Analyses and tests of the B-1 aircraft structural mode control system

3 1176 00156 1985 " NASA-CR-144887 1 . 9800006814 NA S A C o n t r ac t o r R e por t 144 88 7 •3 , : ANA L YSES AND TESTS OF T HE B-1 AIRCR AF T STRUC T URAL MODE CONTROL SYSTEM J ohn H. W ykes , Thom a s R. By ar , Cary J . M ac M i l ler, and Dav i d C. Greek C ont rac t NAS4-25 19 '- J a n ua r y 19 8 0 * !

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Anal g ses and tests o_ the B-1 aircraft structural mode control system.

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NASA Contractor Report 144887 ANALYSES AND TESTS OF THE B-1 AIRCRAFT STRUCTURAL MODE CONTROL SYSTEM John H. Wykes, Thomas R. Byar, Cary J. MaeMiller, and David C. Greek Rockwell International El Segundo, California Prepared for Dryden Flight Research Center under Contract NAS4-2519

N I A ivv o - i 73

National Aeronautics and SpaceAdministration TABLE OF CONTENTS Page SU_tIARY I IKI'RODUCTION 1 FLEXIBLEAIRCRAFT EQUATIONSOF MOTION 3 FLEXIBLEAIRCRAFTANA L YSES MODEL 22 DynamicAnalysis System 22 Free-FreeVibrationModal Data 24 AerodynamicData 27 Control-Surfaces InertiaReaction Forces 38 Active Control Systems 42 COMPARISONSOF ANALYSES AND FLIGHT-TESTRESULTS 42 FORWARD SMCS SENSOR PACKAGE RELOCATION 58 TRUNCATEDANALYTICALMODELS 78 ANALYSISOF SNCS VANEAERODYI_.IICINTERFERENCE EFFECT 90 IMPACT OF SHCSON SE L ECTED LOADS 101 Backgroundof Using SMCS in the B-I FatigueAnalysis I01 Gust Loads Analysis Description 102 Gust Loads StructuralModel 103" Gust Loads Aerodynamics 104 Gust Loads Control Systems 106 Load Method Discussion 106 Gust StatisticalLoad Calculations II0 Load Phasing IIi ConditionMatching 112 -_ Example ConditionLoad Results 113 SMCS VANE EFFECT ON INLET / ENGINE CHARACTERISTICS 154 Test Description 155 Flight-TestInstrumentation 159 i TABLE OF CONTENTS - (Continued) Page Test Results 165 OscillatingVanes 167 Static Vane Deflections 170 Maneuvers 170 Engine Throttle Transients 195 Summary 205 SUMMARY OF SMCS FLIGHT TEST RESULTS 205 SMCS Performancein Time-History-Data Format 206 SMCS Performancein PSD-Data Format 213 SMCS and Handling Qualities 213 SMCS High-GainTests 218 Crew Evaluationsof SMCS Effects 220 Overall Ride Quality 220 Flight-PathControl Tasks 225 Non-Flight-Path Control Tasks 225 Readabilityof Instrumentsand Displays 225 Reaching / Using Controls 226 Crew Fatigue 227 Motion Sickness 227 Physical Discomfort 228 AdditionalRide Quality Observations 228 Handling Qualities 229 APPENDIX - NOMENCLATURE. 230 General 230 Ride Quality Equationsof Motion Related 234 Load Equationsof Motion Related 247 Engine / InletRelated 250 REFERENCES 253 ii LIST OF ILLUSTRATIONS Figure Title Page 1 B-I aircraftwith wings swept aft. 2 2 Sign conventionfor rotationaland linear rates and accelerations. 15 3 Sign conventionfor coefficientsand control surface deflections. 16 4 Structuralmode deflections,slopes,and generalized forces sign convention. 17 5 Angle-of-attack(_) and sideslip Q S ) definitions. 19 6 Euler angle definitionsand rotation sequence. 20 7 Dynamic analysis system. 23 8 Typical elastic axes and mass point locations. 25 9 Fuselage elastic axis refinement. 26 i0 Typical symmetricmode vector plot. 29 ii Typical antisymmetric mode vector plot. 30 12 Typical structuralmode deflectiongrid points. 31 13 Panellingand box grid for Doublet Lattice aerodynamics program. 32 14 Typical low-frequencyadjustmentmade to analytical frequency-dependent aerodynamicdata. 35 15 Typical aeroelasticflexible-to-rigid ratio data for aerodynamiccoefficientsas a function of participatingstructuralmodes. 39 16 Typical control surface inertiareaction generalized forces. 41 17 Pitch axis SCAS analyticalmodel. 43 18 Yaw axis SCAS analyticalmodel. 44 19 Roll axis SCAS analyticalmodel. 45 20 Vertical SMCS analyticalmodel. 46 21 Lateral SMCS analyticalmodel. 47 22 Comparisonof flight test and analyticaldata, frequencyresponse of normal load factor at FS 571.5 (225)due to SMCS vane deflection, SCAS off, SMCS off. 49 23 Comparisonof flight t e st and analyticaldata, frequencyresponse of normal load factor at FS 571.5 (225)due to SMCS vane deflection, SCAS on, SMCS off. 50 24 Comparisonof flight test and analyticaldata, frequencyresponse of normal load factor at FS 571.5 (225)due to SMCS deflection,SCAS on, SMCS on. 51 iii LIST OF ILLUSTRATIONS- (Continued) Figure Title Page 25 Comparisonof analyticalSMCS actuatormodels to flight and simulator test data. 53 26 Effect of test-derivedSMCS actuatormodel, frequency response of normal load factor at FS 571.5 (225) due to SMCS vane deflection,SCAS on, SMCS on. 54 "'27 Comparisonof flight test and analyticaldata, frequencyresponse of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off. 55 28 Comparisonof flight test and analyticaldata, frequency responseof lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off. 56 29 Comparisonof flight test and analyticaldata, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on. 57 30 Effect of SMCS lateralgain on power spectral density of lateral load factor at pilot station FS 746.8 (294). 59 31 SMCS sensor package locationsand coupling characteristics. 60 32 Effect of SMCS sensor package locationon lateral load factor at FS 571.5 (225)due to differentialvane deflection,analyticaldata. 61 33 First fuselage vertical bendingmode damping versus SMCS gains for original and relocated forwardsensor package. 63 34 SMCS vertical axis performancewith relocated forward sensor package, frequencyresponse of normal load factor at FS 515.6 (203)due to SMCS vane deflection, case I. 65 35 SMCS lateral axis performancewith relocatedforward sensor package, frequencyresponse of lateralload factor at FS 515.6 (203)due to SMCS differential vane deflection,case i. 66 36 SMCS lateral axis performancewith relocatedforward sensor package, frequencyresponse of lateralload factor at FS 746.8 (294)due to SMCS differential vane deflection,case I. 67 37 SMCS lateral axis performancewith relocatedforward sensor package, frequencyresponse of lateral load factor at FS 515.6 (203)due to SMCS differential vane deflection,case 2. 68 iv LIST OF ILLUSTRATIONS - (Continued) Figure Title Page 38 SMCS lateralaxis performancewith relocatedforward sensor package, frequencyresponse of lateral load factor at FS 746.8 (294)due to SMCS differential vane deflection,case 2. 69 39 SMCS vertical axis performancewith relocated forward sensor package, frequencyresponse of normal load factor at FS 515.6 (203)due to SMCS vane deflection,case 2. 70 40 SMCS lateral axis performancewith relocatedforward sensor package, frequencyresponse of lateral load factor at FS 515.6 (203)due to SMCS differential vane deflection,case 3. 72 41 SMCS performancewith relocatedsensor, PSD of vertical load factor at FS 746.8 (294) - pilot station, case i. 73 42 SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 746.8 (294) - Pilot station, case i. 74 43 SMCS performancewith relocatedsensor PSD of vertical load factor at FS 746.8 (294) - pilot station,case 2. 75 44 SMCS performancewith relocated sensor,PSD of lateral load factor at FS 515.6 (203) sensor location, case I. 76 45 SMCS performancewith relocated sensor, PSD of lateral load factor at FS 746.8 (294) pilot station, case 2. 77 46 SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 515.6 (203) - sensor location, case 2. 79 47 SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 746.8 (294) - pilot station, case 3. 80 48 SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 515.6 (203) - sensor location, case 3. 81 49 SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 746.8 (294) - pilot station,case 4. 82 50 Normal load factor due to horizontaltail frequency response comparisonsof full and truncateddynamic analyticalmodel, SCAS on. 91 51 Typical vane-inducedinterferenceforces. 93 52 Effect of SMCS vane aerodynamicinterferences, frequencyresponse of normal load factor at FS 571.5 (225) due to SMCS vane deflection,SCAS off, SMCS off. 95 V .

LIST OF ILLUSTRATIONS (Continued) Figure Title Page 53 Effect of SMCS vane aerodynamicinterferences, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off. 96 54 Effect of SMCS vane aerodynamicinterferences, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on. 97 55 Effect of SMCS vane aerodynamicinterferences, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off. 98 _6 Effect of SMCS vane aerodynamicinterferences, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off. 99 57 Effect of SMCS vane aerodynamicinterferences, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on. I00 58 Loads analysis SIC point locationsand Doublet Lattice geometry. 105 59 Control systems for gust loads. 107 60 Structuralstations for gust loads analysis. 108 61 Load calculationequations. 109 62 SMCS off, wing bending moment frequencyresponse - WS 985 cm (387.6in). 115 63 SMCS off, wing bending moment responsepower spectrum WS 985 cm (387.6in.). 116 64 SMCS off, wing bending moment exceedances- WS 985 cm (387.6in.). 117 65 SMCS off, forebodybending moment frequencyresponse FS 1377 cm (542 in.). 118 66 SMCS off, forebody bendingmomentresponse power spectrum - FS 1377 cm (542 in.). 119 67 SMCS off, forebody bendingmoment exceedances FS 1377 cm (542 in.). 120 68 SMCS off, forebody bending moment frequencyresponse FS 2367 cm (932 in.). 121 69 SMCS off, forebodybending moment responsepower spectrum - FS 2367 cm (932 in.). 122 70 SMCS off, forebody bending moment exceedances- FS 2367 cm (932 in.). 123 71 SMCS off, normal load factor frequencyresponse - aircraft CG. 124 vi LIST OF ILLUSTRATIONS(Continued) Figure Title Page 72 SMCS off, normal load factor responsepower spectrum - aircraft CG. 125 73 SMCS off, normal load factor exceedances aircraft CG. 126 74 SMCS off, normal load factor frequencyresponse - pilot station. 127 75 SMCS off, normal load factor responsepower spectrum- pilot station. 128 76 SMCS off, normal load factor exceedances pilot station. 129 77 SMCS off, delta horizontalstabilizerfrequencyresponse. 130 78 SMCS off, delta horizontalstabilizerresponsepower spectrum. 131 79 SMCS off, delta horizontalstabilizerexceedances. 132 80 SMCS on, wing bending moment frequencyresponse - WS 985 cm (387.6in.). 133 81 SMCS on, wing bendingmoment response power spectrum - WS 985 cm (387.6 in.). 134 82 SMCS on, wing bending moment exceedances- FS 985 cm (387.6 in.). 135 83 SMCS on, forebodybendingmoment frequencyresponse - FS 1377 cm (542 in.). 136 84 SMCS on, forebodybending moment response power spectrum - FS 1377 cm (542 in.). 137 85 SMCS on, forebodybendingmoment exceedances FS 1377 cm (542 in.). 138 86 SMCS on, forebody bending moment frequencyresponse FS 2367 cm (932 in.). 139 87 SMCS on, forebodybending moment responsepower spectrum - FS 2367 cm (932 in.). 140 88 SMCS on, forebodybending moment exceedances- FS 2367 cm (932 in.). 141 89 SMCS on, normal load factor frequencyresponse - aircraftCG. 142 90 SMCS on, normal load factor response power spectrum - aircraft CG. 143 91 SMCS on, normal load factor exceedances- aircraft CG. 144 92 SMCS on, normal load factor frequencyresponse - pilot station. 145 93 SMCS on, normal load factor response power spectrum - pilot station. 146 94 SMCS on, normal load factor exceedances- pilot station. 147 95 SMCS on, delta horizontalstabilizerfrequencyresponse. 148 96 SMCS on, delta horizontalstabilizerresponse power spectrum. 149 vii I, IST OF ILLUSTRATIONS - Continued Fi gu r e Title Page 97 SMCS on, delta horizontalstabilizerexceedances. 150 98 SMCS on, delta mode controlvane frequencyresponse. 151 99 SMCS on, delta mode controlvane responsepower spectrum. 152 i00 SMCS on, delta mode controlvane exceedances. 153 I01 B-I flight test aircraft,wings swept 65 degrees. 156 I02 B-I air inductionsystem. 157 103 B-I inlet subsonicdiffuser flow area distribution. 158 104 B-I inlet / engineaerodynamicinterfaceplane instrumentation. 160 105 B-I flight-testAIP data acquisitionand signal conditioningsystem. 161 106 In-flightcalibrationcycle, total pressures at aerodynamicinterfaceplane, flight 1-5, 64 sps. 163 I07 Representativediscriminatoroutput,AIP total- pressure CBW data, flight i-i0. 164 108 AIS / SMCS investigation- flight 2-19,Mach 0.83. 168 I09 AIS / SMCS investigationflight 2-19,Mach 0.83 / 16,000feet, = 3 degrees, vortex ingestionin no. 2 inlet during sideslipoperation,PLA = max. 169 ll0 AIS / SMCS investigation, flight 2-19, _ = 3 degrees, vortex ingestionin no. 1 inlet during sideslip operation. 171 Iii AIS / SMCS investigation, flight 2-19, ff = 3 degrees, vortex ingestionin no. 2 inlet during IDLE-INT throttle transient. 172 If2 Flight 2-33 test matrix and AIP signalsduring Mach 0.85 operationwith SMCS vanes deflected 20 degrees, no. 2 inlet, RB = 7 degrees, RC = 5 degrees 173 113 No. 2 inlet, effect of sideslip angle on steady-state inlet characteristicsat SMCV= 20 degrees, flight 2-33. 174 114 No. 2 inlet, effect of sideslip angle on steady-state inlet characteristics at various SMC vane angles, flight 2-33. 175 .

