NASA Contract Report 177959

C. R. McGehee

Design Verification and Fabrication

of Active Control Systems

for the DAST ARW-2 High Aspect

Ratio Wing, Part 1 /x;-_

The BOEING Co.

Wichita, KS 67277

Contract NAS 1-16010

January 1986

[};ASA-CR- 177959-P t- I) DE SZGN yiPIFIC&Z_u[_ R88-I_!01

AND FABRICATIO_ 'OF AC_ZVE CONT_CL SYGf/,IS

I:OE THE DAST ARW-2 HIGH ASPECT }_AT!O ,if{G,

P_RT I (Boeinq Co.) 334 P CSCL UIC Unclas G]/08 0 117222

Date for general release - January 1988

National Aeronautics and

SpaceAdministration

LangleyResearchCenter

Hampton. Virginia 23665
I .

2.

3.

3.1

3.2

3.2.1

3.2.2

3.2.3

3.2.4

3.2.5

4.

4.1

4.2

4.3

4.4

5.

5.1

5.2

5.3

5.4

5.5

6.

6.1

TABLE OF CONTENTS

INTRODUCTION

SUMMARY

INTEGRATED DESIGN METHODOLOGY

Ffnal Integrated Design Cycle Completion

ARW-2 Performance and Synthesis Criteria

Active Central System Guidelines

ARW-2 Performance Objectives

SystemSynthesis Criteria

Critical Design and Test Conditions

Flight Operation

AEROELASTIC AND DYNAMIC ANALYSIS

Structural Stiffness Update

Aerodynamics for Wing Load Analysis

Aerodynamics for Flutter Analysis

Gust and Maneuver Loads

STABILITY AND CONTROL ANALYSIS

QSE Stability Derivative Estimation

Speed Brake Requirements

Neutral Point

Minimum Trim Drag

Comparison of QSE and Flexible Equations of Motion

ACTIVE/AUTOMATIC FLIGHT CONTROL

Performance and Synthesis Criteria

25

31

41

41

41

41

43

45

46

5O

51

51

51

52

52

61

61

65

65

65

77

81

82
6.2

6.2.1

6.2.2

6.2.3

6.3

6.3.1

6.3.2

6.3.3

6.3.4

6.3.5

6.3.6

6.4

6.4.1

6.4.2

64.3

64.4

65

66

66.1

66.2

66.3

67

TABLEOFCONTENTS(Continued)

Servoactuator Models

Initial Servoactuator Model

Servoactuator BenchTests

Final Servoactuator Model

Flutter Suppression SystemSynthesis

Integration Design Study

System Criteria

Final Flutter Suppression System

Flutter Suppression System Performance

FSSSensitivity

SystemsCompatibility

Load Alleviation SystemAnalysis

Gust Load Alleviation System

ManeuverLoad Alleviation system

Load Alleviation SystemPerformance

Simulated Gust Analysis

Relaxed Static Stability SystemAnalysis

Automatic Flight Control Systems

Primary Longitudinal AFCS

Lateral Directional AFCS

Backup Longitudinal AFCS

Horizontal Tail Authority Evaluation

83

83

89

99

I06

I06

llO

ll2

129

136

162

162

164

168

176

176

194

199

199

201

210

221
6.8

6.8.1

6.8.2

6.8.3

6.8.4

6.9

.

7.1

7.2

7.3

7.3.1

7.3.2

7.4

7.4.1

7.4.2

7.4.3

7.4.4

7.4.5

7.4.6

7.4.7

7.4.8

7.4.9

TABLE OF CONTENTS (Continued)

Control System Compatibility

Altitude Hold System Compatibility

System Gain and Phase Margins

Flying Qualities

Airplane Stability Derivative Sensitivity

Computer Implementation of Ground Based Control

Systems

ELECTRICAL/ELECTRONIC DESIGN AND MODIFICATION

Design Philosophy

ACS Electronics Box Design

Flutter Suppression Systems Sensors

Sensor Performance Requirements

Sensor Selection

Circuit Design

Signal Conditioning

FSS Signal Shaping

FSS Parameter Scheduling

GLA Signal Shaping

GLA Gain Scheduling

RSS Signal Shaping

RSS Gain Scheduling

Function Generator

Pseudo Random Noise Generator

221

225

230

230

234

234

245

245

261

261

261

264

264

272

277

279

279

285

285

285

285

289
7.4.10

7.5

7.5.1

7.5.2

7.6

7.7

7.8

7.8.1

7.8.2

7.8.3

8.

8.1

8.2

9.

9.1

9.1 .l

9.1.2

9.2

9.2.1

9.2.2

9.2.3

9.2.4

TABLE OF CONTENTS (Continued)

Servo Electronics

ACS Modifications

Pre-Delivery Modifications

Post-Delivery Modifications

ACS Power

ACS Electrical Wiring

Special Test Equipment

ACS Tester

Electronics Card Tester

Additional Ground Support Equipment

COMPONENT FABRICATION

Electronic Components

Spares

FLIGHT WORTHINESS TESTING

Component Functional Tests

Servoactuator System

Electronic Components

Flight Assurance Test Results

Initial Performance

Electromagnetic Compatibility

Temperature/Altitude

Vibration Tests

290

291

291

291

292

292

309

309

309

309

313

313

313

319

319

319

320

320

320

321

323

324
9.2.5

9.2.6

I0.

II.

APPENDIXA:

APPENDIXB:

APPENDIXC:

APPENDIXD:

APPENDIXE:

TABLEOFCONTENTS(Continued)

Interface Test

Final Performance

FLIGHTRESTRICTIONS

CONCLUSIONADRECOMMENDATIONS

REFERENCES

*PART2

APPENDICES

DASTARW-2AERODYNAMIC,NERTIAANDTHRUSTDATA

DASTARW-2OUTBOARDAILERONSERVOVALVERELOCATION

STUDY

DASTARW-2FINALFLUTTERSUPPRESSIONSYSTEM

PERFORMANCE DATA

DASTARW-2SENSITIVITYANALYSISDATA

DASTARW-2AUTOMATICCONTROLSYSTEM(ACS)AND

AUTOMATICFLIGHTCONTROL SYSTEM (AFCS) ROOT LOCI

324

326

329

331

332

17

55

lOl

136

176

*Published under separate cover.
LIST OF FIGURES

F IGURE

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2

3

10

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20

21

DAST ARW-2 GENERAL ARRANGEMENT 26

DAST ARW-2 OPERATIONAL FLIGHT ENVELOPE 27

FUNCTIONAL BLOCK DIAGRAM OF THE REMOTELY PILOTED

RESEARCH VEHICLE 29

DAST ARW-2 INTEGRATED DESIGN METHODOLOGY 32

DAST ARW-2 ACS/AFCS SENSORS AND CONTROL SYSTEMS 33

EQUIPMENT INSTALLATION 34

DAST ARW-2 ACS DESIGN AND FLIGHT TEST CONDITIONS 35

DAST ARW-2 WING VERTICAL BENDING MOMENT REDUCTION WITH ACS 36

DAST ARW-2 COMPARISION OF WING ROOT VERTICAL BENDING MOMENT

PSD AND RMS FOR TURBULENCE AND RANDOM INBOARD AILERON

EXCITATION .37

CONTROL SURFACE MAXIMUM DISPLACEMENTS AND RATES 38

DAST ARW-2 INTEGRATED DESIGN METHODOLOGY - THIRD ITERATION 42

DAST ARW-2 LIMIT GUST VELOCITY FOR ACT SYSTEM DESIGN 44

DAST ARW-2 MANEUVER INPUT FOR ACT SYSTEM DESIGN 44

DAST ARW-2 STABILITY MARGIN REQUIREMENTS 45

SELECTED FLYING QUALITIES REQUIREMENTS, MIL-F-8785 CLASS IV,

CATEGORY B 47

DAST ARW-2 SHORT-PERIOD REQUIREMENTS - CATEGORY B FLIGHT PHASE 48

DAST ARW-2 OPERATIONAL FLIGHT ENVELOPE AND DESIGN CONDITIONS 49

DAST ARW-2 AERODYNAMIC PANELS 53

ITERATIVE PROCEDURE FOR MATCHING QSE STABILITY DERIVATIVES 54

AERODYNAMIC PANEL EFFECTIVENESS SCALERS 55

WING CONTROL SURFACE EFFECTIVENESS SCALERS FOR LOAD ANALYSIS

EQUATIONS OF MOTION 55
F I GURE

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23

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25

26

27

28

29

30

31

32

33

34

35

36

37

38

39

40

41

LIST OF FIGURES (CONTINUED)

WING CONTROL SURFACE EFFECT IVENESSS SCALERS FOR FLUTTER

ANALYSIS EQUATIONS OF MOTION

ARW-2 STEADY STATE (Ig) WING LOADS AT THE MANEUVER DESIGN

CONDITION

ARW-2 STEADY STATE (Ig) WING LOADS AT THE GUST LOAD

DES IGN CONDITION

ARW-2 WING MANEUVER LOADS - WITHOUT ACS

ARW-2 WING LOADS DUE TO GUST - WITHOUT ACS

DAST ARW-2 WING CONTROL SURI_ACE CORRECTION RATIOS

DAST ARW-2 INBOARD AILERON RIGID STABILITY DERIVATIVES

DAST ARW-2 OUTBOARD A ILERON R IG ID STAB IL 173'DER IVAT IVES

INBOARD AILERON ELASTIC CORRECTION FACTORS

OUTBOARD A ILERON ELAST IC CORRECT ION FACTORS

INBOARD AILERON QSE LIFT COEFFICIENT

INBOARD AILERON QSE PITCH MOMENT COEFFICIENT

OUTBOARD AILERON QSE LIFT COEFFICIENT

OUTBOARD AILERON QSE PITCHING MOMENT COEFFICIENT

ENGINE RPM REQUIRED FOR STEADY STATE FLIGI-CI"WITH SPEED BRAKE

VAR IAT IONS

DAST ARW-2 NEUTRAL _INTS

CRUISE CONDITION PITCH COEFFICIENT

DAST ARW-2 CENTER OF GRAVITY VARIATION WITH FUEL USAGE

DAST AWR-2 TRIM DRAG VARIATION WITH C.G. POSITION

COMPARISON OF QSE AND FLEXIBLE BASIC AIRPLANE CHARACTERISTIC

ROOTS

56

57

58

59

6O

62

63

64

66

67

68

69

70

71

72

73

74

75

76

78
FIGURE

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43

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46

47

48

.49

5O

51

52

53

54

55

56

57

58

59

60

61

LIST OF FIGURES (CONTINUED)

COMPARISON OF QSE AND FLEXIBLE I_£'DELRSS SYSTEM ROOT LOCUS AT

TI-E GLA DESIGN CONDITION 79

COMPARISON OF QSE AND FLEXIBLE MODEL RSS SYSTEM ROOT LOCUS AT

-MAXIMUM DYNAMIC PRESSURE CONDITION 80

CONTROL SURFACE MAXIMUM DISPLACEMENT AND RATE REQUIREMENTS 84

SYMMETRIC OUTBOARD AILERON DISPLACEMENT AND RATE REQUIREMENTS 85

SERVOACTUATOR REQU IREMENTS 86

DAST ARW-I SERVOACTUATOR FREQUENCY RESI:::_)NSE(ANALYSIS AND TEST) 87

DAST ARW-I ACTUATOR DISPLACEMENT FREQUENCY RESPONSE TEST DATA 88

SERVOVALVE TRANSFER FUNCTION 90

DAST ARW-I TEST AND ANALYSIS DATA COMPARISON 91

DAST ARW-I AND DAST ARW-2 HYDRAULIC ACTUATOR MODEL WITH SURFACE

COMPLIANCE AND ACTUATOR LEAKAGE 92

DAST ARW-2 SERVOACTUATOR PARAMETERS 93

DAST ARW-2 OUTBOARD SERVOACTUATOR WITH SERIES 30 SERVOVALVE

ROOT LOCUS 94

DAST ARW-2 OUTBOARD SERVOACTUATORWITH SERIES 30 SERVOVALVE

FREQUENCY RESPONSE 95

HYDRAULIC RESEARCH SERVOVALVE MODEL ARW-25 (SFECIAL) 96

DAST ARW-2 OUTBOARD SERVOACTUATOR WITH MODEL AR-25 SERVOVALVE 97

SERVOAC'FUATOR DOMINANT MODES COMPARISON W ITH SERIES 30 AND

AR-25 SERVOVALVES 98

OUTBOARD SERVOACTUATOR RESPONSES (AR-25 SERVOVALVE) 100

OUTBOARD SERVOACTUATOR ANALYSIS AND TEST RESPONSE

CHARACTER IST ICS 101

DAST ARW-2 OUTBOARD SERVOACTUATOR WITH MODEL AR-25 SERVOVALVE 102

ARW-2 OUTBOARD SERVOACIIJATOR PERFOIRMANCE 103

8
FIGURE

62

63

64

65

66

67

68

69

70

71

72

73

74

75

76

77

78

79

8O

81

82

LIST OFFIGURES(CONTINUED)

ARW-2INBOARDSERVOACTUATOR PEFOI_MANCE 104

MAXIMUMSURFACEDISPLACEMENTS ANDRATES 105

OUTBOARDAILERONRATEVERSUSHINGEMOMENT 107

DASTARW-2FLIGHTCONDITIONSANALYZED 108

FLUTTERBOUNDARYFORTHEFINAL (ITERATION3) FSSCONFIGURATION

WITHOUTANTISYMMETRI C GAINSCHEDULING 109

.DASTAWR-2SECONDITERATIONFSSBLOCKDIAGRAM 111

DASTARW-2FLLrI'TERSUPPRESSIONTESTANDDESIGNENVELOPE 113

DASTARW-2FLLFFI'ERSUPPRESSIONSYSTEMGAINANDPHASE

REQUIREMENTS 114

M IL-F-9490D STAB IL ITY MARG INS 115

CONTROL SURFACE DEFINITION FI'NAL.DESIGN ITERATION 116

SENSOR POSITION SELECT ION METHODOLOGY 118

SYMMETRIC SENSOR ZERO ROOT LOCI AND FLUTTER MODE DAMPING 119

ANTISYMMETRIC SENSOR ZERO ROOT LOCI AND FLUTTER MODE DAMPING 120

FLUTTER SUPPRESSION SYSTEM FILTER SYNTHESIS PROCEDURE 122

FINAL DAST AI_W-2 FLLFFTER SUPPRESSION SYSTEM 123

SYMMETRIC FREQUENCY RESPONSE OF FSS SHAPING FILTER 124

ANTISYMMETRIC FREQUENCY RESPONSE OF FSS SHAPING FILTER 125

DAST ARW-2 PHASE-GAIN ROOT LOCUS. OF SYMMETRIC FSS,

NOM INAL SYSTEM 126

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS,

NOMI NAL SYSTEM 127

ANTISYMMETRI C SCHEDULER 128

DAST ARW-2 SYMMETRIC ELASTIC MODE DAMPING AND FREQUENCY 130
LIST OF FIGURES (COI'iTI I,JL3ED)

]0

FIGURE

83

84

85

86

87

88

89

9O

91

92

93

94

95

96

97

98

99

100

DAST ARW-2 ANT ISYMMETRIC ELASTIC MODE DAMPING AND FREQUENCY 131

DAST ARW-2 SYMMETRIC FLUTTER MODE DAMPING,

ALTITL_E: 15,000 FEET 132

DAST ARW-2 ANTISYMMETRIC FLUTTER MODE DAMPING,

ALTITUDE: 15,000 FEET 133

DAST ARW-2 SYMMETRIC MODE DAMPING, MACH: 0.86 134

DAST ARW-2 ANT ISYMMETRIC MODE DAMPING, MACH: 0.86 135

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.8 AND 2,000 FEET ALTITUDE 137

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.83 AND 3,000 FEET ALTITUDE 138

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.83 AND 12,000 FEET ALTITUDE 139

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.86 AND 4,250 FEET ALTITUDE 140

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.91 AND 8,000 FEET ALTITUDE 141

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS WITH

PARAMETER SCHEDULING AT MACH 0.91 AND 8,000 FEET ALTITUDE 142

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS WITH

PARAMETER SCHEDULING AT MACH 0.80 AND 2,000 FEET ALTITUDE 143

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS WITH

PARAMETER SCHEDULING AT MACH 0.83 AND 3,000 FEET ALTITUDE 144

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS,

NOMINAL SYSTEM AT MACH 0.83 AND 12,000 FEET ALTITUDE 145

DAST ARW-2 PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS WITH

PARAMETER SCHEDULING AT MACH 0.86 AND 4,250 FEET ALTITUDE 146

DAST ARW-2 OUTBOARD AILERON DISPLACEMENT RESPONSE TO INBOARD

AILERON INPUT 147

DAST ARW-2 OUTBOARD AILERON RATE RESPONSE TO INBOARD AILERON

INPUT 148

DAST ARW-2 FSS SENSOR ACCELERATION RESPONSE TO INBOARD AILERON

INPUT 149
LIST OF FIGURES (CONTINUED)

FIGURE

101

102

103

104

105

106

107

108

109

110

111

112

113

114

115

116

117

118

119

120

INBOARD AILERON DISPLACEMENT RESPONSE TO GUST GENERATOR INPUT 150

ROOT LOCUS OF SYMIVETRIC AND ANTISYMMETRIC FSS WITH SERIES 30

SERVOVALVE 152

PHASE-GAIN ROOT LOCUS OF SYMMETRIC AND ANTISYMMETRIC FSS

WITH AR-25 SERVOVALVE 153

PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS - h_MINAL SYSTEM 154

PHASE-GAIN ROOT LOCUS OF SYMMETRIC FSS - EFFECTS OF

UNCOMPENSATING ACTUATOR 155

PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS - EFFECTS OF

UNCOMPENSATING ACTUATOR 156

PHASE-GAIN ROOT LOCUS OF ANTISYMMETRIC FSS - NOMINAL SYSTEM 157

SYMMETRIC FSS SENSITIVITY TO SURFACE-ACTUATOR MODE FREQUENCY

CHANGE 158

HINGE MOMENT SENSITIVITY OUTBOARD SERVOACTUATOR GAIN AND PHASE

RESPONSE @ 23.8 HZ 160

EFFECT OF HINGE MOMENT ON ANTISYMMETRIC FSS, UNCOMPENSATED

ACTUATOR 161

DAST ARW-2 SYMMETRIC SYSTEM COMPATIBILITY 163

WING ROOT VERTICAL BENDING MOMENT FOR GLA SYSTEM 165

STAB IL IZER RATE REQUIREMENT COMPARI SON 166

GLA STABILIZER ROOT LOCUS COMPARISON MAXIMUM DYNAMIC PRESSURE

CONDITION 167

GLA LOAD ALLEVIATION SYSTEM BLOCK DIAGRAM 169

FUSELAGE VERTICAL ACCELERATION AND INCREMENTAL WING ROOT

VERTICAL BENDING MO_,iENTPSD/RMS WITH GLA SYSTEM 170

WING ROOT VERTICAL BENDING MOMENT/GUST COMPARISON AT GLA TEST

AND DESIGN CONDITIONS 173

WING ROOT VERTICAL BENDING MOMENT FOR MLA SYSTEM 174

MLA INBOARD AILERON SYSTEM BLOCK DIAGRAM 175

LOAD FACTOR AND WING ROOT VERTICAL BENDING MOMENT RESPONSES

FOR DES IGN MANEUVER 177

11
LIST OF FIGURES (CONTINUED)

"12

FIGURE

121

122

123

124

125

126

127

128

129

130

131

132

133

134

135

136

137

138

139

140

GUST DESIGN LIMIT WING SHEAR, VERTICAL BENDING MOMENT AND

TORS ION 180

MANEUVER DESIGN LMIT WING SHEAR, VERTICAL BENDING MOMENT

AND TORSION 181

MINIMUM MARGIN OF SAFELY FOR CRITICAL COMPONENTS 182

INBOARD A ILERON RANDOM EXC ITATI ON SYSTEM 183

WING ROOT VERTICAL BENDING MOMENT FREQUENCY DISTRIBUTION

COMPARI SON OF RANDOM GUST AND INBOARD AILERON RANDOM

EXCI TAT ION 185

PREDICTED INCREMENTAL WING ROOT VERTICAL BENDING MOMENT

REDUCTION WITH GLA FOR MAXIMUM AVAILABLE It_BOARD AILERON

AND EQUIVALENT GUST 186

MAXIMUM PREDICTrED TOTAL WING ROOT VERTICAL BENDING MOMEIxrT

WITH INBOARD AILERON RANDOM EXCITATION 187

WING ROOT VERTICAL BENDING MOMENT PSD/RMS PLOTS FOR INBOARD

AILERON RANDOM EXCITATION 188

WING ROOT VERTICAL BENDING MOMENrT PSD/RMS PLOTS FOR RANDOM GUST 189

GLA II_ST AND DESIGN CONDITIONS ANALYSIS PREDICTIONS 190

VERTICAL ACCELERATION RESPONSE TO RANDOM VERTICAL GUST 191

VERTICAL ACCELERATION RESPONSE TO INBOARD AILERON RANDOM

EXC ITAT ION 192

INBOARD AILERON RANDOM EXCITATION PROBABILIIY DISTRIBUTION

FACTOR 193

PITCHING MOMENT COEFFICIENT VS. ANGLE-OF-ATTACK AT CRITICAL

STABI L ITY CONDI TIONS 195

RSS SYSTEM BLOCK DIAGRAM 196

CLOSED LOOP ACS AND AFCS SHORT PERIOD CHARACTERISTICS 197

CLOSED LOOP ACS AND AFCS PHUGOID CHARACTERISTICS 198

PRIMARY LONG ITIJDINAL AFCS BLOCK DIAGRAM 200

RESPONSE AT LAUNCH CONDITION FOR POSITIVE LOAD FACTOR COMMAND 202

RESPONSE AT LAUNCH CONDITION FOR NEGATIVE LOAD FACTOR COMMAND 203
FIGURE

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142

143

144

145

146

147

148

149

150

151

152

153

154

155

156

157

158

159

160

LIST OF FIGURES (CONTINUED)

RESPONSE AT GLA TEST CONDITION FOR POSITIVE LOAD FACTOR COMMAND

RESPONSE AT GLA TEST CONDITION FOR NEGATIVE LOAD FACTOR COMMAND

RESPONSE AT MLA TEST CONDITION FOR POSITIVE LOAD FACTOR COMMAND

RESPONSE AT CRUISE FOR POSITIVE 0.5g LOAD FACTOR COMMAND

RESPONSE AT CRUISE FOf_F_SITIVE 1.5g LOAD FACTOR COMMAND

RESPONSE AT MAXIMUM DYNAMIC PRESSURE FOR POSITIVE LOAD

FACTOR COMMAND

BASIC AIRPLANE LATERAL-DIRECTONAL MODE CHARACTERISTICS

RUDDER/YAW RATE LOOP PHASE-GAIN ROOT LOCI

RUDDER/ROLL RATE LOOP PHASE-GAIN ROOT LOCI

LATERAL-DIRECTIONAL MODE CHARACTERISTICS WITH AFCS

LATERAL-DIRECTIONAL AUTOMATIC FLIGHT CONTROL SYSTEM BLOCK

DIAGRAM

BACKUP LONGITUDINAL AFCS BLOCK DIAGRAM

BACKUP MODE TRANSITION RESPONSE FROM PITCH UP MANEUVER AT

LAUNCH CONDITION

BACKUP I_DE TRANSITION RESPONSE FROM PITCH DOWN MANEUVER AT

LAUNCH CONDITION

BACKUP MODE TRANSITION RESPONSE FROM PITCH DOWN MANEUVER AT

GLA TEST CONDITION

BACKUP MODE TRANSITION RESPONSE FROM PITCH DOWN MANEUVER AT

MAXIMUM DYNAMIC PRESSURE

STABILIZER REQUIREMENT FOR LIMIT POSITIVE MANEUVER WITH C.G.

POSITION VARIATIONS

TIME RESPONSES TO ASYMMETRIC COI_I'ROL INPLrFSDURING LIMIT

SYMMETRIC MANEUVER

TIME RESPONSE TO ROLL COMMAND INPUT FROM LARGE BANK ANGLE

EFFECTS OF ALTITUDE HOLD LOOPS ON AIRPLANE CHARACTERISTIC

ROOTS

204

205

206

207

208

209

211

212

213

21,t

215

216

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218

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22O

222

223

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226

13
LIST OF FIGURES

FIGURE

161

162

163

164

165

166

167

168

170

171

172

173

1"/4

175

176

177

178

179

180

181

AIRPLANE VERTICAL RESPONSE DUE TO GUST Wlll_ ALTITUDE HOLD

ENGAGED 227

AIRPLANE RESPONSE Wlll-IALTITUDE HOLD ENGAGED IN A

10 DEGREE DIVE • • 228

EFFECTS OF STEADY STALE BANK ANGLE ON ALTITUDE HOLD PERFORMANCE 229

EFFECTS OF CLOSED LOOP SYSTEMS ON AIRPLANE MODES FREQUENCY

AND DAMPING 221

FLIGHT CONDITION - LONGITUDINAL SYSTEM STABILII_' ANALYSIS

MATRIX 232

SHORT PERIOD DAMPING AND FREQUENCY WITH PRIMARY AFCS AND ACTIVE

CONTROL SYSTEMS 233.

CLOSED LOOP SYSTEM SHORT PERIOD DAMPING AND FREQUENCY FOR

BACKUP MODE OF OPERATION 235

ROLL RESPONSE FOR STEP RUDDER WITH CZp VARIATIONS 236

AFCS AND MLA DIGITAL PROGRAM CONSTANTS 237

PRIMARY LONOITUDINAL AFCS FLOW DIAGRAM 238

PRIMARY LAIERAL-DIRECTIONAL AFCS DIGITAL PROGRAM 241

MLA INBOARD AILERON FUNCTION FLOW DIAGRAM 242

LIMIIER FOR PRIMARY AFCS AND MLA 243

DACE BLOCK DIAGRAM 246

LIST OF SYMBOLS FOR ll-IEACS BLOCK DIAGRAM 260

MOTION SENSOR DEFINITION 262

COMPOSITE ACCELEROMETER PERFORMANCE REQUIREMENTS 263

POIENTIOMEIER PERFOIRMANCE REQUIREMENTS 263

PRESSURE TI:V_NSDUCERPERFORMANCE REQUIREMENTS 2_4

SKETCH OF PCB PIEZOTRONICS MODEL 303A03 MINIATURE ACCELEROMETER 265

PCB 303A03 ACCELEROMETER PERFORMANCE SPECIFICATION 266

]4
LIST OF FIGURES(CONTINUED)

FIGURE

183

184

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188

189

190

191

Ig2

193

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196

197

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200

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202

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204

SUNSTRAND MODEL QAI IO0-AA01-12 ACCELEROMETER 267

SUNSTRAND QAI IC0-AA01-12 ACCELEROMETER PERFORMANCE

SPECI F ICATIONS 268

NEI 55FL-134 POTENTIOMETER PERFORMANCE SPECIFICATIONS 269

BELL & ik')WELLTYPE 4-326-0001 PRESSURE TRANSDUCER OUTLINE

D IMENS IONS 270

BELL 4 HOWELL TYPE 4-326-0001 PRESSURE TRANSDUCER PERFORMANCE

SPECI F ICATION 271

UPL INK TELEMETRY COMMANDS 273

UPLINK SIGNAL CONDITIONING CIRCUIT 274

DOWNL INK TELEMETRY SIGNALS 275

DOWNLINK SIGNAL CONDITIONING CIRCUIT 276

ALLOWABLE TOLERANCES FOR FLUT-rER SUPPRESSION SYSTEM FILTERS 278

ANALOG F I LTER MECHANI ZAT I ON 280

PARAMETER SCHEDULING FOR ALL MACH NUMBERS (NOT IMPLEMENTED) 281

BLOCK DI AGRAMOF ANTISYMMETRIC GAIN SCHEDULER 282

PROPOSED SCHEMATIC OF ANTISYMMETRIC GAIN SCHEDULER 283

ALLOWABLE TOLERANCES FOR GUST LOAD ALLEVIATION SYSTEM FILTERS 284

ALLOWABLE TOLERANCES FOR RELAXED STATIC STABILITY SYSTEM FILTERS 266

FUNCTION GENERATOR BLOCK DIAGF,AM

CRESTRONICS MODEL PS 333 DC TO DC CONVERTS PERFORMANCE

SPECI F ICATIONS

DAST ARW-2 ACS W IRING

POWER INpLrI"CONNECTOR (J2)

UPL INK CONNECTOR (J5)

RIGHT WING CONNECTOR (J1)

LEFT WING CONNECTOR (J6)

288

293

294

295

296

298

300

15
LIST OF FIGURES (CONCLUDED)

F IGURE

205

206

207

208

209

210

211

212

213

214

215

216

DOWNL INK CONNECTOR (J4)

NASA AFCS CONNECTOR (J3)

TEST CONNECTOR (J7)

ACS TESTER

ELECTRONICS CARD TESTER

AFCS ELECTRONICS BOX,

ACS ELECTRONICS BOX,

ACS ELECTRONICS BOX,

ACS ELECTRONICS BOX,

FROkq" V IEW

LEFT SIDE - REAR VIEW

TOP VIEW

TOP COVER REMOVED

DAST ARW-2 FLUTTER SUPPRESSION SYSTEM SPARE COMPONENTS

EMC PASS-FAI L SUmmARY

SERVOACTUATOR RERFORMANCE SUMMARY

302

304

305

310

311

314

315

316

317

318

322

325

]6
NOMENCLATURE

ACS

AFCS

AR

ARW-1

ARW-2

b

BCS

BL

BS

CA

CD, CB,CL

C.G.

