

e.G. MOVEMENT RELATIONSHIP

TO AIRCRAFT WEIGHT

52000 -

50000

48000

46000

44000

42000

,oooo38000

36000

34000

32000

_,_ _oooo

=_ 28000

_

I

I

I

I

I

I

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I

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I

W l,lg + B)dy Fuel

Model 988-119 HIA/MedLO

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" " " " I . .AIM.12

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,role Mi on

O's

"" ,Wing Body Fu

I

I

I

iI

i

,¢z

0.18 0.19 0.2 0.21 0.22 0.23 0.24 0.25 0.26 0.27 0.28 0.29 0.3


Chart: C.G. CHART 5/9/94


M/R Model

Parameter Units 988-119HiA/ModLO,

Combat Weight Ibs 41300in. 373Longitudinal C.G. (Body Sta)

Vertical C.C. (from static ground line

Ixx Roll Inertia

lyy Pitch InertiaIzz Yaw Inertia

Ixz Product of Inertia

in.

slug-ft^2

slug-ft^2

slug-ft^2

slug-ft^2

74

63171

123726

204232

1234

144

//gandalf/user/deb9848/agility_dir/mr_dir/988-119_geom 07/16/94 1:43 P]

Aircraft Geometry

Thrust-to-Weight =

Takeoff Gross weight =

Wetted Area =

I.i0 Wing-Loading =


2634.7 Swet/Sref =

56.0

834.9

3.16

Body Geometry



Wetted Area = iii0.0 Volume =

8.56

992.7

Wing Geometry

Area = 834.9

Aspect Ratio = 3.97

Span = 57.57

Mean t/c = 0.05

Sweep Angles




NOTE: ARPITCH = 8.56, ARWE=

wetted Area = 1210.7

Taper Ratio = 0.00



Horizontal Tail Geometry

Area =

Aspect Ratio =

Span =

Mean t/c =

Sweep Angles

Leading Edge =

Quarter Chord =

Trailing Edge =

156.9

2.40

19.40

0.05

42.00

25.81

-37.46

Wetted Area = 313.9

Taper Ratio = 0.00


Engine Geometry







145

//gandalf/user/deb9848/agility_dir/mr_dir/988-119_weights

Weight Statement

Airframe Structure

Fuselage ........................ 6425.Wing ............................ 6262.Canard .......................... 0.

Horizontal Tail ................. 539.

Vertical Tail(s) ................ 0.

Engine Mounts ................... 159.Inlet(s) and Duct(s) ............ 421.Exhaust Duct(s) ................. 0.Pivots .......................... 0


Total 15339

Propulsion System

Engine(s) and Nozzle(s) ......... 3881Engine Start and Control ........ 120Fuel Tanks ....................... 387

Fuel Pumps ...................... 21Fuel Distribution System ........ 249

Air-Refueling System ............ 63

Fuel Inerting System ............ 59Gear Box and Accessories ........ 200

Total 4980

Fixed EquipmentInstruments ..................... 270

Surface Controls ................ 1419

Crew Accomodations .............. 401Armaments ....................... 1179

Avionics ........................ 1725

Electrical System ............... 688

Hydraulics and Pneumatics ....... 423Radar Absorpton Material ........ 0

Auxiliary Power System .......... 182Airconditioning and De-lcing .... 835

Total 7122

Empty Weight .................... 27440

Weight Weight Volume(LBS) Fraction (cu ft)

0.1374

0.13390.0000

0.0115

0.00000.0034

0.0090

0.0000. 0.0000. 0.0263

. 0.0065

. 0.3281

0.0830

0.00260.0083

0.0004

0.00530.0013

0.0013

0.00430.1065

0.0058

0 0304

0 00860 0252

0 0369

0 01470 0091

0.0000

0.00390.0179

0.1523

0.5869

6463

0

50

253

13

03412

245

17

3

181

6

21

4

51

7

35

7039

3317

II.0.

6.

42.

260.

0.

Operational ItemsCrew ............................ 200. 0.0043 3.

Trapped Fuel and Oil ............ 351. 0.0075 7.Gun and Provisions .............. 342. 0.0073 7.

Operational Empty Weight ........ 28333. 0.6060 0.

PayloadAmmunition ...................... ii0. 0.0024 4.

Air-to-Air Missles .............. 690. 0.0148 65.

Air-to-Ground Munitions ......... 4300. 0.0920 97.

Total 5100. 0.1091 166.

Mission Fuel

Wing Fuel ........... 10896. 0.2330 224.Body Fuel ........... 2427. 0.0519 50.External Fuel ....... 0. 0.0000 0.

Design Gross Weight .... ......... 46756. 1.0000 1823.

CG

(ft)

28.27

30.170.00

0.00

0.0049.28

47.350.00

0.00

32.1619.88

28.94

49.28

30.18

30.3230.32

30.3230.32

30.32

49.2845.43

23.5724.34

19.5030.32

21.28

18.1929.19

0.00

0.00

10.31

21.7130.06

11.08

39.80

0.00

0.00

0.00

0.00

30.32

0.00

0.000.00

30.32

07116194

Moment

(ft/ib)

181641188946

00

0

7_4419938

0

039476

6058443904

191259

407411724

630

75491819

17877392

226234

6365

34537

782335745

3670112510

12359

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8611.

154651.

824789.

2216.

13951.0.

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130369.0.

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824789.

1:44

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988-119 High Agility, Moder0te Observ0blesTorget Cross Sectionol Areo Distribution

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988-119 High Agility, Moderate ObservablesCruise Drag Polar

THE BOEINGCOMPANY "'_ _48

//gandalf/user/deb9848/agility_dir/mr_dir/988-119_mission 07/16/94 1:42 PM

CD


Initial Final Fuel Time Range

Weight Weight Burned (min.) (n.mi.)