115 Time historiesof total-pressure recovery and stall- margin index, Mach 0.85, with SMCS vane deflected and 0 and 20 degrees, no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33. 177 116 Dynamic circumferential and radial distortioncomponents, Mach 0.85, with SMCS vane deflected0 and 20 degrees, no. 2 inlet, RB = 7 degrees, flight 2-33. 178 viii LIST OF ILLUSTRATIONS - (Continued) Figure Title Page 117 Dynamic total pressure contours during Mach 0.85 179 operationWith SMCS Vane deflected0 and 20 degrees, no. 2 inlet, RB = 7 degrees, RC = 5 de_rees, flight 2-33.........

118 Effectsof SMC vane position on no. 2 inlet steady-state and dynamic characteristics, Mach = 0.85 and alpha = " 2.7 degrees, flight 2-37. 180 119 Effects of sideslip angle on steady-stateinlet characteristics with SMCV= 0 degrees and alpha = 1 degree, flight 2-37. 181 120 Effects of sideslip angle on steady-stateinlet characteristics with SMCV = 20 degrees and alpha = 1 degree, flight = 2-37. 182 121 Effects of SMCVwake on no. 2 inlet dynamic total- pressure contours during sideslip,SMCV = 20 degrees and alpha = 1.4 degrees, flight 2-37. 183 122 Effects of sidesli p angle on steady-stateinlet characteristics with SMCV = -8 degrees and alpha = 1 degree, flight 2-37. 184 123 Effects of SMCV wake on no. 2 inlet dynamic total- pressure contours during sideslip,SMCV = -8 degrees and alpha = 0.9 degrees, flight 2-37. 185 124 Effects of sideslip angle on Steady-stateinlet characteristics with SMCV= 0 degrees and alpha = 2.6 degrees, flight 2-37. 186 125 Effects of sideslip angle on steady-stateinlet characteristics with S_4CV = 20 degrees and alpha = 2.6 degrees, flight 2-37. 187 126 Effects of SMCVwake on no. 2 inlet dynamic total- pressure contours during sideslip,SMCV = 20 degrees and alpha = 2.6 degrees, flight 2-37. _ 188 127 Effects of sideslip angle on steady-stateinlet 1 characteristics with SMCV = 0 degrees and alpha = 5.8 degrees, flight 2-37. 189 _ 128 Effects of sideslip angie on steady-stateinlet characteristics with SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37. 190 129 IndividualAIP probes, total-pressure recovery array, SMCS vane deflected 20 degrees,Mach 0.85, _ = 6 degrees, flight 2-37. 191 130 Variations in circumferential and radial distortion componentsduring sideslip operationwith the SMCS vane deflected+20 degrees,Mach 0°85, _ = 6 degrees, flight 2-37. 192 ix I, IST OF ILLUSTRATIONS- (Continued) F i gure Title Page 131 Effects of SMCV wake on no. 2 inlet dynamic total- pressure contours during sideslip,SMCV= 20 degrees and alpha = 5.8 degrees, flight 2-37. 193 132 Effects of SMCVwake on no. 1 inlet dynamic total- pressure contours during sideslip,SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37. 194 133 Effectsof angle-of-attack maneuvers on steady-state inlet characteristics with SMC vane at 20 degrees, flight 2-42. 196 134 Effectsof an angle-of-attack maneuver on steady-state inlet characteristics with no. 1 engine at intermediate, flight 2-42, SMCS vane deflectionangle = 0 degrees. 197 135 Effects of an angle-of-attack maneuver on steady-state inlet characteristics with no. 2 engine at intermediate, flight 2-42, SMCS vane deflectionangle = 0 degrees. 198 136 Effects of an angle-of-attack maneuver on steady-state inlet characteristics with no. 2 engine at intermediate, flight 2-42, SMCS vane deflection angle = 0 degrees, 199 137 Effects of sideslip angle on steady-stateinlet characteristics with SMCV= 13 degrees and alpha = 3.3 degrees, flight 2-38. 200 138 Effects of SMCVwake on no. 2 inlet dynamic total- pressure contoursduring sideslip,SMCV = 13 degrees and alpha = 3.3 degrees, flight 2-38. 201 139 Steady-'state inlet characteristics during no. 2 engine transients,SMCV = 20 degrees, alpha = 5.5 degrees, and beta = +4 degrees, flight 2-38. 202 140 Steady-stateinlet characteristics during no. 2 engine transients,SMCV= 20 degrees, alpha = 3 degrees, and beta = +4 degrees, flight 2-38. 203 141 Steady-stateinlet characteristics during no. 2 engine transients,SMCV= 20 degrees and -8 degrees, alpha = 1 degree with positive sideslip,flight 2-38. 204 142 Typical dynamicresponse near crew station due to turbulenceduring low-altitude,high-speedflight. 207 143 SMCS performancein turbulenceM = 0.70 alt = 305 m (I,000ft) AGLA= 65°. 208 144 First fuselage symmetricstructuralmode damping from horizontaltail pulse excitations. 210 X LIST OF ILLUSTRATIONS- (Concluded) Figure Title Page 145 Effect of SMCS vertical gain setting on first fuselage symmetricstructuralmode damping at various aircraft weights. 211 .- 146 First fuselage antisymmetric structuralmode damping from forced SMCS vane oscillations. 212 147 Vertical SMCS performancein turbulenceas shown by power spectraldensity of vertical load factor at pilot station, FS 747 (294). 214 148 Lateral SMCS performancein turbulenceas shown by power spectral density of lateral load factor at pilot station,FS 747 (294), single-peakresponse. 215 149 Lateral SMCS performancein turbulenceas shown by power spectral density of lateral load factor at pilot station,FS 747 (294),double-peakresponse. 216 150 SMCS impact on short-periodand dutch roll frequencies. 217 151 Ride quality ratings for varying degrees of turbulence. 221 152 Effects of turbulenceon flight path control tasks. 221 153 Effects of turbulenceon tasks other than flight path control. 222 154 Effects of turbulenceon readabilityof instruments and displays. 222 155 Effects of turbulenceon reachlng / usin g controls. 223 156 Effects of turbulenceon crew fatigue. 223 157 Effects of turbulenceon tendency for motion sickness. 224 158 Effects of turbulenceon physical discomfort. 224 xi LIST OF TABLES Table No. Title Page I General FlexibleVehicle Equationsof Motion, Time Domain 5 II Equationsof Motion for LongitudinalRigid-Bodyand SymmetricStructuralModes, FrequencyDomain I0 I[l Equationsof Motion for Lateral-Directional, Rigid-Body and AntisymmetricStructuralModes, FrequencyDomain 12 IV Typical AnalyticalStructuralMode Data 28 V Longitudinal-Symmetric Aero CoefficientsFrom Doublet Lattice Program 37 VI SMCS Vane EffectivenessIncludingInterference Effects, Longitudinal-Symmetric Case, FrequencyDomain 94 VII Gust Loads Model StructuralDegrees of Freedom 103 VIII Load Comparison,SMCS Off Versus SMCS On 114 IX SMCS Vane Effects on Inlet / Engine Characteristics, Flight Test Investigation 166 X SMCS High-GainTests Summary 219 x ii ANALYSES AND TESTS OF THE B-I AIRCRAFT STRUCrURALMODE CONTROL SYSTIg4 John H. Wykes, Thomas R. Byar, Cary J. MacMiller, and David C. Greek Rockwell International, North American Aircraft Division - _y An 18-month programwas conductedt o c o mpileand document f o r publication informati o npertainingto analysesand flight tests of the B-I StructuralMode Control System (SMCS). This is the sec o nd phase of a continuingeffort; results from the first phase study are d o cumentedin reference i. This report covers the followingtopics: (I) Flexible aircraft equationsof motion (2) Descriptionof flexibleaircraft analysesm o del (3) Comparisonof analysesand flight-testperformanceresults of the SMCS (4) A stmmaryof the study of the forward_4CS sensor package rel o cati o n (5) Truncated analyticalmodels used in simulationeff o rt (6) An analysisof the SMCS vane interference effects (7) ImFact of SMCS on selectedloads (8) Flight-testresults o f the SMCS vane effects o n inlet / engine characteristics " - (9) Stmmary of SMCS flight-testresults INTROII /CT ION The B-I aircraft is one of the first vehicles t o includea contr o lcon- figuredvehicle (CCV) concept in the early design phases. The aircrafthas a requirementto provide a specifiedlevel of ride quality for the crew. This requir e menthas been met on the B-I thr o ugh the use of an automaticcontrol system (_CS) whose main externalfeature is a set of vanes (near the crew station)which are canted down 30 degrees from the horizontal. (See figure i.)

N Structural mode cont ro I vanes Figure 1. - B-1 aircraft with wings swept aft.