Ch

Active Control System

Automatic Flight Control System

Aspect Ratio

Aeroelastic Research Wing Number i

Aeroelastic Research Wing Number 2

Wing Span, Inch

Backup Control System

Buttock Line, Inch

Body Station, Inch

Mean Aerodynamic Chord Length, Feet

Actuator Coefficient, In 3

Force Coefficients-Drag, Sideslip and Lift, Respectively

Center of Gravity

Control Surface Hinge Moment Coefficeint

ChO,ChS6MAX Control Surface Hinge Moment Coefficient at Zero

and at Maximum Trailing Edge Dow_ Deflection, Respectively

CI,CN,CM

c(E)

c(R)

CM/4

Moment Coefficient-Roll, Yaw and Pitch, Respectively

Elastic (QSE) Coefficient

Rigid Body Coefficient

Pitching Moment Coefficient at 1/4 Chord

(25 Percent MAC)

Control Surface Mean Chord Length, Inch

Force or Moment ({7)_Coefficient Derivative with Respect

to Motion Variable (_)

_7
NOMENCLATURE

DACE

DAST

dB

D.C.

DEQ

DOF

DT

EOM

EMC

FPS

FS

FSS

g

GLA

GW

H

HM

Hz

i

IEQ

Drone Active Control Electronics

Drones for Aerodynamic and Structural Testing

Decibel

Direct Current

Equivalent Viscous Damping of Hydraulic Fluid Trapped

Between Servo Valve and Actuator (In-Lb)/(Rad/Sec)

Degree of Freedom

Time Increment, Sec.

Equations of Motion

Electromagnetic Compatibility

Feet Per Second

Flutter Suppression

Flutter Suppression System

Gravitational Constant, g : 32.17 Ft/Sec 2

Gust Load Alleviation

Gross Weight, Lbs.

Altitude, Feet

Altitude Rate, FPS

Hinge Moment, In-Lb

Hertz (Cycles per Second)

Current, Amperes

Equivalent Rotary Mass Moment of Inertia of Hydraulic

Fluid Trapped Between Servo Valve and Actuator,

In-Lb-Sec 2

18
NOMENCLATURE

I

S

i w

Ixx, Iy,l Z

I

XY

J

KA

KCAS

KF

Kp

KS

EQ

KV

l

L

mA

MAC

MARS

MC

MLA

MMO

mV

MX

Rotary Mass Moment of_Inertia of Control Surface Relative To

Hinge Line, In-Lb-Sec _

Wing Incidence

Moment of Inertia About X, Y and Z Axis, Respectively,

Slug-Ft 2

Product of Inertia with Respect to X and Y Axis, Slug-Ft 2

Square Root of Minus One (_/_T_

Servo Valve Drive Amplifier Gain, mA/Volt

Knots Calibrated Airspeed

Servoactuator Position Feedback Gain, Volt/Deg

Servoactuator Pressure Feedback Gain, Volt/psi

Equivalent Rotary Spring Rate of the Actuator Shaft and OneThird Estimated Control Surface Spring Rate in Series,

In-Lb/Rad

Servo Valve Flow Gain, In3/Sec/mA

Rotary Actuator Vane Length, Inch

Random Gust Scale Length, foot

Milliamperes

Mean Aerodynamic Chord, Inch

Mid Air Recovery System

Cruise Mach Number

Maneuver Load Alleviation

Maximum Operating Mach Number

Millivolt

Vertical Bending Moment (+ Tip Up), In-Lb

19
NOMENCLATURE

My

NZ

A NZ

P

Pa

PCS

PL

PS

psi

psf

PSD

Q

Q

q

Qo

Qc

QSE

R

rD

RMS

RPM

RPRV

Torsional Moment (+ Leading edge Up), In-Lb

Normal Load Factor, g

Incremental Load Factor, Nz-1, g

Roll_Rate About the Longitudinal Axis, (+Right Wing Down),

Deg/Sec

Ambient Pressure, psi or psf

Primary Control system (Primary AFCS)

Actuator Load Pressure, psi

Hydraulic Supply Pressure, psi

Pounds per Square Inch

Pounds per Square Foot

Power Spectral Density

Hydraulic Fluid Flow Rate, In3/Sec

Pitch Rate About the Drone Lateral Axis (+ Nose Up), Deg/Sec

Dynamic Pressure, psf

Servo Valve No-Load Flow Rating, In3/Sec

Impact Pressure, psi or psf

Quasi-Static Elastic

Yaw Rate About the Drone Vertical Axis (+ Nose Right),

Deg/Sec

Rotary Actuator Vane Radius, Inch

Root Mean Square

Revolutions per minute

Remotely Piloted Research Vehicle

20
NOMENCLATURE

PRS

r S

RSS

S

S

SL

SS

T

TE

TTL

T1/2,T2

u

Uo

v

V

Vc

VD

VDC

Vf

VMO

Vv

VZ

Radians per second

Rotary Actuator Shaft Radius, Inch

Relaxed Static Stability

Surface Area, Ft 2

Laplace Transform Variable, Rad/Sec

Sea Level

Control Surface Area, In 2

Thrust, Lb

Trailing Edge

Transistor-Transistor Logic

Time to Half and Double Amplitudes, Respectively, Second

True Airspeed, Ft/Sec

Reference Airspeed, Ft/Sec

Volume of Hydraulic Fluid Trapped on One Side Servovalve

and Actuator, In 3

Volts

Cruise Speed, Ft/Sec

Dive Speed, Ft/Sec

Direct Current Voltage

Flutter Velocity, Ft/sec

Horizontal Tail Volume Coefficient

Maximum Operating Velocity, Ft/Sec

Vertical Tail Volume Coefficient

Vertical Shear (+ Up), Pounds

21
W

WBL

WRVBM

WS

#

#

},

61A'6OA

6STAB

E

C

Cd

Cp

_'sp

Cv

7/p

r/w

NOMENCLATURE

Velocity Along Drone Normal Reference Axis (+ Down), Ft/sec

Wing Buttock Line, Inch

Wing Root Vertical Bending Moment (MX Wing Root), In-Lb

Wing Station - Distance Outboard Along Rear Spar from

Centerline of Body, Inch

Drnne Vertical Displacement from a Reference Altitude

(+ Down), Feet

Vertical Acceleration (+ Down), Ft/Sec2 or In/Sec2

Angle of Attack (+ Nose Up), Degree

Sideslip Angle (+ Right), Degree

Bulk Modulus, Lb/In 2

Glide Angle (+ Down), Degree

Inboard and Outboard Aileron Deflection, Respectively,

(+Trailing Edge Down), Degree

Symmetric Stabilizer Deflection from Steady-State Trim

Position (+ Trailing Edge Down), Degree

Downwash Angle, Degree

Damping Ratio

Dutch Roll Mode Damping Ratio

Phugoid Mode Damping Ratio

Short Period Mode Damping Ratio

Equivalent Second Order Damping Ratio of the Servovalve

Dynamic Pressure Effectiveness

Normal Wash Effectiveness

22
NOMENCLATURE

r/fb _r/wg ,r/ht

#

(7

_)

a) N

(an d

_np

"&_n

sp

ton

V

Aerodynamic Effectiveness Scalers (N( ) : Np( ) : NW( )

for the Forward Body, Wing and Horizontal Tail, Respectively

Drone Pitch Angle (+ Nose Up), Degree

Microvolt - 10 -6 Volts

Mass Density of Air, Slug/Ft 3

Real Part of Complex Numbers

Root Mean Square (RMS)

Roll Convergence Mode Time Constant

Roll Attitude and Rate, Respectively (+ Right Wing Down),

Degree

Heading Attitude and Rate, Respectively (+ Nose Right),

Degree

Frequency, Rad/Sec

Natural Frequency, Rad/Sec

Dutch Roll Mode Natural Frequency, Rad/Sec

Phugoid Mode Natural Frequency, Rad/Sec

Short Period Mode Natural Frequency, Rad/Sec

Equivalent Second Order Undamped Natural Frequency of the

Servo Valve

23
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24
1. INTRODUCTION

A preceding study was performed under National Aeronautics And Space

Administration (NASA) Contract NASI-14665 (Reference i) to establish an

integrated design methodology and apply it to accomplish the integrated

design of a high aspect ratio wing with active controls for flight test

on a BQM-34E/F (Firebee II) drone. The wing will be used to study

aeroelastic effects and to assess and validate the design and predicted

performance of active control systems.

This study, a follow-on to the above, was performed under NASA Contract

NASI-16010 with the objective of accomplishing the final design and

hardware fabrication for four active control systems compatible with and

ready for installation in the NASA Aeroelastic Research Wing No. 2

(ARW-2) and Firebee II drone flight test vehicle.

The four active control systems enclosed in the Drone Active Control

Electronics (DACE) unit are:

Flutter Suppression (FS)

Maneuver Load Alleviation (MLA)

Gust Load Alleviation (GLA)

Relaxed Static Stability (RSS)

This development effort also includes the definition of a ground based

Primary Control System (PCS) and the final design and hardware

fabrication of a Backup Control System (BCS) to be packaged and

compatible with an existing NASA design electronics unit.

NASA is pursuing studies of aeroelastic effects on wing loads and

integration of active control systems to reduce structural weight and

improve efficiency of transport aircraft. Under the on-going Drones

for Aerodynamic and Structural Testing (DAST) project, the ARW-2 is the

second wing to be evaluated. The overall effort involves wing and

hardware fabrication, analytical studies, and wind tunnel and flight

test evaluation. The DAST ARW-2 configuration general arrangement is

shown on Figure 1. Figure 2 shows the flight envelope.

The first wing to be designed for the Firebee II drone, ARW-1 was

designed by Ryan Aeronautical under NASA Contract NAS1-13451 to exhibit

flutter within the aircraft flight envelope. Preliminary and final

design of the Flutter Suppression System (FSS) was accomplished by

Boeing Military Airplane Company (BMAC) under Contracts NASI-14028

(Reference 2) and NAS1-14675 (Reference 3). The ARW-I configuration

went into flutter and crashed during a flight test. The test mishap was

attributed to nonlinear wing characteristics coupled with a gain

implementation error in the flutter system.

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25
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FIGURE 2

DAST ARW-2 OPERATIONAL FLIGHT ENVELOPE

27
The DAST ARW-2 test vehicle is air launched from a B-52B aircraft,

controlled from a ground based cockpit and recovered by air snatching

the recovery parachute with a helicopter. The vehicle will be flown as

a Remotely Piloted Research Vehicle (RPRV) similar to the illustration

shown on Figure 3 and described in Reference 1.

This report provides historical documentation and describes the

completion of the third iteration of the Integrated Design Methodology

initiated and discussed in Reference I. Completion of the FSS, MLA, GLA

and RSS systems synthesis, design of the electronic and mechanical

components required to mechanize the systems on the drone, results of

additional servoactuator actuator analysis, and the results of the

flightworthiness tests conducted prior to delivery of the system

components to NASA are contained herein.

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3O
2. SUMMARY

This is the final report on Contract NASI-16010 for the integrated

design of a high aspect ratio wing with active controls for flight tests

on a BQM-34E/F drone.

The integrated design methodology shown on Figure 4 was used to develop

a high aspect ratio wing design with integral critical control systems

to derive substantial benefits in terms of reduced wing structure weight

and stiffness. The active control systems are used for gust and

maneuver load reduction and flutter control to allow reductions in

structural strength and stiffness from conventional wing designs. The

third and final cycle of the interactive design method of Figure 4 was

completed during this study.

Performance goals and criteria were applied to individual systems and

the systems collectively to assure that vehicle stability margins,

flutter margins, flying qualities and load reductions were achieved.

Figure 5 shows the Active Control System (ACS) and Automatic Flight

Control System (AFCS) sensor and control surface locations and

identifies their use for system applications. The control system

equipment installation arrangement is shown on Figure 6.

L

Individual ACS system design and flight test conditions are shown on the

flight envelope on Figure 7.

Quasi-Static Elastic (QSE) and scaled and unscaled elastic structure

Equations of Motion (EOM) were used in the analysis and synthesis of the

ACS and AFCS systems. The QSE EOM were the primary model for evaluation

of airplane stability. The scaled elastic structure EOM matched basic

airplane dynamics to QSE characteristics and were used to determine wing

loads. The unscaled elastic structure EOM were used for flutter system

analysis.

Load alleviation systems were defined which reduce structural loads over

the critical inboard section of the wing. Load alleviation systems

include both maneuver and gust reduction systems. Wing vertical bending

moment reduction capability of the load alleviation system is shown on

Figure 8. Eighteen percent reduction in critical wing root vertical

bending was achieved with load alleviation control. The wing is gust

load critical over the entire span.

The load reduction capability of the load alleviation system will be

verified by flight tests. A system was implemented to generate random

excitation of the inboard aileron with a power frequency content similar

to wind turbulence in order to evaluate the gust load alleviation

system. Comparisons of the power spectral density and root mean square

characteristics of wing root vertical bending moment due to vertical

turbulence and random inboard aileron excitation are shown on Figure 9.

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A Relaxed Static Stability system (RSS) was defined to provide flight at

aft reduced trim drag, but unstable center of gravity positions. The

RSS provides stability and handling characteristics with C.G. shifts aft

to 33 percent mean aerodynamic chord (MAC) over the vehicle flight

region. This provides stability 18 percent aft of the neutrally stable

position at the critical flight condition.

A Flutter Suppression System (FSS) was synthesized to control symmetric

and antisymmetric flutter modes occurring in the high dynamic pressure

and Mach region within the vehicle flight envelope. Dual accelerometers

mounted on the front spar at wing buttock lines 82 and 92 and.dual

accelerometers mounted on the rear spar at buttock lines 84 and 92 are

combined and filter shaped to drive the outboard ailerons.

Longitudinal and lateral-directional primary automatic flight control

systems were synthesized to provide adequate vehicle handling qualities

for a ground based operator. Backup automatic flight control systems

provide minimal control capability from a chase plane.

Maximum displacement and rate requirement for the control surfaces used

for ACS and AFCS functions are summarized on Figure 10. The values on

Figure 10 are within the design capability of the vehicle surface

actuation systems at the critical design cond4tions.

CONTROL SURFACE

OUTBOARD AILERON

INBOARD AILERON

SYMMETRIC STABILIZER

(FROM TRIM)

DISPLACEMENT

(DEGREES)

_+15

20 T.E.D.

10 T.E.U.

RATE

(DEG/SEC)

640

50

17.2

FIGURE 10

CONTROL SURFACE MAXIMUM DISPLACEMENTS AND RATES

The major requirements accomplished under this contract are:

a. The design methodology to define the ARW-2 high aspect ratio wing

and supporting active control systems, developed and partially

completed under Contract NAS1-14665, was finalized. The finalized

ACS systems include a flutter suppression system, a relaxed static

stability system, and gust and maneuver load alleviation systems.

be

Co

Primary and backup AFCS were finalized.

The ACS and AFCS hardware was procured and/or fabricated, assembled

and tested.

38
In addition to this final report, the following items were delivered

under this contract:

a. ACScomponenthardware, electronics, ground support equipment, and

spare parts.

b. BCSprinted circuit assemblies.

c. ACSassembly, installation, checkout, maintenance and operation

procedure documentation.

d. Drawings and schematics of the ACSlayout and assembly.

e. Flight assurance and EMCtest procedures and test report.

39
(This page intentionally left blank)

4O
3. INTEGRATED DESIGN METHODOLOGY

The integrated design methodology to define the DAST ARW-2 wing and the

ACS is illustrated on Figure 4. Under the previous contract study, two

major design iterations and the initiation of a third were conducted as

discussed in Section 3 of Reference I. The previous contract study had

defined the ACS and a wing design. The last stress analysis results in

the previous contract study identified small modifications to the wing

design used in the last performance analysis. The changes were

incorporated in the final wing design but final performance analyses

were not conducted. With the possible exception of the FSS, the small

changes in the wing design were not expected to affect system

performance.

3.1 Final Integrated Design Cycle Completion. A final (third) system

performance analysis cycle conducted under the present contract isshown

on Figure 11. This final interactive cycle was initiated during the

previous study. The part of the final iteration conducted during the

previous study is indicated by the solid path of Figure 11. After the

final stress analysis of the previous study, the detailed structural

design and jig shape were defined because it was apparent that the

iterative design procedure had converged. The completion of the final

iterative cycle conducted during the present study is indicated by the

dashed paths shown on Figure 11.

The final performance analysis cycle included updated aerodynamic

modeling reflecting wind tunnel test data for the 0.237 scale ARW-2

model. The change in aerodynamics had a significant impact on final

wing loads as discussed in subsequent sections of this report.

In a task related to the efforts of this contract, the wing skin design

defined in the last contract study was revised by NASA to make the

torsional characteristics agree with that used in the mathematical

description of the wing design.

3.2 ARW-2 Performance and Synthesis Criteria. The ARW-2 design

criteria including design guidelines, performance objectives and active

control system synthesis criteria are presented in the following

paragraphs.

3.2.1 Active Control System Guidelines. The wing shall be designed to

use active control systems for control of gust and maneuver loads and

flutter to allow maximum reduction in structural strength and stiffness

from conventional wing design. This integrated design analysis effort

shall be performed in an iterative manner.

The active control system design techniques used should be applicable to

transport aircraft design.

41 I:_I:ECED_ PAGE BLANK NOT FILME_
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The wing structural design shall be such that it is necessary to have

the active control systems functioning in the normal operating flight

envelope. Minimization of trim drag through the use of relaxed static

stability is a goal.

System reliability and redundancy should be considered as appropriate

for unmanned, limited time operations.

3.2.2 ARW-2 Performance Objectives

3.2.2.1 Flutter Suppression System. The wing shall be designed to be

free of divergence and flutter at all speeds up to 120 percent of the

design dive speed (I.2VD) with the FSS operating. At the design dive

speed, Vn, at least one and one-half percent equivalent structural

damping Fatio (_ = .015) must exist for all flutter critical structural

modes. The wing without the FSS shall exhibit flutter within the flight

envelope.

3.2.2.2 Load Alleviation. Wing design limit maneuver and gust loads

shall be met with the MLA and GLA functions of the ACS operating. The

ACS system shall be load critical and reduce wing bending moment

approximately 20 percent. Design limit maneuver wing loads will be

determined for maneuvering load factors of +2.5 and -1.0 per FAR Part

25.

The atmospheric turbulence model to be used for ACS synthesis will be

Gaussian, stationary and isotropic random gust with the Von Karman

spectral density as specified in MIL-A-OO8861A. Characteristic gust

length of 2500 feet will be used for altitudes above 2500 feet.

The structure will be designed to meet limit gust velocity values from

MIL-A-OO8861A as shown on Figure 12 for the cruise speed (Vr). The

limit design response is obtained by multiplying the limit _ust velocity

by the response per unit gust as defined in MIL-A-OO8861A. Due to

limited hydraulic flow capacity on the drone test vehicle, the design

gust velocity at the flutter critical design dive condition will be 12

ft/sec.

Longitudinal maneuvers will be ramp style maneuvers as defined in

MIL-A-OO8861A for a transport. This maneuver is illustrated on Figure

13. The structure shall be sized to meet requirements of gust and

maneuver loads independently.

3.2.2.3 Relaxed Static Stability. Thegoal of the RSS system is to

provide longitudinal stability and flying qualities for the ARW-2 drone

with an aft center of gravity for minimum trim drag. A requirement is

for the drone to be stable at launch without the RSS system. If these

two requirements are incompatible, emphasis shall be on vehicle

stability at launch and theminimum trim requirement shall be relaxed.

A droppable mass ballast shall not be used to achieve minimum trim drag.

43

D_
• RANDOMTURBULENCE

- VON KARMAN SPECTRUM

e MIL-A-O08861A

70 - /--(62 FPS MAXIMUM @ 7,000 FEET)

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4O

,,-i,3O

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FIGURE 12

DAST ARW-2 LIMIT GUST VELOCITY FOR ACT SYSTEM DESIGN

• MANEUVER COMMANDS

- ±2.5g MAXIMUM

• MIL-A-OO8861A

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FIGURE 13

DAST ARW-2 MANEUVER INPUT FOR ACT SYSTEM DESIGN

44
The RSSsystem shall provide the intent of Level 1 flying qualities as

d_fined in MIL-F-8785B(ASG)and summarizedin Paragraph 3.2.3.3, as

appropriate for the DASTresearch mission. The intent is to achieve

required longitudinal drone stability with the RSS system and use the

longitudinal AFCS as the drone control loop.

3.2.2.4 Primary AFCS. The primary AFCS shall provide the intent of

Level 1 flying qualities in each axis as defined in MIL-F-8785B (ASG) as

appropriate for the DAST research mission with the existing BQM-34E/F

control surface authority. The flying qualities are summarized in

Paragraph 3.2.3.3.

3.2.2.5 Backup AFCS. The backup AFCS shall be designed to provide

minimum flying qualities necessary to return the drone to a safe Mid Air

Recovery System (MARS) recovery in case of failure in the primary AFCS

ground computer loop. Minimum flying qualities for the backup AFCS will

be at least Level 3 as outlined in Paragraph 3.2.3.3,

3.2.2.6 ACS and AFCS Normal Operation. Normal operation shall be with

all of the ACS and the primary AFCS operating. The AFCS and ACS shall

become active (engage) three seconds after separation from the launch

aircraft.

3.2.2.7 ACS and AFCS Abnormal Operation. The ACS systems shall remain

operative when the AFCS switches to the onboard backup mode.

For failure of the GLA, MLA or FSS functions, the RSS function shall

remain engaged and flight restrictions shall be applied.

For loss of the RSS function the automatic recovery sequence shall be

initiated.

The backup AFCS must hold the drone in a recoverable attitude for up to

20 seconds in case of loss of control signal from either the ground

pilot or the F-104 backup pilot.

3.2.3 System Synthesis Criteria.

3.2.3.1 Stability Margins. All ACS and AFCS functions shall be

synthesized to meet MIL-F-9490D gain and phas e margins at the design

dive speed (Vn) as summarized on Figure 14. At 1.2V D all modes shall be

stable at nomTnal gain and phase.

MODE

f < 0.06 HZ

MANEUVER

STRUCTURAL

GAIN MARGIN

±3.0 dB

±4.5 dB

±6.0 dB

PHASE MARGIN

±20 o

+_.300

+450

FIGURE 14

DAST ARW-2 STABILITY MARGIN REQUIREMENTS

45
3.2.3.2 Reliability. Reliability and redundancy of the ACS shall be

such that the overall probability of vehicle abort and loss is not

significantly increased and is consistent with unmanned, limited time

operation. System redundancy shall be utilized as required to achieve

this goal.

3.2.3.3 Flying Qualities. The ACS and AFCS systems shall meet selected

flying qualities of MIL-F-8785B (ASG) for a Class IV aircraft, Category

B flight phase. The selected longitudinal and lateral-directional

flying quality requirements to be used for design are summarized on

Figure 15 and Figure 16.

3.2.3.4 Structural Safety Factor. A free flight safety factor of 1.5,

consistent with manned transport criteria per FAR Part 25, shall be used

for structural design.

3.2.3.5 Operational Environment. The active control system electronics

shall be designed for the altitude, temperature and vibration

environment specified in NASA-DFRC Process Specification 21-2

commensurate with 60,000 foot maximum test altitude.

3.2.3.6 Service Life. Design operational life of the active control

system electronics shall be 100 hours, assuming a maximum of 50 flights

over a five year period.

3.2.3.7 Electrical Power. Electrical power available for the wing

control surfaces required by the ACS is approximately 43 amperes without

modifications to the electrical system. The drone vehicle has 28 VDC

electrical power.

3.2.3.8 Ground Computer and Telemetry. The AFCS and ACS shall be

compatible with the ground computer and telemetry capabilities presented

in Reference 1.

3.2.4 Critical Design and Test Conditions. The ACS and AFCS critical

design conditions are shown on Figure 17. This figure shows the

airspeed and Mach number for V_, Vn and 1.2V n that are representative of

a transport aircraft. A fligh_ te_t placard_was set at Mach 0.86

because wind tunnel test data on the ARW-2 high aspect ratio

supercritical wing was not available above Mach 0.81 to support the

stability and control analysis.

The critical maneuver load design condition is the high angle of attack

(+HAA) condition at sea level for maximum gross weight with testing at

10,000 feet. The difference in wing loads between the sea level

condition and the test condition should be small because the upper left

hand corner of the V-n diagram for sea level and 10,000 feet are almost

identical. The lift coefficient at buffet does not change significantly

with Mach number in the 0.3 to 0.4 range.

46
PHUGOID MODE DAMPING

SHORT PERIOD DAMPING

DUTCH ROLL DAMPING

DUTCH ROLL FREQUENCY(RAD/SEC)

DUTCH ROLL FREQUENCY -

DAMPING PRODUCT (RAD/SEC)

ROLL MODE TIME CONSTANT (SEC)

SPIRAL STABILITY-TIME TO

DOUBLE AMPLITUDE (SEC)

LONGITUDINAL STATIC STABILITY

LEVEL 1

_p > 0.04

0.3 < _sp < 2.0

_d > 0.008

cond > 0.4

_dmnd > 0,15

TR< 1.4

Ts > 20

6s7V > 0

LEVEL 3

T2 > 55 SECONDS

{sp > 0.15

_d > 0.02

mnd _ 0.4

T R < I0.0

TS>4

6s/V > 0

FIGURE 15

SELECTED FLYING QUALITIES REQUIREMENTS, MIL-F-8785 CLASS IV, CATEGORYB

47
I00

_LUES OF n/_

OUTSIDE THE RANGE SHOWN ARE DEFINED

-- BY STRAIGHT-LINE EXTENSIONS

I

2

_nsp

n/cL

lO.O

lO

_SP "

RAD/SEC

0.085

O.038

0.I

1.0 I0

n/cz - g's/RADIAN

lO0

FIGURE 16

DAST ARW-2 SHORT-PERIOD REQUIREMENTS - CATEGORY B FLIGHT PHASE

48
ALTITUDE

- I000 FT.

60

50

40

30

20

lO

/-'IDESIGN CONDITIONS

1 .Og

GROSS WEIGHT

2200 LBS 2.5g I

2500 LBS._ I

..J

°/

0

0 .2 .4 .6 .8 1.0

MACH NUMBER

1.2

FIGURE 17

DAST ARW-2 OPERATIONAL FLIGHT ENVELOPE AND DESIGN CONDITIONS

49
The critical design and test condition for GLA is VC (350 KCAS) at 7000

feet altitude.