Mach Altitude CL CD

(feet)

0. 0.248 0.0151

0. 0.248 0.0151

0. 0.248 0.0151

40226. 0.287 0.0186

20000 0.090 0.0118

20000 0.090 0.0118

20000 0.715 0.0768

20000 0.079 0.0116

0.702 0.0717

0.272 0.0185

0.317 0.0209

0.248 0.0151

635. ibs

Warm-up and taxi46756 46486. 270 0 i0.00 0.0 0.000

Warm-up and taxi

46486 46369. 117 9 0.25 0 0 0.000

Warm-up and taxi46369 45936. 432 4 0.25 0 0 0.300


45936 45310 626 5 0.32 2 1 0 945 0. 0.044 0.0111


45310 43700 1609 7 1.62 i0 8 0 844 42903. 0.241 0.0167


43700 39845 3855 5 88.07 672 9 0 800


39845. 39296 548 7 5.55 50 0 0 880

Drop 4300.00 ibs of expendables39296. 34996 0.0 0 00 0 0 0 880


34996. 34904 91 5 0 28 3 0 0 880


34904. 34361 543 6 5 55 50 0 0 880


34361. 34194. 166 9 0 55 3 0 0 880 20000


34194. 33331. 862 5 1 37 9 9 0 830 48575


33331. 30552. 2779 5 79 31 640 1 0 840 49934


30552. 29763. 788 3 20.00 69.3 0 314 0

Total Mission Fuel = 12693. ibs Reserve Fuel =

Power Net Fuel Error

Setting Thrust Flow Code

0.040 1331. 1620. 0

1.000 33264. 28289. 0

2.000 54610. 103772. 0

2.000 65309. 136132. 0

2.000 17925. 34701. 0

0.317 2698. 2570. 0

0.235 5190. 5927. 0

0.235 5190. 5927. 0

0.863 39185. 19615. 0

0.232 5122. 5871. 0

0.806 39185. 18207. 0

0.806 12449. 24121. 0

0.376 2112. 2057. 0

0.054 1833. 2354. 0

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Model 988-118, Medium Agility - Low Observables

As a medium agility, low observable vehicle Model 988-118 is a moderate gross weight single place,subsonic delta wing design powered by two turbojfan engines of 30,830 pounds augmented thrusteach. The external general arrangement, shown on drawing ASC 988-118-1, is characterized by themoderate leading edge sweep of thirty-eight (38) degrees, nested lower forebody inlet apertures, _ullspan trailing edge elevons, and upper body mounted thrust vectoring exhaust nozzles. The wingleading edge incorporates large powered slats that are used to augment maneuver performance.

Control effectors include the yaw thrust vectoring exhaust nozzles with _+45degrees of deflection,Yaw Vane pairs on the upper and lower surface integrated with the nozzle, and four elevons per semi-span. Elevons are single panel at the most inboard position,with the two outboard panels being spliton the wing reference plane.

Inlets are integrated/nested with the lower forebody, and exhaust nozzles are located side-by-side onthe upper aft fuselage. Yaw Vane pairs are integrated with the nozzles and on the lower aft body.

The interior layout, shown on inboard profile drawing ASC 988-118-2, accommodates the crew,subsystems, weapons and propulsion system within a low profile body shape. The forebody isconventional in arrangement and includes avionics, crew station, gun system, and subsystems.Center body contents are inlet system, weapons bay, and main landing gear. The aft body providesengine and exhaust system accommodation.

Propulsion system installation features consist of the nested external compression fixed ramp inlets,each feeding a long vertical offset inlet diffuser running over the weapons bay to an engine face.

The exhaust system includes augmentor spray bars, and a fully offset duct to nozzle exit plane. Theduct turns through a circular bearing/rotation plane to direct exhaust gas out the nozzle aperture. Asignificant and challenging risk issue is presented here in this concept of making the rotating nozzlesystem augmentor.capable. A discussion of this issue is contained in Section 7.0, Areas of HighTechnical Risk.

Nozzle concept is that of a variable throat SERN type that utilizes the upper aft deck as the expansionsurface.

All fuel is contained in the main wing panel outboard of the side of body. Provision would be made toprotect get-home fuel in each wing tank.

151

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s " I z 0 •

GROUP WE]GI rr STATEMENT

MISSION: Multi-role Mission


WNG

BODY

MAIN GEAR

NOSEGEAR

AIR INDUCTION

ENGINE SECTION

YAW VANES

TOTAL STRUCTURE

ENGNES

AMADS

ENGINE CONTROLS

STARTING SYSTEM

FUEL SYSTEM

VECTORING NOZZLES

TOTAL PROPULSION

FUGHT CONTROLS

APU

INSTRUMENTS

HYDRAULICS

ELECTRICAL

AVIONICS

ARMAMENTFURNISHINGS & EQUIPMENT

AIR CONDITIONING

ANTI-ICE

LOAD AND HANDLING

WEIGHT

(LB)

6931

7618

1288

330

1145

293

294

17899

3352

200

40

80

1020

1412

6104

1380

210

270

506

618

1700

204

371

658

10

10

TOTALRXEDE-CXJlPMENT 5938

_E_'vfPTY 29941

CFEW

CREW EQUIPMENT

OIL & TRAPPED OIL

TRAPPED FUEL

GUN INSTALLATION

LAUNCHERS__JECTORS

AMMO CASES

NE_EXP USEFUL LOAD

OPERATING 14_GHT

215

40

100

214

252

70O

90

1611

31552

4990

110

14248

50900

BO_BS_ISSlLESEXPENDABLE

FUEL

NOSE STATION 0 IN

WING MAC 260 IN

LEMAC 373 IN

BODY LENGTH 710 IN

BODY STATION PERCENT MAC

499

417

490

225

341

574

638

452

574

501

375

536

488

654

574

542

621

180

532

396

200

236

221

246

130

417

350

457 32.2%

175

175

546

488

236

393

236

347

451 30.1%

GROSS _

394

236

488

455 31.6%

SpreadSheet: 988-118 WT. STATEMENT 15 4 5 / 6 / 9 4

54000

52000

50000

48000

46000

44000

.ooo

4000038000

36000

34000

32000

30000

0.27

Nose @ BS 0LEMAC @ BS 372.69

MAC Length = 261.1 In.AC @ 31.85% MAC

Model 988-118 MedA/LO

l laslc MuI If-role M ssio_/,_ . .&

_.-"Wln_ Fuel

'r"_M',, ....ul.#l

" ""•--- Amino

AIM-120's _"• ":'_

.-"Wllg Fuel

0.275 0.28 0.285 0.29 0.295 0.3 0.305 0.31 0.315 0.32 0.325 0.33

C.G. LOCATION - %MAC

0.335 0.34

Chart: C.G. CHART 5/9/{) 4


M/R Model


Combat Weight

Longitudinal C.G. (Body Sta)

Vertical C.C. (from static ground line)

Ixx Roll Inertia

Ibs

in.

in.

slug-ft^2

slug-ft^2lyy Pitch InertiaIzz Yaw Inertia slug-ft^2

Ixz Product of Inertia slug-ft^2

MedA/LOd

43770

449.5

82

73135

159377

249068

1539

156

//gandalf/user/deb9848/agility_dir/mr_dir/988-118_geom 07/16/94 2:28 P}

Aircraft Geometry

Thrust-to-Weight =

Takeoff Gross Weight =

Wetted Area =

I.I0 Wing-Loading =


2982.9 Swet/Sref =

48.3

1119.0

2.67

Body Geometry



Wetted Area = 1300.0 Volume =

8.35

1055.4

Wing Geometry

Area = 1119.0

Aspect Ratio = 3.52

Span = 62.76

Mean t/c = 0.05

Sweep Angles




NOTE: ARPITCH= ii.i0, ARWE=

Wetted Area = 1682.9

Taper Ratio = 0.00



Engine Geometry







157

//gandalf/user/deb9848/agility_dir/mr_dir/988-118_weights 07/16/94 2:29

Weight Statement Weight Weight Volume(LBS) Fraction (cuft)

Airframe Structure

Fuselage ........................ 7848.