A sM_stantial sa v ingsin weightwas achieved with thisapproach as compared to direct material stiffening. T he details of systom re_luiranents had to be determined from a production (long-life) point of view , which has not been done before for a system of this type. Extensive wind tunnel tests of the vane characteristics were conducted. Analytical models of the flexible air- craft and control systems were developed to analyze requirements and t o investigate stability and performance. Component parts were tested to the .- requirements in the laboratory. Flight tests of the SMCShave been conducted , and comparisons with analytical predictions have been made. Because of all of this , it has been recognized that the B - 1 offers an excellent opportunity for much needed further evaluation of such a system as the SMCSto insure the optimun use of these systems for future applications.

The overall objective of this research area is to compile and document information about the conceptual design , development , and flight tests of the B-1 SMCSand its impact on ride quality. Since the B-1 is the first aircraft to have a system such as the SMCSdesigned for production and long-service use , it is expected that the reports prepared will add to the technology base for design of future large military or civil aircraft. The specific overall objectives are to: (1) Investigate the improvements in total dynamic response of a flexible aircraft and the potential benefits to ride qualities , handling qualities , crew efficiency , and reduced dynamic loads on the pr imary structures (2) Evaluate the effectiveness and performance o£ the SMCS , which uses snell aerodynamic surfaces at the vehicle nose to provide damping to the structural modes The majoreffortof the phasei study(ref.I) was to compile, edit,and preparefor publication as a N A SAcontractor reportthe existing-information on the B-I SMCSconceptual designand development. The majoreffortof the presentphaseII studyis to reporton the analyses and flighttestsof the SMCS;existing information has been augmented by someadditional limited anal y ses and flight-test datareductions.

FLEXIBLE AIRCRAFT EQUATIONS OF M orION The equations of motionof the flexible B-1 formthe foundation of much that is to followand so it is appropriate to discussthemfirst. T he treat- ment of theseequations is suchthat theyare not developed hereinfromfirst principles.Thereare manytextbooks thatdo thistask;references 2 and 3 are typicalof these. The equations, however, are presented in sufficient detailto be readil y related to the textbook treatments.The equations of motion oi,this section of the report were developedto serve the purposes of ride quality, terrain-following, and handling qualities evaluations;the equationsof motion relatedto the loads analyses are discussed in a subse- quent section. An att e mptwas made to include the main aircraft charac- ter(sties importantto these several types of studies;however, as specific studieswere conducted,minor modificationswere often made. It is not intended to catalog all of these iterations;where importantto the studies reported herein, they will be discussed.

The discussionstouch upon the key featuresof these equationsalong with explanationsof the form of the data where it is felt that this would be help- ful. The authors have electedto stay away from matrix notation in these _ initialdiscussionsin order to show as much informationabout the dynamic modeling as possible. The appendix containsa list of symbolsused in defining the equationsof motion of this section.

The equationsof the motion in tables I, II, and III are written in a body axes system (figures2 and 3) where the X-axis passes through the center of gravity and is parallel to the vehicle fuselagereferenceaxis (FRL). To help those who are more familiar with stabilityaxes notation, it shouldbe observed that all of the aerodynamiccoefficientsbut two appear the same in either the stabilityaxes system or the body axes system. The two that are differentare the normal force coefficient,CN (bodyaxes), versus lift coef- ficient, CL (stabilityaxes), and the chord force coefficient,CC (bodyaxes), and drag coefficient,CD (stabilityaxes). For small angles of attack,CN nearly equals CL and CC nearly equals CD.

All accelerationsand velocitiesare defined positiveas indicatedin figure 2. One exceptionis noted r e lati v e to the definitionof the vertical load factor. As a concessionto stabilityand control and flight test conven- tion, a positiveload factor,nz, is associatedwith a positive CN. In con- trast, the definitionof Z-axis acceleration(consistent with the positiv e definitionsof figure 2) is associatedwith a positive CZ in figure 3 (Cz = -CN).

Ride quality analyses have b een conducted on the simulator using time - domain equations while frequency - domain analyses have been conducted using digital computing equipment. The equations in the time domain are presented in table I and the frequencydomain in tables II and III.

The structuralflexibilityof the air vehicle is defined in terms of free v ibrationmodes of the structure (oftenreferred to as normal modes).