The FSS is designed to be flutter free to 1.2 VD at constant altitude

and at constant Mach number. Testing shall be within the flight test

envelope bounded by the placard at Mach 0.86 and VD = 440 KCAS.

The primary RSS design and test condition is the Mach 0.8 cruise

condition at 46,800 feet. The AFCS is required throughout the flight

envelope.

The design Cl gross weight is 2350 pounds. This allows use of 150

pounds of fu_l to launch and stabilize at the 46,800 foot test condition

for load testing.

3.2.5 Flight Operation. The BQM-34E/F drone uses a Continental

YJ69-T-406 engine with a rated static thrust of 1840 pounds at sea

level. Characteristics of the drone engine are such that operation

below 80 percent RPM is not recommended.

The DAST ARW-2 vehicle shall be launched from an inboard pylon on a B-52

aircraft.

The drone shall be installed on the launch vehicle pylon in an attitude

so that it will have a -O.5g load factor at separation.

A single launch condition will be used for ACS ahd load testing.

Altitude:

Mach Number:

Center-of-Gravity:

Airspeed:

Gross Weight:

15,000 feet

0.4

20 percent MAC

100 KCAS

2500 Pounds

The ACS system shall be inoperative during captive flight. The AFCS and

ACS systems will be designed to be inoperative for three seconds after

separation from the launch aircraft.

The control surfaces added to the ARW-2 wing for the flutter

suppression, maneuver load alleviation and gust load alleviation system

shall be held in the faired position during separation of the drone from

the launch aircraft. This requires that the hydraulic power supply unit

and the servoactuator feedback electronics be turned on.

The DAST ARW-2 vehicle shall use an all movable horizontal tail surface

as elevons for both pitch and roll control and a rudder on the vertical

tail surface for yaw control. The ACS shall use wing control surfaces

and tail surfaces as required. Capability shall be provided in the ACS

electronics to use the wing control surfaces for roll control during

separation, if required.

The ARW-2 vehicle shall be recovered by a helicopter using a Mid Air

Recovery System (MARS). The RSS and AFCS shall remain active until

the drag chute opens using a storage battery for electrical power.

5O
4. AEROELASTICANDDYNAMICANALYSIS

The final analysis cycle was conducted with small modifications to

structural stiffness and revised aerodynamics that match wind tunnel

test data of the 0.237 scale model. The revised aerodynamics, which

involved matching QSEstability derivatives and changes in control

surfacemodeling and effectiveness, significantly increased wing loads

at the gust load design condition. The revised structural stiffness did

not significantly affect wing loads. The airplane model also reflected

a relocation of fuselage ballast to shift the C.G. forward 12 percent

MACat maximumgross weight.

Twodifferent methods of incorporating aerodynamic effects into the

flexible EOMwere utilized. Onemodel was used for wing load analysis

and this model provided a close match of structural dynamic EOMto

elasticallycorrected wind tunnel data. The other model was used for

flutter analysisand provided an accurate description of airplane

structural vibration modedynamics.

The revised airplane model predicted a 10 percent increase in the

critical load parameter of wing root vertical bending moment (WRVBM) at

the design gust condition and a 0.5 percent increase in WRVBM for design

maneuver. The previous analysis had predicted that the inboard section

of the wing was maneuver load critical; but due to the increased gust

load, the entire wing became gust load critical.

The increased gust load consequently resulted in a increased load level

with active gust load alleviation system, and a structural resizing

iteration cycle to increase wing strength was anticipated. However,

because the critical inboard section of the wing was previously maneuver

critical and because a large amount of conservatism had existed in the

previous design iteration stress analysis, a resizing iteration cycle

was not required. Additional stress analysis was completed which

included the load carrying capability of previously excluded wing

section redundant structural elements. The refined analysis indicated

positive safety margins at all critical stress points.

4.1 Structural Stiffness Update. As a result of the last stress

analysis in the previous study, some changes to the size of the front

and rear spars were required. Also, skin thickness at the attachment to

the spar and rib was revised. These changes were included in the wing

design, but had not been included in the analysis of wing structure and

ACS performance.

4.2 Aerodynamics for Wing Load Analysis. A technique for scaling

lifting surface theory aerodynamics to match test data was developed

during the last iteration as outlined in Section 5.5.2 of Reference 1.

In the previous iteration cycle, flexible EOM contained aerodynamic

effectiveness scaling which approximately matched rigid 0.237 scale

51
model tunnel data. For this analysis cycle, the EOM for wing load

analyses were scaled to match QSE values for CM_, C,. , and C.. of

Reference 4. This match provided close agreem_Y1t o_QSE andm_l_TA_le

EOM characteristics.

Aerodynamic scalers were defined for the forward body, wing and

horizontal tail panels of Figure 18 using the iterative procedure shown

on Figure 19. The scalers for the QSE match at each design or test Mach

number are shown on Figure 20.

The wing control surface degrees of freedom in the loads analysis

equations of motion were scaled through the use of scalers that match

small surface planar doublet elastic Ci_ to the QSE w_C_e" The control surface effectiveness factors for the ARW-2 surfaces assumed to be

approximately the same as those for the smaller wind tunnel model

surfaces. The effectiveness scalars for each Mach number are shown on

Figure21.

4.3 Aerodynamics for Flutter Analysis. Equations 9f motion for flutter

analyses were developed using planar doublet aerodynamic influence

matrices without scaling to match rigid body stability derivatives to

either wind tunnel or QSE data. However, the wing control surface

degrees of freedom were scaled to match the rigid CIR values defined in

the loads analysis EOM. The effectiveness scalers _'_r flutter EOM are

shown on Figure 22. Flutter boundary definitions and flutter

suppression system analysis are presented in Paragraph 6.3.

4.4 Gust and Maneuver Loads. Total wing loads without the load

alleviation systems and steady state ig loads are shown on Figures 23

through 26 for gust and maneuver at the respective design conditions

Wing loads with the load alleviation systems are shown in Section 6.4.

52
J

J

J

J

J

I

J

Z..J

"_Z

0..

Q

00_

r-_ t_j

>O_Z

"_0.

m,"

0

i,

CO

L_J

O_

LL

t_

,.-I

L_I

Z

Z

>.,

m,,."

I

m-"

I'--

G3

53
SCALE

HORIZONTAL STABILIZER

TO MATCH QSE

CM_sTAB

SCALE

W ING

TO MATCH QSE

SCALE

FORWARD FUSELAGE

TO MATCH QSE

CM_

FIGURE 19

ITERATIVE PROCEDURE FOR MATCHING QSE STABILITY DERIVATIVES

54
MACH

NO. P'FB nWG nHT

0.35

0.40

0.60

0.70

O. 80

O. 86

0.91

0.855

O. 868

O. 880

1.035

O. 785

1.021

i. 000

1,047

1.020

1.064

0.933

1.150

1.040

1.000

0.937

O.938

0.969

O.945

0.971

0.954

O.950

FIGURE 20

AERODYNAMIC PANEL EFFECTIVENESS SCALERS

WING CONTROL

MACH INBOARD OUTBOARD

NO. AILERON AILERON

0,35

0.40

0.60

O. 70

O. 80

O. 86

0.852

0.882

0.745

0.931

0.571

0.518

0.643

0.667

0.645

0.800

0.500

0.514

SURFACE

FIGURE 21

EFFECTIVENESS SCALERS

EQUATIONS OF MOTION

FOR LOAD ANALYSIS

55
WINGCONTROL

MACH

NO.

O.40

0.60

0.70

0.80

0.86

INBOARD

AILERON

0.886

0.813

0.803

0.720

0.552

ii

OUTBOARD

AILERON

0.691

0.726

0.697

0.655

O.554

FIGURE 22

SURFACE EFFECTIVENESS SCALERS

ANALYSIS EQUATIONS OF MOTION

FOR FLUTTER

56
WING

SHEAR

(LBS)

1400 l w i I 70

• MANEUVER DESIGN CONDITION

• MACH: 0.35

• ALTITUDE: SEA LEVEL

• GROSS WEIGHT: 2500 LBS.

• C.G.: 25% MAC

1200 _ ' I 60

I000 50

800 _ 40

_ \ _V Z

600 30

"x x 1o.3 -/'_',

400 l '_ 20

_,/---My x 10-2 '_

-

'_ 10

200 _ _ _ .-.- "_,

0 0

0 20 40 60 80 I00

WING BUTTOCK LINE (INCHES)

WING

MOMENT

(IN-LBS)

FIGURE 23

ARW-2 STEADY STATE (lg) WING LOADS AT THE MANEUVER DESIGN CONDITION

57
1400

1200

i000

800

WING

SHEAR

(LBS)

600

400

200

0

\

\

-My x 10 -2 --"_/_\_ \ --Vz

%

%

Mx x 10-3.

\

\

\ \ •

_, '\

\

%

J "% \

%

%

\

%

%

• G ST DESIGN CONDITION

• MACH: 0.60

• ALTITUDE: 7000 FEET -

• GROSS WEIGHT: 2500 LBS.

• C.G.: 20% MAC

b ,_ ,m,, I

0 20 40 60 80

WING BUTTOCK LINE (INCHES)

140

120

100

8O

WING

MOMENT

(IN-LBS)

6O

4O

2O

FIGURE 24

ARW-2 STEADY STATE (lg) WING LOADS AT THE GUST LOAD DESIGN CONDITION

58
WING

SHEAR

(LBS)

2500

2000

1500

i000

50O

\

\

_ -3_- VZ MX x 10

\

\ "

, ,,.

%

_ -

10-2 . " -My x -/ _ ',,

1 I I I

• MANEUVER DESIGN CONDITION

• MACH: 0.35

• ALTITUDE: sEA LEVEL

• GROSS WEIGHT: 2500 LBS.

• C.G.: 25% MAC

"'-'--- -__-_o

200

160

120

8O

4O

WING

MOMENT

(IN-LBS)

-500

0 2O 40 60

WING BUTTOCK LINE (INCHES)

80 i00

FIGURE 25

ARW-2 WING MANEUVER LOADS - WITHOUT ACS

59
WING

SHEAR

(LBS)

3OOO

2500

2000

1500

i000

500

0

\

\

\

\

\

\

\

\

\

\

\

20

I f I I

• GUST DESIGN CONDITION

• MACH: 0.60

• ALTITUDE: 7000 FEET

• GROSS WEIGHT: 2500 LBS.

• C.G.: 20% MAC

I

\

40 60

WING BUTTOCK LINE (INCHES)

8O

240

200

160

120

8O

4O

0

100

WING

MOMENT

(IN-LBS)

FIGURE 26

ARW-2 WING LOADS DUE TO GUST - WITHOUT ACS

6O
5. STABILITYANDCONTROLANALYSIS

An objective of this final design cycle iteration, prior to hardware

implementation of the control systems, was to verify performance with

airplane mathematical models containing elastically corrected wind

tunnel test data for the 0.237 scale model. The only control systems

evaluated with the latest aerodynamicdata in the previous contract

study were the lateral-directional systems discussed in Section 6.6.2.

The data used in the previous contract study are documentedin Reference

1.

The elastically corrected, 0.237 scale model tunnel data documentedin

Reference 4 was significantly diff@rent than the data available in the

previous analysis. C.G. neutral point_ at the higher dynamic pressure

conditions were significantly forward of the previous predictions. The

neutral point varies over the flight range from 38._ percent MACat Ig

cruise to 14.5 percent MACat the maximumdynamic pressure condition.

Wing control surface tunnel data was corrected for revised size as well

as elastic corrections.

Speedbrake requirements for steady state flight with minimumengine RPM

constraints were established. Speedbrakes are required at launch and

the maneuverload test condition.

Cruise trim drag was estimated for the range of C.G. positions to be

contro|led by the RSSsystem. The method of applying, aerodynamic

_ffectiveness scalers in the elastic structure EOMto match QSE

aerodynamic coefficients as discussed in Section 4 was validated by the

close agreementof QSEand elastic structure EOMdynamics.

5.1 QSEStability Derivative Estimation. Quasi-static elastic

aerodynamic data for the 0.237 scale model were generated by FLEXSTAB,a

system of digital computer programs for predicting flexibility effects

on airplane stability and control characteristics. QSEdata for the

DASTARW-2are contained in Boeing Document,D3-I1542-I, DASTARW-2

Piloted Simulator Aerodynamic Data based on 0.237 Scale Model Wind

Tunnel Test Results, 7 March 1979.

Wind tunnel data for the wing control surfaces were not available at the

time the QSEdata were generated. After QSEdata had been generated,

wind tunnel tests were conducted with wing control surfaces. These test

surfaces were smaller than the presently sized surfaces and elastic

corrections for wing control surfaces were obtained from the digital

programs that generated the elastic structure EOM. Twocoefficients

for each stability derivative were used to arrive at the QSEwing

surface data. First, rigid control derivatives were estimated

analytically for bothwind tunnel test size surfaces and actual surface

sizes. Ratios of actual to tunnel data were established from the

estimate. These ratios are shownon Figure 27.

Rigid wing control surface derivatives with the scale factor of Figure

27 applied are shown on Figures 28 and 29.

61
1.8 CL_IA

l,i

mY

r'_

..J

L_

0

..J

L_

Z

Fr-_

Z

F=

mY

L_

C_

C_

I

F--

CZ_

r-_

0

F--

r_"

v

1.6

1.4

1.2

1.0

0.8

0.6

0.4

0.2

CL_oA

CMaoA

0 •_1 I I I ! I

v

0 0.4 0.5 0.6 0.7 0.8

MACH NO.

CM_IA

I

0.9

FIGURE 27

DAST ARW-2 WING CONTROL SURFACE SIZE CORRECTION RATIOS

62
• TOTAL 2 SIDES

m WIND TUNNEL SIZE CORRECTED

R

CMalA

0.004

0.003

0.002

0.001

._ I I I f

0.6 0.7 0.8 0.9

MACH NO.

CR

L_IA

0"017 I

0.016

0.015-

0.014-

0.013-

0.012-

j I t

0.7 0.8 0.9

MACH NO.

FIGURE 28

DAST ARW-2 INBOARD AILERON RIGID STABILITY DERIVATIVES

63
• TOTAL2 SIDES

• WINDTUNNELSIZE CORRECTED

MACHNO.

0.7 0.8

I I

0.9

I

R

CM_OA

-0.001

-0.002

-0.003

-0.004

-0.005 -

R

CL6oA

O. 004 - ,

0.003 -

O. 002

0.001 -

I

OoL_o. 6

I I I

0.7 0.8 0.9

MACH NO.

FIGURE 29

DAST ARW-2 OUTBOARD AILERON RIGID STABILITY DERIVATIVES

64
Elastic correction factors for the actual surface sizes were computed as

functions of dynamic pressure and Mach numbers. The elastic correction

factors for the inboard and outboard lift and pitching moment

coefficients are shown on Figures 30 and 31. Both correction factors

above were applied to the rigid test derivatives to obtain the QSE

surface data shown on Figures 32 through 35.

5.2 Speed Brake Requirements. Minimum thrust levels of the DAST

vehicle engine requires that speed brakes be utilized at some test

conditions. Aerodynamic drag data with speed brakes obtained from

NASA Data Sheets for Wind Tunnel Test 833, June 1980 and thrust data

generated by NASA DFRC are shown in Appendix A. Percent engine RPM

to maintain steady-state flight test conditions as functions of speed

brake setting is shown on Figure 36. Steady-state flight can be

achieved at all design test conditions without speed brakes except

MLA and launch. Fifty degrees speed brake is required at the MLA

test condition and 55 degrees speed brake is required in the event

a steady-state launch condition is to be maintained.

5.3 Neutral Point. Neutral points for the DAST ARW-2 design and test

conditions, shown on Figure 37, were computed from the QSE aerodynamic

data included in Appendix A and obtained from Boeing Document, D3-11542-1.

Figure 37 shows the neutral points for lg trimmed flight at each condition

and also the neutral point for 1.2g flight at the cruise condition.

Stability analysis was conducted at the 1.2g cruise condition due to the

unstable .characteristics of the pitching moment versus angle of attack

at this point _s shown on Figure 38.

The neutral points at some flight conditions are significantly forward

of those obtained with the previous aerodynamic data estimates and the

flight C.G. range was moved forward accordingly. The neutral point at

launch is 29.7 percent MAC. A launch C.G. of 20 percent is selected to

provide a stable air launch condition.

The C.G. range with fuel burn is from 20 to 32.8 percent MAC as shown on

Figure 39. Neutral points at the maximum dynamic pressure condition and

the 1.2g cruise condition of 14.5 and 16.7, respectively, and neutral

points within the C.G. range at other flight conditions provide a broad

range of stability characteristics both stable and unstable, which

should be beneficial in flight test verification of the RSS system.

5.4 Minimum Trim Drag. DAST ARW-2 cruise trim drag, computed from the

QSE data of Boeing Document D3-11542-1, is shown as a function of C.G.

position on Figure 40. Minimum drag is well back of the selected flight

C.G. range and is not attainable due to the highly unstable pitch

characteristics at some flight conditions and available stabilizer

authority. However, significant reduction in drag is obtained by

operating at aft, unstable C.G. positions compared to operating at the

most critical neutrally stable C.G. The RSS system, used to provide

65
1.5 -

/

MACH

NO°

0.35 &

O. 40

0.60

O. 70

0,80

0.86

l

i00

I I I I

200 300 400 500

DYNAMIC PRESSURE (PSF)

FIGURE 30

INBOARD AILERON ELASTIC CORRECTION FACTORS

l

600

I

700

66
1.0

%

I I

100 200

I I I I

300 400 500

DYNAMIC PRESSURE (PSF)

600

0.35

AND

0.40

0.60

O. 70

0.80

0.86

I

7OO

FIGURE 31

OUTBOARDAILERON ELASTIC CORRECTION FACTORS

67
E

CL_IA

TOTAL

TWO

SIDES

0.014

0.013

0.012

]

0.6

MACH NO.

FIGURE 32

INBOARDAILERON QSE LIFT COEFFICIENT

68
E

CM_IA

TOTAL

IWO

SIDES

FIGURE 33

INBOARD AILERON QSE PITCHING MOMENT COEFFICIENT

69
o.oo4....................................................................... --..........................

r

0.003

0.002 ' ; ": ...... :................

E

CLCoA 0.001

TOTAL

TWO

SIDES

0

-0.001

-0.002

...... ,;........ f ....... _"': ....-_'-:" _";-" --- : _-_ ....:_-"- " ,t" : 650 PSF

I".......................................

FIGURE 34

OUTBOARDAILERON QSE LIFT COEFFICIENT

70
-0.001

E

CMaoA

TOTAL

TWO

SIDES

-0.002

-0.003

-0.004

-0.005

FIGURE 35

OUTBOARD AILERON QSE PITCHING MOMENT COEFFICIENT

71
100

HIGH

ALTITUDE

CRUISE

C_-

c_

I._

Z

Z

ILl

I--

Z

C_

Q,.

9O

GLA

MINIMUM ENGINE SPEED

8O

LAUNCH

O0 I0_ 20i 30 40 50i 60l

SPEED BRAKE POSITION (DEGREE)

PIGURE 36

ENGINE RPM REQUIRED FOR STEADY STATE FLIGHT WITH SPEED BRAKE VARIATIONS

72
FLIGHT CONDITION

CONDITION

MANEUVER DESIGN

MANEUVER TEST

LAUNCH

GUST DESIGN

GUST TEST

HIGH ALTITUDE

lg CRUISE

1.2g CRUISE

MAXIMUM q

MACH NO.

0.35

0.42

0.40

0.60

0.70

0.70

ALTITUDE

(FEET)

SEA LEVEL

10,000

15,000

7,000

15,000

50,000

NEUTRAL

POINT

(PERCENT

MAc)

30.0

28.9

29.7

27.2

28.3

21.1

O.80

0.80

0.86

46,800

46,800

15,000

38.6

16.7

14.5

FIGURE 37

DAST ARW-2 NEUTRAL POINTS

73
• MACH: 0.80

• ALTITUDE: 46,800 FEET

• GROSS WEIGHT: 2350 LBS.

0.14-

0.10

CM_ MAC

1.2g FLIGHT

06 .

6

0.02__ . "

I I I I I I I I I

-2 0 2 4 6

(DEG)

FIGURE 38

CRUISE CONDITION PITCH COEFFICIENT

74
GROSS

WEIGHT

(LBS)

2600

2400

2200

2000

]800

MAX. GROSS

"-W'EI_T'_--

_2500 LBS. _

DESIGN CL,,_GROS_S_S --.--

""WE'I"G'HT -_--2350-'LBS"_-. --

HINIMUM GROSS__W.WEIGHT= 2200 LB.B..S.

LAUNCH CRUISE RSS

20% MAC 27.4% MAC 32.8% MAC

I r

5 20 25 30

CENTER OF GRAVITY (PERCENT MAC)

I

35

i

40

FIGURE 39

DAST ARW-2 CENTER OF G_VITY VARIATION WITH FUEL USAGE

75
0

CO

O

°.

'-m

l, I .._I

i.i

I.I..O

Ooe)

0.04

CO

°°

"_" I--"

..i.--.4

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76
feedback controlled stability at the aft, unstable C.G. positions is

discussed in Section 6.5.

5.5 Comparisonof QSEand Flexible Equations of Motion. Both QSEand

elastic EOMwere used in the analysis and design of the ACS. As

discussed in Section 4, the method of aerodynamically scaling the

elastic EOMwas modified to provide close agreementof the elastic EOM

to QSEEOM. A comparison of QSEand elastic EOMcharacteristic roots at

the design flight conditions is shownon Figure 41. At mostconditions,

a very close agreementexists. At the conditions where the least

agreementexists, closing the system feedback loops reduces the

difference in placement of the roots as shownon Figures 42 and 43.

77
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78
e MACH: 0.60

• ALTITUDE: 7000 FEET

• C.G.: 27.4% MAC

QSE

_- ELASTIC

POINT "1" - OPERATING GAIN

(PREVIOUS RSS SYSTEM)

I I I I 35.0

30.0

25.0

SHORTPERIOD

_ 11 .... + 0

20.0

jm, RAD/SEC

15.0

10.0

(_, RAD/SEC

FIGURE 42

COMPARISON OF QSE AND FLEXIBLE MODEL RSS SYSTEM ROOT LOCUS AT

THE GLA DESIGN CONDITION

79
• MACH: 0.86

• ALTITUDE: 15,000 FEET

• C.G.: 27.4% MAC

.... QSE

_ELASTIC

POINT "I" - OPERATING GAIN

(PREVIOUS RSS SYSTEM)

I I I I

SHORT PERIOD /

/

!

/

l

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/

35.0

30.0

25.0

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jm, RAD/SEC

15.0

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c_, RAD/SEC

FIGURE 43

COMPARISON OF QSE AND FLEXIBLE MODEL RSS SYSTEM ROOT LOCUS AT

MAXIMUM DYNAMIC PRESSURE CONDITION

8O
6. ACTIVE/AUTOMATICFLIGHTCONTROLSYSTEMSSYNTHESIS

Active Control Systems (ACS)were synthesized as part of the integrated

design of the DASTARW-2wing using conventional analytical techniques.

This synthesis is described in detail in Reference 1. The ACSincludes

Gust Load Alleviation (GLA) and ManeuverLoad Alleviation (MLA) to

minimize wing structural strength, flutter suppression to suppress

flutter up to 20 percent above design dive speed, and Relaxed Static

Stability (RSS)to minimize trim drag.

In addition to the ACS, primary and backup Automatic Flight Control

Systems (AFCS)were synthesized to provide longitudinal and lateraldirectional dynamic stability. The primary AFCS,which will be

implementedon the ground based computer, provides the intent of

MIL-F-8785BLevel 1 flying qualities. The backup AFCSprovides

qualities necessary to return the drone to a safe recovery area in

case of failure in the primary AFCSground computer loop. The AFCS

for the lateral-directional axis provides Dutch roll damping, turn

coordination, roll attitude commandand conventional maneuver

responses. The primary pitch AFCSand the RSSsystem provide

angle of attack and vertical acceleration limiting through acceleration

responses to operator pitch commandsand also computesrequired MLA

inboard aileron commands. The backup pitch AFCSlimits angle of attack

and produces a glide angle commandsystem.

The synthesis conducted during the third design cycle iteration

completed in this study and documentedherein verified the structural

and control system design with wind tunnel test data from the 0.237

scale ARW-2model.

Brief summariesof synthesis and performance criteria used to design the

ACSand AFCSand evaluate performance are presented in Paragraph 6.1

with more detailed presentation in Paragraph 3.2.

Servoactuator models were developed and synthesized with a goal to

provide an analytical model that would more closely predict hardware

test results. Becauseof someunpredictable nonlinear characteristics

such as friction, changes in bulk modulus and minimumfeedback

relationships, the bandwidth predicted by analysis is difficult to

realize with hardware; however a model was developed and finalized

during testing. The model definitions, analysis and test results are

presented in Paragraph 6.2.

Paragraph 6.3 discusses the Flutter Supression System (FSS) and its

improvement in flutter speed. The FSS provides structural damping that

meets FAR Part 25 to 1.2 Vn up to 15,000 feet altitude and 20 percent

margin above the placard t_st speed.

81
Synthesis and performance evaluation of the gust and maneuverload

alleviation systems are presented in Paragraph 6.4. Results indicate

that the MLAand GLAsystems reduce maximumwing design loads by

approximately 18 percent.

The use of an RSSsystem to provide stability and Flying qualities at

reduced trim drag is presented in Paragraph 6.5. Although minimumtrim

drag cannot be achieved, the RSSsystem permits flight at aft unstable

C.G. positions to take advantage of reduced trim drag.

The primary and backup AFCSwere reevaluated using revised aerodynamics

and physical properties and the results are discussed in Paragraph 6.6.

The longitudinal primary AFCSwas revised to be morecompatible with

larger variations in aerodynamiccharacteristics at different flight

conditions.

Paragraph 6.7 discusses the evaluation of control authority of the DAST

stabilizer configuration. Analysis indicates more positive stabilizer

authority is desirable.

Control system compatibility is discussed in Paragraph 6.8 and analysis

indicates that system capability is satisfactory and no airplane mode

destabilization exceeds requirements.

Computerimplementation of the ground based control systems is discussed

in Paragraph 6.9. Flow diagrams and difference equations are presented

in this discussion.

6.1 Performance and Synthesis Criteria. The GLAand MLAsystems were

to be critical systemsminimizing structural strength and stiffness by

reducing wing root bending momentby approximately 20 percent.

The ARW-2wing without FSSwas to exhibit flutter within the flight

envelope. With FSSthe flutter boundary was to be extended to 1.2 VD.

The goal of the RSSsystem was to provide stability and flying qualities

as required to operate the drone with center of gravity at the minimum

trim drag condition. A stable launch condition was required with an

inoperative RSSsystem.

Flying quality criteria were per MIL-F-8785Bfor a Class IV, Category B,

aircraft. The ACSand AFCSwere to provide the intent of Level 1 flying

qualities in the primary modeand Level 3 flying qualities in the backup

mode(primary modeinoperative).

The wing control surfaces were to be typical of a transport design,

trailing edge surfaces with similar effectivenesses and locations. All

ACSsystems were to be compatible and capable of operating

simultaneously as one system.

82
Minimumstability margins taken from MIL-F-9490Dare given on Figure 14.

The design gust velocity envelope and ramp type maneuverused to design

the GLAand MLAsystems are presented on Figures 12 and 13. The design

peak gust level was 62 ft/sec except the design gust level for the FSS

synthesis was 12 ft/sec peak (6 ft/sec RMSat 2_level). Design limit

wing maneuverloads were determined for maneuvering load factors of +2.5

and -1.0 per FARPart 25.

ElectriCal power available for the ACSwing control surface

servoactuators was 43 amperesat 28 VDC.

Control surface maximumdisplacement and rate requirements are shownon

Figure 44. Analysis was conducted to determine maximumcontrol activity

in randomturbulence to determine these displacement and rate

requirements. AVon Karmanspectrum with a characteristic gust scale

length of 2500 feet was used to represent the randomatmospheric

turbulence characteristics. Aileron RMSdisplacement and rate per unit

gust at the design and test conditions were determined. The results of

the PowerSpectral Density (PSD)symmetric analysis for the outboard

ailerons at the FSStest condition are shownon Figure 45. The

contributions from antisymmetric are negligible and are not shown.