Wing ............................ 9069.Canard .......................... 0.

Horizontal Tail ................. 0

Vertical Tail(s) ................ 0.

Engine Mounts ................... 184Inlet(s) and Duct(s) ............ 473

Exhaust Duct(s) ................. 0Pivots .......................... 0


Total 19347

Propulsion System

Engine(s) and Nozzle(s) ......... 4489Engine Start and Control ........ 120Fuel Tanks ...................... 435

Fuel Pumps ...................... 23Fuel Distribution System ........ 267

Air-Refueling System ............ 63

Fuel Inerting System ............ 63Gear Box and Accessories ........ 200

Total 5660

Fixed EquipmentInstruments ..................... 270

Surface Controls ................ 1206

Crew Accomodations .............. 401Armaments_ ...................... 1802

Avionics ........................ 1725

Electrical System ............... 688Hydraulics and Pneumatics ....... 496

Radar Absorpton Material ........ 0Auxiliary Power System .......... 182

Airconditioning and De-Icing .... 963.Total 7732.

Empty Weight .................... 32740.

0.1452

0.16780.0000

. 0.0000

0.00000.0034

0.0088

0.00000.0000

0.0263

0.00650.3580

0.08300.0022

0.00810.0004

0.00490.0012

0.00120.0037

0.1047

0.0050

0.0223

0.00740.0333

0.0319

0.01270.0092

0.00000.0034

0.0178

0.1431

0.6057

Operational ItemsCrew ............................ 200.

Trapped Fuel and Oil ............ 405.Gun and Provisions .............. 342.

0.00370.0075

0.0063

Operational Empty Weight ........ 33687. 0.6233

78

910

0

02

57140

42

14

297

18

3-56

17

2

24

-21

7 .

30.70.

60.

33.17.

12.

0.6.

48.

284.

0.

3.

8.

7.

0 .

CG

(ft)

33.49

40.880.00

0.00

0.0054.62

52.55

0.000.00

35.56

22.0437.56

54.6232.85

33.5333.53

33.5333.53

33.53

54.62

50.57

25.3434.70

20.63

33.5323.53

20.18

31.500.00

0.00

10.9825.61

36.99

11.08

44.08

0.00

0.00

Moment

(ft/ib)

262844.

370760."0.

0.0.

10056.24_58

00

504837768

726770

2451864435

14594775

89582012

2116

8193286270

6841

41840

8276

6042740593

1387315622

0

010574

198045

2216.17870.

0.

0 .

PayloadAmmunition ...................... ii0. 0.0020 4. 0.00 0.

Air-to-Air Missles .............. 690. 0.0128 65. 0.00 0.

Air-to-Ground Munitions ......... 4300. 0.0796 97. 33.53 144192.Total 5100. 0.0944 166. 0.00 0.

Mission Fuel

Wing Fuel ........... 15262. 0.2824 314. 0.00 0.Body Fuel ........... 0. 0.0000 0. 0.00 0.External Fuel ....... 0. 0.0000 0.

Design Gross Weight .... ......... 54049. 1.0000 1791. 33.53

158

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LF_C_

CALC

c";Co;IAPPD.

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R.Engelbeck

_PO8 IOIO_

988-118 Medium Agility, Moderate 0bserables


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0_0

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73

988-118 Medium Agility,Moderate ObseroblesCruiseDrag Polar

THE BOEINGCOMPANY "_ 16o

//gandalf/user/deb9848/agility_dir/mr_dir/988-118_mission 07/16/94 2:27 PM


Initial Final Fuel Time Range Mach Altitude

Weight Weight Burned (min.) (n.mi.) (feet)

CL CD

---%

O_

Warm-up and taxi54049. 53737. 312.1 10.00 0.0 0 000 0. 0 220 0.0128

Warm-up and taxi53737. 53601. 136.3 0 25 0 0 0 000 0. 0 220 0.0128

Warm-up and taxi

53601. 53101. 499.8 0 25 0 0 0.300 0. 0 220 0.0128

Acceleration from Maoh 0.300 to Mach 0.935

53101. 52391. 710.1 0 31 2 1 0 935 0. 0 039 0.0095

Climb from 0.0 ft. to 44111.9 ft. at 216.7 ft/seo

52391. 50469. 1921.8 1 73 ii 7 0 847 44112. 0 218 0.0146


50469. 46146. 4323.5 87 96 672 1 0 800


46146 45514. 632.1 5 55 50 0 0 880

Drop 4600.00 ibs of expendables

45514 40914. 0.0 0 00 0.0 0 880


40914 40830. 83.7 0.28 3.0 0 880


40830 40203. 626.7 5.55 50.0 0.880


40203 40045. 158.6 0.55 3.0 0.880 20000. 0.613 0.0517


40045 38991. 1053.7 1.49 10.9 0.841 49738. 0.243 0.0161


38991 35810. 3180.6 79.97 639.1 0.832 49984. 0.283 0.0182


35810 34916. 894.3 20.00 68.8 0.311 0. 0.220 0.0128

Total Mission Fuel = 14533. ibs Reserve Fuel = 727. Ibs

41201. 0 259 0.0163

20000. 0 078 0.0101

20000. 0.078 0.0101

20000. 0.624 0.0534

20000. 0.069 0.0100

Power Net

Setting Thrust

0.040 1538. 1873

1.000 38452. 32702

2.000 63128. 119958

2.000 75423. 157065

2.000 19804 38343

0.323 3035 2888

0.234 5971 6828.

0.234 5971 6828.

0.696 45297 17937.

0.231 5899 6770.

0.673 45297 17306.