Figure 4 defines the sign conventionsused in associationwith.these normal TABLE I. - GENERAL FLEXIBLE VEHICLE EQUATIONS OF MOTION, TIME DO~~IN [Total Vehicle, Body Axes, Units: ft, lb, rad, sec] FORCE EQUATIONS, RIGID-BODY MODES

- ~~~&H - (rr:ov:cyXco5r)6cv- (:~h.)5A - (t~j[(CN~TF"tJ

i3 - -', + Poc -~~),8 + ~?:,cosesln<p +(~ s..,\{C~ (J + CY.6Ibw)~ + Cyr('b..,')r- + CYf(bw'f~

v. V) \m"l."J (0() ~~ ~ z.\i.J'

m n

+ CYp(~)p + Cyp(~fp + [[CY71lni+ CYni(~il + [(CyPn.n~(3 + Cyo .. 6r- + C'ys..Sr--

.-. .=. I •

ANrlt.YMI1ET~IC SYMMETRIC

+ Cy~~ +CY6~S: + CY&:v0~v+CY6~~CV +e;,OAoh.+c,-~6A.} +(~,jj&' +(~~(s\nf')~v+ (m:'I..~)8A

+ (~~)~(yv,)v"J

TABLE I. - Continued STRUCTURAL _DE EQUATIONS Symmetric modes, i --1 to n _ . _ .. _ _ _ C _ _ n J= . l \ MI \ MI Antisy_etric modes, i = 1 to m m n

+ _ , _v J _ _ + +

J = l ,J = l

_ T " A m 0 "

TABLE I. - Continued b,_]B _TEQUATIONS, RIGID-BODY MODES _ C m_(_ --" + \ ly / L ( o 0 .

n \- Iy - +\_ , ( i , .

_ w _ .r _, Syl.lrl i _ T R I C S ' Y _ 11"t £ T R t¢" . i_ _. \ I x / l x / TABLE I . - Continued LOAD-FACTOR EQUATIONS (ACCELERO_TER READINGS) _ . x X

R OMi _" 7

_I AT ANy | r_ is-_-IN THE SAME "rATION )4 . . Y .i2 !

D II ::_E C TION AS -I- Cr. I TATION X , Y, ;z I'l) , I S + IN TH E SA M E DI I _ E C T ION AS - i - C 7 ROTARY-RATE EQUATIONS (RATE GYRO READINGS) n [_,T AN y TATION _(,y] i =l _ -- _ L i ._,g _ .

P = P + L e, r _; ( '. _._ - ,_._R,_ .o0_ s )

i,, I LSTAT ION

rA T ANy X ' Y_ ]_J

I -- = I _ + _ - = - _ Y' _ " rA T ANY Z _ _' 7 _i ( AN TI S Y T' I I _ ETRIC N O DES ) L S T A T I ON _,Y,_] i =l TABLEI. - Concluded EULERANGLES COS_ EARTH-AXES VELOCITIES

_-- N_CoseCOS_ + V y (sIn_ slnocos_- cos_sln_) + V_(co%_slnocos_r. sln_ s,n_)

V. . E--_coses,n_ . ',_(cos}cos_ +s, nc}_nes,mf)+V_(cos}s,nesmg-sln{cos_r)

x4= -_s_n e + v r s,n_cose+v , cos}cose

A convenient equivalent expression for Vze is in terms of altitude rate of change and _ and 8.

I_ = V. (s,n e cos0ccos /s -cosesln{s,rt p - cosecos_cos,_sln0c)

_o TABLE II. - EQUATIONS OF MOTION FOR LONGITUDINAL RIGID-BODY AND ¢_ Sl q v_4ETRIC STRUCTURAL MODES, FREQUENCY DOMAIN [TotalVehicle, Body Axes, Units: ft, Ib, rad, sec] NORMAL-FORCE EQUATION, RIGID-BODY _K)DE PITCHING-MOMENT EQUATION, RIGID-BODY MODES TABLE II. Concluded SY HMETRICSTRUCTURALMODE EQUATION,i = 1 to n

,<,.+ i < ...>

t , _ I% 7 t, i < ,U"

n H x # , / NORMAL LOAD FACTOR (INCREMENT FROM 1 g TRIM) t_ , _ F A r A ., _x.¥, lz - J i - , i L _ TAT I ON Note: The variables are complex in form but have been written as _, q, _i, 6k, andwg as a space-saving notation.

[( ) ()R + i ( ) _ ( )g] representsreal and imaginaryparts of unsteady aerodynamics coefficientform.

These equationsrepresentperturbations from a 1 g trimmed flight condition.

• F _a N TABLE III. - EQUATIONS OF MOTION FOR LATERAL-DIRECTIONAL, RIGID-BODY, AND ANTISY_4_ETRICSTRUCRJRAL .MODES, FREQUF_NCYDOMAIN [TotalVehicle, Body Axes, Units: ft, Ib, rad, sec] SIDE-FORCE EQUATION, RIGID-BODYMODE !

TABLE III. - Continued ROLLING-MOMENT EQUATION, RIGID- BODY ~IODE

ANTISYMMETRIC STRUCTIJRAL MODE EQUATION, i = 1 to m

TABLE III. - Concluded LATERAL LOAD FACTOR ST A TAT ANy _oN x , y , :_] i: l Note: The variablesare complex in form but have been written as _, r, p, q i, 8k, and Vg as a space-savingnotation.

+ [( ) ()R i ( ) ()i] representsreal and imaginaryparts of unsteady aero- dynamics in coefficientfom . These equationsrepresentperturbationsfrom 1 g trimmed flight conditions.

+Y +q +vz , a +X + r +P ' + Z _ Bqdy axes Figure 2. - Sign conventionfor rotationaland linear rates and accelerations.

+Cn + _I_ +N .6r +Y +C N +N .._ +C_ - +L ' = Differental angle 6H +T between left and \ right panels +Cz + Z +X +6cv, Symmetric deflection +Z +_cv, Differential deflection Deflection magnitude and sense referenced to right panel; right +Cm panel trailing edge down, left +M panel trailing edge up.

Body axes Figure 3. - Sign conventionfor coefficientsand control surfacedeflections.

Symmetric Modes Side view

+_fx

+ X Body axis _.d d, ing point [ + C r l i AntisymmetricModes +_ / i _lizing point +C7 / .,

< +x ' --_'x cG _+4 'x

Top view Looking forward Figure 4. Structuralmode deflections,slopes,and generalized forces sign convention.

modes. There are a number of advantagesto using the modal approach in contrast to the direct-influence coefficientapproach for such studiesas referred to herein.

(i) It can describe the static as well as dynamic characteristics of the flexiblevehicle in a consistentmanner within the same format.

(2) Both inertiaand aerodynamicloadings are accounted for in the modal generalizedforce data for both the static and dynamic cases.

(3) The modal data are amenable to a number of simplification schemes.

(4) The approach integratesbest with control system design requirements in the handling and ride qualitiesdesign areas.

Consider first, the most extensivelydetailed equationsof motion in the time domain as describedin table I. These equationsneed be used in this detailed form shown only when large-scalemaneuveringis studiedas in the terrain-following problem. More simplifiedequationswill serve other analyses purposes.

These equationswere developedassuming that the angles of attack, _, and sideslip angles, 8, (figure5), would be small (lessthan i0 degrees),but the vehicle orientationin space as defined by the Euler angles 9,@ , and • (fig- ure 6) would be unrestricted(exceptfor @ = 90 degrees).

Aerodynamicdata indicatedare preliminaryestimatesof those required.

These data are shown in derivativeform exdept where it is anticipatednon- linear characteristics with _, _, or control deflectionoccur. As an example, the normal force curve was expected to be nonlinearwith _ so the normalforce coefficientis expressedas CN(_) insteadof the more familiar linear form CN_- The control surfaces explicitlyshown are those anticipatedbeing required by either the Stabilityand Control AugmentationSystem (SCAS)or SMCS. Other control-surfaceinputs are shown in general form as functionsof _k for the kth surface.

Because of the requirementto control structuralmotion at the frequencies of the lower free-vibration modes, it is necessary to consider unsteady aero- dynamic effectsof the control motion as well as the inertiareaction forces of these surfaces. In the normal force-equation format of _.able I, the unsteady aerodynamicsare shown by the notion CN_k_ k + CN. & for the kth 6 k k

E_

Body axes Figure 5. - Angle-of - attack(_) and sideslip (8) definitions.

Vertical plane Airplane plane of symmetry l + X-Body I' I I I I Horizontal plane m I + Y. Bod \ I l \ I \ \ \ + X-Earth + Z-Body + YrEarth I + Z-Earth Rotation sequence indicated by numbers at arrow heads Figure 6. Euler angle definitionsand rotation sequence.

2O control surface. The control-surfaceinertiareaction force in this equation has the form (mk _k) 6"k- Similar aerodynamicand inertiareaction-force terms may be recognizedin the moment and structuralmode equations.

Terms in table I involving4, 4, r, and p are generallyof small conse- quence but have been includedto be consistentwith the frequency-domain equationsof tables II and III where unsteady aerodynamics must be considered.

The desirabilityof this consistencywill become apparent as these equations " are discussed in subsequentparagraphs.

The B-I engine gyroscopicmoments, IR_Rr and IRO_Rq , shown in the equations have not proved to be significantin handling qualitiesor terrain-following studies. It had been anticipatedthat they might have been significantin large-scalemaneuvering. They have, however,been left in the equationsof table I.

The equationsof table I includethe ability to change speed. It is assumed that if significantchanges in Mach numbersare to be realized,these changes will be reflectedin the appropriateuse of Mach carpet data for the aerodynamicsrequired. Velocity changeswill show up directly in Vo, while combined altitude-velocity changeswill appear indirectlyin qo, (1 / 2 pV2).

The gust representation shown in table I is in the aerodynamictransfer- function form. Tilegust excitation,Wg and Vg, would come from random signal generationsources, shaped and scaled to reflectthe desired gust power spectraldensity and intensity.

The longitudinaland lateral-directional rigid-bodymotions are coupled during large-scalemaneuvers throughthe inertiaterms and engine gyroscopic effects. The symmetricand antisymnetricstructuralmode motions are coupled by terms representingthe dihedral effect due to symmetricwing bending. For small motions about a trim condition,the rigid-bodymode equationscan be decoupledthrough eliminationof the inertiaand gyroscopiccoupling. The structuralmode equationscan be decoupledby using trimmed airplane static -: symmetricstructuralresponse parametersat fixed values to determinethe effectivedihedral due to symmetricwing bending.

The load factor and rotation rates of the large-scalemaneuveringflexible air v ehicle (as read by accelerometers and gyros mounted on the fuselage structure)are presented in table I.

Euler angle equationsand earth axis velocitiesare given in table I and can be used in terrain-following studiesto determinethe vehicle attitude and locationwith respect to the earth's surface.

The frequency-domain equationsof motion for the flexibleairplane are given in table II for the longitudinal-sy_netric case and table III for the lateral-directional-antisymmetric case. These equationsare uncoupledand representmotion perturbationsfrom wing-level1 g trimmed flight. Compari- son of these frequency-domain equationswith the uncoupledtime-domain equationswill help in identifyingequivalentterms.

As shown, the vehicle response aerodyn.amics (that is derivativesassociated with response parameterssuch as a, _, q, q) are quasisteady, while the control surfaces and the gust are sho_cn as functionsof the forcing frequencyin the form [( )R + i ( )I]- This format has been convenientand sufficiently accurate for.preliminary ride qualitiesand structural-mode control analyses.

Digital programs are availableat Rochcell;however, that will also accept vehicle responseaerodynamicsfrom unsteady aerodynamictheories as a function of frequency,permittingmore refined ride quality and S_ICS stabilityanalyses.

The equationdescribingthe normal load-factorresponse at any location in the flexibleaircraft is presentedin table II. The similarequation for the lateral load factor is given in table III. These responsesare used in ride qualityanalyses.

FLEXIBLE AIRCRAFT ANALYSES bDDEL This sectiondescribes how the data were obtained to implementthe flexible aircraft equationsof motion used for ride quality analyses. It will not be the purpose of this presentationto providea complete set of data used in all analyses discussed;but it will be the intent to provideunderstandingof the data used.

DYNAMIC ANALYSIS SYST_ As a basis of understandingthe contentsof this sectionbetter, as well as topics of other sections,the chart of figure 7 is presented. Shown is -- the complete dynamicanalysis system supportingthe developmentof flexible vehicle dynamicanalysis models for control system developmentand ride quality analyses at RocMcell International's North American Aircraft Division, E1 Segundo. The path through this system as employed in developingthe SMCS is as follows. (No supersonicanalyseswere conductedduring SMCS developmentso the Mach-box program capabilitywas not used.)

Starting at the top of the chart, it is seen that the processbegins with a definitionof the vehicle geometry,basic wind tunnel correlated aerodynamics,structuralstiffness,and mass characteristics being provided to I I I il - Basic aerodynamics Vehi c le geometry I S tructural definition '' _ ' , STAR 6 stiffness' vibration analysis El, GJ'SIC I-- J T IMaSs characteristics,_ Modes, frequencies processing J Modal I _ Normalized modes, Doublet lattice II lJ M>IMaCh box l_lized

m<] l I

D y n amics, J _ ' I_ , Genera l ized

Techno l og_ J G e n era li ze d j aero f o rces

I I I

system Dynami c Flutter analysis response analysis C o ntr ol _ FH251 / FH255 H STAR 4 Control system requirements Flutter dynamic response , stability characteristics • _ I al -, Figure 7. - Dynamic analysis system.

the DynamicsTechnology(,roup. As indicated,the stiffnesscan he used in the EI-iU format or in the form of structuralinfluencecoefficients(SIC).

The ,lassand stiffness&ata enter the STAR 6 program, and vibrationanalyses of the whole vehicle are accomplished. The output of the program, then, is whole veh i cle vibrationmode shapes and frequencies.

These mode-shapedata are next manipulatedto produce normalizedmodes.