The servoactuator requirements are summarizedand presented on Figure

46. These requirements are based on estimated hinge momentsand PSD

analysis. Also shownon this figure are NASAactuator test results for

comparison to the requirements.

6.2 Servoactuator Models. Math models were developed and used during

previous NASAcontracts. Test results did not comparefavorably with

results obtained from analysis. Theseanalyses results are documented

in References 2 and 3 and the test results are documentedin Appendix

B. A model is required that can predict the implemented servoactuator

performance, however, becauseof someunpredictable nonlinear

characteristics such as friction, changes in bulk modulus and pressure

feedback relationships, the bandwidth predicted by analysis is difficult

to realize with hardware. The effort to derive a workable model is

described by the following paragraphs. The final result eliminates

pressure feedback and limits the servoactuators to a 50 Hz bandwidth

which provides reasonably predictable test results.

6.2.1 Initial Servoactuator Model. The closed loop servoactuator

dominant frequency modepredicted by analysis was approximately 110 Hz,

however, the dominant modeobtained during hardware testing was

approximately 70 Hz. A comparison of analysis and test results using

the DASTARW-1parameters is shownon Figure 47. Further testing

_ndicated that hydraulic line lengths had little effect upon the closed

loop bandwidth as shownby the frequency responses of Figure 48 although

hydraulic frequencies and loop stability are affected by line lengths.

83
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84
PSD

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o SYMMETRIC

• MACH: 0.86

• ALTITUDE: i5.,000 FEET

• FSS, GLA AND RSS ON

• 12 FT/SEC PEAK VERTICAL GUST

2_

OUTBOARD AILERO/N/_wX_ - PEAK

//_' DISPLACEMENT

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FREQUENCY (HZ)

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DISPLACEMENT

(DEG)

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(DEG/SEC)

FIGURE 45

SYRHETRIC OUTBOARD AILERON DISPLACEMENT AND RATE REQUIREMENTS

85
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88
A low frequency dominant mode was always evident during servoactuator

testing but not identified in analysis. A study was made in an attempt

to determine this anomaly. A servo valve represented by the transfer

function as developed in Reference 4 is shown on Figure 49 and includes

a first order lag associated with the servovalve closed loop.

Servovalve parameters were selected and used in this transfer function

to provide a good curve-fit to servoactuator test data and the results

are shown on Figure 50.

A baseline actuator was synthesized from ARW-1 data and these techniques

were used to develop the ARW-2 servoactuator math model as shown by the

block diagram on Figure 51 and the parameters listed on Figure 52.

The parameters of a Moog Series 30 servovalve were used to synthesize

the servoactuator as shown by the root locus on Figure 53 and the

resulting frequency response is shown on Figure 54.

A wider bandwidth servovalve selected for analysis in an effort to

improve servoactuator bandwidth did not increase the actuator bandwidth

appreciably but did improve performance because the first order lag

associated with the servovalve closed loop crossover frequency was moved

from 640 to 1250 radians. A Hydraulic Research Model AR-25 servo valve

was selected for hardware implementation because it is a direct

replacement for the Moog Series 30 and its crossover frequency is 1000

radians. The AR-25 servovalve frequency response and transfer function

are shown on Figure 55. The resulting servoactuator frequency response

is shown on Figure 56 and indicates that a nearly flat response to 80 Hz

can be realized.

6.2.2 Servoactuator Bench Tests. Testing was accomplished on a

breadboard of the DAST ARW-2 outboard and inboard aileron servoactuator.

The breadboard included hydraulic line lengths, load inertias and

electronics equivalent to the ARW-2 design.

The objectives of the bench testing were to verify the synthesized

servoactuator model and select the final gains required for good closed

loop stability. A favorable test and analysis comparison of the

servoactuator dominant modes would ensure confidence that the analysis

model could accurately predict the hardware response. The surface

actuator mode isdetermined during test by closing only the position

loop and increasing the loop gain until the system is just unstable.

The servo valve mode is then determined by closing the pressure loop

with position loop closed at nominal gain and increasing pressure

feedback gain until system again goes unstable. A summary of these

results is shown on Figure 57 which compares analysis and test results

using both the Series 30 and Model AR-25 servovalves.

Extensive closed loop testing was accomplished, however desired

performance could not be realized with the pressure feedback loop

89
Qv

Vc

Where:

KvKA

((S/mc) +I)(($2/_02) +(260/m O) +I)

QV = Servovalve no-load flow, inch3/sec

Vc = D-C valve flow command, volts

KV = ServRvalve no-load flow gain constant,

inchO/sec/ma

KA = Current drive gain constant, ma/volt

mO = Naturalfrequency of spring-mass resonance of

torque motor, rad/sec

mC = Crossover frequency of spool position closed

loop, rad/sec.

FIGURE 49

SERVOVALVE TRANSFER FUNCTION

9O
PHASE (DEG)

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_ PRESSURE LOOP

KAKvKP = 0.0430 IN3/SEC/PSI

\

\

-2500 -1500 -1000

_, RAD/SEC

3500

3000

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\ _ FLUID MODE

\\ SERVOVALVE J

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FIGURE 53

DAST ARW-2 OUTBOARD SERVOACTUATOR WITH SERIES 30 SERVOVALVE ROOT LOCUS

94
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97
MODE

SURFACE-ACTUATOR

SERVOVALVE

Ai,IALYSISFREQUENCY

SERIES 30

76-82 HZ

120-160 HZ

AR-25

95-100 HZ

350-400 HZ

TEST FREQUENCY

SERIES 30

72-76 HZ

115-120 HZ

AR-25

80-92 HZ

350-355 HZ

FIGURE 57

SERVOACTUATOR DOMINANT MODES COMPARISON WITH SERIES 30 AND AR-25 SERVOVALVES

98
closed. It was suspected that the pressure loop was influenced by

actuator friction which caused a decrease in amplitude with an increase

in frequency until at some frequency the effect of friction became less

evident resulting in an increase in gain which caused stability problems

at the actuator dominant frequency mode. The effect of friction can be

seen from the responses shown on Figure 58. These results show that by

increasing the break frequency of the washout in the pressure feedback

loop, friction is less effective in reducing output amplitude at lower

frequencies, however, the peak responses at the actuator dominant mode

are very sensitive to input command amplitude indicating that other

non-linearities in the servoactuator may be affected.

Pressure feedback reduces bandwidth and depends upon a lightly damped

position loop to extend the bandwidth. Pressure feedback also causes

the actuator performance to be more sensitive to changes within the

servo loop. For example, friction, hydraulic bulk modulus, supply and

load pressures can destabilize servovalve performance as described in

Reference 4.

6.2.3 Final Servoactuator Model. Pressure feedback was eliminated and

notch filters were added to limit the actuator bandwidth. Position loop

gain was adjusted to give optimum performance with phase and gain

margins greater than +_45 degrees and +6 dB, respectively. The curve

labeled "Test" shown on Figure 59 is the resulting frequency response

for the outboard servoactuator. Similar results were obtained for the

inboard servoactuators'.

The solid curve on Figure 59 represents the result of a curve-fit to the

test data. During the flight worthiness testing, frequency responses

were again obtained and the outboard actuator response is shown by the

curve identified as "Final Implementation" which compares favorably with

the other responses shown on this figure. The transfer function for

this curve-fit shown on this figure was used as the actuator model to

update the flutter analysis described in Section 6.3.

The frequency responses on Figure 60 show the effect of including the

band limiting notch filters and only a 17-degree increase in phase lag

at 20 Hz results. This additional phase shift does not reduce the

flutter suppression stability margins below the specified requirements.

Figures 61 and 62 show that the outboard servoactuators are no-load

velocity limited above 10 Hz and the inboard servoactuators are no-load

velocity limited above 2 Hz.

Maximum surface displacements and rates are shown on Figure 63. The

static and maximum hinge moments for the outboard ailerons at the FSS

condition are 225 in-lb and 280 in-lb respectively at zero degrees and

15 degrees deflection.

99
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The loaded rate capability of the outboard ailerons is shown on Figure

64 to be 615 degrees per second at zero degrees deflection and 370

degrees per second at maximum deflection of 15 degrees. A peak

displacement of approximately 4.25 degrees and a peak rate of

approximately 200 degrees per second are required at 18 Hz can be

determined from the results shown on Figure 45. From this data the

maximum no-load rate requirement of 955 degrees per second was selected

for testing; however, a no-load rate in excess of 1000 degrees per

second was obtained. Similar rationale was used for all actuators which

verify that the servoactuators meet rate and stability requirements as

implemented for all flight conditions. These rates and displacements

should however be verified during flight testing.

6.3 Flutter Suppression System Synthesis. Analysis was conducted in

this study to finalize the synthesis of a FSS for the DAST ARW-2 drone

configuration. Analysis was conducted previously in iterations one and

two of the integrated design study (Reference 1), which identified a

wing ballast configuration that produced a flutter boundary 9 or 10

percent below design dive speed VD. Preliminary symmetric and

antisymmetric flutter suppression systems were synthesized during the

integrated design cycle_ for this configuration. A summary of the

integrated design study results is presented in Paragraph 6.3.1. The

conditions analyzed during this synthesis are shown on Figure 65.

Final synthesis was conducted using the mathematical models described in

Paragraph 6.3.3. EOM were updated to reflect the 0.237 scale ARW-2

model wind tunnel test results and contain unscaled fuselage and wing

structural elastic modes in addition to the scaled rigid body modes.

The update of these EOM and finalization of the FSS were accomplished in

the third iteration of the integrated design as documented herein. A

performance evaluation was conducted on the final FSS and is described

in Paragraph 6.3.4.

6.3.1 Integration Design Study. This paragraph summarizes results of

iteration one and two of the integrated design study documented in

Reference 1. In this study preliminary design of a flutter suppression

system for the DAST ARW-2 drone was accomplished.

6.3.1.1 Configuration. During the course of the integrated design

study, several configurations were modeled and analyzed. The final

configuration included a 2 pound mass added to the leading edge of the

rib at Wing Buttock Line (WBL) 98.9. This configuration exhibited

symmetric and antisymmetric flutter modes that were similar in nature

and produced an open loop flutter boundary approximately 10 percent

below design dive speed (Vn) , as shown on Figure 66. The outboard

aileron selected for flutteF suppression was sized as described in

Reference 1 to have an 18 inch span.

I06
1200 -

AILERON

RATE

(DEG/SEC)

I000 -

800 -

600 -

400-

m

200-

0

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NOTE: THIS CURVE DOES NOT \

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BY 20°/SEC AT 4,5 DEG.

AMPLITUDE AT 18 HZ.

i I

100 200

HINGE MOMENT (FT-LBS)

300

FIGURE 64

OUTBOARD AILERON RATE VERSUS HINGE MOMENT
MACH

0.60

O.70

O.80

ALTITUDE

(FEET)

7,000

12,000

15,000

18,000

25,000

12,000

15,000

18,000

25,000

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18,000

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(FEET)

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12,000

15,000

16,000

17,000

18,000

20,000

25,000

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8,000

12,000

15,000

18,000

25,000

6,000

8,000

10,000

12,000

15,000

18,000

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FIGURE 65

DAST ARW-2 FLIGHT CONDITIONS ANALYZED

I08

I
ALTITUDE

(FEET)

SYMMETRIC UNSCALED

ANTISYMMETRIC UNSCALED

FLIGHT M D

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MACH NO.

FIGURE 66

FLUTTER BOUNDARY FOR THE FINAL (ITERATION 3) FSS CONFIGURATION

WITHOUT ANTISYMMETRIC GAIN SCHEDULING

109
6.3.1.2 FSS Definition. The FSS system synthesized as described in

Reference 1 to suppress the symmetric and antisymmetric flutter modes is

shown in the block diagram on Figure 67. The system uses the difference

of vertical accelerations at WBL 84 on the rear spar and WBL 82 on the

front spar on the left and right wing panels summed to drive 18-inch

span outboard ailerons through appropriate shaping filters. Both

shaping filters had variable gains and filter first order break

frequencies to adapt the system to the flight conditions at which the

system must operate.

6.3.1.3 Actuator Dynamics.

analysis was:

The servoactuator model used through the

6.583 x104 (DEG/DEG)

OCMD (S +401)(S 2 +1681.8S +1.028 x106)(S 2 +654.7S +1.597 x106)

The actuator dynamic representation was to be updated "during design

cycle iteration three in an effort to more closely match the dynamics

observed during hardware tests.

6.3.1.4 Results. The results of the integrated design study (Reference

i) showed that the FSS system actively suppressed the flutter mode which

was composed primarily of first wing vertical bending and first torsion

elastic modes. The third iteration included updated EOM based on the

0.237 scale ARW-2 model wind tunnel test results and an updated actuator

model that is documented herein.

6.3.2 System Criteria. Criteria used during the FSS synthesis guided

the form of the final system. The criteria included constraints on the

system, such as type of sensors, size of surfaces, complexity of shaping

filters and implementation and synthesis criteria which set performance

goals such as stability margins and model damping.

6.3.2.1 Constraints. Constraint criteria ultimately affect performance

of a system but do not specify any particular system performance. All

constraints imposed were either directly or indirectly attributable to

the integrated design study results. Constraints that were a direct

result of the integrated design study were intended to minimize

duplication of analysis already performed or those considered adequate

and frozen by interation two such as surface size, while the indirect

constraints were to improve the system performance.

The control surfaces were selected in iteration two (Reference 1) to

have a span of 18 inches with approximately 23 percent chord. With

conventional airplane design, the control surfaces are usually a

constraint, however, using integrated design methodology, the control

surfaces selected for FSS were adequate to suppress flutter throughout

the flight envelope.

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Other constraints were considered during the finalization of the FSS

system_ Actuator compensation to reduce phase shifts induced when the

pole-zero cancellation is imperfect was minimized and used only after

sensitivity analysis verified that an anticipated pole-zero mismatch had

minimal effect on FSS performance. Sensitivity analysis is described in

Paragraph 6.3.5.

The control laws should have as large as possible roll off to reduce

coupling with high frequency modes and to reduce susceptibility to high

frequency noise. This was accomplished with shaping filters in the

servoactuator drive electronics as described in Paragraph 6.2.3.

The control laws should have as low gain as possible at low frequencies

to reduce coupling with rigid body and filter modes. This was verified

by compatibility analysis as described in Paragraph 6.8.

6.3.2.2 Synthesis Criteria. Synthesis criteria specify desired

performance of the system. The synthesis criteria used in the FSS

development are discussed in the following paragraph.

The FSS should provide mode damping so that the wing is free from

divergence and flutter at all speeds up to 120 percent of the design

dive speed, Vn. These requirements, minimum flight altitude of 10,000

feet and the _est envelope are shown on Figure 68. The gain and phase

requirements are shown on Figure 69. The FSS should exhibit MIL-F-9490D

stability margins at VD as given on Figure 70. The FSS should not

degrade damping of any mode to below a damping ratio of 0.01 (except the

flutter mode) and should not significantly reduce damping of any mode

with damping ratio below 0.01. The FSS should be capable of operating

in 6 feet/second RMS (at 2_level) random turbulence with 12 feet/second

peaks.

6.3.3 Final Flutter Suppression System. The form of the integrated

design FSS was used as the starting point in the final system synthesis.

The EOM used in the analysis reflected the 0.237 scale ARW-2 wind tunnel

test results with scaled structural modes and unscaled rigid body modes.

The integrated design study results including sensor types and the

systems criteria are discussed in Reference 1 and summarized above.

6.3.3.1 Configuration. The final DAST ARW-2 wing configuration is

shown on Figure 71. The 2 pound ballast is located at the leading edge

of the rib at WBL 98.9 and the outboard 18 inch, 23 percent chord,

trailing edge control surface is located between WBL 89.7 and WBL 107.7.

The accelerometer positions are shown on Figure 6 to illustrate the

physical relationship of the various components.

The mathematical model included all the above components and fuselage

and empennage structural elastic modes. These modes account for wingbody and wing-wing coupling which might affect the flutter modes.

112
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FIGURE 69

DAST ARW-2 FLUTTER SUPPRESSION SYSTEM GAIN AND PHASE REQUIREMENTS

114
MODE

FREQUENCY _ 0.06 HZ

0.06 HZ _ FREQUENCY _ FIRST STRUCTURAL MODE

ALL STRUCTURAL MODES

GAIN

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FIGURE 70

MIL-F-9490D STABILITY MARGINS

775
INBOARD AILERON,

INBOARD SEGMENT

FOR MLA

INBOARD AILERON,

OUTBOARD SEGMENT

FOR MLA

PERCENT SURFACE CHORD

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BL

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BL

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OUTBOARD AILERON

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BL

89.7 I

BL

107.7

FIGURE 71

CONTROL SURFACE DEFINITION - FINAL DESIGN ITERATION

116
6.3.3.2 Servoactuator Model. The transfer function for the

servoactuator was defined through analysis and bench testing described

in Paragraph 6.2. The simplified transfer function of the servoactuator

for the no-load condition is:

6AIL _ 4.046 x lO14(S2 +28.6S +477.52 )

6CMD (S +I80)(S 2 +251S +3142)(S 2 +229S +477.52)(S 2 +286S +477.52 )

Servoactuator dynamics vary with hinge moment as discussed in Paragraph

6.3.5.2.4. The no-load transfer function was used during the synthesis

study. The effect of maximum resisting and maximum aiding hinge moment

variation in actuator dynamics of FSS performance was evaluated for a

purely analytical actuator model.

6.3.3.3 Sensor Location. Sensor type and control surface size and

location were fixed in the integrated design study (Reference 1) leaving

sensor location and orientation to be defined if affected by the updated

EOM. Zero locus techniques were used to select these two parameters to

satisfy two primary goals. Coupling with the flutter mode was to be

maximized while minimizing adverse coupling with other structural and

rigid body modes.

To establish sensor locat_ons, the selection process must be an integral

part of the 9verall FSS synthesis as shown on Figure 72. This method

was used in an iterative manner to establish sensor locations based upon

closed-loop system results.

During the integrated design analysis described in Reference 1, the

difference of two vertical acceierometer signals on the front and rear

spar at the WBL 82 and WBL 84, respectively, was selected to synthesize

the FSS. Initial third iteration FSS synthesis with updated EOM

utilized these locations, however, both symmetric and antisymmetric

flutter requirements could not be met and the sensor position

methodology of Figure 72 was implemented.

Effects of varying locations and spanwise orientation of the two

vertical accelerometers along the front and rear spar were evaluated

using the above describedmethodology with results shown on Figures 73

and 74. These zero root loci, subsequent closed loop flutter analysis

described in Paragraph 6.3.3.4 and sensor location sensitivity described

in Paragraph 6.3.5.1 show that vertical accelerometers located at WBL 82

on front spar and WBL 84 on rear spar offer best results for symmetric FSS

and vertical accelerometers located at WBL 92 on both front and rear

spars offer best results for antisymmetric FSS. These sensor locations

were used to finalize the FSS system. The front spar location WBL 80

was not selected, although it appears, as shown on Figure 73, to be a

better selection, because this zero location retards movement of the

flutter mode to the left more than the selected WBL 82.

117
INTEGRATION

STUDY SENSOR

POSITION

SELECTION INTEGRATION I

STUDY FSS

SYNTHESIS

SENSOR

UPDATE I

SENSOR

POSITIONS

FIGURE 72

SENSOR POSITION SELECTION METHODOLOGY

118
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WBL 86

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jm (RAD/SEC)

130

X FLUTTER

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MODE

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6.3.3.4 Filter Definition. The FSSfilter configuration was finalized

using accelerometer locations described in Paragraph 6.3.3.3 and an

actuator transfer function which was curve-fit to the laboratory bench

test results described in Paragraph 6.2.2.

The flutter suppression filters were synthesized using root locus

techniques in an iterative method shown on Figure 75. The synthesis

began with definition of filters at the flutter suppression design

condition, Mach 0.86 at 15,000 feet. The filters were then evaluated at

other conditions and adjustments made as necessary to meet the

requirements.

The region in the flight envelope in which the FSS is required to

stabilize the flutter modes is shown on Figure 68 by the shaded area.

The flight envelope was selected as shown on this figure by the crosshatched area. The placard set at 0.86 Mach during the integrated design

study could not be increased to 0.91Mach because of thrust limitations.

The final filters developed in the third design iteration shown on

Figure 76 utilize washout filters of S/(S + 2) to eliminate low

frequency and steady-state commands to the control surfaces.

6.3.3.4.1 Synthesis. Synthesis was initiated using the form of the

iteration two filters with an analytical model of the servoactuator.

Servoactuator compensation was included during the synthesis procedure.

Sensor locations were selected using this FSS model and these sensor

locations were verified with the final filters and actuator. After

laboratory bench tests had been run, a transfer function which had been

curve-fit to the laboratory results was included in the synthesis

procedure in place of the analytical model. The inclusion of the

dynamics of the servoactuator determined by test led to the synthesis of

the final FSS filters shown on Figure 76. The actuator transfer

function is also shown on this figure and the FSS is shown as

implemented to receive sensor signals from both wings.

The frequency response of the symmetric and antisymmetric shaping

filters with and without the actuator is shown on Figure 77 and Figure

78 respectively. Root locus plots for the symmetric and antisymmetric

system at 0.86 Mach, 15,000 feet, are shown on Figures 79 and 80

respectively. Additional root locus analysis verifying FSS stability

throughout the flight envelope was conducted and has been presented at

the design reviews and critical conditions are included in Appendix C.

6.3.3.4.2 Parameter Scheduling. Scheduling of gains or time constants

was not required to meet the gain and phase margins at the design flight

condition for either symmetric or antisymmetric modes. Symmetric

scheduling is not required at the 0.91Mach condition. Gain scheduling

is required at the 0.91 Mach condition for the antisymmetric FSS. The

gain schedule is shown on Figure 81. Scheduling is not required for the

flight envelope and is not implemented for flight test.

121
GENERATE

AEROELEASTIC

E.O.M.'s AT

PARTICULAR

FLIGHT

CONDITIONS

SYNTHESIZE

SYMMETRIC AND

_NTISYMMETRIC

FSS FILTERS

CHECK

PERFORMANCE

AT FLIGHT

CONDITION

__F MEET

ICATION?

FAIL ]

YES

ADDITIONAL

FLIGHT CONDITION

.,fTO ANALYZE?

FINAL FSS

FILTER

CONFIGURATION

FIGURE 75

FLUTTE.R SUPPRESSION SYSTEM FILTER SYNTHESIS PROCEDURE

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6.3.4 Flutter Suppression System Performance. Analysis was conducted

to evaluate performance and sensitivity of the flutter suppression

system and compatibility of the system with gust load alleviation,

maneuver load alleviation, relaxed static stability and automatic flight

control systems. Performance evaluation consisted of analysis to

establish the ability of the FSS to stabilize the flutter mode and not

degrade stability of the other structural modes throughout the flutter

envelope to a damping ratio of less than 0.01, except when the

unaugmented model damping ratio is below 0.01. The flutter suppression

system definition used for the performance evaluation is presented in

Paragraph 6.3.3.4.

Analysis was also conducted to define sensitivity of the FSS to various

expected system variations and these sensitivity studies are described

in Paragraph 6.3.5.

Effects of the FSS on rigid body modes and effects of the GLA, MLA, RSS

and AFCS on the flutter modes were examined to determine if the systems

were compatible. System compatibility is discussed in Paragraph 6.8.

All systems are compatible.

6.3.4.1 Stability. Analysis was conducted to determine stability

characteristics of the DAST ARW-2 vehicle with the FSS over the entire

flutter envelope. This analysis included determination of damping and

frequency of the flutter mode and other structural elastic modes with

the FSS operating and the evaluation of stability margins of the system.

A list of all flight conditions analyzed is presented on Figure 65..

6.3.4.1.1 Damping and Frequency Evaluation. Closed loop damping and

frequency were determined for each of the rigid body and structural

elastic modes with the FSS engaged. The resulting flutter boundaries

with the systems operating are shown on Figure 66. Both systems exceed

design goals. The higher closed loop symmetric flutter boundary is due

to the higher loop gain.

Damping ratios and frequencies with the FSS on and off are presented for

the symmetric and antisymmetric systems at Mach 0.86 and 15,000 feet

altitude on Figures 82 and 83 respectively. This data shows that the

FSS does not reduce mode damping ratio to below O.01or degrade damping

of modes with damping ratios below 0.01.

Flutter mode damping ratios versus Mach number at a constant altitude of

15,000 feet is shown for the symmetric and antisymmetric systems on

Figures 84 and 85, respectively. These plots illustrate the violence of

the flutter modes and the capability of the FSS to stabilize the modes

to 1.2 Vn- Flutter mode damping ratio versus altitude at Mach 0.86 is

shown for both on Figure 86 and 87.

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6.3.4.1.2 Stability Margins. Analysis was conducted to demonstrate the

FSS meets stability margin requirements shown on Figure 70. Figures 79

and 80 show the nominal FSS has a greater than ±6.0 dB gain margins and

±45 degrees phase margins at the design condition. Nominal refers to

a system without gain or parameter scheduling. When compared to the

phase and gain requirements shown on Figure 69 and the flutter suppression envelope shown on Figure 68, Figures 88 through 92 show that the

symmetric subsystem does not require gain or time constant scheduling,

while Figure 93 shows the antisymmetric subsystem requires gain

scheduling to meet the gain and phase margin requirements at the

0.91Mach, 15,000 feet condition. Figures 93 through 97 show that the

antisymmetric system meets the phase and gain requirements shown on

Figure 69 with the implementation of the scheduler shown on Figure 81.

6.3.4.2 Control Surface Requirements. Analysis was conducted to

determine maximum control surface activity in random turbulence to

verify that the servoactuator system components selected in iteration

two design study were adequate for the updated FSS. The results of this

analysis are discussed in Section 6.1 and shown on Figures 45 and 46.

6.3.4.2.1 A Simulated Gust Excitation Generator. Outboard control

surface response to a random noise generator is included in the Drone

Active Control Electronics (DACE) unit to simulate a gust environment as

described in Paragraph 7.4.9. The generator output signal is input to

the inboard ailerons for airplane excitation. The outboard ailerons

will respond to inboard aileron displacements and rates when the FSS is

engaged and the frequency response results are shown on Figures 98, 99,

and 100 which show very little outboard aileron response to inboard

excitation below 5 Hz. Figure 101 shows very little inboard response to

the random noise generator input above 5 Hz, therefore no inputs to FSS

at the flutter frequencies are expected.

6.3.5 FSS Sensitivity. Analysis was conducted to determine sensitivity

of the FSS to variations of system parameters. The sensitivity studies

included sensor location, notch filter changes in Q and frequency,

servoactuator dynamics, hinge moment effects, parameter scheduling and

phase and gain changes.

Results of the sensitivity studies show that the FSS provides stability

in the flight envelope as implemented without scheduling and will

provide stability to 1.2 times the design dive speed with gain

scheduling of the antisymmetric filter. The flight conditions analyzed

are shown on Figure 65 and the critical phase-gain root loci are shown

on Figures 79 and 80, on Figures 88 through 97 and in Appendix C.

See Appendix D for sensitivity results.

The airplane is relatively insensitive to small changes in stability

derivatives and the ACS tends to increase frequency and damping of the

fuselage mode.

Hinge moments change gain and phase of the servoactuators and should be

further evaluated during ground and flight test.

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150
6.3.5.1 Sensor Location. The effects of varying sensor locations are

discussed in Paragraph 6.3.3.3. The results of the sensitivity study

indicates that sensor location is relatively critical and the sensors

should be installed within 0.5 inch of the analyzed locations. Figures

72 and 73 of Appendix D summarize this sensitivity study.

6.3.5.2 Servoactuator Dynamics. Several sensitivity studies were made

to determine how system stability is affected by parameter changes

within the actuator loop. These studies were conducted using analytical

models and included servoactuator loop sensitivity to changes in servo

valve models, surface-actuator mode compensation pressure feedback gain

reduction, and hinge moment.

Sensitivity studies were not made on the final actuator model derived

from testing; however, the results are valid and are presented as

insight to guide future actuator evaluations.