0.673 13877 26891.

0.375 2406 2334.

0.053 2059 2672.

Fuel Error

Flow Code

L_;" IIL

E# "18G

IL'IGG

#6" L#L

06 "_OL

8# " #E8

6g "g8_

#0"866

£I '_EII

60"gIL

8g'Z96

88 "_

II 9

L8 L

O0 6

66

O0 6

LE L

O0 6

O0 6

00"6

_I'SLL£E

8I'6I£9

0£'#£

0L'0EZ

I E D.3,O"_

00"009

O0"OOG

O0"OGG

00"009

00"0_9

00"008

O0"OG#

00"009

00"0001

00"0_9

O0 00El

O0 006

O0

O0

08 9

O0 L

O# #

O0 6

O0 "9

00"6

00"6

00"6

00"00£

00"00_

00"09

00"0_#-

e_ Tnbe_

O0 0

O0 0

O0 0

O0 0

O0 0

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00"0

00"0

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00"0

00"0

00"0

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00"0

00"0

00"0

00"0

00"0

00"0

00"0000#

00"0

ePn_T_I_

00"0

00"0

00"0

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7=0 C=r_ai &_

Comparisons

Air-to-Ground Designs

The Air-to-Ground designs for high and medium agility collapsed into the same design when themaneuvering devices were added. The Aircraft Design Synthesis results discussed in section 5.6concluded that when the maneuvering flaps were added, the aircraft thrust-to-weight required to meetboth levels of agility were exceed by the 6.5g sustained turn requirement. Further, the resultingreduction in aircraft thrust-to-weight made the thrust required to meet the lateral sideforce agilitymetrics the driver in aircraft engine size. Tripling the aircraft engine size to meet the lateral side forceagility matric would have a huge impact on aircraft weight and cost. The recommend approach for thenext design cycle would be to add large aerodynamic sideforce generators to the designs.

The flying wing configurations in general are extremely vulnerable to spiraling weight growth as thedesign matures since wing area growth is constrained by the LO philosophy, carder suitabilitygeometric constraints, and limited center-of-gravity flexibility.

Air-to-Air Designs and Multi-Role Design Drivers

Designs with significant Air-to-Air capability were driven to high T/W levels because of the maneuverrequirements. All the designs had low wing loading (W/S) because of their instantaneous turnrequirements. The combination of low wing loading and the resulting wing spans were judged to benear flutter boundaries between 3.5 and 4.5 aspect ratio.

The Impact of Carder Suitability

Adding the carder suitability features to the otherwise identical USAF customer added 14 to 17% tothe aircraft empty weight. The low wing Ioaclings, relatively high aspect ratio, and high aircraft thrust-to-weight ratios kept the single-engine rate of climb, catapult, and recovery performance boundaries.The biggest issue for.carder suitabilitywas the general size and weight of the aircraft and the adverseimpact it has on deck handling. Some issues remain conceming the impact of large inertias on therotation rates required to meet the 10 ft. sink requirement during a catapult launch, and the rotationrates required to accomplish a bolter. These issues were not addressed by the simplistic cardersuitability methods used to size the configurations.

The Impact of Observables Design Philosophy

The observables design philosophy as implemented here was to minimize the number of edges andsurfaces on the aircraft. One major impact of this philosophy is reduced maximum liftcapabilitybecause the flap system of tailless designs must be used for maneuver and trim requirements. Thisimpacts the instantaneous turn capability and carder launch and recovery speeds.

The Impact of Agility

Our design intent was to embrace the agility requirements from the outset of the study. Agility drovethe layout of the aircraft, the control system philosophy, and control surface sizing. In the case of theAir-to-Ground designs, agility would drive the propulsion system size unless an alternative sideforcegenerator concept were utilized.

The use of yaw thrust vectodng was key to the achievement of the lateral sideforce agility levels. Theuse of yaw vectoring was selected over conventional tails because of its effectiveness at high angles-of-attack and low speeds. In addition, thrust vectoring would probably neutralize the issue of

163

departure. Removal of the vertical tails on all configurationswas done to offset the impact in weightfrom the thrust vectoring system by elimination of the structural weight and drag penalties of thevertical tail. The added benefit of the elimination of the vertical tail is the reductionin side signature.Use of vertical tails would have required that they be canted to keep the side signature down resultingin the cross couplingof the yaw and pitch axis. The use of yaw vectoring without tails would eliminatethis undesirable cross coupling.

164

Design Interactions

Observables vs Agility

Design for agilitytends to favor concepts with more/larger control effectors, and low inertias. Designfor observabilitytends to drive the number of surfaces that produce a radar return down. Minimizingthe number of surfaces drives the designer to aerodynamically inefficient deltoid wings of low aspectratios. This aerodynamic efficiency drives the wing size up to partially offset the efficiency loss. Thelargerwing in turn makes the aircraft largerand heavier. The result is an aircraftdesignthat hasrelatively large inertias and fewer control eflectors. Design emphasis on low observables will be adetriment to agilityat a given level of maneuverability and missionperformance.

Carrier Suitabilityvs Agility

The Navy has traditionally been more stringent in the specification of maneuvering requirements thatutilize as much of an aircraft flight envelope as possible. The Navy expects their pilotsto fly to theedge of this envelope and consequently drives the designer to provide Level 1 flying qualities to themaximum limits of the operational envelope. The Navy requiries highdeparture resistance at highangles-of-attack sufficient to prevent loss of control while maneuvering close to and possible throughportions of the flight envelope where control authority traditionallybegins to diminish. The Air Forcewill typically accept limitersto avoid approaching CLmax boundaries throughoutthe maneuveringenvelope. The Air F-16 employs an angle-of-attack limitingschedule which shrinks the left boundaryof the energy maneuverability envelope significantlybeyond corner speed. Unique maneuverdevices are normallyfound on naval aircraftto ensure maximum maneuvering performance over a fullflight envelope. These devices usually take advantage of an already unique low speed, high liftsystem such as the maneuver flap. All of these features are positive contributionsto the agility of anaircraft.

A carrier suitable design is constrained in both weight and size withinthe operating limitationsof anaircraft carrier. Wing fold weight increases with span and tends to drive the wing span of the aircraftdown. Minimizing"the span minimizes roll inertia for any given weight helpingthe rollagilityof theaircraft. However, the decreased span also has an adverse effecton the roll control power necessaryto start and stopthe aircraft roll. The increase in aircraftweight to handle the structural loads andadditional equipment associated with carrier based operations overwhelm any positive aspects of theNavy designs resultingin aircraft designs less agile than their Air Force counterpartswith the samemaneuver and mission performance.