For tile B-I ride quality analyses,both the s y _netricand antisymmetricmodes were normalizedto a I_int on the nose of the vehicle. Generalizedmass data consistentwith the normalizedmodes are produced. Finally_modal deflection data are developedthrough interpolati o n programs along selectedstreamwise strips for input into the aerodynamicprograms.

For tile ride quality related studies reported herein, the Doublet Lattice.

Program has been used to provide theoretical aerodynamic generalized forces as required. These generaI ized aerodynamic forces are in dimensional form.

Programs have been developed which process the data from dimensional form to the coefficient form required by the dynamic analyses programs. These dynamic analyses programs employ the frequency-domain equations of motion discussed earl ier.

The 1:tt-251 program provides dynamic response results for the longitudinal- symmetric case , while the FH-255 program provides dynamic response results for the lateral-directional-antisymmetric case. Both of these programs can accept either frcxtuency-dependent or quasistead y data. Active controls can be included. Frequency responses due to gust or control forces may be obtained.

When gust inputs are employed , ride quality parameters are output and control- system deflections and rate responses are obtained in power spectral density form. Stability analyses are performed using the characteristic determinant frequency evaluation technique of reference 4.

FREE-FREE VIBRATION MODAL DATA The flexible aspects of the aircraft have been treated in the modal _ format as opposed to the direct-influence-coefficient approach. Ninety per- cent of the ride quality analyses performed have been accomplished using free-free vibration modes which were obtained using an EI-GJ description of the vehicle stiffness; more recent modal data have been obtained using structural influence coefficients. The details of the EI-GJ approach are discussed here.

Figure 8 shows the typical distribution of mass points on the elastic axes assumed. This is an earlymodel; figure 9 shows a refinementof the fuselage elasticaxis made at a later date. Each analysis includeda flexible . _FS 2461 (969) FS 4018 (1582) Nace IIe , points Z Typical control point motions calculated from EI-GJ approach WL 72.4

I

Uk -81 (-32) FS 3942 (1552) Figure 8. - Typical elastic axes and mass point locations.

I-o Elastic axis ., 'D f E A _ FRP

I -- L._ _1 (w,o)

Elastic axis coordinates* Point Fus Sta WL A O 38.6 (15.2) B 406 (160) 34.0 (13.4) C I003 (395) 143.8 (56.6) D 1283 (505) 123.2 (48.5) E 2477 (975) 80.0 (31.5) F 3937 (1550) 154.9 (61.0) G 4013 (1580) 50.8 (20.0) H 4430 (1744) 57.7 (22.7) *Straight lines between these points Figure 9. - Fusela g e elasticaxis refinement.

wing, fuseIage, horizontal tail, and vertical tail. Also included were flexibly mounted engines / nacelles. The resuItant mode shapes consisted of elastic axis deflections and rotations illustrated in figure 8.

Because the analyses were primarily oriented toward ride quality at the crew statio n, the free-free vibration modes were normalized at the most forward mass point at the nose of the aircraft. While most modes show a high degree of coupling among vehicle components , table IV identifies the main component ; (where this is possible) and lists the associatedfreque.ncies.In the analyses discussedherein, I0 symmetricmodes and 12 antisy_netric modes were used.

Discussionsto follow later in this section describe the rationaleused in selectingthe modes shown.

When ground vibration test (GVT)data became available,they were usod to upgrade the modal data. Symmetricorthogonalmodes were successfullyex- tracted from such tests. First, the effects of the soft support system were removed from the measured orthogonalset of modes. These data were next used, with proper fuel weights included,to analytically obtained orthogonalfree- free modes at the desiredweight condition. It was not possible,however, to directly extract an orthogonalset of antisymmetric modes from G V F data.

To obtain usable, consistent, antisymnetric modal data reflecting test results, a technique was used of adjusting local stiffness data until a successful approximation oF the measured data were obtained analytically . Then , as in the s_nmetric case , the effects of the soft suspension were deleted and the desired fuel weights adeed analytically to obtain orthogonal antisymmetric modes. The data of table IV are typical of those based on the GVT data ob- tained as described. It is to be noted also that structural damping was extracted; these data were obtained using the oscillation decay method.

Figures i0 and II present typicalsymmetricand antisymmetric mode vector plots.

The computerprograms used to determinethe aerodynamicdata require vibrationmode shapes to be definedon a grid system whose chords are parallel to the free stream. Thus the basic mode shapeswere interpolated to find point deflectionsalong strean_isechords for all lifting surfaces. (See fig- : ure 12.) The mode-deflection data on the grid system shown were used as input to the Doublet Lattice Program where interpolations to the Doublet Lattice grid system were made and required slope data determined within the program.

AERODYNAqICDATA The Doublet Lattice aerodynamictheory was used to obtain most of the aerodynamicsdue to the flexible structure. In addition,extensivewind TABI,Ii IV. - TYP I CAl,ANA L YTICA L ST R UCTURAl, MODE I)ATA Wt = 119,296.8 kg _(263 000 lb) A = 65° Structural a b lode Mode Frequency no. description Hz damping gs Symmetric 1 Wing, first bending 2.22 0.062 2 Fuselage,first bending 2.84 .094 3 Fi'rst nacelle 3.26 .024 4 IIorizontal tail, f_rst bending 4.19 .028 5 Wing, fore and aft bending 4. 2 3 .05 2 6 Fuselage, second bending 6.28 .016 7 Wing, s econd bending 7.57 .0 22 8 llori z ontal tail, fore and aft bending 8.31 .064 9 Fuselage, third bending 11.15 .055 10 llorizontal tail , first torsion 27.35 .042 Antisymmetric 1 First nacelle I.73 .145 2 Wing, first bending 2.41 .054 3 Horiz tail, first bending 3.51 .043 4 llorizontal tail, first bending 3.96 .025 5 tlori z ontal tail, fore and a f t bending 4.14 .049 6 Wing , fore and aft bending 4.20 .031 7 Fuselage, first bending 5.58 .032 8 Wing , second bending 6.96 .031 9 Vertical tail, first bending 7.21 .019 10 Fuselage, second torsion 9.72 .078 11 Fuselage, second bending 10.30 .022 12 Vertical tail, first torsion , 35.34 .020 aDeterminedfrom ground vibration tests Fore and aft motion Lateral motion Vertical motion Mode l _ Frequency = 2.23Hz Figure I0. Typical symmetricmode vector plot.

Fore and aft motion Lateral motion Vertical motion • \

\

\ _ / M o de l Frequency = 1.73 Hz t :- Figure II. - Typical antisymmetric mode vector plot.

3O ..

Sb t. 3 ' ..

' ..

;>..1 •• Figure 12. - Typical structural mode deflection grid points.

Horizontal tail location, WL 320 (126) WL 63.5 (25) I FS 0 Horizontal tail pivot, FS 4018 (1582) Figure 13. - Panellingand box grid for Doublet Lattice aerodynamicsprogram.

t_mel data were also generatedt o o btain static,rigid-bodyforce and moment- coefficientdata along with pressure-distribution data. These two data sources were correlated (to be described)and used as input to the flexible aircraft analyses.

T he panelling,box grid, and control-surfaces setup for the Doublet Lattice Program are shown in figure 13. The wing and fuselage forebodyhad five main panels with a total of 151 boxes. T he horizontaltail had two panels and a total of 64 boxes; the horizontaltail was _ut in, also, as an all-movablecontrol surface. The vertical tail had two panels and 73 boxes.

In addition, the vertical tail had an end plate at its base consistingof one panel with 30 boxes. As shown in figure 13, the lower rudder control surfacewas also modeled.

For longitudinal-symmetric data, includingcontrol effectiveness and gust data, the wing / forebodyand horizontaltail were run as shown in figure 13.

This same wing / forebodyarrangementwas used along with the full empennage (horizontal tail, vertical tail, and end plate) to obtain lateral-directional- antisymmetric, whole-vehicledata. Vertical tail gust data, lower rudder control effectiveness, and differentialhorizontaltail effectiveness data were obtained by runningthe empennageplus end plate as an entity. F o r fuselage side gust and generalizedaerodynamicforces,a modified slender body theory was used which made use of wind tunnel developedside-force distributiondata.

It is to be noted that the SMCS vanes were not modeled for the Doublet L attice Program. The aerodynamicsfor the vane were obtained from wind tunnel tests as described in referencei. None o f the vane-relatedaerodynamics were frequencydependent. The reducedfrequency (k) determined o n the basis of an assumed frequencyof i0 Hz (62.8 rad / sec),the highest frequencyrange that vane is expectedto be effectiveat M = 0.85, is k = _--_= (02.8)(2.46) = 0.0812

: ZVo (2](9Sl)

Compared to k for the wing under the same circumstances k- (62.8)(15.23) _- 0.503

(2)(9Sl)

This is a relativelylow reduced frequency,and the aerodynamicunsteadiness effectsare judged to be acceptablysmall. In retrospect,this appears to have been a valid judgement. In generatingthe structural-mdde generalized force coefficients,the vane force was ass u ned to act at a point; i.e., FS 581.66 (229). Thus C_. = ¢iFS 581.66for symmetricmodes is an example. ' 1_cv CN_cv In developingthe Doublet Lattice paneling and box patterns,lifting- surface spanload distributions were compared to the wind tunnel-related data.

The gap between the vehicle centerlineand the first row of chordwiseboxes of the horizontaltail is one of the devices used to obtain matched data sets for the horizontaltail. It was reasoned that with good matches of the lifting-surfacespanloadings, the computermodel developedwould give valid answers for both rigid body and structuralmodes.

The point of vi_¢ was adopted that the wind tunnel-related airloaddis- tribution for the rigid vehiclewere the most accurate data available. Thus, the rigid vehicle aerodynamicscouplings into the structuralmodes (Cni _ and Cni6 data are examples)were computed using these distributionsrather than Doublet Latticetheory, for the zero-frequency case. These data, in turn, were used to scale the frequency-dependent data produce d by the Doublet Lattice Program. The Cnin.j and Cni_j data were used directlyas generatedby the programwithout scalingfor both the synraetric and antisy_netric modes.

As an example of how the wind tunnel data and the Doublet Lattice frequency- dependent gust aerodynamicdata were brought into agreement,consider the typical e_ampleof figure 14. Shown is the pitching-moment coefficientdata due to a unit vertical gust velocity. The magnitude trends versus the reduced frequency,k, were assumed basicallycorrect and all adjustmentsto match wind tunnel based data were made at low frequenciesas illustrated. The data magni- t_le of the real componentat zero frequencywas adjusted to match the wind tunnel based data and then faired into the basic real curve at low reduced frequencies. For the case illustrated, the gust coefficientat zero frequency was determinedfrom angle-of-attack data, CmWg (Cm_) / Vo. This matching of the angle-of-attackand gust velocity data is essentialto obtainingvalid power spectraldensity responsesdue to gust data at low frequencies; a mis- match will produce a load factor response at zero frequencywhich is not there in the real world. In the real world, a stable aircraftwill weathervaneinto the resultantvelocity due to the combinedmotion and gust velocity and have no load factor at zero frequency.

As mentionedwhile discussingthe analyses flow, the data producedby the Doublet Lattice Programmust be reduced to the coefficientform of the equationsof motion discussed in the previous section. In order to do this, the data are processed in the followingmanner.

0 Real l Symbols denote [] Image[ calculated data point O.OO4 0.002 C r r_ g 1 , ft / se c 0 __ .6 k .8 1 . 0 " _ I . _I 6 -0.002 ,--_r]F -- ..

-- ,ow-freqoency adjustment to

fit wind-tunnel-test based data -0.004 Gust reference point is FS 2649 (I043) Figure 14. - Typical low-frequency adjustmentmade to analytical frequency-dependent aerod_lamicdata.

The program outputs data normalizedto air density (2p) and to frequency (o_ 2) as illustrated here using the dimensionallift force due to plunging Lh (for the whole vehicle), [Lh / 2 P _]. The divisionby two is because the data generated in the program are for half of a vehicle, p is assumed to be unity by the program. Velocity is determinedby the data of Mach number and velocity of sound at the altitude selected. Frequency,o_,is input at a number of selected values.