6.3.5.2.1 Servovalve Models. Three servovalve models were evaluated

which included a Moog Series 30, a Moog Series 31, and a Hydraulic

Research Model AR-25. A servovalve with a wide bandwidth is desirable

for improved system performance. The Series 31 was discarded because of

cost and circuit complexity. The AR-25 servovalve was selected because

it is a direct replacement for the Moog Series 30 but has a wider

bandwidth as described in Paragraph 6.2. FSS performance improved

resulting from the increased servovalve bandwidth even though the closed

loop actuator bandwidth did not increase appreciably. Figures 102 and

103 compare servovalve results.

6.3.5.2.2 Surface-Actuator Mode Compensation Filter. During synthesis,

a filter was added to the symmetric FSS to cancel the surface-actuator

mode (_n) and improve FSS performance. The form of this filter is:

S2 + 26(an S +_n

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(S + _n )2

System sensitivity to removing and varying filter parameters was

evaluated. The effects of removing the actuator compensation

destabilized airplane modes and these modes were difficult to stabilize.

Figures 104 and I05 show these effects on symmetric FSS. Removing the

antisymmetric compensating filter decreased the lead phase margin to 40

degrees as shown by the result presented on Figure 106 as compared to

Figure 107.

Since removing the compensating filter caused adverse effects, implying

that the filter should remain in, an evaluation was made to determine

the effects of changing parameters within the filter. The compensating

surface-actuator filter mode frequency was changed plus and minus 20

percent of the nominal value. The results are summarized on Figure 108

for the symmetric FSS and indicate that the FSS is relatively

insensitive to filter parameter changes. The filters are implemented

with the highest grade of components which are relatively insensitive to

environmental changes. Figures 88 and 89 showing the root locus of

this sensitivity analysis can be found in Appendix D.

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6.3.5.2.3 Pressure Feedback. Pressure feedback was used in the

servoactuator closed loop synthesis prior to removal during bench

testing. The effects of pressure feedback gain sensitivity on FSS

stability were analyzed by reducing the gain by 30 percent and little

decrease in stability was noted, however the antisymmetric lag phase

margin was reduced to 40 degrees. The results are shown on Figures 93

thru 96 in Appendix D.

Pressure feedback affects the closed loop actuator response in several

ways as was evident during bench testing. The nonlinearities associated

with actuator friction are much more prominent when pressure feedback is

used. The pressure feedback loop measures the force required to

overcome friction thereby decreasing the loop gain at low frequencies.

The effects of friction diminish with increase in frequency effectively

returning the loop to nominal gain. Therefore when enough pressure

feedback is used to damp the actuator at the resonant frequency, the

loop gain at low frequencies is inadequate.

6.3.5.2.4 Hinge Moment Effects. The hinge moment changes due to

aerodynamic forces during flight affect the gain and phase margins of

the servoactuator loop. These results are summarized by Figure 109 as

the hinge moment is varied and shown in Appendix D on Figures 97

through 104. Figure 110 is included here to show that the antisymmetric

FSS does not require compensation to cancel the surface-actuator mode as

did the symmetric FSS. The FSS was synthesized using the no-load hinge

moment actuator transfer function. Following servoactuator sensitivity

analysis, conclusions are summarized as follows:

• Series 31 or AR-25 Servovalves may be used

• Symmetric FSS requires Actuator compensation

Relatively insensitive to +_20 percent frequency change of surfaceactuator mode

Uncompensating antis3n_metric FSS reduces phase margin to +_40

degrees

Thirty percent reduction in pressure feedback gain reduces

antisymmetric FSS phase margin to +_40 degrees

Increased hinge moments decrease servoactuator gain resulting in

FSS gain reduction

6.3.5.3 Parameter Scheduling. The ARW-2 is thrust limited to 0.87 Mach

and scheduling below these speeds is not required. Sensitivity to

variations in antisymmetric gain scheduling was not performed because

the antisymmetric subsystem meets the required gain and phase margins

within the realizable flight envelope without gain scheduling.

Parameter or gain scheduling is not required for the symmetric axis for

any Mach number within the scope of the analysis performed. FSS

scheduling is discussed in Paragraph 7.4.3.

159
HINGE MOMENT PHASE GAIN

NO LOAD

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HINGE MOMENTSENSITIVITY

OUTBOARD SERVOACTUATORGAIN AND PHASE RESPONSE @ 23,8 HZ

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scheduling at least +45 degrees phase margin and +6 dB gain margin at

the design condition-of 0.86 Machand 15,000 feet-altitude.

Figures 82 and 83 showthe open and closed loop symmetric and

antisymmetric modedamping and frequencies at nominal gain and phase.

These figures show that the FSSmeets the dampingrequirements that no

degradation in structural modedampingshall be below a dampingratio of

0.01 except whenthe unaugmentedmodedamping is less than 0.01.

Comparabledata at other flight conditions are shownon Figures 38

through 47 found in Appendix C. Figure 84 and 85 show the symmetric

and antisymmetric flutter modedampingas a function of Machnumberat a

constant altitude of 15,000 feet. Symmetric and antisymmetric flutter

modedampingas a function of altitude and a constant Machnumberare

shownon Figures 86 and 87 respectively. Similar data is presented in

Appendix C on Figures 48 through 59 at other analyzed Machnumberand

altitudes.

6.3.6 SystemsCompatibility. Flutter suppression system compatibility

with the load alleviation systems, the relaxed static stability system

and the AFCSwere verified by determining the effects of the FSSon

rigid body modes, and the ACSand AFCSsystems on the flutter mode.

System compatibility was verified by linear system analysis and six

Degree of Freedom(6 DOF)simulation'analysis. Structural dynamic and

QSEEOMwere used in the analysis to determine the effect of each system

coupling. A summaryof FSSsystem compatibility study is shownon

Figure 111 by presenting the effect on modedamping with various system

loop closures.

The block diagram shownon Figure 174 was linearized and used to

evaluate compatibility of the flutter suppression system. System

compatibility is further discussed in Paragraph 6.8.

6.4 Load Alleviation SystemsAnalysis. Changesin the mathematical

model discussed in Section 4 dictated a reevaluation of the load

alleviation systems to determine if the changes affected the load

reduction capability of the systems. The load alleviation systems

defined during the previous design cycle were evaluated for the

following performance objectives:

To achieve the original goal of 15 percent reduction in critical

wing root vertical bending

• To makethe wing loads maneuvercritical

To achieve the design loads established by performance of the load

alleviation during the previous design cycle

162

Incremental wing load reduction with the gust load alleviation was

comparable to that predicted in the former analysis. However, steady

state loads at the gust design condition were greater than previously

predicted.
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The maneuver load alleviation system provided less incremental load

reduction than the former prediction, but 1 g loads did not increase at

the maneuver design condition. With load alleviation system active,

total maneuver design loads were less than gust design loads; therefore,

wing loads are not maneuver critical as desired.

The load alleviation system provides an 18 percent reduction in total

critical wing root vertical bending moment. Although total wing loads

with load alleviation system active are greater than the former

prediction, positive margins of safety are maintained at critical stress

points.

A method and supporting data are presented in Paragraph 6.4.4 to utilize

pseudorandom excitation of the inboard aileron to evaluate wing load

reduction capability of the gust load alleviation system.

6.4.1 Gust Load Alleviation System. The GLA system defined during the

previous design cycle was evaluated for gust load reduction capability

with the revised mathematical model.

The revised mathematical model produced larger wing root vertical

bending moment (WRVBM) in the basic airplane as shown on Figure 112.

The GLA system provided approximately the same incremental reduction in

WRVBM as shown on Figure 112; however, because steady state (Ig)

airplane loads were i,ncreased, the design load values established in the

previous design cycle were not achieved.

The stabilizer loop of the previously defined system produced excessive

stabilizer rate requirements at high frequencies which were not

contributing to the load reduction capability of the system. These rate

requirements exceeded the capability of the stabilizer actuation

hardware. During this design cycle, a first order filter of 90/(S + 90)

was added to attenuate high frequency feedback signals. This filter did

not affect the low frequency feedback response which provides the load

reduction function of the system. PSD/RMS plots of stabilizer rate

response to random gust of Figure 113 show the benefit of adding the

stabilizer filter. The additional stabilizer filter reduces peak rate

requirement for the design gust level from 188 degrees/second to 97

degrees/second.

The GLA system contains gain scheduling to reduce outboard aileron and

stabilizer deflections at dynamic pressures greater than that at the GLA

design condition. The previous system which attenuates the feedback

signal inversely proportional to dynamic pressure did not have adequate

phase margin in the stabilizer loop at the off design maximum dynamic

pressure condition as shown on Figure 114(A). The gain schedule was

revised to provide additional gain reduction in both the outboard

aileron and stabilizer loops at high dynamic pressures. The gain

reduction provided adequate phase margin at maximum dynamic pressure as

shown on Figure 114(B).

164
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• GROSS WEIGHT: 2500 LBS.

• PRIMARY AFCS, RSS AND GLA SYSTEMS CLOSED

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A block diagram of the revised GLAsystem is shownon Figure 115.

Neither of the system changesabove affect the load reduction capability

of the GLA at the GLA design condition.

Power spectral density/root mean square plots of wing root vertical

bending, fuselage vertical acceleration and GLA surface displacements

and rate at the GLA design condition are shown on Figure 116. GLA

system performance is based on the Von Karman gust spectrum with a scale

length of 2,500 feet. The peak gust amplitude is 62 feet/second at the

GLA condition.

The GLA design condition is on the V curve (350 KCAS) at 7,000 feet

where the maximum gust occurs. Sinc_ flight tests are limited to

altitudes above 10,000 feet, a test condition of 350 KCAS at 15,000 feet

was selected. Airplane response to gust at the test condition should be

very similar to that at the design condition as indicated by wing root

vertical bending PSD/RMS plots shown on Figure 117.

6.4.2 Maneuver Load Alleviation System. The MLA system defined during

the previous design cycle was evaluated for maneuver load reduction

effectiveness with the revised mathematical model.

The MLA system consists of the GLA system and symmetric inboard aileron

and an additional stabilizer signal. The inboard aileron and additional

stabilizer signal are controlled by inputs from the longitudinal primary

AFCS. With the revised mathematical model for this analysis, maneuver

loads for steady state (lg) flight were approximately the same as before

the revision. Wing control surfaces were less effective than the

"previous predictions; therefore, the MLA system did not provide as much

maneuver load reduction. Figure 118 compares WRVBM to that of the

previous design cycle, The MLA provides less incremental load reduction

than previously predicted; however, basic airplane maneuver loads had

not increased, and as a result, maneuver loads were less than gust

loads. Changes in the model which affect MLA performance are explained

in Section 4.

During this design cycle, two modifications were made to the MLA inboard

aileron and stabilizer command system. Previously, the inboard aileron

command was proportional to the load factor command input to the primary

AFCS. The aileron command was revised to make the control surface

proportional to actual load factor response for load factor commands

greater than +O.5g incremental and equal to load factor command

otherwise. This change was made to assure that the MLA system response

corresponds to actual large load factors when, at some flight conditions

and C.G. positions, airplane response may not exactly equal the

Command. The other modification changed the signal gain relating

stabilizer to inboard aileron from 0.28 to 0.11 to adjust for different

pitching moment effects of the inboard aileron in the revised model.

The revised MLA inboard aileron and stabilizer control system is shown

in the block diagram on Figure 119.

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• GROSS WEIGHT: 2500 LBS.

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PEAK AMPLITUDE : 62 FPS

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171
• MACH 0.60

• ALTITUDE: 7000 FEET

• GROSS WEIGHT: 2500 LBS.

m PRIMARY AFCS, RSS AND GLA SYSTEMS CLOSED

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Figure 118 also shows WRVBM for a selected flight test condition at

10,000 feet. The design condition occurs at the minimum speed and

altitude that a 2.5g maneuver can be executed. The test condition for

positive maneuvers was selected as the minimum speed that a 2.5g

maneuver can be executed at 10,000 feet.

Maneuver time responses of load factor, WRVBM, and control surfaces with

the MLA are shown on Figure 120.

6.4.3 Load Alleviation System Performance. Gust loading produced

maximum WRVBM for both basic airplane and with load alleviation systems

engaged. The reduction in critical WRVBM with the load alleviation

systems closed is 18 percent; however, maximum loads with the systems

engaged are greater than the design loads predicted in the previous

design cycle.

Gust and maneuver loads along the wing span with load alleviation

systems closed are shown on Figures 121 and 122. Vertical bending

moment along the entire span is greater for gust; consequently, the wing

is not maneuver critical as desired.

Since the design loads of the previous design cycle were not achieved, a

structural resizing iteration was anticipated. A stress analysis

exercise using the critical combination of wing loads from Figures 121

and 122 indicated that positive margins of safety could be obtained at"

critical stress points by eliminating some conservatism in the.previous

stress analysis. The previous stress analysis did not include the load

carrying capability of effective areas of the skin forward of the front

spar. A comparison of minimum margins of safety of critical components

for the previous and present stress analyses is shown on Figure 123.

6.4.4 Simulated Gust Analysis. The difficulty and time required to

find large amplitude gust conditions and DAST flight time limitations

make it desirable to have an alternate method of verifying the gust load

reduction capability of the GLA system and verifying analytical modeling

of the airplane and load dynamics.

Gust load reduction is provided by outboard aileron and stabilizer

response. The inboard aileron can be used to generate airplane load

responses similar to that produced by random gust. Random excitation of

the inboard aileron produces WRVBM response having PSD characteristics

similar to that due to random gust disturbances. Desired PSD

characteristics are obtained by passing a signal from a shift register

noise generator through a set of first order lag and washout filters

selected to give the desired PSD shape.

Implementation of the inboard aileron random excitation generator is

shown on Figure 124. The shift register noise generator provides a good

representation of white noise out to approximately 10 percent of the

clock frequency. The noise generator has a repetitive cycle time of

approximately 174 minutes.

176
• MACH 0.35

m ALTITUDE: SEA LEVEL

• GROSS WEIGHT: 2500 LBS.

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LOAD FACTOR AND WING ROOT VERTICAL BENDING MOMENT RESPONSES FOR DESIGN MANEUVER

177
• MACH 0.35

• ALTITUDE: SEA LEVEL

• GROSS WEIGHT: 2500 LBS.

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INBOARD AILERON AND OUTBOARD AILERON RESPONSES FOR DESIGN MANEUVER

178
• MACH 0.35

• ALTITUDE: SEA LEVEL

• GROSS WEIGHT: 2500 LBS.

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AILERON RANDOM EXCITATION

188
• MACH 0.70

• ALTITUDE: 15,000 FEET

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187
e FtACH0.70

• ALTITUDE: 15,000 FEET

• GROSSWEIGHT: 2350 LBS.

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189
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FIGURE 126

PREDICTED INCREMENTAL WING ROOT VERTICAL BENDING MOMENT REDUCTION WITH GLA FOR

MAXIMUM AVAILABLE INBOARD AILERON AND EQUIVALENT GUST

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The noise filter on Figure 124 shapes the generated white noise to

roughly match the PSD of the WRVBM due to gust as shown on Figure 125.

Limitations on the number of filters that could be incorporated in the

hardware package determined how close a match that could be produced.

The intent was to place a large percentage of the power of WRVBM due to

inboard aileron in the same frequency range as that due to gust. As

indicated on Figure 125, the system adequately accomplishes that intent.

Performance of the GLA system can be verified by relating test data with

inboard aileron excitation toanalysis data with inboard aileron

excitation and gust. All baseline GLA testing should be conducted with

the primary AFCS and RSS engaged. Performance of the GLA system can

also be verified by comparing the incremental improvement beyond that

obtained with these systems. The predicted reduction in WRVBM with the

maximum available inboard aileron excitation and an equivalent level of

random gust is indicated on Figure 126. The reduction in WRVBM with

aileron excitation is 16.5 percent and 14.3 percent with random gust.

With the similarity in predicted percent reduction and the similarity of

load power as a function of excitation frequency, a close match of

inboard aileron excitation test data to analysis data should

substantiate that the gust load increment will be as predicted.

With maximum available inboard aileron excitation, peak loads are

predicted to be well below the design load limit as shown on Figure 127.

PSD/RMS plots of WRVBM due to inboard aileron random excitation and gust

are shown on Figures 128 and 129, _espectively. Predicted steady-state

(lg) values of WRVBM and incremental RMS (or peak) values per unit RMS

(or peak) inboard aileron and gust excitation are shown on Figure 130.

RMS (or peak) ratios of fuselage vertical acceleration are also shown on

Figure 130.

Gain of the inboard aileron random excitation filter was initially

scaled for required RMS values on the six DOF airplane dynamics

simulation discussed in Reference i and later verified by hardware tests

discussed in Section 9. Time responses and RMS data of random

excitation and airplane motion were obtained from the simulation.

Random gust in the six DOF simulation was also produced by the shift

register pair on Figure 124 combined with a set of four first-order

linear filters selected to closely approximate the frequency content of

the Von Karman gust spectrum.

Typical time responses of input excitation and airplane vertical

acceleration are shown on Figures 131 and 132. The ratio of vertical

acceleration to input excitation agrees well with PSD/RMS analysis

results shown on Figure 130.

The measured probability distribution function of the inboard aileron

random excitation is compared to the Gaussian distribution on Figure

133. The inboard aileron random excitation provides a reasonable match

to Gaussian characteristics as indicated on Figure 133.

184
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6.5 Relaxed Static Stability System Analysis. RSS System defined in

the previous design iteration was evaluated for the capability to

maintain stability at all flight conditions and C.G. variations with

fuel burn.

The RSS system permits flight at aft, unstable C.G. positions to take

advantage of reduced drag associated with reducing or eliminating down

trim tail loads by placing the C.G. aft.

The RSS permits flight at an aft C.G. position of 33 percent MAC over

the entire flight envelope. The neutral point of the trim cruise

condition is 38.6 percent MAC as shown on Figure 37; however, a large

forward shift in neutral point at other flight conditions limits the

capability of the RSS to stabilize the aircraft at C.G. positions aft of

33 percent MAC. The largest static instability that the RSS stabilizes

is -18.5 percent MAC at the maximum dynamic pressure condition.

At the DAST 2 launch condition, the static neutral point was determined

to be 29.7 percent MAC. The launch C.G. was selected by NASA to be 20

percent MAC, providing a stable launch condition. The weight to C.G.

relationship for the selected fuel burn sequence is shown on Figure 39.

The cruise drag versus C.G. position shown on Figure 40 indicates that

minimum trim drag is well back of the 33 percent aft C.G. limit.

The critical condition for static stability are indicated by the CM

versus _ curves on Figure 134. At the maximu_ dynamic pressure

condition the CM curve has an unstable slope at trim angle of attack.

At the cruise condition, an unstable CM slope occurs during positive

maneuvers. Neutral points for maximum"dynamic pressure and a 1.2g

maneuver at cruise are 14.5 and 16.7 percent MAC, respectively. These

two conditions are the most critical for RSS performance.

A revision of the RSS was required to meet stability criteria at the

critical conditions. Without the primary AFCS engaged, which provided

some stability, a gain increase of 1.7 was required to meet minimum gain

margin. The critical condition which dictated the 1.7 gain increase is

maximum Q flight condition, 33 percent MAC C.G. position with the backup

AFCS engaged. A root locus for this condition is shown on Figure 199 of

Appendix E.

A low frequency lag-lead filter (S + .I)/(S + .05) was added to the RSS

feedback loop to double the gain at zero frequency. At the aft C.G.,

1.2g maneuver cruise condition with the primary AFCS and also the GLA

system closed, the low frequency gain increased the stability of an

unstable real characteristic root which approached a zero near the

origin as shown on Figures 180 and 182 of Appendix E.

The revised RSS system is shown in the block diagram on Figure 135.

Short period and phugoid damping and frequency from the evaluation of

RSS and other appropriate systems with the structural dynamic equations

of motion are shown on Figures 136 and 137. The weights of Figure 136

and 137 are the most typical weights for each test condition.

194
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MAXIMUM DYNAMIC PRESSURE

MACH: 0.86

ALTITUDE: 15,000 FEET

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MACH: 0.80

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PITCHING MOMENT COEFFICIENT VS. ANGLE-OF-ATTACK AT CRITICAL STABILITY CONDITIONS

195
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Short period characteristics at the weight and C.G. extremes obtained

with the QSE equations of motion are shown on Figures 166 and 167.

6.6 Automatic Flight Control Systems. The primary and backup AFCS

systems defined during the previous study were re-evaluated using

revised aerodynamic and physical properties.

The longitudinal primary AFCS was revised to be more compatible with

larger variations in aerodynamic characteristics at different flight

conditions. An altitude hold mode was added to the primary longitudinal

AFCS. The altitude hold mode and other AFCS functions are compatible

except during banked turns where an offset in altitude occurs as

discussed in Paragraph 6.8.1.

Simulation time response evaluation still indicates a possibility of

exceeding positive load factor limits if the backup system is engaged in

a descent condition at high dynamic pressure conditions. To eliminate

the possibility of exceeding limit load factoG a circuit which will ramp

in the initial stabilizer step command change should be considered for

the longitudinal backup AFCS.

The lateral-directional AFCS control laws required no change in the

present design cycle. Signal synchronization of primary and backup

lateral-directional AFCS modes was added.

6.6.1 Primary Longitudinal AFCS. The primary function of the system is

to accept operator inputs in the form of a load factor command and

process the input and load factor feedback to generate a stabilizer

command to provide desirable airplane response. Additional functions of

the system are to provide a MLA system command when that system is

engaged and to provide a positive angle of attack limit. A block

diagram of the revised system is shown on Figure 138. The changes to

the system are discussed below.

With aerodynamic characteristics used in the previous design iteration,

steady-state load factor response agreed closely with load factor

command. The close agreement in response and command previously allowed

a load factor limit feature to be implemented by limiting the command.

Due to the wider variations in aerodynamic characteristics in the

present iteration, the response did not always agree exactly with the

command; and consequently, the limits on the command had to be removed

to provide for limit maneuvers in the cases when the response was less

than the command.

With the previous close agreement of response to command, a low gain

load factor error signal was integrated and summed into the RSS

integrator to provide further agreement of response to command. With

the present differences in response and command, an input produced an

initial response that for some conditions was significantly different

199
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than the command and the integrator produced a sluggish ramp up to the

input level. It also produced a sluggish decay after the input was

removed.

The load factor error integrator was eliminated in the present iteration

because a quick rising response that did not correspond exactly with the

command was considered to be more desirable than one with a sluggish

ramp response after an initial buildup.

A 1/cos(@) load factor term which previously eliminated the operator

having to apply a conventional column input for a steady-state turn was

eliminated. With the difference in response and command in some cases,

the column input with the 1/cos(_) term, although small, might be

opposite that of a conventional column command in a turn.

In the revised system, a third order filter duplicating the RSS filter

was eliminated and the pseudo pitch rate command signal generated by the

AFCS was summed into the RSS feedback loop upstream of the RSS filter

instead of downstream. This change eliminated duplicating the complex

filter but required one more uplink signal than the previously defined

configuration.

A slight modification which simplified the angle of attack limiter

reduced the amplitude of response transients when the angle of attack

limit value was exceeded.

Another change to the primary longitudinal AFCS added an altitude hold

mode option to the system as shown on Figure 138. Compatibility of the

altitude hold mode is discussed in Section 6.8.1.

Typical airplane motion time responses for symmetric maneuvers with the

primary AFCS and other appropriate systems at flight conditions spanning

the flight range are shown on Figures 139 through 146. These responses

illustrate the following characteristics:

e Variation of responses with center of gravity

Variation of responses with use of the load alleviation systems

Effects of the angle of attack limiter on responses

A flow diagram for the ground based primary AFCS digital program is

included in Section 6.9.

6.6.2 Lateral-Directional Automatic Flight Control System. The

lateral-directional AFCS defined in the previous design cycle was

evaluated with the revised QSE mathematical model. Revisions to the QSE

201
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RESPONSE AT LAUNCH CONDITION FOR POSITIVE LOAD FACTOR COMMAND

202
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ANGLE 0 iii;i_; i!ili_!i!';!Lil,i_!li'_!:;il;

ili!: i',_ !i!i ii_! _!! !_iil;i;;_H!!!I!]_

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POSITION 0 ;i:;I __:lii::_]i lil_,_l_:i':,_!

:'"'iIi Z.:_ :ili'iii:.

( )

iilllil;i::ii_'"'' ,_!Ui

(A) MLA/GLA OFF

ORIGIN_L PAGE I_

OF POOR QUAL/Ty

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(B) MLAIGLA ON

FIGURE 140

RESPONSE AT LAUNCH CONDITION FOR NEGATIVE LOAD FACTOR COMMAND

203
• MACH: 0.70

• ALTITUDE: 15,000 FEET

• C.G.: 27.5%MAC

• PRIMARYAFCSANDRSSON

OPERATOR

COMMAND

(g's)

INCREMENTAL

LOAD FACTOR

(g's) •

ANGLE

OF

ATTACK

(DEG.)

PITCH

ANGLE

(DEG.)

STABILIZER 4

POSITION 0

(DEG,) -4

-8

(A) MLA/GLA OFF

H, .!:-)

iiii_i_i ii

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_+

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[;_i+i_ti++_!I: iililqi+_:i::ilI!_:i 'i +

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, ;! :: Hf!!::f_+HI;:_I::,:, : ....... + !,

ii!i_!ilil+;qiii_i_i+luhF::l._;i:._:', '+

(B) MLA/GLA ON

FIGURE 141

RESPONSE AT GLA TEST CONDITION FOR POSITIVE LOAD FACTOR COMMAND

204
OPERATOR

COMMAND

(g's)

INCREMENTAL

LOAD FACTOR

(g's)

ANGLE

OF

ATTACK

(BEG.)

PITCH

ANGLE

(BEG.)

STABILIZER

POSITION

(BEG.)

• MACH: 0.70

• ALTITUDE: 15,000 FEET

• C.G.: 27.5% MAC

m PRIMARY AFCS AND RSS ON

4

2

0

-2

-4

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ihl iiil_tm_:l :!:ti:.i:::tl I :

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(B) MLA/GLA ON

FIGURE 142

RESPONSE AT GLA TEST CONDITION FOR NEGATIVE LOAD FACTOR COMMAND

205
• MACH: 0.42

• ALTITUDE: 10,000 FEET

• PRIMARY AFCS, RSS AND MLA/GLA ON

OPERATOR

COMMAND

(g's)

INCREMENTAL

LOAD FACTOR

(g's)

ANGLE

OF

ATTACK

(BEG.)

PITCH

ANGLE

(BEG.)

STABILIZER

POSITION

(BEG.)

4

2

0

-2

-4

4

2

0

-2

-4

8

4

0

-4

-8

2O

10

0

-10

-20

8

4

0

-4

-8

+

I

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l l l l l I I ! k

i:! i:]:

i i/!i.

(c)

C.G.: 33% MAC

FIGURE 143

RESPONSE AT MLA TEST CONDITION FOR POSITIVE LOAD FACTOR COMMAND

206
OPERATOR

COMMAND

(g's)

INCREMENTAL

LOAD FACTOR

(g's)

ANGLE

OF

ATTACK

(BEG.)

PITCH

ANGLE

(BEG.)

STABILIZER

POSITION

(BEG.)

• MACH: 0.80

• ALTITUDE: 46,800 FEET

• C.G.: 33% MAC

• PRIMARY AFCS AND RSS ON

ORIGINAL PAGE IS

OF POOR QUALITY

1.0

0.5

0

-0.5

-I.0

1.0

0.5

0

-0.5

-i.0

4

2

0

-2

-4

20

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!_it! i I

:: mi " i i

i i

;: i !

_ ...._;-- --_:_

I,,L i

!<:i: ;

i;-L'4

(B) MLA/GLA ON

FIGURE 144

RESPONSE AT CRUISE FOR POSITIVE O.5g LOAD FACTOR COMMAND

207
OPERATOR

COMMAND

(g!s)

INCREMENTAL

LOAD FACTOR

(g's)

ANGLE

OF

ATTACK

(BEG.)

PITCH

ANGLE

(BEG.)