Air-to-Air vs Air-to-Ground Operational Mission Roles

Aircraft designed to the primarily subsonic Air Interdictionmission without any stringent supersonic orAir-to-Air maneuver requirements will typically have large low bypass engines and low aircraft T/W foroptimal cruise performance. Aircraft designed to meet the challenging Air-to-Air maneuverrequirements will typically have aircraft T/W greater than 1.0 and low bypass ratios. The keytechnology used in all the agility designs in this study is thrust vectoring. The benefit of thrustvectoring for agility is more effective on the Air-to-Air designs than the Air-to-Ground designs becauseof the greater T/W of the Air-to-Air designs.

165

Wave Drag Levels are Achievable

The specific excess power and sustained turn requirements are ambitious. These requirementsrepresent a ten percent improvement in maneuver capability over F-15 and F-14 fighter capability. Toobtain this maneuver capability the design philosophyfor wave drag is to work the cross-sectional areadistribution as hard as possible to minimize the transonicdrag rise and supersonic drag levels.Reduced wave drag will help minimize the engine size required for maneuver and minimize the fuelconsumed during supersonic cruise on the defensive counter air mission. Although ideal L-V Haackarea distributionsare targeted, figure shows a 30 to 44 percent conservatism in the final designs.This conservatism placed the Boeing designs comfortably within the demonstrated levels achieved bypast designs.

166

O

II

1:3

OI

04

II

,,,..pE3

O

0.7

0.6

0.5

0.4

0.3

0,2

0.1

Sym Aircraft1 XF4D-12 XF-91

O 3 XF-100A4 XFY-15 D558116 X-3

2O

40

01

b

(ref. area)

Tail

I1= _ =1

I I I I

4 5 6 7

Sym

D

Aircraft7 F-100D 14 F-102A8 F-4E 15 RA-5C9 F8U-3 16 F-105D

10 F-11A 17 F-5A11 F-8D 18 F-10412 Mirage3G19 F-106A13 B-58,

13A B-58+POD

n 24

[] 27

Sym

A

Sym

A

Aircraft20 F-14A21 F-B111A22 LFAX-4:23 LFAX-9

Aircraft24 F-15A25 F-1626 YF-1727 F-18

1%

All

25 26[] []

13A

14 Z_15 06

13/122 AlS

/t23

_Sears-Haack

minimum drag body

I I I

9 10 11

I

12

Fineness ratio (2/d)

Mission observablesagility model no.

CDo @ M = 0.8

CDo @ M = 1.2Sref

A=

ACD= - estimatedLength

AmaxdY/d

ACD= - Sears-Haack

ACD_ est / ACD_ ideal

Multi-role

Low Mediummedium high988-118 988-119

0.0098 0.01160.0278 0.0340

1119.0 834.970.0 71.0

0.2877 0.2634

58.83 53.1970.0 71.0

9.44 9.51

7.41 7.470.20 0.20

1.44 1.32

Lowmedium988-114

Air superiorityMediumhigh988-115

0.0101

0.02311429.1

68.0

0.273260.3

68.0

9.307.31

0.21

1.30

0.0127

0.03161014.6

63.0

0.304467.14

63.0

8.967.03

0.22

1.38

Wave Drag Sanity Check

167

spb-7-7-re-Ad9

6.0 Critical Assessment

Comparisons

Configuration Design

An assessment summary, figure 6.0-1, has been made for Models 988-115, -118 and -122/-123. Thegeneralized elements consider long term program issues such as growth capability in mission typeand payload size, as being critical to establishing design acceptability. If constrained to the singlemissionpayloads the practicality of these designs is suspect.

168

ASSESSMENT SUMMARY

...I.

03(D

ELEMENT

Strengths

AIR SUPERIORITY

MODEL 988-115

Control Effector Mix

- Vectoring Nozzle

- Canards

- Tails

- Yaw Vanes

MULTIROLE

MODEL 988-118

Control Effectors

- Vectoring Nozzle

directional control power:all altitude

- Yaw Vanes

AIR INTERDICTION

MODEL 988-112/-123

• Control Effector Mix for Side

Force

- Vectoring Nozzles- Yaw Vanes

• Payload/Radius capability

Weakness

Suitability

Achievability• Cost

• Supportability• Effectiveness

• Signature

• IR missile FOV/FOR

• Size drives affordability• Limited internal stores

carriage

• External stores capable

- Conformal

- Pylon mounted

• Difficult to operate on carrier

• Driven by technical issues• Obtainable & sustainable

• Close in high probability• Vulnerable to threats - many

edges

• IR missile FOV/FOR

• Limited internal stores

carriage volume

• External stores capable on

Wing- Conformal

- Pylon mounted

• Driven bytechnology• Obtainable & sustainable

• Expanded capability• Reduced vulnerability-

cleanerdesign

IR missile FOV/FOR

Limited by growth

incorporated for internal

payload

• Internal sores capability drives

bay size & vehicle• External stores not desired

alternate

• Driven bytechnologies used• Obtainable & sustainable

• Broad capability• Low levels areinherentin

basic design

Sheet1

.,,jo

Agility LevelObservables Level

Model Number

Maximum Negative PsAcceleration

Maximum Rate of ClimbInstantaneous Load FactorInstantaneous Load FactorInstantaneous Load FactorInstantaneous Load Factor

Sustained Load FactorSustained Load FactorSustained Load FactorSustained Load FactorSustained Load FactorSustained Load FactorSpecific Excess PowerSpecific Excess PowerSpecific Excess Power

Mach No0.60.90.60.90.91.20.60.90.60.91.20.91.20.60.90.6

Medium Medium High Medium HighLow Low Moderate Low Moderate

Altitude-ft Requirement 988.114 988-114A 988-115 988-1t8 988-11915000 -100 -16.34 -16.34 -59.37 127.57 -53.5440000 60 54.39 54,39 51.53 52.68 51.4150000 500 6603.28 6603.28 6620.42 6557 6586.7820000 9 9 9 9 9 9300003000010000100002000020000200003000030000

00

10000

969

4.47

6.855

9OO1300650

96.62

94.518.557.7

5.655.76

971.61436.88718.98

6.62

4.518.557.7

5,655.76

971.61436.88718.98

99

6.379

4,388.297,9

5.475.82

964.871439.25713.71

7.02

4.768.857.875.855.84

962.061440.38711.58

6.4

4.48.047.525.3

5.54978.091481.73723.28

Specific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess PowerSpecific Excess Power