Having this information,table V illustrateshow the frequency-dependent coefficientdata are developedas a functionof frequency(mr reduced frequency, k). Longitudinal-sy n metrical coefficientsare obtained using the information of table V; lateral-directional-antisymmetric coefficientdata are obtained in a similarmanner.

Reflected in table V is the sign conventionbuilt into the Doublet Lattice Programat Rockwelland the sign conventionassumed for the equationsof motion of tables I, II, and III. The or_l°y differenceof importanceto the understandingof the derivationsof table V is the fact that lift force, L, of the Doublet Lattice Program is of opposite sense to the equationof motion normal force, N, and its associatedcoefficient,CN. Otherwise,the pitching moment, M, and structuralmode generalizedforces,Qi' are identicalin definition.

As an aid in understandinghow table V was assembled,the following example is given for the derivationof CN.

For no pitching (@ =@= 0 )

Vo Vo

therefore: qoSw CN _= - [Lh] h force units IMAG IMAG

v°2T CN= -

o!

IMAG INAG TABLE V. LONGITUDINAL-SYMVIETRICAERO COEFFICIENTS FROM DOUBLET LATTICE PROGRAM The bracketed quantity [ ] here and in table V comes directly from the Doublet Lattice Programas a function of frequency.

One of the questions to be resolved in using a modal approach to aircraft flexibility modeling is the one of how many modes to use. The technique used to help make this decision for the described analyses is discussed here.

First , the following criteria were developed as guides in the decision process.

The modes contributing to main a_roelastic impacts on all short-period and Dutch-roll characteristics were to be included. The modes contributing to main aeroelastic impacts on control effectiveness were to be included. The modes contributing most to flexible fuselage motion at the pilot station and active control sensors were to be included.

Figure 15 is typical of the data generatedto assist this mode selection process. For the longitudinal-syn_netric case used as an illustration, some 25 whole-aircraftnormalizedmodes were input to the Doublet LatticeProgram and run at a frequencyof n / 2 (a frequencyjudged to be in the region of short period and Dutch roll frequency). These data then were reduced to aerodynamic coefficient form and entered into a programwhich calculatesquasisteady flexible-to-rigid (F / R) ratios (refer to 'TruncatedAnalyticalModels") for all of the key aerodynamicderivatives (CN a andCn_ are used as illustrations here). These F / R ratios are calculatedas one mode after the other is elimi- nated. As a result, it is possible to identify the individualmodes contribut- ing most to a given derivative'saeroelasticimpact.

After the preceding procedure is accomplished for all important deriva- tives , plots similar to figure 15 are assembled and inspected as a whole. Thus , those modes making important contributions to all derivatives are selected for r et ent ion.

As far as fuselagemotion was concerned,as many modes as possible,having fuselagemotion as a main component,were selected. Usually modes reflecting up to the third fuselage-bending and the second fuselage-torsion mode could be selected.

• 2 - CONTROL-SURFACES INERTIA REACTION FORCES The inertia reaction forces of the control surfaces are importantinputs to the stabilityanalyses of active control systems. The detailed final form of these inputs for the B-I control surfaces are indicatedin the equationsof motion of the first part of this report. The basic approach that was used in developingthese expressionswill be discussedhere, but each control-surface input will not be developedin detail.

For constant Mach number and altitude l.O 0.8 CN o l _ _-- ....

F / R %, %..-_ 0.6 \ \

l

l

l

l 0.4 | l - . /i I / " L_....-- - - / CM_x 0.2 0 I * I | | I I a s i I I I I I I I I I I I I I a 0 4 8 12 16 20 24 Number of modes "L Figure 15. Typical aeroelasticflexible-to-rigid ratiodata for aerodynamic coefficientsas a function of participatingstructuralmodes.

Figure 16a shows that when a typical control surface is acceleratedin the positive sense of the deflection,a mass reaction force and moment are developedat the center of the mass of the control surface. This force and moment are reacted into the basic aircraft structureat the control surface hingelineas a force and moment as shown. The force acting on the aircraft is the force shown,andthe moment acting about the aircraft center of gravity is (_m_ + I )6" where _ is the distance between the control-surface hingeline _E and the aircraft center of gravity.

The generalizedforce acting on a typical structuralmo'dedue to control" surface accelerationis illustratedin figure 16b where the specific example of a s_etric mode is used. Before getting i1_to the specific example,con- sider the concept of a generalizedforce. A generalizedforce has units of work, that is m - N (foot-pounds). In this case, it is either force multiplied by mode deflectionat the force applicationpoint, moment multipliedby mode slope at the moment applicationpoint, or both. A positive generalizedforce would act to increasethe deflection (or virtualwork) of the structuralmode.

Looking at the example in the figure, it can be seen, for the example shown, that the reaction force causes a generalizedmode force incrementof -_i_Lm_ " and the reactionmoment a generalizedmode force incrementof +¢iILI_L6".

The precedingdevelopmenthas proved an adequate representation for small control surfaces such as the B-I SMCS controlvane and lower rudder control surface. It is not accurate enough,however, for large control surfacessuch as the all-movablehorizontaltail. In this latter case, it was necessaryto break up the mass characteristics of the surfaceonto a distributedgrid system.

Using the distributed:m_ssdata and the previouslydeveloped logic, generalized forces were developed for rigid body and structuralmodes. These data, de v el- oped using distributedmasses, were input to the digital programwhich imple- ments the equationsof motion of table II using an equivalentpoint mas.s representationof the data.

During the B-I development,checks of the pitch SCAS, with the aircraft on the ground resting on its landing gear, revealed a structuralmode-coupling instabilitywhen excitedwith sharp horizontaltail Control inputs. This • _- instabilitycould be analyticallyduplicatedby using generalizedcontrol- surface inertia reaction forces developedemployingthe distributedmass approach but could not be duplicatedusing the single-point mass representation.

The inertiareaction forces are particularlyimportantto the stability of SMCS with the aircraft on the ground. The inertia reaction forces have an opposite s_se to theaerodynamic forces of the SMCS vanes. The stabilityof the system is establishedby these aerodynamicforces. Thus, if the aerodynamic forces disappear,the feedback sense is effectivelyr_versed,producingan t 4O Reaction force and moment acting on primary structure at H_ Reaction force and moment about surface CG ° _ Primary I ' +(_ Aircraft H_

'I

structu , [ _l_ . _.._ l CG of contro l l l surface mass,m Force at aircraft CG = +m{_" Moment about aircraft CG = -(_m{ + IF L )i_" (pitch axis assumed for illustration) (a) Rigid-body generalized forces Symmetric mode i \ Reaction View from X 4 /// _ / _'(t s 'CG Reaction" force _iFL left side moment \_ 7L Z Note definitions of +(_li and +_iI'L.

q , 'm / / 5 . Structural generalized forces = - ( { _i H_

(b)Structural generalized forces Figure 16..- Typical control surface inertia reactiongeneralizedforces.

instability. On the B-I, a switch on the landing gear prevents operating the SMCS while on the ground, precludingany inadvertantdamage due to this potential instability.

ACTIVE GONTROL SYSTt_ / IS ..

Two types of active control systemswere included in the analysesper- taining to this study. One type, SCAS, is associatedwith control of whole- vehicle (short period and Dutch roll) modes of motion. The second type, SMCS, has the functionto control fuselage structuralmotion to improve ride quality.

The block diagrams and analyticalm o deling of the SCAS are given in figures 17 through 19 and S_ES in figures 20 and 21. Flight condition-dependent gains are shown for M = 0.85 at SL. These figures indicatethe type of sensors, compensations, gains, and actuatormodeling assumed for each of the indicatedsystems. The control-surface deflectionequationsare cast in a form directly usable by the Rockwell response analyses programs. That is to say, the overall gain is indicated,system dynamics are representedby numerator and denominatorroots of polynomials,and vehicle motions are defined as measured by the appropriatesensors.

COMPARISONSOF ANALYSES AND FLIGHT-TESTRESULTS The ride quality indices,_z and Hy (refer to reference 5 for detail definitions),for the vertical and lateral axes, respectively,cannot be measured directly in flight. One difficultyis the fact that the parameters are obtained from weighted power spectral density curves of crew-station accelerations,and the weighting can presentlyonly be done as a postflight operation. Secondly,it is very difficultto preciselyfly a specification vehicle weight at the specification Mach number and altitude. For the B-l, the approach has been to demonstratethat the analyticalmodel can duplicate flight-testresultsand then proceed to use the verified model in the required - _ ride qualityanalyses. It is the intent of this task to present the data matches which provided the verificationand give an evaluationof the factors affecting the matches.

The flight-testdata obtained for matching purposes were frequency responsesof load factors at FS 571.5(225)due to SMCS vane inputs. Both vertical and lateral load-factordata were obtained. The flight condition flown was M = 0.85 at 762 meters (2500 feet). The vehicleweight was approxi- mately 119 296 kilograms (263 000 pounds). Wing sweep was 65 degrees. Data were taken with all control systems inoperative; then, with only the SCAS • 8 p. A o R ,K i " E a

G, _o _ < p ( s_ .o4S+9 c o _ . a ( , 5'z.14. 1 o

k_ %, I'2.1 ., ,, L " _ 4,9 _ 0 1 AT _ + 57 ,14- S _| O F. S . 2 ; <. 49 5.k .

W | LTFJ _ i N OTCH M PEMSATI O SE P -. V O A C T UA _

: q H

RAD I N OR .HAL L __ GA IN / // 3 --__-- '2- ( ' Y_ ;% ' 7" ° _6_ r'i' 1 3 " 2 8 _ _ . S . 2' ¢ 4 9 mAD 1 19 9 < 3 . ; _ 3 "L_ ; =1 ..J F o r flight c o ndition o f M = 0.85 at SL Figure 17 . - Pitch axis SCAS analyticalmodel .

~ ~ Fl XEO F'OR ALL

g

J=LIGHT CONDITIONS 6.0

~ ..... ~ .44-

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• _I RA D For flight condition of M = 0.85 a t SL Figure 21. - Lateral SMCS analyticalmodel.

operating;and finally,with both SCAS and SMCS (cockpitgains 1.5) o perating.

The forcingamplitudeand frequencysettingswere set manuall y in the cockpit.

The measured data were processedto obtain both magnitude and phase charac - teristics.

The analyticalstructural-mode characteristics for the data-match analyses were obtained from test and analysis sources. The symmetricstructural modes were obtained directly from GVT. It was not possible to obtain a set of orthogonalmodes for the antisymmetriccase directly from the GVT. Instead, the analyticalstructuralmodel was adjusted in stiffnessuntil the frequency and mode-shapecharacteristics were as close to the observed characteristics as possible. In supportof these tests, the basic vehicleweight characteristics (no fuel) wePe identified. For the specificdata-matchinganalyses, fuel loadingswere determinedfrom measur e ments made during the flight-testperiod when the frequencyresponseswere being executed.

The pitch SCAS characteristics used in the analyseswere as described in figure 17; the yaw and roll SCAS were as described in figures18 and 19, respectively. The vertical SMCS descriptionis given in figure 20; and the lateral SMCS descriptionis shown in figure 21.