STABILIZER

POSITION

(BEG.)

• MACH 0.80

• ALTITUDE: 46,800 FEET

e C.G.: 33% MAC

• PRIMARY AFCS AND RSS ON

4

2

0

-2

-4

4

2

0

-2

-4

8

4

0

-4

-8

20

10

0

-10

-20

8

4

0

-4

-8

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FIGURE 145

RESPONSE AT CRUISE FOR POSITIVE 1.5g LOAD FACTOR COMMAND

208
I..,R._GINAL PAGE Ig

_]_ I_K). R QUALIT_

OPERATOR

COMMAND

(g's)

INCREMENTAL

LOAD FACTOR

(g's)

ANGLE

OF

ATTACK

(BEG.)

PITCH

ANGLE

(BEG.)

STABILIZER

POSITION

(BEG.)

• MACH: 0.86

• ALTITUDE: 15,000 FEET

• C.G.: 33% MAC

• PRIMARY AFCS AND RSS ON

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.(A) MLAIGLA OFF

FIGURE 146

RESPONSE AT MAXIMUM DYNAMIC PRESSURE FOR POSITIVE LOAD FACTOR COMMAND

209
lateral-directional mathematical model consisted of a changeto the

yawing momentof inertia associated with the shift in center of gravity

and correcting a sign error on the control derivative CY6 RUD"

Figure 147 showsbasic airplane lateral-directional characteristics at

three flight conditions. The first two conditions which are low dynamic

pressure cruise conditions have the most unstable Dutch roll

characteristics and the third condition which is the maximumdynamic

pressure condition has the largest Dutch roll frequency. With the

present mathematical description, Dutch roll dampingwithout the AFCSis

not as unstable as previously indicated; however, it is still unstable

at the cruise conditions and does not meet minimumdampingcriteria at

any of the flight conditions.

Root loci of the individual rudder feedback loops, both of which provide

Dutch roll stability, are shownon Figures 148 and 149 for the high

altitude cruise condition. Lateral-directional airplane characteristics

with the AFCSare shownon Figure 150 for thethree flight conditions.

Dutch roll dampingwith the AFCSis also greater than with the previous

model.

The function characteristics of the lateral-directional AFCSrequired no

change in the present design cycle; however, circuits were addedto the

roll axis feedback loops and commandchannels to synchronize the primary

and backup modesignals. The revised system is shownin the block

diagram on Figure 151. A flow diagram of the ground based primary AFCS

digital program is included in Section 6.9.

6.6.3 Backup Longitudinal AFCS. The backup longitudinal AFCSis shown

in the block diagram on Figure 152. Nochangeswere madethat affected

the functional characteristics of the system during this design

iteration. The block diagram has been revised to correct an error in

gain distribution shownon Figure 7-72 of Reference 1. Also, the block

diagram has been revised to show a signal going to the RSSintegrator

instead of a separate integrator as shownpreviously.

Time responses for flight subsequent to switching from primary modeto

backupmodeare shownon Figures 153 through 156 for various flight

conditions. The responses are transitions from attitudes or maneuvers

established in the primary modeto backupcontrol with no backup

operator commands. Figures 154 through 156 showresponses for backup

recovery from negative maneuversat three flight conditions. Recovery

from an initial pitch downattitude produces positive load factors that

for certain combinations of pitch attitude and high values of dynamic

pressure can exceed the wing design structural limit. Since the GLA/MLA

system will not be engagedin the backup mode, limit load factor is

approximately +2g (+Ig incremental). Figure 154 showsresponses for

recovery from a pitch downattitude of 15 degrees at the launch

condition. At this low dynamic pressure condition, the backup control

210
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211
• MACH: 0.70

• ALTITUDE: 5000 FT.

• RUDDER/ROLL RATE LOOP CLOSED

• DIFFERENTIAL STABILIZER/ROLL RATE LOOP CLOSED

NO.

I I I

GAIN dB

0.595 -4.5

0.0 (OPERATING GAIN)

I:ll I 4:5

-8.0

o (RAD/SEC)

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I

2.0

14.0

12.0

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jco (RAD/SEC)

6.0

_ 4.0

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4.0

FIGURE 168

RUDDER/YAW RATE LOOP PHASE-GAIN ROOT LOCI

212
• MACH: 0.70

• ALTITUDE: 50000 FEET

• RUDDER/YAW RATE LOOP CLOSED

• DIFFERENTIAL STABILIZER/ROLL RATE LOOP CLOSED

I.. i I I ") ..

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2 1.000 0.0 (OPERATING GAIN)

3 1.680 4.5

14.0

12.0

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PHASE ANGLE = 0°

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jo_ (RAD/SEC)

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0

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FIGURE 149

RUDDER/ROLL RATE LOOP PHASE-GAIN ROOT LOCl

213
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• MACH: 0.40

m ALTITUDE: 15,000 FEET

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INCREMENTAL

LOAD FACTOR

(gs)

STABILIZER

POSITION

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BACKUP MODE TRANSITION RESPONSE FROM PITCH UP MANEUVER AT LAUNCH CONDITION

217
• l,ta,CH: 0.40

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BACKUP MODE TPJIIINSITION RESPONSE FROM PITCH DOWNMANEUVER AT LAUNCH CONDITION

218
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219
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AT MAXIMUM DYNAMIC PRESSURE

220
system produced a +1.5g (+O.5g incremental) load factor. At the high

dynamic pressure GLAtest condition, recovery from 15 degrees pitch down

produced a load factor of 2.2g as shownon Figure 155. At the maximum

dynamic pressure condition, recovery from only 4 degrees initial pitch

downproduced a load factor of 2.4g as shownon Figure 156. These

results are similar to those obtained with the previous mathematical

model.

6.7 Horizontal Tail Authority Evaluation. The DASTstabilizer

configuration provides +7 degrees (trailing edge down) and -12 degrees

displacement of the individual stabilizer surfaces which are moved

symmetrically for pitch control and antisymmetrically for roll control.

Twofeatures of the DASTII configuration increase the requirement for

positive stabilizer. More positive stabilizer is required to trim out

the positive pitching momentdue to moving the center of gravity aft for

minimumtrim drag. Also, positive stabilizer instead of conventional

negative stabilizer is required for positive maneuverswith the maneuver

load alleviation system to counteract the pitching characteristics of -

the wing control surfaces. These stabilizer requirements are

illustrated on Figure 157 which shows load factor and stabilizer

responses for positive maneuverswith MLAas the C.G. is movedaft. At

the aft C.G. limit, symmetric stabilizer is very near the positive

limit. Any significant roll commandat the 33 percent C.G. position

will drive one stabilizer surface into the positive limit. Symmetric

stabilizer is required for aircraft stability at aft C.G. conditions and

saturating one surface essentially reduces the gain of the stabilizer

feedback systems by one-half. This alters the load factor to operator

commandrelationship which could possibly result in a response exceeding

the design limit if operator Corrective symmetric action does not

accompanythe antisymmetric commandsas illustrated on Figures 158 and

159. Saturating one surface could also potentially create an

instability since one-half gain is below the minimumgain margin

designed into the systems.

Analysis results indicate that more positive stabilizer authority is

desirable. Without increased positive authority, extreme caution should

be used when applying antisymmetric commands during large positive

maneuvers at low dynamic pressure conditions with aft C.G. positions and

the load alleviation system engaged.

6.8 Control System Compatibility. Operational and stability

requirements dictate that some of the ACS/AFCS systems operate together.

Various analysis tasks were conducted to assure that systems operating

together were compatible. The following four combinations of system are

required to operate togethe r during some phase of flight tests:

• RSS and PCS

• RSS, PCS and GLA/MLA

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Operations of the PCS systems will at times include the altitude hold

mode.

Tasks to verify system compatibility involved linear analysis and

evaluation of the system combinations on the non-linear 6 DOF airplane

simulation. The functional capability of each individual system was

verified with other appropriate systems closed. These effects are

included in the sections of this report which discuss the individual

systems. Other tasks to determine tangible performance such as

stability margins and handling qualities are discussed in the following

subsections.

6.8.1 Altitude Hold System Compatibility. The Altitude Hold Control

System shown on Figure 138 supplied by NASA was evaluated for control

effectiveness and compatibility withthe other control systems.

The altitude hold system is a very low gain system and does not

significantly affect airplane characteristic roots as shown on Figure

160. Since the system had an insignificant effect on airplane

characteristic roots, the effects of the system were not included in the

evaluation of stability margins or frequency and damping discussed in

the following sections.

Typical airplane vertical translation due to random gust with the

altitude hold system is shown on Figure 161. Peak vertical displacement

is on the order of 3 feet per feet/second RMS vertical gust at the GLA

design condition. Figure 162 shows airplane response at cruise when the

altitude hold system is engaged in an initial 10 degree dive. The

airplane settles out at the altitude at the time the altitude hold was

engaged in approximately 30 seconds.

The altitude hold system provides good control for straight flight but a

steady-state altitude offset occurs in turns as shown on Figure 163.

The altitude offset occurs because the altitude hold system must

generate a steady-state error signal to compensate for a steady-state

pitch rate feedback signal to the stabilizer through the RSS system. If

it becomes desirable to eliminate the altitude offset in turns, it is

suggested that the ground based digitally programmed altitude hold

system provide an additional function to the signal transmitted to the

RSS filter which, computed as a function of roll angle and airplane

velocity, will compensate for steady-state turning pitch rate.

225
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EFFECTS OF STEADY STATE BANK ANGLE ON ALTITUDE HOLD PERFORMANCE

229
6.8.2 System Gain and Phase Margins. Gain and phase stability margins

of each individual loop of each system were determined for each

combination of systems that will be closed during some phase of flight

testing.

The FSS had an insignificant effect on the airplane basic translation

and rotation modes as discussed in Section 6.3. The stability margins

of the FSS on airplane structural dynamic modes are presented in Section

6.3.

All ACS/AFCS systems, excluding the FSS, did not significantly affect

airplane structural modes as indicated on Figure 164; therefore, the

stability margins of these systems were determined using GLSEEOM. Those

system combinations and the various flight conditions and C.G. extremes

where stability margins were determined are shown on Figure 165.

Phase and gain root loci for each element on Figure 165 are shown in

Appendix E.

All system loops meet gain criteria of +4.5 dB at all conditions.

Except for the maximum dynamic^pressur_aft C.G. condition, all loops

meet the phase criteria of +30 v. At the maximum dynamic pressure-aft

C.G. condition, the RSS and-PCS had only approximately 25 degrees phase

lag when these two systems were the only symmetric systems engaged.

These system phase margins characteristics are shown on Figures 196 and

197 of Appendix E. With the GLA/MLS system engaged, the RSS and PCS

loops meet phase margin criteria. No testsat the maximum dynamic

pressure condition without GLA/MLA engaged are anticipated.

6.8.3 Flying Qualities. Airplane dynamics were evaluated for flying

quality characteristics with each combination of systems that may be

closed.

The lightest damped closed loop system characteristic roots in the short

period frequency range are compared to MIL-F-8785C Level 1 short period

damping and frequency criteria on Figure 166. Time response evaluation

on the 6 DOF simulation indicates that these characteristic roots are

dominant during typical maneuvers. The closed loop systems produce high

values of short period frequency, slightly exceeding the criteria in

some cases and just below maximum criteria in the other cases. Short

period frequency increases with increasing gain in the RSS and PCS

feedback loops and are highest for forward C.G. positions. The gain

values were selected to provide gain margins of the divergent real roots

resulting from unstable pitching characteristics at aft C.G. positions.

Except for operation of the RSS and PCS only at the maximum dynamic

pressure condition, the systems meet minimum damping ratio criteria. At

this condition, the GLA provides criteria level damping ratios with some

increase in short period frequency as indicated on Figure 166.

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Short period frequency and damping with the BCS engaged are compared to

Level 3 criteria on Figure 167. With the relaxed criteria for backup

mode of operation, the systems meet frequency and damping criteria at

all flight conditions and C.G. extremes.

Each combination of closed loop systems meets phugoid damping criteria.

No lightly damped or unstable roots occur in the low frequency (phugoid)

frequency range as indicated in the root loci plots on Figures 189

through Figure 204 of Appendix E.

6.8.4 Airplane Stability Derivative Sensitivity. ACS and AFCS

sensitivity to variations in longitudinal stability derivatives CM_,

CLa, CMQ, CDU and CLu and lateral-directional derivatives Cl#, Cl#,

CN# and CNR were evaluated on the 6 DOF simulation of airplane

dynamics. During the last iteration, sensitivity to stability

derivative variations was determined by computing airplane

characteristic roots with +_20 percent change in the derivatives. The

stability derivatives were varied individually at each of six flight

conditions listed on Figure 166. No unstable or even lightly damped

characteristics were produced with stability derivative variations; and

for step control surface inputs, changes in airplane responses with the

stability derivative variations were minimal.

As indicated on Tables 7-1X and 7,XV of Reference 1, the most lightly

damped Dutch roll response occurred with variations in CIR. Airplane

response with variations in Clp at the most lightly damped condition

indicated in Reference 1 is shbwn on Figure 168.

6.9 Computer Implementation of Ground Based Control Systems. The

following control systems will be implemented on the ground based

computer:

o MLA Inboard Aileron and Stabilizer Functions

• Primary Longitudinal AFCS

Differential Stabilizer/Roll Rate Loop of Primary LateralDirectional AFCS.

With the restriction that the GLA/MLA systems should always be engaged

at high dynamic pressure conditions, simulation and analysis results

indicate satisfactory performance of the ground based systems using the

expected iteration time increment of 0.025 second.

Difference equation constants which are to be computed using the final

ground station/drone data transmittal iteration time increment (DT) are

given on Figure 169. Flow diagrams for implementation of the ground

based systems are shown on Figures 170 through 173.

234
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235
• MACH: 0.80

• ALTITUDE: 46800 FEET

m GROSS WEIGHT: 2350 LBS.

• SYSTEM RSS, GLA, PCS AND ALTITUDE HOLD ENGAGED

• +i ° STEP RUDDER FOR 1 SECOND

ROLL

ANGLE

(BEG.)

25

-25

, , I

ROLL

ANGLE

(BEG.)

ROLL

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(BEG.)

I I 1 I I I I I I I I ! I I I I I I i

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FIGURE 168

ROLL RESPONSE FOR STEP RUDDER WITH C£p VARIATIONS

236
CC

YAZ : EXP (-0.5 * DT)

XAZ : (YAZ - 1.0 + 0.5 * DT)/(O.25 * DT)

WAZ : (i.0 - YAZ-O.5 * YAZ * DT)/(O.25 * DT)

YIO = EXP(-IO. * DT)

XlO = (YIO - 1.0 + 10. * DT)/IO./DT

WIO : (I.0 - YIO - i0.* YIO * DT)/IO./DT

GLL : 25.0 * DT

FIGURE 169

AFCS AND MLA DIGITAL PROGRAMCONSTANTS

237
II

C LONGITUDINAL PRIMARY AFCS

C INPUT DATA

C Al (OPERATOR COLUMN COMMAND- INCREMENTAL G'S)

C DP (DYNAMIC PRESSURE - PSF)

C ALFDG (ANGLE OF ATTACK - DEG)

C AZI (INCREMENTAL NORMALACCELERATION - G'S)

C U (TRUE VELOCITY - FPS)

C ISH (ALTITUDE HOLD ENGAGE SIGNAL -

C ISH = 1, ALT. HOLD ON; ISH : O, ALT.. HOLD.OFF

C ZH (ALTITUDE - FT)

C ZHD (ALTITUDE RATE - FT/SEC)

C THTDG (PITCH ANGLE - DEG)

READ DATA INPUT /

AZSC : AI

AZM = 0.0015 * DP * (6.5 - ALFDG) + AZ1

AM = AZM +AZSC * (AZM - AZ1)

CALL LIMIT(AM, AZSC, 0.0)

AZSC : AM

()

FIGURE 170

PRIMARY LONGITUDINAL AFCS FLOW DIAGRAM

238
T

I AZSC= 0.0

p

I AZ ,AZSCI

AXI = -AZSC * 1844./U

AZSl = AZS2 * YAZ + AZSCP * WAZ + AZSC * XAZ

AZS2 : AZSl

AZS3 : 200.* AZSC/DP + (O.5-100./DP) * AZSl

EAZ : AZS3 - AZI

GDP = 400./DP

CALL LIMIT (GDP, 3.0, 0.0)

AC = -GDP * EAZ

DAI : AC - AIIP

CALL LIMIT (DAI, GLL, -GLL)

AC : AIIP + DAI

AIIP = AC

AZSCP : AZSC

FIGURE 170 (CONTINUED)

PRIMARY LONGITUDINAL AFCS FLOW DIAGRAM

239
AIH : O.025*(THTDG-THTP)+O.024 *

(ZHD-ZDP)+O.OO6*(ZH-RH)

CALL LIMIT (AIH, 2.0, -2.0)

AHX = 0.001 * (ZH - RH)

AC : AC + AIH

T

AIH : 0.0

AHX : 0.0

RH : ZH

THTP : THTDG

ZDP : ZHD]

_._.

C OUTPUT VARIABLES

C AC (STABILIZER PROPORTIONAL COMMAND- DEG.)

C AXI (INPUT TO RSS FILTER CIRCUIT - DEG/SEC)

C AHX (INPUT TO RSS INTEGRATOR - DEG/SEC)

C AZML (INPUT TO MLA INBOARD AILERON SYSTEM - G'S)

FIGURE 170 (CONCLUDED)

PRIMARY LONGITUDINAL AFCS FLOW DIAGRAM

240
C LATERAL-DIRECTIONAL PRIMARY AFCS DIGITAL PROGRAM

C INPUT DATA

C DSTI (OPERATOR DIFFERENTIAL STAB. COMMAND- DEG

C P (ROLL RATE - DEG/SEC)

C DSS (DIFFERENTIAL STABILIZER COMMAND

C SYNCHRONIZATION SIGNAL - BEG)

C ISP (PRIMARY AFCS ENGAGE SIGNAL -

C ISP : I, PCS ON; ISP : O, PCS OFF)

SL = DSTI - 0.1 * P

IF(ISP.L_

T

ISLI : DSS- 3.5" DSTI I

I SLI = SLI + (SL + SLP) *DT/2.01

I SLP DSTO==SLSLI + 3.5 * DSTI

I C OUTPUT DSTO (DIFFERENTIAL VARIABLE STABILIZER COMMAND . DEG)

•FIGURE 171

PRIMARY LATERAL-DIRECTIONAL AFCS DIGITAL PROGRAM

241
NOTE:

Place downstream of longitudinal primary AFCS program.

C

C

C

C

C

C

C

MLA SYSTEM INBOARD AILERON SYSTEM

INPUT DATA (GENERATED OR USEDPREVIOUSLY IN

LONGITUDINAL PRIMARY AFCS PROGRAM)

AZML (COLUMN COMMAND- INCREMENTAL G'S)

DP (DYNAMIC PRESSURE - PSF)

AZl (INCREMENTAL NORMALACCELERATION , G'S)

AC (STABILIZ ND - DEG)

T

I AZIA : AZML

AZIA = 1.05 * AZ1

ASL = 2425./DP

CALL LIMIT (ASL, 13.33, 0.0)

AZIA : AZIA * ASL

GMS : GMSP * YIO + AZIAP * WIO + AZIA * XIO

GMSP : GMS

AZIAP = AZIA

CALL LIMIT (GSM, 20.0, 0.0)

AC = AC + 0.Ii * GMS

242

c

C

C

OUTPUT VARIABLES

GMS (-INBOARD AILERON COMMAND - DEG)

AC (STABILIZER PROPORTIINAL COMMAND - DEG)

FIGURE 172

MLA INBOARD AILERON FUNCTION FLOW DIAGRAM
SUBROUTINE LIMIT (VAR, VUL, VLL) /

C PARAMETER IDENTIFICATION

C VAR - VAR-IABLE TO BE LIMITED

C VUL - VALUE OF UPPER LIMIT

C VLL - VALUE OF LOWER LIMIT

T

p

VAR = VUL I

RETURN I i

I _TU_I

T

VAR = VLL

RETURN

END

FIGURE 173

LIMITER FOR PRIMARY AFCS AND MLA

243
(This page intentionally left blank)

244
7. ELECTRICAL/ELECTRONIC DESIGN AND MODIFICATION

This section details design requirements and the methods used to comply

with these requirements for the DAST ARW-2 Active Control System (ACS)

electrical and electronic components.

The ACS electronics include motion feedback sensors, uplink and

downlink telemetry signal conditioning, mechanization of the flutter

suppression, gust load alleviation and relaxed static stability filters

and a function generator for inputing commands to the outboard aileron

servoactuators for flutter suppression system testing, a random noise

generator to apply commands to the inboard aileron servoactuator

feedback loops and servovalve drive amplifiers. Figure 174 shows the

"as-delivered" functional block diagram of the ACS, however,

modifications were made after delivery to NASA Langley. The changes

made to the ACS were a result of a "hardware-in-the-loop" (HITL)

simulation at NASA Langley which produced airplane responses different

from the responses obtained in the ACS simulation performed at Boeing.

The changes are covered in detail in Paragraph 7.5 of this document.

Excitation voltages required by the ACS accelerometers and servoactuator

position and pressure feedback transducers were furnished by DC to DC

converters, Figure 175 defines symbols used on the block diagram shown

on Figure 174.

7.1 Design Philosophy. The basic philosophy was to provide as much

flexibility as possible to ease maintenance and to provide required

performance within the constraints of reasonable cost, minimum physical

size and the expected drone flight environment.

The number of circuit cards and components was kept to a minimum. The

components are military grade mounted on two-sided mica ply etched

circuit cards with plated through holes. Card edge connectors are not

used in order to decrease the possibility of receiving intermittent

signals during vibration.

Filter networks were isolated to allow critical break frequencies to be

Selected At Test (SAT). A built-in test generator with selectable mode

and signal amplitude telemetry triggering was provided. Parameter

changes are implemented by SAT component provisions for infrequent

changes, variable gain controls for rapid changes between flights and

inflight logic-controlled gain changes and parameter time constant

scheduling.

The sensors and output signals are analog and are compatible with the

existing telemetry system.

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FIGURE 174

DACE BLOCK DIAGRAM (UPLINK CONTROL SIGNALS)

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INPUT/OUTPUT MNEMONICS

AAC Antisymmetric Acceleration Command

AFII Antisymmetric Filter Test Input I

AFOI Antisymmetric Filter Test Output I

AFO2 Antisymmetric Filter Test Output Z

AF03 Antisymmetric Filter Test Output 3

ALT HOLD Altitude Hold Input

A_P Ambient Pressure Input

BS250 Body Station 250 Acceleration Input

BUSS Backup Symmetric Stabilizer Input

DBS250 Body Station 250 Acceleration Downlink

DPP Dynamic Pressure Oownlink

DLRSI RSS Integrator Output Downlink

DLTSS Total Symmetric Stabilizer Downlink

DPO Dynamic Pressure Output

DKGLA Scheduled GLA Gain Downlink

FSEXC Flutter Suppression Excitation Command

GAOl GLA Aileron Loop Test Output 1

GL01 GLA Test Output 1

GL02 GLA Test Output 2

GS01 GLA Stabilizer Loop Test Output 1

GS02 GLA Stabilizer Loop Test Output 2

IBAC Inboard Aileron CommandDownltnk

IMPP Impact Pressure Input

KBLA Scheduled GLA Gain

LFSXX Left Front Spar (WBLXX)

LRSXX Left Rear Spar (WBLXX)

LWFSAC Left Wing Flutter Suppression Actuator

Conanand

LWFSO Left Wtng Flutter Suppression Output

(Oo_vnlink)

MLCMD MLA Command Input

PPR Pseudo Pitch Rate

PPR01 Pseudo Pitch Rate Test Output !

RF01 Relaxed Static Stability Filter Test

Output

RFSXX Right Front Spar (WBLXX)

RRSXX Right Rear Spar (WBLXX)

RSEO Relaxed Static Stability Error

Output

RWFSAC Right Wing Flutter Suppression

Actuator Command

RWFSO Right Wing Flutter Suppression

Output (Downltnk)

SAC Symmetric Acceleration Command

SFII Symmetric Filter Test Input I

SF02 Symmetric Filter Test Output I

SF02 Symmetric Filter Test Output Z

SF03 Symmetric Filter Test Output 3

SPR Sensor Pitch Rate Input

SPR01 Sensor Pitch Rate Test Output I

INPUT/OUTPUT MNEMONICS

TALT HOLD Altitude Hold Test Input

TAMP

TBUSS

TIMP

TIBCO

TMLCMD

TBS250

TLFSXX

TLRSXX

TRFSXX

TRRSXX

TRSID

TPPR

TSPR

TSS

Ambient Pressure Test Input

Backup Symmetric Stabilizer Test Input

Impact Pressure Test Input

Inboard Aileron Command Test Output

MLA Command Test Input

Body Station 250 Acceleration Input

Test Input Left Front Spar (WBLXX)

Test Input Left Rear Spar (WBLXX)

Test Input Right Front Spar (WBLXX)

Test Input Right Rear Spar (WBLXX)

RSS Integrator Test Output

Pseudo Pitch Rate Test Input

Sensor Pitch Rate Test Input

Total Symmetric Stabilizer Output

INBOARD AILERON MNEMONICS

LIBIBS (L11) Left Inboard Aileron, Inboard Segment

LIBOBS (LIO) Left Inboard Aileron, Outboard Segment

RIBIBS (R11) Right Inboard Aileron, Inboard Segment

RIBOBS (RIO) Right Inboard Aileron, Outboard Segment

XXXDP Differential Pressure, Downlink

XXXDTI Differential Pressure, Downlink Test Input

XXXPOSI Position Test Input

XXXPOSO Position Test Output

XXXTIC Test Input Command

XXXVLO Valve, Low Side

OUTBOARD AILERON MNEMONICS

LOBS (LO) Left Outboard Segment

ROBS (RO) Right Outboard Segment

XXCII Port CI Test Input

XXCI2 Port C2 Test.lnput

XXDP Differential Pressure Downlink

XXDPR Differential Pressure Test Output

XXPOS Position Oownlink

XXPOSI Position Test Input

XXPOSO Position Test Output

XXTIC Test Input Command

XXVLO Valve Low Side

FIGURE 175

LIST OF SYMBOLS FOR THE ACS BLOCK DIAGRAM

260
7.2 ACS Electronics Box Design. The box is required to house the ACS

electronics and mate with an existing BQM-34E/F mounting tray. All

connectors are required to be mounted on the front of the box for

accessibility. Space is available beneath the mounting tray for

mounting of the ACS electronics power supplies.

The electronics box is a Boeing designed and Fabricated aluminum box

(Drawing Number 35-35100-1) that mates with an existing mounting tray in

the drone fuselage at Body Station 212.8. Electronic components used

are military qualified or commercial grade meeting military

environmental specifications. The electronics box is 17.75 inches long,

9.4 inches wide and 6.3 inches high with extending mounting flanges

making an overall length of 22.5 inches. The box weight is 24 pounds.

The circuit card sizes are 4.475 inches high and 9.05 inches wide, and

3.3 inches wide and 4.55 inches high. The box includes provisions for

18 circuit cards and all interface connectors are mounted on the Front.

Figures showing the box design size are presented in Paragraph 8.1. All

power is supplied by externally mounted DC to DC converters. EMI

filters are included on all power lines. All cable connectors include

EMI backshells to ensure electromagnetic compatibility. Boeing

Document D3-1_115-2, EMC Qualification Report for the Active Control

System DAST (ARW-2), 12 January 1983 verifies FSS electronics

compliance with MIL-STD-461A.

7.3 Flutter Suppression Systems Sensors. The DAST ARW-2 flutter

suppression system requires wing accelerometers as the aircraft motion

sensors for feedback signals. The gust load alleviation system requires

a fuselage accelerometer as the aircraft motion sensor for feedback

sensors and pressure transducers are required for actuator load

telemetry data. The following paragraphs outline performance and

installation requirements of these sensors as well as the design

philosophy and approach.

7.3.1 Sensor Performance Requirements. The ACS requires nine motion

sensors with type, orientation and location as defined on Figure 176.