1000

1.4 30000

1140.6 1140.6 1140.38 1140.190.9 100001.2 10000 600 998.04 998.04 998.04 998.04 998.040.6 20000 450 486.46 486.46 482.73 481.7 489.670.9 20000 800 838.9 838.9 837.73 837.38 857.71.2 20000 650 707.31 707.31 784.85 749.13 789.581.4 20000 600 742.55 742.55 817.34 807.19 795.040.9 30000 550 554.13 554.13 553.05 553.22 566.331,2 30000 500 583.91 583.91 630.31 607.46 633.55

600 711.26 711.26 755.14 747.13 742.15

1169.86

Page 1

Chart5

Instantaneous Turn Rate

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9

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Flight Condition

I

0.6

[] Requirement

I988-114

[] 988-114A

n988-115

I988-118

__988-119

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GO

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1200

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0.6 0.9 0.6 0.9 1.2 0,6 0.9 1.2 1.4 0.9

Flight Condition

1.2 1.4

Page 1

,-.k

TOGW (Ibs)

•"_-- Air-to-ground

A-6E

58,600

958-123N

80,910

• Multi-role

A-12 F-16

60,000 ---> 37,50080,000

16,285

6,846

0

988-119 988-119N

48,801 54,704

27,886 32,033

14,205 15,923

4,990 4,990

F-18

51,900

F-15A

50,000

Air-to-air

F-14

74,349

988-115 988-115N

59,549 67,397

31,746 36,367

24,026 27,192

1,800 1,800

Empty weight (Ibs) 25,980 34,961 23,050 26,768 41,353

Max. internal fuel (Ibs) 15,939 33,596 10,860 11,050 16,200

Max. internal payload (Ibs) 0 9,100 0 0 0

T/W .317 .34 .64 1.10 1.10 .61 .96 1.13 1.13 .72

W/S 111 50 45 _ 50 125

AR 5.311 3.5 3.75 3.0

Engines J52__2.)8A1B Adv(2a_)cedA/G

F:(42_2

(dry)

59

3.97

59

3.97

Ad:2a_)ced

(NA)

130

4.0

F! 4

82

3.0

F-%

58

3.8

(2)Advanced

(A/A)

58

3.8

(2)Advanced

(A/A)

132

2.58 _ 7.28

F! lo

Span (ft) 53 75 70.27 It 31.0 57 61 40.4 42.8 63 67 38.2 _ 64.1

Sref (sq ft) 528.9 1,618 1,317 300 830 931 400 608 1,032 1,167 565

- - 5.0 6.77 5.0 5.2 6.4 6.4 4.8Sustained load factor@ M = 0.6, H = 10,000 ft

6.77

Historical Comparisons

spb-8-7-re-Ad7

8.0 Fliaht Research Needs Assessment

The Value of Agility in Combat Effectiveness

The single largest uncertainty in designing an agile fighter is quantifying the value of agility in terms ofcombat effectiveness. In the absence of quantifiable measures of merit, the militaryhas beenreluctantto impose specific agilityrequirements on the aircraft designer. The aircraft industry isreluctantto develop methods or design aircraft to meet requirements their customer has notspecificallycalled out. Over the lastfew years, programs likeX-29, X-31, HARV and VISTA haveproducedresearch aircraftthat are arguably more agile than currentfighter aircraft. On those programswhere more than one aircraft exist (Skethe X-31), 1 vs 1 flightcombat simulationsand to beconducted with one research vehicle simulating a conventional fighter with the same basic flightcharacteristicsas the fully functionalresearch aircraft. In thisway the impact of agilityon combateffectiveness can be isolated from other flight characteristics. Issues concerning the impact of agilityon combat effectiveness in the M vs N scenario would best be quantified through flight combatsimulationsusing a collection of research aircraft against current inventory fighters. This would helpquantify the effect of number of adversaries has on the value of agility.

Control Effectors

Research needs to continue to develop new and creative methods to control aircraft. New challengessuch as lateral controlof tailless aircraft and concern for the signature characteristics of controls beingdeflected are only a subset of the research that needs to be conducted.

Wind tunnel tests need to be conducted to quantify the benefits of Tiperons against the heavy weightpenalties of attachment. Porous leading and trailing edge devices promise low RCS and reducedmechanical complexity, but more windtunnel testing is needed before the concept is proven anddesign information developed to effectively implement the concept into aircraft design. Leadingedge blowing, leading edge suction, and tangential wing blowing need more windtunnel databasedevelopment to prove the concepts, quantify the benefits, and determine the blowing requirementsand weight penalties. A database of windtunnel data needs to be developed for drag rudders as afunction of deflection and wing plantorm. A windtunnel database of the effectiveness of articulatingforebody strakes, nose strakes, chines, and other forebody shapes needs to be developed. All typesof forebody jet blowing, slot blowing, and suction need wind tunnel research to quantify theireffectiveness and how these concepts are impacted by forebody shape.

The primary attribute of all the designs produced in this design study is the radical amount of yawthrust vectoring used. Continued development of all thrust vectoring schemes with the objective ofproving the use of thrust vectoring as a primary control should be pursued vigorously.

Aerodynamics

Low signature requirements are driving the aircraft designer to simple tailless designs like the B-2 andA-12 configurations. This design philosophy would benefit from research into devices to counter theinherent inefficiencies of low aspect ratio wings with lotsof wetted area. Windtunnel testing of vortexflap concepts with sharp and semi-sharp leading edges need to be tested with various planformvariations, especially leading edge sweep. Variable camber or mission adaptive wing designs wereproven in the MAW and AFTI F-111 program but research into how to implement the concept withcomposite materials in an environment of emphasized low signature, lower weight, cost limitationsneeds to be researched. Success in CFD research into predictingtransition has direct impact on thesuccessfuldesign of natural laminarflow shapes. Natural laminar flow is one of the few technologiescapable of reducingparasite drag, a very important component of drag for aircraft with large wingsurfaces likethe B-2 and A-12. Research into methods to maintain a smooth surface in a dirty serviceenvironment and manufacturing issues need to be addressed for passive laminar flow concepts. Inaddition, reducing the maintenance requirements, weight, and cost of active laminar flow conceptsshould continue to be researched. Flight Testing samples, detailed CFD, and windtunnel testing ofporous lifting surface technology would help quantify the benefits of reduced shock strength andinstallationdrag penalties on the Mach capability of potential designs.