Comparisonof the analyticaland flight-testresults of the normal load factor of FS 571.5 (225) _ ' requency response due to synmetricSMCS vane deflec- tions (SMCSused as an excitationsystem) for the conditionof no active con- trols (basicaircraft)are sho_cn in figure 22. The three response peaks are (starting with the lowest frequency)first fuselagebending, second fuselage bending, and third fuselagebending, respectively;there is some influenceon the second peak from the wing second bending. The quality of the match is consideredexcellent. To obtain this match, however, three adjustmentswere made to the originalmodal characteristics: (I) the wing first bending-mode frequencywas reduced from 7.57 to 7.00 Hz, (2) the fuselage third-bending frequencywas reduced from 11.15 Hz to 8.60 Hz (thislatter change is substan- tial, and no reason has been found to explainwhy the original mode was off), and (3) the structuraldamping (gs) for the fuselage third bending was changed from 0.055 to 0.025. Having made these adjustmentsfor the basic aircraft .

responses,no additionaladjustmentswere made to the analyseswith control systems operating.

Figure 23 displays the data matches for the frequencyresponsewith the SCAS operating. Comparingthe first peak-responsemagnitudeof this figure with the previous figure, it can be seen that the SCAS excites this peak some.

Again, the analytical-to-test data match is excellent.

Figure 24 shows the data matches for the frequencyresponsewith both the SCAS and SMCS operating. The effectiveness o f the SMCS in damping the first fuselage-bending mode is demonstratedby these data. The data match is Note: 8cv is control surface deflection. The flight-test data measurements of the forcing command were analytically processed to remove effects of actuator dynamics,which were measured , in order to permit com- parisons with analytical results on this and similar subsequent figures.

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Phase

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ang]e 150 I \ , _\ deg / j / _ / ' S l I I I I I I _ // _ O_ / 2 4 6 8 _'_0

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% I Frequency (f) - Hz - 50: - " Flight-test data Analytical data g rad L 0 2 4 6 8 I0 Frequency (f) - Hz M = 0.85, air = 762 m (2500 ft) Wt = 119 296kg (263 000 lb), . A= 65 ° Figure 22. - Comparisonof flight test and analyticaldata, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off.

lOO !

Phase I

o

deg I angle ,50 / _ / ,__l_ 2 // 0 / " a _ , / 2 4 6 8 1 Frequency (f) Hz -50 Flight-test data Analytical data 12 - g , d 0 I l I I I n n 0 2 4 6 8 lO _ Frequency (f) - Hz M = 0.85, al t = 762 m (250 0 ft) Wt = 119 2 9 6kg (263 000 lb), A = 65 ° Figure 23 . - Comparisonof flight test and analyticaldata, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off.

5O 15 o Flight-testdata

)

SMCS vert gain 1.5

Analytical data 9

n z g 4 / \ \ 7 " * 0 '_'I " I I I I I , , a i 0 2 4 6 8 I 0 Frequen c y (f) - Hz M = 0.85, alt = 762 m (2500ft) Wt = ll9 296k g (263 000 Ib), .A= 65 ° Figure 24. - Comparisonof flight test an d analyticaldata, frequency responseof normal load factor at FS 571.5 (225) due to SMCS vane deflecti o n , SCAS on, SMCS on.

• 51 excellentfor the first two response peaks and fair for the third peak. The lack of better fit for this third peak was initiallyattributedto SMCS act- uator modeling in this higher frequencyrange; however, using the flight-test- derived actuatormodel of figure 25 did not substantiallyimprove the match to the flight-testdata in the 8 to I0 Hz frequencyrange as shown in figure 26.

The comparisonof the frequencyresponse lateral load factor at FS 571.5 (225)due to differentialSMCS vane deflectionsfor the basic vehicle is shown in figure 27. The comparisonis fair; the frequenciesof the three peak responsesare duplicatedby the analyses,but the amplitudeof the low-frequency peak is off by a considerableamount. The phase angle is matched reasonably well in the midfrequencyrange only; specificreasons for why the data do not match better are not known. The peaks are identified(startingwith the low- frequency peak) as the wing fore and aft mode with a large fuselage side-bending component,fuselage first side bending, and second fuselagetorsion. This last peak in the analysiswas obtainedby dropping the frequencyfor this mode from 9.72 to 7 Hz; although this mode was the only logicalone to adjust, there is no reason known for the noted discrepancy. Having made this logical adjustmentfor the basic vehicle description, no other adjustments were made when the control systemswere operated.

The data comparisonof figure 28 are for the case of SCAS operating.

The data are similarto the basic aircraftresponse of figure 27. Again, the agre_nentof analysis to test data is only fair.

In figure 29, the frequencyresponse comparisonsare made for the case of SCAS and SMCS operating. The trend of the analyses and the flight-test data are similar; i.e., the first and second peaks are attenuatedbut the peak around 7 Hz is increasedby the SMCS operation.

The implications of these analytical / flight-test data comparisonsare importantto the B-I ride quality verification. The data imply that the vehicle analyticalstiffnessand mass characteristics, whole-vehiclecontrol- surface aerodynamics, SMCS vane aerodynamicsand inertiareaction forces,and SCAS and SMCS modeling are fairly accurate. Thus, the ride quality charac- teristicscan be calculatedwith considerableaccuracy at specification or any other set of flight conditions. The Iongitudinal-sy_netric aircraft characteristics have been more accuratelydescribedthen the lateral-directional- antisymmetricset.

The aircraft ride quality characteristics have been calculatedusing the describeddata set and have been presented in reference I, pages 56 and 57.

Flight c o ntrol s i mulat o r Flight-test-derived frequency response .........

Analyti c al m o del a c tuator , K + 50 3.0- K = 1.93 deg / volt -154.4 (s-205.5) Fitted transfer function (S + 82.8 ± j95.6) 0 I I I I I I I I ! I 0 2 4 6 8 lO Frequen c y (f) - Hz 0 2 4 6 8 lO 0 _ J w I I I i i J , ' P ha se -20 , _ ' __'w"A A : angle ____A deg -40 -60 Figure 25. - Comparisonof analyticalSMCS actuator models t o flight and simulatortest data.

Phase 1O0 l angle k deg : _" 2 4 6 8 lO - SO Frequency (f) - Hz I "' -- Flight test data , ( 50 ------Analytical data _-_--_) actuator _cv 8 ......Analytical data, -154.4 (S +82.8 .+-J95.6) actuator, I n z (S-205.5) Test derived g rad I _*, 4 X' t _..

!

0 2 4 6 8 I0 Frequency (f) - Hz M = 0.85 , alt = 762 m (2500 f t) Wt = 119 296kg (263 000 lbs),A= 65 ° Figure 26. - Effect of test-derivedSMCS actuatormodel, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on.

F\ 3 00 / I l Phase 20 0 I x "" angle I \ / _._z,_.

I i \ 1 00 u I \ . deg / _ \ " 0 I . I * I I 2 4 6_ 8 J O -100 - Frequency (f) - Hz - 200 - _ " ....

Flight test data Analytical data m

In ]

g I rad , _1 _ I 4 I \ ; I _J I I I 2 I l

, \

/I _ 0 _ * a_ , a , * * a , 0 2 . 4 6 8 0 Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = I19 296kg (263 000 lbs),A = 65 ° Figure 27 . - Comparisonof flight test and analyticaldata, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off.

Figure 28. - Comparisonof flight test and analyticaldata, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off.

200 '

l o o _ r\

-. Pha s e 0 / "'_ - _- _ _ . , / , _'._ angle 2 4 6_ L ' / -lO

de_ -100 _re_uen_ If) H z "7" _ J

- 2 00 " " " " -300 Fl i ght test data ) data Analyt i cal _SMCS 1at gain 1 . 5 _Y I 6 [ _ cv I I I g rad 4 !

t I t I | I i \ 2 I i I I \ 0 2 4 6 8 0 Frequen c y (f) - Hz M = 0. 8 5, alt = 76 2 m ( 2 500 ft) Wt = 119 2 9 6kg (263 000 ]b),A= 65 ° Figure 29. Comparisonof flight test and analyticaldata, frequency response of lateral load factor at FS 571.5 (225) due to SMCS vane deflection,SCAS on, SMCS on.

FORWARD SMCS SENSOR PACKAGE RELOCATION On the whole, the SMCS has worked well in impr o vingthe ride quality o f the low-altitude, high-speedflight regime operatingat cockpitsetting gains of i.5 in the vertical and 1.5 in the lateral. The lateral axis performance, however, has been below that of the vertical.

As part of the ongoing investigation to determinehow to improvethe lateral SMCS performance,the data shown in figure 30 were obtained. The data shownare power spectraldensity (PSD)plots of the pilot stationlateral accelerationresponse due to turbulencemeasured during flight 1-20 while flying at M = 0.80 at about 305 meters (i000 feet) altitude. The data show that the SMCS, with the lateral gains set at i.5, significantly reduces the key peak response at 4.5 Hz, slightlymodifies a second peak at 6 Hz, and increasesthe response significantlyat 7 Hz. The net effect is an improvement in ride quality but not a large one. When the pilot increasedthe lateralgain to 2.2, therewas a dramatic increase o f the approximately7 Hz responseto a level which was felt by the crew to be not acceptableeven though the 4.5 and 6 Hz responseswere further reduced.

The comparisonof the flight-testand analyticallateralacceleration frequencyresponsesdue to vane excitationshown in figure 29 d_nonstrated that the anal y ticalmodel could reproducethe essence of 7 Hz increasedresponse.

A study of the analysis results indicatedthat the two peak responsesat 4.5 and 6 Hz are due to fuselage side bending componentsin these modes while the responseat 7 Hz is the fuselage second torsionalmode. Figure 31 will assist in showing how the responsephenomenonobser v ed occurs. Key in the analysis is the locationof the SMCS sensor packageat fuselagestation FS 571.5 (225), WL 142.24 (56) and BP 60.96 (24). Both the vertical and lateralaccelerometers are located in this package. As the vanes are differentially deflected,a side force and a torque are created. The lateral accelerometersees lateral accelerationdue to both the side force and torque. When the lateralaccelera- tion signalsare sent through the SMCS, the side bending-ir_luced signalsare properly phased but the torsion-induced signalsare adverselyphased, resulting in a reducedgain margin of the 7 Hz mode. It is also importantto note that the vertical accelerometeralso sees the torque-induced motion, and undesirable SMCS symmetricalvane motions are caused by lateralSMCS operation. Data, however, have shown this not to be a large influence.

Analyses shown in figure 32 indicatethat observed ad v erse torsion coupling,as well as the coupling into the vertical axis, could be eliminated or attenuatedby relocatingthe SMCS sensor package close to the fuselage centerlineand near the elasticaxis. Since there would be no lateralmoment arm, the couplingto the vertical axis would actually result in the torsional signal phasingbeing favorable.

SMCS off SMCS on, vert gain 1.5,1at gain 1.5 • o ... .. SMCS on, vert gain 1.5,1at gain 2.2 Flight 1- 2 0 M = 0. 8 0 , air = 305 m (lO00 ft) A= 65 ° Figure 30 . - Effect of SMCS lateral gain on p o wer spectral density of lateral load factor at pilot stationFS 746.8(294).

SMCS nz sensor package torsion I A n iIon at F$ 571.5(2251 ' Ytors old location ' 60.96 c £ -_ (24. in.), i142.24 c m __Estimated (56 in.) //' elastic axis _-_ 16 . 05 c m (6.32 in.)

I 9.

.62 in.)

I SMCS I SMCS vane : sensor package Resultant 1 at F S 515 , 6(203) For c e vectors " " due to vane new location deflection Looking toward rear from nose Figure 3 1. - SMCS sensor package locationsand couplingcharacteristics.

SMCS off 8 _k _ SMCS on - " ny 1 old sensor loc at i o n

I

lat gain 3.0 g 6 rad SMCS on - new sensor location 4 FI lat gain 3.0

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%

eO eS •If • O 2 4 6 8 lO Frequency (f) - Hz I. • I I I .... I I ,I 0 I0 20 30 40 50 G 0 Frequen c y (w) - rad / se c M-- 0.85, a lt = 7 6 2 m ( 2 500 ft) Wt = ll9 387kg (263 000 Ib), A = 65 ° Figure 32. - Effect of SMCS sensor package locationon lateral load factor at FS 571.5 (225)due to differentialvane deflection,analyticaldata.

The advantagesof moving the SMCS packagemay be sLm_narized as follows- (i) Higher lateralgains could be used to improvelateral ride quality.

(2) Higher effectivevertical gains (due to a farther forward sensor location)would improvevertical ride quality.

(3) The coupling of lateral vane inputs to the vertical axis would be eliminated.

As substantiated by both pilot co_ent and flight-testdata, the original SMCS considerablyimprovedthe ride quality. However, the gains in the system were limitedto values below those originallyintendedby the coupling described herein and not by system maximum capability. Based on the describedadvances, action was taken to move the forward SMCS sensor package to the new location at FS 515.6 (203),WL 19.36 (7.62) and BP 16.05 ( 6 .32). This relocationwas first accomplishedon A / C-I and then on A / C'2; this modificationwas never accomplishedon A / C-3. A / C-4 has the sensor package at the new location.

SMCS stabilitytests were initiallyperformedat high altitude to evaluate the forwardSMCS sensor package relocationeffects. The flight condition / configurationwas M = 0.85, altitude 6096 meters (20 000 feet) and weight 119 297 kilograms (263 000 pounds).