The vertical accelerometers in the wings should have their sensitive

axes vertical when the wing is in lg flight at Mach 0.86, 15,000 feet.

The accelerometers shall have the following characteristics:

Amplitude Frequency

Sensors Range Range

Accuracy

(% of Applied Accel.)

1,2,3,4, +jOg 0.5 - 1000 Hz 3

5,6,7,8

9 _ 5g 0.5 - 200 Hz 4

261
SENSOR

i

1

4

6

7

"8

9

TYPE

Accelerometer

Accel erometer

Accelerometer

Accelerometer

Accel erometer

Accel erometer

Accelerometer

Accelerometer

Accelerometer

LOCATION

Left Wing

Left Wing

Left Wing

Left Wing

Right Wing

Right Wing

Right Wing

Right Wing

Fuselage

(Center

Line)

ORIENTATION

Vertical

(51Deg)

Vertical

(51 Deg)

Vertical

(51Deg)

Vertical

(±1Deg)

Vertical

(±1Deg)

vertical

(±1.Deg)

Vertical

(±I Deg)

Vertical

(±i Deg)

Vertical

(51Deg)

BS

Rear Spar

Rear Spar

Front Spar

Front Spar

Rear Spar

Rear Spar

Front Spar

Front Spar

250

WBL WL

84 Center of

Rear Spar

92 Center of

Rear Spar

82 Center of

Front Spar

92 Center of

Front Spar

84 Center of

Rear Spar

92 Center of

Rear Spar

82 Center of

Front Spar

92 Center of

Front Spar

FIGURE 176

MOTION SENSOR DEFINITION

262

9
The sensors are required to be flightworthy. All sensors are required

to meet these requirements while operating in the BQM-34E/F drone

environment defined by NASA - DFRC Process Specification 21-i and 21-2.

Figure 177 contains the accelerometers performance requirements.

PARAMETER

Range

Natural Frequency

Size

Damping Ratio

Resolution

Hys te res i s

Threshold

Cross-Axis Sensitivity

Temperature Range

l,i

REQUIREMENT

WING FUSELAGE

-+lOg

1000 Hz

0.3 xO.5 Inches

±5g

200 Hz

1.25 xl.75 Inches

0.6 Minimum

O.02g Minimum

O.O01g Minimum

O.O01g Minimum

O.05g/g Maximum

-40°F to 200°F

FIGURE 177

COMPOSITE ACCELEROMETER PERFORMANCEREQUIREMENTS

The potentiometer performance requirements are presented on Figure 178.

The potentiometers are mounted on the rotary servoactuator shafts for

inboard ailerons as shown on NASA Drawing LD-316091. The potentiometers

are mounted on a potentiometer mounting extension for outboard ailerons

as shown on Boeing Drawing 35-34617. Size and resolution were the

primary requirements.

Case Diameter

TYpe Resistive Element

Usable Angle of Rotation

Output Gradient

Center Tap Required

iii

O. 5 Inch

Conductive Plastic

-+20 Degrees Minimum

+i.0 Volt for 10 Degree

Rotations, Minimum

No

FIGURE 178

POTENTIOMETER PERFORMANCE REQUIREMENTS

263
The pressure transducer performance requirements are presented on Figure

179. Twelve transducers are required, two for each servoactuator to

measure C1 and C2 port pressures.

Pressure Range

Rated Exciation

Input Impedance

Sensitivity

Frequency Responses

0 to 1500 psi

10 VDC

350 OHM Nominal

40 MV ±5%

(Open Circuit at Rated

Excitation)

>1000 Hz

FIGURE 179

PRESSURE TRANSDUCER PERFORMANCE REQUIREMENTS

7.3.2 Sensor Selection. A sketch of the wing accelerometer selected is

shown on Figure 180. The manufacturer specifications are listed on

Figure 181. Performance requirements of the ARW-2 were met by the

sensors used on ARW-I.

More space was available for mounting transducers in the Fuselage,

allowing the use of a servo type accelerometer for measuring fuselage

acceleration. Manufacturer's specifications, availability and NASA's

recommendations from ARW-1 led to the selection of the Sundstrand Model

QA-1100-AA01-12 accelerometer. Figure 182 shows a sketch of the

accelerometer and the manufacturer's specifications are listed on Figure

183.

The potentiometers selected are the New England Instrument Company Model

55 FL1-134. These potentiometers are single turn 3000 precision

conductive plastic type presenting nearly infinite resolution. The 1⁄2

inch diameter was selected to be compatible with the space available.

The manufacturer specifications are listed on Figure 184.

The pressure transducers selected are the Bell & Howell Type 4-326-0001.

These transducers are a four-arm active strain gage Wheatstone design

with rugged construction offering highly reliable service. Figure 185

shows the outline dimensions for the Type 4-326-001 pressure transducer

and the manufacturer's specifications are listed on Figure 186.

7.4 Circuit Design. The electronics provide signal conditioning for

the analog signals from the sensors, accepts uplink discrete telemetry

commands and provides downlink analog and discrete signals to be

compatible with the existing telemetry system. The primary function of

the electronics is signal shaping. The filter network required

flexibility to allow critical break points to be selected by resistor

changes at designated places in the circuitry. A built-in sweep and

doublet generator is required to provide FSS test signals to assist in

264
t

9/32 In.

Hex

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FIGURE 18()

Diameter

0.05 In.

- SKETCH OF PCB PIEZOTRONICS MODEL 303A03 MINIATURE

ACCELEROMETER

265
Sensitivity, Nominal

Resolution

Resonant Frequency, MTD.

Frequency Range, _5%

Overload Recovery

Discharge Time Constant

Amplitude Linearity

Range for _5V

Output Impedance

Output Bias

Transverse Sensitivity

Strain Sensitivity

Temperature Sensitivity

Vibration Maximum

Shock, Maximum

Structure

Size, Hex x Height

Connections

Case Material

Power Supply Voltage

Power Supply Current

10 mV/g

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2 to 20mA thru Current Regulation Diode

FIGURE 181

PCB 303A03 ACCELEROMETER PEFORMANCE SPECIFICATION

266
T

1.50 In.

Max.

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t

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Hole

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FIGURE 182 SUNDSTRAND MODEL QA1100-AA01-12 ACCELEROMETER

267
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268
Electrical Specifications

Typical Absolute Linearity

Theoretical Electrical Angle

Resistance

Resistance Tolerance

Output Smoothness

Dielectric Withstanding Voltage

Insulation Resistance

Mechanical Specifications

Starting Torque, Single Cup

Mechanical Angle

Weight, Single Cup

Shaft Runout

Pilot Runout

Lateral Runout

Shaft Radial Play

Shaft End Play

_00 v

5000 Ohms

I

+10%

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0.2 oz-in Maximum

360 o Continuous Rotation

0.25 oz Maximum

.002 in/in of Shaft Length

.001 in T.I.R. Maximum

.001 in T.I.R. Maximum

.001 in T.I.R. Maximum

.005 in T.I.R. Maximum

Figure 184

NE1 55FL-134 POTENTIOMETER PERFORMANCE SPECIFICATIONS

269
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• Dia. Max. I_

17

2.126 In.L----_

<>WIRING DIAGRAM'

PIN FUNCTION 0.76 In / F 2"lOaxl'n'-_-_'O00_1/--_atlngc_nende_°r(SR A - Input Hex PCS)064-8-45 B - Output itting Per 4-326- "

C + Output MS-33656-4

D .+ Input

FIGURE 185- - BELL & HOWELL TYPE 4-326-0001 PRESSURE TRANSDUCER OUTLINE

DIMENSIONS

270
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271
determining proper FSSoperation during flight. A randomnoise

generator has been provided to simulate vertical gusts and eliminate the

need to search for actual gust inputs. Parameter changes are to be

implementedby designated resistor changes for infrequent changes,

variable gain controls for rapid changesbetweenflights and in-flight

logic-controlled gain changes and parameter gain scheduling.

The ACSelectronics are designed with all circuitry on cards. Whenthe

cards are removedfrom the box, no electronics are left in the box

except for EMIfilters and wire harness.

7.4.1 Signal Conditioning. The uplink commandsand downlink signals to

be received and transmitted by telemetry are conditioned in the

electronics to be compatible with the drone primary telemetry system.

The uplink commandsare required to be compatible with the existing

telemetry system. The telemetry commandsare discrete 0 to 5 VDC

Schottky TTL open-collector outputs.

The uplink commandsconsist of discrete signals and are defined on

Figure 187. Figure 188 presents a typical uplink signal conditioning

circuit. The circuitry will accept 0 to 5 VDCSchottky TTL opencollector telemetry outputs. The pull-up resistors are provided in the

ACSbox as shownon Figure 188 to improve noise immunity.

Static and impact pressure signals are included in the electronics and

are required for gain scheduling. Access to these signals is through

the AFCSinterface cable.

The downlink signal requirement is to be compatible with an existing

telemetry system. The telemetry system accepts _5 VDCanalog signals

and 0 to 5 VDCdiscrete signals.

The downlink signals consist of 13 discrete signals defined on

Figure 187 and 25 analog signals defined on Figure 189. Figure 190

presents a typical downlink signal conditioning circuit. Additional

inverting and noninverting circuits are provided on the downlink

circuit cards for use as required. Provisions are provided to allow

adjustment of offset and gain for the downlink signals. Multiple

inputs are included for versatility.

A single card is used for signal conditioning outputs of the eight-wing

accelerometers, PCB303A03, and the single fuselage accelerometer,

Sundstrand QA1100-AA01-12. The card provides the constant current

sources required by the PCBaccelerometers and the acceleration mixing

circuit to provide the necessary vertical motion signals required by the

ACS. A two radian washout is provided in each of the wing circuits to

block the DCoffset of the PCBaccelerometer drive power electronics.

272
Launch

Start Doublet

Start Swee p

FSS Excitation Amplifier

FSS Excitation

Gust Excitation 1

Gust Excitation 2

GLA Gain 1

GLA Gain 2

RSS Gain 1

RSS Gain 2

FSS Gain 1

FSS Gain 2

On/Off

On/Off

On/Off

Hi/Low

Symmetric/Antisymmetric

On/Off

On/Off

On/Off

On/Off

On/Off

On/Off

On/Off

On/Off

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

5 VDC/O VDC

Figure 187

UPLINK TELEMETRY COMMANDS

273
TELEMETRY ,>

INPUT

+5 VDC_

TEST >

INPUT

4.75K

'4.75K

DISCRETE

_(DOWNLINK)

TO ACS LOGIC

FIGURE 188

UPLINK SIGNAL CONDITIONING CIRCUIT

274
Measurement

Left Outboard

Aileron Command

Right Outboard

Aileron Command

Inboard Aileron

Command

Left Outboard Aileron

Position

Right Outboard Aileron

Position

Left Inboard Aileron

Outboard Segment

Position

Right Inboard Aileron

Outboard Segment

Position

Right Inboard Aileron

Inboard Segment

Left Outboard

Differential Pressure

Right Outboard

Differential Pressure

FSS Excitation

Symmetric

Stabilizer SUM

Sy_metrlc

Stabilizer RSS

Symmetric

Stabilizer GLA

eac t - epseudo

_BS 250

Left Wing WBL 82

Front Spar

Left Wing WBL 84

Rear Spar

Right Wing WBL 82

Front Spar

Right Wing WBL 84

Rear Spar

Left Wing WBL 92

Front Spar

Left Wing WBL 92

Rear Spar

Right Wing WBL 92

Front Spar

Right Wing WBL g2

Rear Spar

Parameter

Identi- Signal

fication Range Level Factor

LWFSO ±15 Deg ±5 VDC 3.0 Oeg/V

RWFSO ±15 Deg ±5 VDC 3,0 Deg/V

IBAC +20 Deg ±5 VDC 4.0 Oeg/V

-20 Deg

LOPOS ±15 Deg ±5 VDC 3.0 Deg/V

ROPOS ±15 Deg ±5 VDC 3.0 Oeg/V

LIOPOS +20 Deg +5 VDC 4.0 Deg/V

-10 Deg -2.5 VDC

RIOPOS +20 Deg +5 VDC 4.0 Deg/V

-10 Deg -2.5 VDC

RIIPOS +20 Deg +5 VDC 4,0 Deg/V

-20 Deg -2.5 VDC

LODP ±1500 PSI ±5 VDC 300 PSI/V

RODP ±1500 PSI ±5 VDC 300 PSI/V

FSEXC ±2 Deg ±1 VDC 2 Deg/V

OLTSS ±13 Deg ±5 VDC 2.60eg/Y

DLRSI ±13 Deg ±5 VDC 2.6 Deg/V

DLGS ±13 De9 ±5 VDC 2.6 Oeg/V

TPHTS +60 Oeg/Sec +S VDC 12 Oeg/Sec/V

D6S250 ±2 g ±S VDC 0.4 g/V

DLFS82 ±20 g ±5 VDC 4 g/V

OLRS84 ±20 g ±S VDC 4 g/V

DRFS82 ±20 g ±S VDC 4 g/V

DRRS84 ±20 g ±5 VDC 4 g/V

OLFSg2 ±20 g ±5 VOC 4 g/V

OLRSg2 ±20 g ±5 VOC 4 g/V

DRFSg2 ±20 g ±5 VDC 4 g/V

ORRSg2 ±20 g ±S VDC 4 g/V

FIGURE 189

DOWNLINK TELEMETRY SIGNALS

275
+V

-V

R1

SIGNALINPUT_"_

R1

R2 R4

TO

TELEMETRY

GAIN=

R2R3 +R2R 4 +R3R 4

RIR3

FIGURE 190

DOWNLINK SIGNAL CONDITIONING CIRCUIT

276
7.4.2 FSS Signal Shaping. The requirements are summarized below. A

functional block diagram of the ARW-2 flutter suppression system is

presented on Figure 174. The tolerances on gain and phase as a function

of frequency are presented on Figure 191.

Provisions for three discrete inflight gain changes are required. The

gains required are nominal, half nominal and twice nominal.

The symmetric FSS sensor-to-surface sinusoidal steady-state phasing is

as follows:

WING MOTION (_up, _down)

Left Wing

Right Wing

Rear Spar 84_ g25

Front Spar 82f g2f

Rear Spar 84_ 92_

Front Spar 82_ 92f

Aileron Trailinq Edge

TED

TED

The antisymmetric FSS sensor-to-surface steady-state sinusoidal phasing

is as follows:

WING MOTION (fup, _down) Aileron Trailinq Edge

Left W_ng Rear Spar 84_ 92_

Front Spar 82_ 92f TED

Right Wing Rear Spar 84_ 92_

Front Spar 82f 92_ TEU

The flutter suppression system consists of wing accelerometers driving

the ailerons through appropriate symmetric and antisymmetric shaping

filters.

The primary function of the FSS electronics is signal shaping. Each

shaping filter input and output is brought out to the circuit card

connector to facilitate adding a filter if required anywhere in the

electronic circuitry. The filters are mechanized by a second order over

a second order with the general transfer function:

Eo (S) : K(S2 + 26_ nl S + nl 2)

Ei (S 2 + 262wn 2 S +_n22

volts/volt

The second order blocks are implemented in state variable form with

voltage out of operational amplifiers being proportional to first or

second derivatives of the output, or proportional to the output itself.

The state variable design approach facilitates relatively easy gain and

277
3.0

2.0

1.0

MAGNITUDE

(dB) 0

-I .0

-2.0

-3.0

10.0

5.0

PHASE 0

(DEG)

_\\\\\\\\\\\\\\\_\\_

_\\\\\\\\\_\\\\\\\\

__\\__\\\\\\\\\\

FIGURE 191

ALLOWABLE TOLERANCES FOR FLUTTER SUPPRESSION SYSTEM FILTERS

278
break point changes throughout the filters. Figure 192 presents a

typical filter circuit as implemented in the electronics. The resistor

values are calculated as shown on the figure.

The symmetric and antisymmetric filters are mechanized using the same

basic circuit described above and shown on Figure 192. The same circuit

boards are used for all filters. The transfer functions are given on

Figure 174.

The uplink and summing card provides final gains for the FSS system.

The inflight gain select and the FSS engage functions are mechanized on

this circuit card and receive commands from the uplink logic. The gains

on the symmetric and antisymmetric circuit card are adjustable and are

shown by the symmetric gains between connections 11J2-6/11J2-11 and

12J2-6/12J2-11 on Figure 174.

7.4.3 FSS Parameter Scheduling. Analyses stipulate that parameter

scheduling as a function of Mach number is not required within the

flight envelope due to thrust limit actions. Antisymmetric gain

scheduling is required to achieve FSS stability to 1.2V d, which lies

outside the flight envelope. The gain schedule requirea is shown on

Figure 193 with block diagram and proposed schematic of the scheduler

shown on Figures 194 and 195, respectively.

7.4.4 GLA Signal Shaping. The requirements are summarized below. A

functional block diagram of the ARW-2 gust load alleviation system is

presented on Figure 115. The tolerances on gain and phase as a function

of frequency are presented on Figure 196.

Provisions for three discrete inflight gain changes are required. The

gains required are nominal, half nominal and twice nominal.

The sensor-to-control steady-state phasing is as follows:

BS250 Aileron Trailing Edge Stabilizer Trailing Edge

UP TEU TED

The gust load alleviation system consists of a single servo-type

accelerometer located at fuselage BS 250 (wing center section) driving

the outboard ailerons and stabilizer through appropriate shaping

filters.

The shaping filters are mechanized using the state variable filters

identical in form to those used in the flutter suppression system.

transfer functions are given on Figure 174.

The

The uplink and summing card provides final gains for the GLA system.

The inflight gain select and GLA engage functions are mechanized on this

circuit card and receive their command from the uplink logic.

279
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281
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ACCELERATION

CARD

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CIRCUIT

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SCALING FAND OFFSET

X

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FIGURE 194

BLOCK DIAGRAM OF ANTISYMMETRIC GAIN SCHEDULER

282
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283
MAGNITUDE

(dB)

PHASE

(BEG)

3.0

2.0

1.0

0

-I .0

-2.0

-3.0

10.0

5.0

-5.0

_\\\\\\\\\\\\\\\\\\\"

-I0"_.0 I0.0

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FREQUENCY (HZ)

FIGURE 196

ALLOWABLE TOLERANCES FOR GUST LOAD ALLEVIATION SYSTEM FILTERS

284
7.4.5 GLAGain Scheduling. The GLAgain scheduler is implementedon

the scheduler circuit card. The accelerometer signal is multiplied by

the gain schedule derived from the dynamic pressure computing circuit

located on the scheduler card. After being scheduled, the signal is

split for individual processing in the aileron and stabilizer filters.

The maximumon-card scheduled gain is 10.0 as comparedto the analytical

gain of 13.3. The additional 1.33 factor is picked up downstreamof the

scheduler.

7.4.6 RSSSignal Shaping. The requirements are summarizedbelow. A

functional block diagram of the ARW-2relaxed static stability system is

presented on Figure 135. The gain and phase tolerances as a function of

frequency are presented on Figure 197.

Provision for three discrete inflight gain changes is required. The

gains required are nominal, half nominal and twice nominal.

The sensor-to-control surface steady-state phasing is as follows:

Pseudo Sensor Stabilizer

Pitch Rate Pitch Rate Trailing Edge

+(nose up) -(nose down) TEU

The relaxed static stability system consists of a pseudo pitch rate

command and sensor pitch rate feedback signals driving the stabilizer

through appropriate shaping filters.

The shaping filters are mechanized using the state variable filters

identical in form to those used in the flutter suppression system. The

transfer functions are given on Figure 174. The uplink and summing card

provides final gains for the RSS system. The inflight gain select and

RSS engage functions are mechanized on this circuit card and receive

their commands from the uplink logic.

7.4.7 RSS Gain Scheduling. The RSS gain scheduler is implemented on

the Scheduler Circuit Card. The pitch rate error signal is divided by

the dynamic pressure signal (derived on the scheduler card). After

being scheduled, the signal is passed to the stabilizer filter.

7.4.8 Function Generator. The requirements are summarized below. An

aileron square wave doublet and sine wave sweep excitation is required

in the onboard ACS electronics box. The design requirements are as

follows:

• Sinusoidal Sweep

- Frequency Range

- Duration

10 to 40 Hz

10 Seconds

285
3.0

2.0

1,0

MAGNITUDE

(dB) 0

-I .0

-2.0

-3.0

10.0

5.0"

PHASE 0

(BEG)

-5.0

-10.0

_\_\\\\_\\\\\\\\\\\\\_

_\\\\\\\\\\\\\\\\\\\\\\_\_\

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0.1 1.0

FREQUENCY (HZ)

I0.0

FIGURE 197

ALLOWABLE TOLERANCES FOR RELAXED STATIC STABILITY SYSTEM FILTERS

286
- Type of Sweep

- Amplitudes

Antilog

2 and 4 Degrees Aileron

• Square Wave Doublet

- Frequency

- Duration

- Amplitudes

25 Hz

I Cycle

2 and 4 Degrees Aileron

Provisions are required to allow the sweep and doublet amplitudes to be

changed by circuit modification. One uplink command is used to select a

"high" or "low" gain for both sweep and doublet.

The excitation generator is required to have capability to select

symmetric or antisymmetric inputs to the control surfaces. Automatic

reset capability is required when the excitation is disengaged.

The function generator is included in the ACS electronics to provide

inputs to the aileron for testing the flutter suppression system. The

functional block diagram shown on Figure 198 indicates the theory of

operation of the function generator.

The Voltage Controlled Oscillator (VCO) is free running at the start

frequency initially set. The output of the multiplier is approximately

zero and the analog switch output is zero. The "start logic" will

enable the log sweep generator. At the same time, the analog switch is

opened and the integrator begins to ramp up and limits at 10 VDC in 0.15

seconds. The sweep function causes the VCO frequency to increase until

the stop frequency is reached.• At that time, the reset logic is enabled

and sweep logic is inhibited causing the integrator to ramp down and the

sweep is reset to zero and the analog switch is closed. The frequency

is swept according to:

afo

f- - z_

a - t

a = fiT

fl - fo

•6 = (}o sin

where:

fo = start frequency

fl = end frequency

287
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LJ..CO Q

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288 ;"
f : output frequency

T = sweep time

6 = control surface angle

6 o = maximum control surface angle

= phase angle

The sweep enable signal must be present to allow a complete sweep. Only

one sweep can occur for each time the sweep enable signal is applied.

The sweep will immediately stop after the amplitude ramps down when the

sweep enable signal is removed.

A square wave doublet is generated when the doublet start signal is

applied. The period of the doublet is determined by the clock. Only

one doublet can occur for each time the doublet start signal is applied.

The doublet cannot be terminated before completion once it has been

initiated.

An amplitude select signal will cause the amplitude of either the sine

wave or the doublet to be increased by a factor of two. Both the sine

wave and doub|et can be initiated at the same time. A logical "one" (+5

VDC) enables all function generator commands. The sine wave sweep star_

and stop frequencies may be varied +5.0 Hz. The sweep duration is

variable from five to 20 seconds. The soft start and stop provided by

the nominal integration ramp time of 0.5 seconds, which can be adjusted

over the range 0.7 seconds up and 0.3 seconds down to 0,3 seconds up and

0.7 seconds down. The square doublet is not variable and can be changed

only by circuit modification.

The function generator provides a method to input commands to the FSS

during flight testing. The function generator when used in conjunction

with the FSS gain commands can provide open loop (gain off) or closed

loop (gain on) modes of operation.

7.4.9 Pseudo Random Noise Generator. The random noise generator is

implemented to provide a method of flight test verification of the gust

load alleviation system. The random noise generator is shown on Figure

124. The output of the noise generator is passed through a shaping

filter and applied to the inboard aileron servoactuator. The motion of

the inboard aileron surfaces produces wing root bending moment power

spectral densities similar to those produced by actual vertical gusts.

The power spectral densities are compared in Paragraph 6.4.4.

The random noise generator is mechanized with two finite length shift

registers whose output sequences are combined to produce a sequence of

length equal to the product of the length of each individual shift

289
register. The randomnoise generator is capable of producing 174

minutes of uncorrelated noise with the existing 100 hertz clock.

To begin the sequence, each shift register is loaded with a non-zero

initial value. The initial value is loaded whenthe doublet commandis

depressed. The initial value is constant, therefore the sequencealways

begins at the samepoint each time the shift register is loaded.

The digital sequence is passed through a shaping filter mechanizedas a

state-variable filter. The random-noise then passes through a variable

gain amplifier to scale the magnitude of the noise which is then applied

to the inboard aileron input.

7.4.10 Servo Electronics. The requirements are summarizedbelow. The

final gains were set experimentally. The servoactuator loops are

essentially the samefor all surfaces. Variance in loop gains and notch

filter location is the only difference betweenthe servoactuators.

Six wing control surfaces are poweredby hydraulic servoactuators. Each

of the servoactuators utilize position feedback. The position feedback

signal comesfrom a potentiometer mountedto measureactuator shaft

angular position. The position signal is fed back through gain only.

A load pressure signal is formed from the outputs of the strain gage

bridge and pressure transducers plumbed into the hydraulic lines between

the servovalve control ports and actuator. The load pressure sigoal is

sent to the drone telemetry system for inflight monitoring.

Notch filters were addedto reduce the effects of actuator and fluid

modeson the servoactuator stability. The outboard valve driver

required three notches located at 402 Hz, 90 Hz and 72 Hz. The outboard

segmentof the inboard aileron required two notches located at 285 Hz

and 72 Hz. The inboard segmentrequired a single notch at 72 Hz.

The servoamplifier is required to accept two input commands,one from

the Drone Active Control Electronics (DACE)and one from the ground

support equipment. The outboard aileron commandfrom the DACEincludes

FSSsymmetric and antisymmetric filter outputs, closed loop GLA/MLA

filter output and function generator output.

The inboard aileron commandfrom the DACEincludes the open-loop MLA

output and the randomnoise generator output. The feedback signals to

the servoamplifier are provided by the surface position potentiometers.

Provision for surface position nulling, feedback gain adjustment and

forward path gain adjustment are included.

Pressure feedback is not required for servoactuator stabilization.

Pressure feedback is provided for control surface load monitoring via

the drone telemetry system. The pressure feedback signal processing

electronics is located on the valve driver circuit card.

290
7.5 ACSModifications

7.5.1 Pre-Delivery Modifications. Washoutsare added in front of and

behind the FSSshaping filters eliminating the propagation of DCoffsets

through the shaping filters and servoactuators.

Addition of the altitude hold input to the RSSintegrator to the Primary

Control System (PCS)was madeafter the circuit boards had been

populated.

The 0.7 gain increase in the RSSsystem whenthe BackupControl System

(BCS) is engagedwas madeafter the circuit boards had been populated.

7.5.2 Post-Delivery Modifications. During hardware-in-the-loop

simulation at NASA-Langley,a numberof changeswere made. These

changes are not reflected in the data package sent with the DACE.

Changeswere madein RSSand GLAstabilizer feedback loops. The

changesmadewere as follows:.

a.

b.

GLA stabilizer gain was incorrect. Investigation yielded two

items:

(1) GLA Filter #3 had a gain 2.8 V/V more than that indicated by

the block diagram in the operations manual. R101 on the 39-

28047 circuit card was changed from 280K to lOOK to reduce the

excess gain. This change brought the implementation in line

with the block diagram of the operations manual.

(2) The block diagram of the operations manual showed the

stabilizer deflection to a lg input to be 0.8740 . The

deflection to a lg input called out by the analysis should be

0.7980 . R46 on the 39-28031 circuit card was changed from

1.21 Meg to 1.10 Meg to reduce the excess gain. This change

brought the implementation in line with the analysis block

diagram of the Preliminary Design Review (PDR).

RSS Gain Scheduler was found to limit at the MLA condition (0.42M

at 10,000 feet, approximately 127 psf).

(1) R53 on the 39-28043 scheduler circuit card was changed from

20K to 4.99K. This change reduces the signal amplitude going

into the divider by a factor of 4.

C2) R95 on the 39-28051RSS filter circuit card was changed from

115K to 459K. This change increased the filter gain by 4,

which results in no net gain change in the RSS path.