175

Propulsion

One of the early findings of this design study was the importance of engine technology level on theaircraft size and fuel requirements. These aircraft designs, relative to their contemporary counterparts,have long design mission radius requirements coupled with high thrust-to-weight ratios required fbrmaneuver. The resulting strong sensitivity to the propulsion weight and fuel consumptioncharacteristics drove the design team to select the GEN 6+ level of technology to keep the aircraft sizedown. Fighter engine thrust matching for cruise radius conflicts with ever increasing demands forfighter maneuverability and acceleration. Efforts to reduce fuel consumption in both cruise andcombat require cycle optimizations at both low and high power ends of the engine thrust spectrum.IHPTET GEN 6 engine technologies including variable cycle engine/control technologies researchshould continue to be strongly supported.

Structures & Materials

Advanced Aluminum-lithium Alloys and Advanced Titanium alloys have had a historyof failures.Advanced Aluminum-Lithium Alloyshave failed due to poor ductility. Weldability and crack growth hasbeen the cause of failure for the advanced Titanium Alloys. Unless significantprogress is made intothese issues and reducing the high cost penalty, these material probably will not see wide spread use.Power metallurgy using current materials stillneeds further development to realize any savings inmanufacturing costs and may not be worth the effort. However, metal matrix composites is potentiallya major benefit.

Expect continued research into composite materials technology like graphite based composites.Research on the cost and ease of use of Boron based composites should be emphasized. Researchinto preventing water contamination of composite materials like Kevlar should be pursued. The use ofadvanced resins could save weight by improving the materials toughness but manufacturing researchwill be criticalto its success.

Manufacturing Techniques

Improved materials are nice, but the next breakthroughswill be in new joining and manufacturingmethods like welding and co-curing. Advanced manufacturing techniques like superplastic forming,T. welding, composite welding, and Z-pinning all have major potential benefits for future fighters. Stillmore time and research is needed to realize cost and weight savings using structural techniques suchas welded joints, Issogrid, Column Core, and Z-pinning.

Structural Design Issues

A couple of the concepts developed for this study were arbitrary limited by our discomfort with thepotential of encountering the structural flutter boundary. Research into establishing the location ofthe flutter boundary in conceptual design would help the designer produce a good design withoutthe highcost of higher order studies currently required to establish a flutter boundary Research intousing the control system in an active flutter suppression system would allow the use of more wingaspect ratio, sweep, and less thickness for improved aerodynamic efficiency. This technology needsa flight demonstration on an unmanned drone to prove the technology. Research into designtechniques for smart structures could help designers realize weight savings with clever designs, andavoid weight penalties with not so clever designs.

Avionics

Research into defining a growthpath of RF and digital processing upgrades for the JIAWG/Pave PillerClass of integrated Avionics would reduce overall development cost while minimizingthe weightgrowth as the system expands. Research into combined multispectralapertures and staring focalplane arrays would help reduce the developmental cost of advanced targeting FLIR, Integrated NavFLIR/IRST/MLD. Advanced multilayerwafer IC on ceramic substrate,planar slotted radiators,MMIC,and component & substrate integration research would help realize a 50% weight reduction for tiled

176

Array Radar. Offboard data management significantlyimpactsavionics system weight on the aircraft.Research into data fusion and into reducing RCS communications apertures and receiver sensitivitywould help realize a 50% weight reduction in avionics. One of the highest risktechnologies that couldbenefit from research is Integrated Sensor Systems (ISS) to produce common RF modules for further

reductions in avionics weight.

177

Vehicle Management System

The Vehicle Management System (VMS) is the integration of a large number of subtechnologies.High payoff areas of research are Phototonics,Improved Hydraulic System concepts, and the allElectric Airplane. Subsystem Utilities IntegrationTechnology (SUIT) research would help realize a50% weight reduction by understanding physical and functional integration, the suitabilityof differentfluids, energy utilization, and advanced packaging.

Crew Systems

One of the ground rules going into this study is that a single pilotwill be able to handle the task loadscurrently being handled by two man crews. Most of the reductionsin pilotworkload would be throughautomation and vastly improved displays. Research into helmet mounted displays, night visionsystems, panoramic displays, and 3-D audio would all contributeto the goal of reducingthe pilotsworkload and improving his situational awareness. Additional research into laser-hardeningtechnologies is necessary to protect the pilots survivabilityand missioneffectiveness.

Weapons

The signature requirements drove the need to carry the design weapons load in internal weaponsbays. These internal bays have substantialweight and volume penalties that drive up aircraft size andweight. Research into low signature weapons to replace the current inventory weapons is stronglyrecommended. The inherent low drag relative to conventional weapons would contributeto smaller,lighter, and cheaper aircraft. Conformal carriage of these reduced signature weapons would reducepylon weight and interference drags further reducing aircraft size and weight. Research into "AllEnvelope" Air-to-Air Weapons combined with aircraft agilitywould significantlyimprove theeffectiveness of fighter in an air-to-air engagement.

Unique Naval Aircraft Technology Requirements

The requirement for Navy aircraft to operate and be based on aircraftcarders severely limitsthe aircraftgeometry and penalizes the aircraft weight. It is very difficultfor the aircraft designer to develop anaircraft design competitive with its contemporary land based counterparts. Research into methods toexpand the design envelope of carder based aircraft would have significant impact on the combateffectiveness of naval aircraft. Improvements in the carder elevators, catapult and recovery systemsare one obvious means of increasingthe capability of aircraft designed for the carriers. Anotherapproach is the use of technology to reduce design margins required to maintain the same or bettersafety levels. One possible example would be the development of an automated carrier landingsystem to reduce risk and aircraft loss in carrier landingin all types weather conditions.

178

Yaw Thrust Vectoring Nozzle with Augmentation

Each of the configuration concepts utilize a thrust vectoring rotating nozzle to produce yaw/directionalcontrol moments. The basic rotating nozzle concept is used in smaller scale on the Pegasus enginein Harrier aircraft. The nozzle is a fixed throat and is actuated at required rates by an air driven motorthrough a chain drive system.

The applicationof a rotating nozzle for dedicated yaw control power will require the development of adrive system capable of generating both rate and response appropriate for precise vehicle controlrequirements. Application of this yaw thrust vectoring concept and mechanization have beenexplored by The Boeing Company under past proprietary study work and is currently in process ofdisclosure proceedings for submital to the U.S. Patent Office.

Integration of a dry power/fixed throat nozzle, although not without risk, is considered to beachievable. However, integration of engine thrust augmentation and providing a functional variablethroat rotating exhaust nozzle introducesa challenging high risk element into the system. No priorwork has been undertaken to describe the approach or the concept(s) that could be utilized toachieve this capability.

The Air Superiority and Multirole type vehicles sized under this study require augmentation in ordertoachieve the stipulated performance. The attractionof this nozzle resides in elimination of thevertical/canted tails used in conventional designs, thereby reducing observable signature levels andusing direct engine thrust for assured yaw control power throughout the flight envelope.

A potential validation path for developing this concept is shown in Figure 7.0-1. This summaryoverview addresses both the nozzle and yaw vane concept development, testing and evaluation.

The YF-23 (ATF Prototype) is considered to be a logicalflight research candidate aircraft for actual fullscale testing and evaluation of the proposed yaw control effector system concept described herein.

Figure 7.0-2 shows how the concept could be employed by modifyingthe existing aircraft aftfuselage. This application could be a phased program that undertakes the research and developmentof a drythrust nozzle initiallyfollowed by a parallel effort to produce the augmented engine variablethroat nozzle.

The expected results of this research and development would show effective and direct comparisonsfor observable signature changes when removing canted tails, flying qualities with vectoring in yawaxis, experience with advanced materials application such as Titanium Matrix Composite (TMC) andAdvanced Carbon-Carbon (ACC) in the exhaust system, and flight control system limitations withpowerful vectoring nozzle integration.

179

Flight Research Needs Assessment

Analysis methods weak

• Conceptual level Products of Inertia• Determining structual flutter boundaries• Non-linear aerodynamics• Engine transient response for bolters and acceleration performance

Control effectors - Continue quest for new ideas

• Wind Tunnel Database Development Quantify benefitsOptimal configurationsPenalties (blowing)

• Tiperons• All blowing and suctiondevices• Forebody strakes/chimes• Fluidic thrust vectoring

• Aerodynamics

EmpericalMethods

• Nonlinear aerodynamics *update DATCOM)• Use CFD to develop design methods (base drag)• Emperically corrected low order panel codes for conceptual design

CFD Methods

• Improve transition prediction• Continue validation of CFD methods• Improve turn-around and ease of use.

180

POTENTIAL YAW THRUST VECTORING RESEARCH PROGRAM

...I.

CO

ACTIVITYy 1 2 3 4 5

iQ 1121314 1121314 1121314 1121314 1121314

• Concept/DesignDevelopment

• Cold Gas Flow T&E

• Design Optimization

• Hot Gas T & E - FAB,Test & Evaluation

Flight Article- Nozzle system

- Design- Fab. & Assembly- Ground testing

- Air Vehicle

- Design- Fab. & Assembly- Nozzle Integration- Ground test

Flight test

Assumptions• YF-23 Aircraft is obtainable, flyable,

and compatible with proposed modificationas shown on Fig. 7.0-2

I DRY AUG.

coPo

Attributes:

• Airframe in flyable storage(mayber in NASA possession)

• Could modify by replacing aftfuselage that incorporatesTVRN (mode of TMC), horizontals

in lieu of V-tails, and yaw vane concept

S = 950 It2

b = 43.604 It = 523.25 in

Cr = 481.2 in

C,_ = 39.6in

_t = 0.08

AR = 2.0

,,- 40 ° J\

\

• Incorporate yaw thrustvectonng nozzle _ 45 degdeflection

• Veetail replaced by horizontal-retain/match plan view projectedarea of vector

Yaw Vectoring Concept on YF-23 as Research Article

spb-7-7-re-C3

I Form ApprovedREPORT DOCUMENTATION PAGE OMBNo.0704-0188

Public reporting burden tol' lh L,,co41eOtion o_ kdo_lon Is HImI, ted to rage 1 ho_r per responme, irrciudhllg the tirrm lot revwng ir_ltruotwts, =eau,¢_tk_g e3dsUng data =_=t ¢er_

_tMrorlg dllfKJ_li_ the _ _ _ _ _ _i_ the _ _ It_o(Irqr_at_r_. SeN_d cofflmsntI re_rdillO th_ 19_t'Q_i¢_i_imlto or any other _ _ _i

oo_ection of ktlom_tto_, Including =u_lon= tot reducing this burden, to Wmh_tgton Headquarters Sen_=es, DIrecmr'ae for Irdom_t_n O_ and Reports, 1215 Jelter=_ Day _"

Highway, Sulto 1204, A._,_.._-_., VA _ and to the Office o_ I_ and Buciget, Paperwork Reductto_ Pro_ (0704-0t88), Wlh/_, DC 20r==03.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

June 1995 Contractor Report (September 1994)

4. TITLE AND SUBTITLE

InvestigationIntothe Impactof Agilityon ConceptualFighterDesign

i6. AUTHOR(S)

R. M. Engelbeck

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

Boeing Defense & Space GroupSeattle, WA 98124

9. SPONSORINGI MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Langley Research Center

Hampton, VA 23681-0001

5. FUNDING NUMBERS

505-68-70-09

NAS1-18762

Task 22

8. PERFORMING ORGANIZATIONREPORT NUMBER

:10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

NASA CR-195079

I1. SUPPLEMENTARYNOTES

Contract Technical Monitor: M.J. Logan, NASA Langley Research Center, Hampton, VA 23681-0001

12a. DISTRIBUTIONI AVAILABILITYSTATEMENT

UNCLASSIFIED " UNLIMITED

Subject Category 05

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13. ABSTRACT (Maximum 200 word_)

The Agility Design Study was performed by the Boeing Defense and Space Group for the

NASA Langley Research Center. The objective of the study was to assess the impact of agility

requirements on new fighter configurations. Global trade issues investigated were the level

of agility, the mission role of the aircraft (air-to-ground, multi-role, or air-to-air), and

whether the customer is Air Force, Navy, or joint service. Mission profiles and design

objectives were supplied by NASA. An extensive technology assessment was conducted to

establish the available technologies to industry for the aircraft. Conceptual level

methodology is presented to assess the five NASA-supplied agili,ty metrics. Twelve

configurations were developed to address the global trade issues. Three-view drawings,inboard profiles, and performance estimates were made and are included in the report. Acritical assessment and lessons learned from the study are also presented.

14.SUBJECTTERMS

Agility,Fighters,ConceptualDesign

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UNCLASSIFIED

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OF THIS PAGE

UNCLASSIFIED

19. SECURITY CLASSIFICATIONOF ABSTRACT

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190

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Investigation Into the Impact of Agility in Conceptual Fighter Design