29l
7.6 Active Control System Power. The power required for the ACS is

obtained from Crestronics, Inc., DC to DC converters utilizing the 28

VDC power onboard the drone. The power levels required and converter

part numbers are:

Voltage Current Part Number

+15 VDC 600 mA

¥ 5 VDC 800 mA

¥24 VDC 130 mA

PS 333-24-30-BCT

PS 333-24-10-BCT

PS 333-24-26-FW

The external DC to DC converters were selected to furnish the excitation

power to the ACS electronics to eliminate as much wasted power within

the electronics box as possible. The manufacturer's specifications are

as listed on Figure 199.

The converters are rated for one ampere continuous and the currents

listed above are the actual requirements. The three converters are

mounted side by side on to 6.0 x 10.0 x 1.0 inch aluminum extrusion.

Dow Corning PC 340 white silicone grease was used for mounting to reduce

the temperature coefficient. The assembled unit has a weight of 5

pounds and is to be mounted beneath the ACS electronics box. A 6.7

microfarad 35 VDC capacitor was installed at the 28 VDC input and one at

each power output.

7.7 Active Control System Electrical Wiring. All drone wiringwas done

by NASA with Boeing furnishing the wing wire. Figure 200 shows the DAST

ARW-2 ACS wiring. All wire furnished for the wiring bundles except the

power cable which is described below was either two, three, or four wire

shielded 22 AWG M27500 MIL specification wire, except the servo valve

wire which was unshielded.

Six primary cable harnesses are required for interfacing the ACS

electronics with the drone. One cable interfaces to the DC to DC

converter power unit, one cable will interface with the drone AFCS box,

one connects to the uplink receiver, one interfaces with the downlink

transmitter, two connects the ACS electronics box to each wing semispan

cable harness through the pressure bulkhead. Provisions are also

included for a ground test cable to permit day-to-day preflight checkout

of the ACS.

Figures 201 through 207 list the connector pin assignments. The power

cable power wire size is 18 AWG and includes one each for each voltage

level. The power returns are wire size 16 AWG and includes one each for

the +26 VDC and five for the +5 VDC and +15 VDC. This power cable

attaches to the terminal strip provided on the converter power unit.

The 28 VDC main ships power also attaches to the designated terminals on

this terminal strip.

292
Model Number: PS333

CaseUnits:

IIN MAX:

CaseDepth:

Mounting Dimension:

1.5 x 3.0 inch Extruded Aluminum

2.00 Amps

3.80 inch

2.50 x 1.20 inch

Rectifier System

FW - Full Wave

BCT- Bridge Center Tap

Regulated Units

Input

EIN MAX - Voltage: Customer may specify any input voltage

between 11 and 28 volts DC. EIN MIN is

the lowest input voltage applied. This

should include the lowest excursion of

ripple.

AEIN - Input Voltage Differential: The normal AEIN is

4 volts. For AEIN greater than 4 volts

derate the input current to

(6)IIN MAX

I Derate =

AEIN + 2

IIN MAX - Input Current: 2.0"amps

IIN MAX shall be derated for AEIN greater than 4 volts.

IIN MAX shall be derated for 70uc operation

(Derate input IIN MAX 1%/°C)

Output

Eo - Voltage: The output voltage is screwdriver adjustable

from -20% to +5% of nominal.

Po - Power: The output power is expressed as:

(O.7EIN MAXIIN ) - I0

Regulation

Line:

Temperature:

Load:

Effi ciency:

O.01%/Ainput volt

(0.02%/oc) -2mV/OC

100% to 10%

For 30 VDC and over: 3% Regulation

For 29 VDC and under:

% Regulation = (O.02E o+O.3)/(E o xlO -2)

(O.75EINIIN) -I o

EINIIN x10 -2

FIGURE 199

CRESTRONICS MODEL PS 333 DC TO DC CONVERTS PERFORMANCE SPECIFICATIONS

293
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7.8 Special Test Equipment.

7.8.1 ACS Tester. The ACS tester is used for conducting quick look

functional tests during preflight and troubleshooting to the card level.

Figure 208 shows a front view of this tester.

When a malfunction in the ACS electronics is identified the ACS tester

will be used to isolate the problem down to the card level. This tester

can be used with the electronics installed in the drone as well as for

bench tests. The ACS tester was also used during initial checkout and

later to check out the electronics at NASA Langley. ACS tester includes

all uplink commands, provisions for externally monitoring of downlink

signals, filter and function generator inputs and wing control surface

positions. Also provisions for external inputs to the system shaping

filters and servo valve drivers are included for open and closed loop

preflight testing.

7.8.2 Electronics Card Tester. The electronics card tester will be

used to isolate faults to individual components on the circuit cards.

This tester was required during electronics checkout and will be

required on site during the integration of the ACS into the drone

vehicle and during electronics checkout and will be required on site

during the integration of the ACS into the drone vehicle and during the

flight tests. The electronics card tester includes all necessary test

points and input capability to test all card types. External power must

be provided. Figure 209 shows the front view of this tester.

7.8.3 Additional Ground Support Equipment. In addition to the two

testers described above, other ground support equipment will be

required. This miscellaneous equipment is tabulated:

• 28 VDC Ground Power

• Hydraulic System Service Cart

• Str_p Chart Recorder

• Transfer Function Analyzer

• Function Generator

• Variable _10 VDC Input

• Digital Voltmeter

• Oscilloscope

309
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(This page intentionally left blank)

312
8. COMPONENTFABRICATION

All components required for installation of the flutter suppression

system in the drone test vehicle except for the hydraulic power supply,

actuators and filter were either fabricated or procured and shipped to

NASA. The active control system electronics were fabricated by Boeing

engineering staff laboratory personnel following drawings prepared by

electronic design personnel.

Components purchased for installation in the drone were selected to

operate within an aircraft environment and, in general, are of the

quality normally utilized in military equipment. Some components were

obtained in plant from Boeing stocks to minimize expenditures for

components that had to be procured.

The following paragraphs discuss fabrication and assembly of the

electronic components required for the DAST ARW-2 active control system

installation.

8.1 Electronic Components. The electronics box is a Boeing designed

and fabricated aluminum box 18.1 inches long, 4.85 inches wide and 4.58

inches high with extending mounting flanges making an overall length of

22.5 inches and an overall width of 5.4 inches. The box weighs 13.75

pounds. The box internal wiring is type AWG-22. Figures 210 through

213 present photographs of the fabricated electronics box. The box

includes provisions for 18 circuit cards. Figure 213 shows 18 cards

installed. Assembly drawings were included with the hardwar_

delivery.

8.2 Spares. Spare components provided to NASA as backup to the DAST

ARW-2 active control system components are shown in Figure 214. The

spare components listed were specified by NASA.

PI_ECEDH'KI PAGE BL_ I_,OTFUU_

313
31 4 

FIGURE 210 

AFCS ELECTRONICS BOX, FRONT VIEW 

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Component

Accelerometer

Accelerometer

Power Supplies

Manufacturer

Sundstrand Data Control, Inc.

PCB Piezotronics, Inc.

Crestronics

Part Number

QA-1100-AA01-03

Model 303A03

PS333-24-26-FW

PS333-24-30-BCT

PS333-24-10-BCT

Various

Spares

1

3

I

1

1

1 of each

card type

FIGURE 214

DAST ARW-2 FLUTTER SUPPRESSION SYSTEM SPARE COMPONENTS

318
9. FLIGHTWORTHINESS TESTING

The active control system electronics box was tested to demonstrate

flightworthiness prior to delivery of the components to NASA. The tests

conducted included component functional tests, interface tests of the

electronics and servoactuators, vibration tests, temperature/altitude

tests and electromagnetic compatibility tests.

Component functional tests are discussed in Paragraph 9.1. Flight

assurance tests, including initial performance, integration, vibration,

temperature/altitude, electromagnetic capability and final performance

tests are summarized in Paragraph 9.2.

The temperature/altitude and vibration tests were conducted per NASADFRC Process Specification No. 21-2. Electromagnetic compatibility

tests were conducted per required portions of MIL-STD-461A as discussed

in Paragraph 9.2.5.

9.1 Component Functional Tests. Functional tests of the active control

system electronics and aileron actuation system components were

conducted to assure satisfactory performance. The servoactuator

functional test results showed two hydraulic fluid modes present in the

system and that changes to the servoactuator compensation and/or the

electronics would be required to give satisfactory performance. All

populated cards in the electronics, including the spare cards, were

tested for proper operation, the filters were tuned and required gains

established.

9.1.1 Servoactuator System. Functional tests were conducted on the

DAST ARW-2 aileron servoactuators before the interface of the

servoactuators with the active control system electronics was

accomplished. The functional test was conducted to determine the

dynamic performance that could be attained with the hardware as a

verification of the linear stability analysis discussed in Paragraph

6.2. Results of the testing accomplished showed good agreement with

the linear analysis having included the two hydraulic fluid modes in the

analysis. The fluid modes are caused by the separation between the

servovalve and actuator which would not be necessary on a full scale

aircraft because sufficient space would exist to install the servovalve

on the control surface actuator.

9.1.1.1 Test Setup. Three complete servoactuators were set up for the

functional tests using a steel bar (dummy load) rotary inertia, as

described in the linear analysis of Paragraph 6.2, and ACS-type valve

driver circuit card to close the feedback loops. An LM148 operational

amplifier with current feedback was used as the servovalve drive

amplifier. Fabrication of the DAST ARW-2 aileron control surfaces and

active control system electronics was not finished when the

servoactuator functional tests began.

319
Geometry of the hydraulic lines between the servovalve and actuator was

set up similar to the installation required in the ARW-2 wing X except

the line to the outboard actuator was given an additional 180v bend to

minimize the size of the layout. Each servoactuator was rotated 900 to

prevent interference between the dummy load inertia and the hydraulic

tubing. The hydraulic lines were set up over a full-size installation

drawing, hence the breadboard length did not differ appreciably from the

wing installation. The pressure transducers were plumbed into the

servovalve control ports with line length between the transducers and

tee fittings in the manifold block about the same as was estimated for

the installation in the drone wing.

Hydraulic power was provided by a laboratory hydraulic power unit with

about 5 gal/min flow capability at 1500 psi. _ 5 micron nominal, 15

micron absolute filter and an MS 28797-1 25 in accumulator were plumbed

into the pressure line between the pump and servovalve pressure ports.

Pressure could be adjusted manually up to the desired 1500 psi supply

pressure.

9.1.2 Electronic Components. All electronic components used were

military qualified or commercial grade meeting military environmental

specifications. Initially, all populated cards were tested for proper

operation. The filters were tuned and required gains established. The

box wiring was verified and all dimensions checked. The cards were

installed function by function to verify operation. A complete

operational check was then performed on the electronic box.

9.2 Flight Assurance Test Results. The Boeing test facilities were

rescheduled twice because DACE functional checkout required more time

than anticipated. A long term conflict of facilities would have

occurred. DACE gain and phase tuning was halted until after

temperature-altitude, EMC and vibration testing had occurred to avoid

the conflict. Prior to final performance testing, DACE gain and phase

tuning was completed. Direct comparison of final and initial test

results is therefore extremely difficult. Test result comparison

between similar tests (e.g. GLA aileron response, etc,) is valid and

shows good agreement. The DACE was inspected for workmanship and then

verification was made that the unit envelope would meet the required

interface dimensions to the BQM-34E/F drone aircraft. The testing

consisted of initial performance, temperature/altitude, vibration,

interface and final performance. The results of these tests are

presented by the following paragraphs. Data sheets may be found in

Boeing Document D3-12115-1, Flight Assurance Report for the Active

Control System DAST (ARW-2), 19 August 1983.

9.2.1 Initial Performance. The DACE only was tested during the initial

performance test. Compatibility between the DACE and the hydraulic

system was verified during the interface test.

320
9.2.1.1 Power Supply Performance. The input power was varied from 24.0

VDC to 32.0 VDC while providing the total load current required by the

DACE. The operating currents for each power supply voltage level were

measured. The results are recorded on Data Sheet 4.1.2.1. The power

supplies successfully passed all tests.

9.2.1.2 FSS. The procedures of Test 4.1.3.1 were used to check the

FSS. Filter gains and phases together with uplink and downlink

functions associated with the FSS were within required tolerances and

are recorded on Data Sheet 4.1.3.1.

9.2.1.3 Function Generator. The procedures of Test 4.1.3.2 were used

to verify operation of the function generator. Results are recorded on

Data Sheet 4.1.3.2. The function generator successfully passed all

requirements.

9.2.1.4 MLA. The procedures of Test 4.1.3.3 were used to determine

proper operation of the MLA. Results are recorded on Data Sheet

4.1.3.3. The MLA successfully passed all requirements.

9.2.1.5 GLA. The procedures of Test 4.1.3.4 were used to determine

proper operation of the GLA. Included in these procedures were checkout

of the GLA scheduler, filter gains and phases together with uplink and

downlink functions. Results are recorded on Data Sheet 4.1.3.4. The

GLA successfully passed all requirements.

9.2.1.6 RSS. The procedures of Test 4.1.3.5 were used to verify the

RSS. The RSS scheduler, filters, uplink and downlink functions were

checked out. Results are recorded on Data Sheet 4.1.3.5. The RSS

successfully passed all requirements.

9.2.1.7 Valve Drivers. The procedures of Tests 4.1.3.6 were used to

check out the left and right outboard valve drivers. Data are recorded

on Data Sheets 4.1.3.6.

The procedures of Tests 4.1.3.7 were used to check out the left and

right inboard and outboard segments of the inboard valve drivers. Data

are recorded on Data Sheets 4.1.3.7. All valve drivers successfully

passed all requirements.

9.2.1.8 Uplink Confirmation. The procedures of Tests 4.1.3.8 were used

to check out the uplink confirmation (downlink)signals. All signals

were within the required tolerances and are recorded on Data Sheet

4_1.3.8.

9.2.2 Electromagnetic Compatibility. The EMC tests were satisfactorily

completed as discussed in Boeing DocJment D3-12115-2. A copy of the PassFai'lSummary is given on Figure 215.

321
TEST

CE03 BB

CE03 NB

CE04 BB

CE04 NB

CS01

C002

CS06

RE02 BB

RE02 NB

RS03

MIL-STD-461A TEST

REQUIREMENTS

By Similarity

By Similarity

PASS

PASS

PASS

PASS

PASS

PASS

PASS

PASS

PASS

FIGURE 215

EMC PASS-FAIL SUMMARY

322
9.2.3 Temperature/Altitude. This test was performed per the requirements

and procedures of Boeing Document D3-12114-I, Flight Assurance Test

Procedures for the Active Control System DAST (ARW-2), 7 May 1981.

Figure 4.5.2 of Boeing Document D3-12114-I presents the test plan that

was followed during testing. The test was conducted in the BMAC

Temperature/Altitude Facility on 9-13 July 1982. Figure 4.5.2 of

Boeing Document D3-12114-I shows the test instrumentation and Figure II

of Boeing Document D3-12115-I shows the DACE and power supplies installed

in the test chamber.

9.2.3.1 Pretest Performance. The DACE was operated in standard ambient

conditions. Performance data were recorded after temperature

stabilization had occurred. The system met all requirements and the

data are recorded on Data Sheet 4.5.3.1.

9.2.3.2 Startup and Operation at -65°F. After completion of the cold

soak at -65°F, the system was turned on. The _15 VDC power supply was

not immediately functional. Turn on voltage was +4.7 VDC. When the

supply temperature reached -46°F, the output vol_ge suddenly came into

tolerance. The dynamic pressure circuit was out of tolerance by 0.13

VDC (13 psf, equivalent), however, this was not deemed consequential

since the GLA and RSS responses, whose schedules are a function of

dynamic pressure, were well within the gain (_+0.5 dB) and phase (_5.0

deg) to tolerances set. Performance results are recorded on Data Sheets

4.5.3.2.

9.2.3.3 Operation at -65°F, 60,000 Feet Altitude. After temperature

stabilization, the DACE was turned on and the chamber altitude adjusted

to 60,000 feet. Again, the dynamic pressure circuit was out of

tolerance. No deviations were noted due to changes in altitude. The

performance results are recorded on Data Sheets 4.5.3.3.

9.2.3.4 Performance at -40°F, 40,000 Feet Altitude. The chamber

temperature was increased to -40°F and the altitude adjusted to 40,000

feet. The DACE successfully passed all requirements except the dynamic

pressure circuit which has been previously discussed. The results are

recorded on Data Sheets 4.5.3.4.

9.2.3.5 Operation at Standard Conditions. The chamber temperature and

altitude were returned to laboratory ambient levels and the performance

tests were repeated. The DACE successfully passed all requirements,

except: (1) the sine wave sweep generator did not sweep in frequency and

(2) the FSS symmetric channel phase exhibited an 8 degree phase lag.

Inspection revealed a large amount of condensation on the circuit cards.

When the condensation had been removed from the foil side of the

function generator circuit card, the sine wave sweep resumed normal

operation. When the condensation had been removed from the foil side of

the symmetric filter card, the 8 degree phase lag was eliminated. Use

of a heat blanket or conformal coating is recommended to prevent the

accumulation of condensation on the circuit cards. The results are

recorded on Data Sheets 4.5.3.5.

323
9.2.3.6 Performance at 160°F, 30,000 Feet. Following a 16 hour

temperature soak at 160°F with power off, the altitude was adjusted to

30,000 feet. Upon reaching the specified altitude, the DACE was soaked

for another 4 hours with power on. Following the 4 hour, power on, hot

soak a performance test was conducted. The DACE successfully passed all

tests, except the GLA scheduler was found to be out of tolerance by 0.24

VDC. The effect on GLA system performance was negligible as all

responses were well within the tolerances set, +0.5 dB for gain and _5.0

deg for phase. Test results are recorded on DaTa Sheets 4.5.3.6.

9.2.3.7 Operation at 160°F, 60,000 Feet. The altitude was adjusted to

60,000 feet. The DACE successfully passed all tests, except for the GLA

scheduler which was out of tolerance but had negligible effect on the

GLA system performance. Test results are recorded on Data Sheets

4.5.3.7.

9.2.3.8 Operation at Standard Conditions. The chamber was returned to

laboratory ambient conditions and the performance tests repeated. All

functions operated properly and values returned close to those of the

pretest. Inspection showed no physical damage had occurred. Operation

was satisfactory. Results are contained on Data Sheets 4.5.3.8.

9.2.4 Vibration Tests. This test was performed per the requirements

and procedures of Boeing Document D3-12114-1. Figures are to be

found in Boeing Document D3-12115-1. Figure 12 presents the

requirement and Figure 13 gives the test input equivalent of this

requirement. The test was conducted in the BMAC Vibration Facility

located in Building 342J on 16-19 July 1982.

Three axes sinusoidal excitation was applied to the DACE at laboratory

ambient temperatures. The DACE was powered and operating during all

testing. No gain or phase distortion was observed during the sweep or

dwell periods. No signal distortion was observed on the oscilloscope

during these periods.

Figures 14, 15, and 16 show the locations of accelerometer placement on

the DACE.

Figures 17, 18, 19, and 20 show the response of the box to the vibration

inputs. Test data is recorded on Data Sheets 4.3.5.

9.2.5 Interface Test. The purpose of the interface test was to verify

electrical compatibility between the DACE and the hydraulic breadboard.

Additional data on the actuators were taken and are presented on Figure

216.

9.2.5.1 Interface Performance Tests. These tests were performed to

verify operation of the DACE after being connected to the hydraulic

breadboard. The results of the tests showed agreement with the data

taken during the initial performance test when the breadboard was

324
ACTUATOR

OUTBOARD AILERON

INBOARD AILERON

OUTBOARD SEGMENT

INBOARD AILERON

INBOARD SEGMENT

BANDWIDTH

(HZ)

49,2

22.0

25.5

HYSTERESIS

(DEGREES)

0.3

0.I

0.0

MAXIMUM

RATE

(DEG/SEC)

955

129

129

FIGURE 216

SERVOACTUATOR PERFORMANCESUMMARY

325
disconnected from the DACE. Oneexception is notable but not

significant. A deviation between initial and interface performance

tests exists in the MLAresponse. In the initial performance test, the

filter implementedwas (20)/(S+20). This filter was determined to be in

error and was subsequently corrected to (125)/(S+125).

9.2.5.2 Actuator Performance Tests. These tests were performed to

verify operation of the hydraulic actuators. Figure 216 summarizesthe

results of these tests. During the testing, no crossta]k was observed

nor were any hydraulic leaks observed.

9.2.5.3 Interface SystemPerformanceTests. Thesetests were performed

to obtain end-to-end data on the systems which utilized the actuators on

the hydraulic breadboard. The data were expected to be the composite of

the appropriate sections in paragraph 4.2.2 and the frequency responses

in Paragraph 4.2.3.1 of Boeing DocumentD3-12114-I. The data taken met

the expected results.

9.2.6 Final Performance. The DACEonly was tested during the final

performance test. Compatibility betweenthe DACEand the hydraulic

system was verified during the interface test. Data Sheets are to be

found in Boeing DocumentD3-12115-1.

9.2.6.1 Test Setup. The test setup was performed per Paragraph 4.6.1.

4.2.6.2 PowerSupply Voltage Check. The procedures of Test 4.6.2 were

used to verify voltage levels of the power supply. The results are

recorded on Data Sheet 4.6. The power supplies successfully passed all

tests.

9.2.6.3 FSS. The procedures of Tests 4.6.3.1 and 4.6..3.2 were used to

check the FSS. The final filter gains and phases were within the

required tolerances and are recorded on Data Sheet 4.6.

9.2.6.4 GLA. The procedures of Tests 4.6.3.5 through 4.6.3.7 were used

to determine proper operation of the GLA. Included in the procedures

were checkout of the GLAscheduler, filter gains and phases together

with the downlink signals. Results are recorded on Data Sheet 4.6. The

GLAsuccessfully passed all requirements.

9.2.6.5 RSS. The procedures of Tests 4.6.3.8 through 4.6.3.11 were

used to verify the RSS. The RSSscheduler, filters and downlink

functions were checked out. Results are recorded on Data Sheet 4.6.

The RSSsuccessfully passed all requirements.

9.2.6.6 MLA. The procedures of Tests 4.6.3.12 and 4.6.3.13 were used

to verify the MLA. Results are recorded on Data Sheet 4.6. The MLA

successfully passed all requirements.

326

I
9.2.6.7 Function Generator. The procedures of Tests 4.6.3.14 and

4.6.3.15 were used to verify operation of the function generator.

Results are recorded on Data Sheet 4.6. The function generator

successfully passed all requirements.

9.2.6.8 Value Drivers. The procedures of Test 4.6.3.16 were used to

check out the left and right outboard value drivers. Data are recorded

on Data Sheets 4.1.3.6.

The procedures of Test 4.6.3.17 were used to check out the left and

right, inboard and outboard segments of the inboard value drivers.

Results are recorded on Data Sheet 4.6.

All value drivers successfully passed all requirements.

9.2.6.9 Uplink Confirmation. The procedures of Test 4.6.3.18 were used

to check out the uplink confirmation (downlink) signals. All signals

were within the required tolerances and are recorded on Data Sheet 4.6.

9.2.6.10 AFCS, Right Wing, Downlink Connector Interface. The

procedures of Tests 4.6.3.19 through 4.6.3.21 were used to check the

signals on the AFCS, right wing and downlink connector. The signal

interface data are recorded on Data Sheet 4.6. The signal interface

successfully passed all requirements.

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328
i0. FLIGHT RESTRICTIONS

The wing structure is designed for positive 2.5g maneuver with the load

alleviation system on at the maneuver design condition and the loads at

the maneuver test condition will be slightly less than the design level.

The test condition maneuvers should be restricted to a maximum of 2.1g

without the load alleviation systems engaged.

The wing structure is sized for a design gust of 62 fps peak with the

load alleviation system on. At the gust test condition, the critical

load level will occur at a gust peak of 40 fps without the load

alleviation system engaged.

Because of the limited phase margin of the primary AFCS and the RSS at

the maximum dynamic pressure condition with the load alleviation systems

off, the dynamic pressure of 412 psf occurring at the GLA test condition

should not be exceeded significantly with the load alleviation systems

disengaged.

PRI_G _'AGE BLANK NOT FILMED

pAGE BLANK NOT

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330
11. CONCLUSIONS AND RECOMMENDATIONS

11.1 Conclusions. The third design cycle iteration analysis has been

completed and results indicate that the ARW-2 wing structure and control

systems meet the design objectives.

System hardware has been fabricated and functional, EMC and

environmental tests have been successfully completed.

Boeing furnished hardware is ready for installation, checkout and flight

test on the DAST ARW-2 vehicle.

11.2 Recommendations. The DAST ARW-2 wing with integrated active

controls should be flight tested to verify the performance and benefits

of active controls in the design of high aspect ratio wings.

It is recommended that the following testing sequence be established to

proceed from the most stable flight conditions to most unstable

conditions:

Recommended sequence of tests

a. Primary AFCS and RSS system verification at launch and load

alleviation test conditions

b. Load alleviation systems evaluation tests

c. RSS system verification at cruise and maximum altitude conditions

d. Flutter system evaluation

It is also recommended that the backup AFCS be verified, perhaps at the

end of a flight, before advancing to the more unstable flight

conditions.

I_.CED_'_ PAGE 'BLAf_ I_IOT _IIL_"I_

331
REFERENCES

I •

•

,

•

o

Boeing Document D3-I1535-I, Integrated Design of a High Aspect

Ratio Research Wing with an Active Control system for Flight Tests

on a BQM-34E/F Drone Vehicle, 25 June 1979. (Available as NASA

CR-166108, Volume I; NASA CR-166109, Volume I, Appendices)

Olen E. Visor and Francis D. Severt, Preliminary Design Study of

Flutter Suppression Control System forthe BQM-34E/F Drone Aircraft

with a Supercritical Wing, NASA CR-145208, 7 July 1976.

C. R. McGehee and G. E. Hodges, Final Design and Fabrication of

an Active Control System for Flutter Suppression on a Supercritical

Aeroelastic Research Wing, NASA CR-165714, June 1981.

Book, Herbert E. Merritt, Hydraulic Control System, John Wiley

& Sons, 1967.

John W. Edwards, NASA TN D-6928, Analysis of an Electrohydraulic

Aircraft Control-Service Servo and Comparison with Test Results,

dated August 1972.

332
Standard BibliographicPage

I. ReportNASANo.CR_I77959,Part i 2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle

DESIGN VERIFICATION AND FABRICATION OF ACTIVE CONTROL

SYSTEMS FOR THE DAST ARW-2 HIGH ASPECT RATIO WING,

PART l

7. Author(s)

C. R. McGehee

9. Per_rming Organization Name and Address

Boeing Military Airplane Company

P.O. Box 7730

Wichita, Kansas 67277-7730

12. Sponsoring Agency Name _d Address

National Aeronautics and

Washington, DC 20546

Space Administration

5. Report Date

January 1986

6. Performing Organization Code

8. Performing Organization Report No.

D500-10897-I

I0. Work Unit No.

11. Contract or Grant No.

NASI -16Ol 0

13. Type of Report and Period Covered

Contractor Report

14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract

A study was conducted under Drones for Aerodynamic and Structural Testing

(DAST) program to accomplish the final design and hardware fabrication for

four active control systems compatible with and ready for installation in

the NASA Aeroelastic Research Wing No. 2 (ARW-2) and Firebee II drone flight

test vehicle. The wing structure was designed so that Active Control Systems

(ACS) are required in the normal flight envelope by integrating control

system design with aerodynamics and structure technologies. The DAST ARW-2

configuration uses flutter suppression, relaxed static stability and gust

and maneuver load alleviation ACS systems, and an automatic flight control

system. Performance goals and criteria were applied to individual systems

and the systems collectively to assure that vehicle stability margins,

flutter margins, flying qualities and load reductions were achieved.

17. Key Words (Suggested by Authors(s))

Flutter Suppression

Relaxed Static Stability

Maneuver Load Alleviation

Gust Load Alleviation

18. Distribution

Until January 1988

19. S_urity Classif.(of this report)

Unclassified S°'Unclassified Security Classif.(of this page) 121"N°'°fPages122"Price332