The First fuselagevertical bending-modedamping obtained from the pitch- pulse transientdata is shown in figure 33 and is seen to be a linear function of the S_4CS gain. All other modes were stable at the indicatedgain conditions.

The first fuselagebending-modedamping with the SMCS forward accelerometer relocated (flight1-41) is comparedwith the results with the accelerometer in its previous location (flight1-7). The mode damping appears to be signifi- cantly larger at the higher gains with the new sensor location.

The lateral bendingmodes were not stimulatedsignificantly by the rudder pulses, so that similardamping characteristicscould not be obtained. However, the lateral SMCS was stable for all values of gain tested (maxim_ cockpit knob settingof 6). Results obtained with the previous sensor location (flight1-7) showed the SMCS to be unstable at a setting of 6.

• Following the high-altitudetest, stabilitytests were conductedat low altitude. The initial flight condition / configuration was M = 0.85, altitude 914.4 meters (3,000feet), A = 65 degrees, and weight 119 297 kilograms (263 000 pounds). The excitationswere horizontaltail and rudder pulses.

First, each axis was tested to a maximum gain for that axis (V-gain= 3.0 and L-gain = 3.0, respectively). Followingthis, tests were conductedwith the Legend: "--O-----O-- _ Flight 1-41 Flight I-7 Mach = 0.85 Mach= 0.85 A = 65 degrees A = 65 degrees Air - 6096 m Alt = 6096 m (20000 ft) (200ooft) Wt = I19 297 kg Wt = 131 544 kg (2 63 ooo Ib) (290 ooo I b) Relocated SMCS Original SMCS sensor sensor 0.3

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: 0 ' !, , o 2 4 6 V e rtica l SMCS gai n (cockpit knob set) Figure 33. - First fuselage verticalbending mode damping versus SMCS gains for original and relocated forwardsensor package.

SMCS operating in each axis simultaneously. The maximum combined setting was V-gain and L-gain = 3.0. The qualityof the response data was such that the d_qlping factor could not be numericallydetermined;however, these qualita- tive data indicatedthat the SMCS was stable.

Followingthe horizontaltail and rudder pulse-stability checks, SMCS response to SMCS vane excitationevaluations were made. Figure 34 contains the vertical-axis-response data. Shown are - the normal load factor at FS 515.6 (203) (sensorlocation) frequencyresponsedue to SMCS sy_netric vane deflectionsfor SMCS off and SMCS on at vertical gains of 1.5 and 1.8.

These data show good performancefor the vertical SMCS and are not signifi- cantly different from the similar data of figure 24 for the SMCS forward sensor packageat the original location.

Figure 35 shows the lateral load factor at FS 515.6 (203) (sensorlocation) frequencyresponsesdue to differentialSMCS vane deflectionsfor SMCS off and on at lateral gains of i.5 and 2.2. As shown, the large 5 Hz (approximate) first fuselage lateralbending mode was significantly attenuatedat the higher gain without the previouslynoted large adverse torsional coupling response at approximately7 Hz.

Figure 36 shows the lateral load factor at FS 746.8 (294) (pilot station) frequencyresponses due to differentialSMCS vane deflectionsfor the SMCS off and on at lateralgains of 1.5 and 2.2. As on the previous plot, the large 5 Hz first fuselage lateralbending mode was significantlyattenuated at higher gains. The 7 Hz response,however, shows a slight increase in magnitude over the zero gain response.

Additionaldata similarto that just described,but at a higher weight condition,are shown in figures 37 and 38. These data indicate less attenua- tion of the 5 Hz mode and more excitationof the next higher frequencymode peak.

Rememberingthat the lateral nominal gain setting is i.5, all of these data indicateda substantialnet improvementin lateralload-factorresponse due to the SMCS.

To further check out the relocatedforward sensor package, tests were conductedat the off-designconditionat M = 0.55, altitude 762 meters (2500 feet),A : 55 degrees. Figure 39 presents the normal load factor at FS 515.6 (203)frequencyresponse due to sy_aetricSMCS vane deflectionfor SMCS off and on at vertical gains 2.5 and 3.0. These data show good perform- ance for the vertical SMCS at this off-designcondition.

SCAS only S C AS + SMCS, vert gain 1.5 ............... SCAS + SMCS, vert gain 1.8 .I I I I I , I I : 0 I0 20 3 0 4 0 5 0 , 6 0 Fr e qu e n c y (w) - r a d / se c Fit 2-20, runs 9 .1, 9 . 2 , 9 .3 M = 0.85, a lt = 9 14m (300 0 ft) Wt = 12 9 502kg ( 2 84 500 l b s), A= 65 ° Figure 34. - SMCS vertical axis performancewith relocatedforward sensor package, frequencyresponse of normal load factor at FS 515.6 (203)due to S_CS vane deflection,case i.

SCAS only SCAS + SMCS, lat gain 1.5 ............ SCAS + SMCS, lat gain 2.2 1 2 g r-_ / \ 4 / . _ .

/ m 0 2 4 6 8 10 Fr e qu e ncy (f) - Hz I , I I I I I I 0 I0 20 30 40 50 60 Frequen c y (w) - rad / s ec f M = 0.85, al t = 914m (3000 ft) Wt = 12 2 018kg ( 2 69 000 I b s), A= 65° Figure 3 5. - SMCS lateral axis performancewith relocated forward sensor package, frequencyresponse of laterall o ad factor at FS 515.6 (203)due to SMCS differential vane deflection,case I.

r SCAS 0nly SCAS + SMCS, lat gain 1.5 ............ SCAS + SMCS, lat gain 2.2 - - I 0 cv I 0 , . .--- '. • _" Ii ' ' " t. I I ! I O 2 4 6 8 I0 Frequency (f) - Hz I I I I I ' I 0 I0 20 30 40 50 60 "_ Frequency iW) - rad / sec M = 0.85, alt = 914m (3000 ft) Wt = 122 018kg (269 000 Ib), A= 65° l:igure36. S_S lateral axis performance with relocated forward sensor package, frequency response of lateral load factor at FS 746.8 (294) due to SMCS differential vane deflection, case i.

SCAS o_ly ' SCAS + SMCS, lat gain .75 SCAS + SMCS, lat gain 1.5 + + + + + +SCAS + SMCS, lat gain 3.0 1 2 CV I[ g rad |'I .

• ..

I .4 0 s i i I I I I I I 0 2 4 6 8 I0 Frequency (f) - Hz I I I I I I I 0 I0 20 30 40 50 60 Frequency (o_)- rad / sec M = 0.85, alt = 487.m (1600 ft) ' Wt = 139 709,kg (308.000 Ib), A = 6 5° Figure 37. - SMCS lateralaxis perf o rmancewith rel o catedf o rward sensor package, frequencyresponse of lateral load factor at FS 515.6 (203) due to SMCS differentialvane deflection,case 2.

SCAS only SCAS + SMCS, fat,gain .75 SCAS + SMCS, lat gain 1.5 + + + + + SCAS + SMCS, fat gain 3.0 lO ny .

.

6cv . .

6 . .

. .

g 4 rad .

4 _" 4 - 2 / , @ 0 ' I I I I I I I I 0 2 4 6 8 lO Frequency (f) - Hz I I I I I , I i "- 0 l0 20 30 40 50! 60 Frequency (0_)-rad / sec M = 0.85, a lt =487m (1600 ft) Wt = 139 709 kg (308 000 Ib), A = 65° Figure 38. SMCS lateral axis performance with relocated forward sensor package, frequency response of lateral load factor at FS 746.8 _294) due to SMCS differential vane deflection, case 2.

SCAS only SCAS + SMCS, vert gain 2.5 ....... SCAS + SMCS, vert gain 3.0 g rad 0 2 4 6 8 0 Frequency (f) - Hz I I I I I I I 0 I0 20 30 40 50 60 Frequency(_) - rad / sec M : 0.55, alt : 762 m (2500 ft) Wt = 132 451 kg (292 000 Ibs),A: 55 ° Figure 39. - SMCS vertical axis performancewith relocatedforward sensor package, frequencyresponseof normal load factor at FS 515.6 (203) due to SMCS vane deflection,case 2.

7 0 Figure 40 shows the lateral load factor at F$ 515.6 (203) frequency response due to differentialdeflectionof the SMCS vanes for the SMCS off and on at lateralgains of 2.5 and 3.7. As shown, the large 5 Hz first lateral side-bendingmode was significantlyattenuatedat the higher gains without the previouslynoted adverse torsionalresponse at approximately7 Hz.

" " In additionto the preceding frequencyresponse data, SMCS performance data in turbulencewere obtained to evaluatethe relocatedforwardSMCS sensor package. Figure 41 is a power spectraldensity plot of the pilot station (FS 746.8 (294))vertical load-factorresponsedue to ttmbulencewith the SMCS off and on at vertical gains of 1.5 and 1.9. As shown, the SMCS was very effective in attenuatingthe first fuselagevertical bending response.

The power spectral densityof the load-factorresponsewas normalized to unit root mean square (RMS) gust intensity COw_) as derived from the angle of attack (_ vane) data. This is n o t an accurate_technique but is the best available in absence of gust boom data. A similar normalization(OVg) has been att e mpted for the lateral axis data using sideslip (B vane) data.

Figure 42 shows the lateral load-factorresponseat the pilot station with the lateral SMCS gain zero but the vertical gains at 0, l.S, and 1.9.

Since there is r_ mechanism for the vertical axis SMCS motion to couple into the lateralaxis, these data indicatethe level of repeatabilityof the lateral data.

Figure 43 containsvertical load-factorresponse data at the pilot station with the SMCS off and on at vertical gains of 1.5 and 1.9, togetherwith lateral gains of i.5 and 2.2, respectively. Comparisonsof these data with the data of figure 41 show that little or no coupling of the lateral axis activity is evident in the vertical axis response.

Figure 44 shows the S_S performancein the lateralaxis at FS 515.6 (203)with the SMCS off and lateral gains of i.5 and 2.2. These data show that the 5 Hz (approximately 30 radiansper second) first fuselage lateral : bending mode is significantly attenuatedat both gain settings. However, at gain setting I.5, the second fuselage lateralbending-modepeak response near 6 Hz r_aains about the same in magnitudebut shifts slightlyupward in fre- quency. At lateral gain of 2.2, a significantincrease in magnitudedevelops with this frequencyshift.

Similar data to figure 44 for the pilot stationare shown in figure 45.

These data show the S Hz mode reductionbut indicate an increasedcoupling with gain increaseof the higher modes.

Attention is directed to the power spectraldensity scales for figures144 and 45. The data of figure 45 are a factor of I00 smaller than figure 44.

SCAS only SCAS + SMCS, fat gain 2.5 ........ SCAS + SMCS, lat gain 3.7 g rad 0 2 4 6 8 10 F re quen c y (f) - H z I, I ! a I m. / _ _ _ e I 0 l0 20 30 40 50 60 Frequency (O J )- rad / se c M = 0.55, alt = 762 m (250 0 ft) Wt = 129 276 kg (2 8 5 000 Ibs) , A = 55 ° Figure 40. - SMCS lateral axis performancewith relocated forwardsensor package, frequencyrespon s e of lateral load factor at FS 515. 6 (203) due to SMCS differentialvane deflection,case 3.

Run ll.l vert gain = 0.0, lat gain = O.0 ------ Run II.2 vert gain = 1.5, lat gain = 0.0 ........ Run ll.3 vert gain = 1.9, lat gain = 0.0 SCAS on 20 0 xlO_ 6 " xlO-5 160 _nz

12o g2

g2 (ft /sec) 2 (m /sec) 2 " rad / sec rad / sec 8 0 8O

|

40 40 0 0 0 10 20 30 40 5 0 6 0 Frequency (_) - rad / sec I I "1 I I • I :_ | , I I I I 0 2 4 6 8 I0 Frequency (f) - Hz M = 0.85, alto152 m(500 ft)AGL Wts 128 369kg(283 000 Ib),A = 65° Figure 41. - SMCS performancewith relocatedsensor, PSD of vertical load factor at FS 746.8(294)- pilot station, case I.

Run II.I vert gain = 0.0, lat gain = 0.0 .... Run 11.2 vert gain = 1 . 5, lat gain = 0.0 .......... Run ll.3 vert gain = 1.9, lat gain = 0.0 i SCAS on 0.28 xlO-6 .- o.28 x]O-5 0.24 0.24 0.20 O. 20 _ny

o.16

0.16 2 2 g g (f t-TTsec) 2 (m /--_ec) 2 "rad / sec rad / sec O. 12 o. 08 o. 08 I .

0. 1 2 !

0.04 0.04 \_ _1 _ / _\

V '

IQQ e - foe 0 0 I I I I l . - 0 ] 0 20 30 40 50 60 Freq u en c y ( u _)- rad / sec I I _L I I I I I • I I O 2 4 6 8 l O Frequen c y (f) - Hz M = 0.85, altu]52 m(500 ft)AGL Wtm 128 369kg(283 0 0 0 Ib),A= 65 ° Figure 42. SMCS performance with relocated sensor, PSD of lateral load factor at FS 746.8(294) pilot station, case i.

-- Run 11.1, vert gain = 0 . 0, 1at gain = 0.0 .... Run 11 . 6, vert gain = 1.5, lat gain = 1 . 5 ............ Run 11.7, vert gain = 1 . 9, lat gain = 2 . 2 SCAS on xl 2oo xlO-5 160 ....

160 _nz 120 g2 g2 (ft /sec) 2 (m/sec) 2 .rad / sec rad / sec 80 8o

'1

40 i

0 0 0 Io 20 30 4 0 50 60 = Frequency (_) - rad / sec ' I I I I I I I I'" I I I 0 2 4 6 8 10 Frequency (f) - Hz M = 0.85, altB152m(500 ft)AGL Wts 128 369kg(283 000 Ib),A = 65° Figure 43. SMCS performancewith relocatedsensor PSD of vertical load factor at FS 746.8(294)- pilot station, case 2.

Run If.l, vert gain = 0.0, lat gain = 0.0 Run ll.4, vert gain = 0.0, lat gain = 1.5 ....... ... Run ll.5, vert gain = 0.0, lat gain = 2.2 SCAS on - - 28 ....

28 x x10-5 % e Q 24" " ee ee ..

20 "' 20 _m. : ' __ L IZ

O _v : ;!

2 16 2 16 g (ft /sec) _ rad / sec 12 P, I (m /see) 2 rad / sec !il l i T,,, i

/ \ I: . _

, ; "...-_ - .-

': "' " "'"

0 0 I',-. I i .....

0 I0 20. 30 4 0 50 60 Frequen c y ((_) - rad / sec I I I, i_ I ,,, I , _ m .... I 0 2 ...4 _ 6_ . 8 I0 Frequency (f) - Hz M = 0.85, alt--152m(500 ft)AGL Wt =128 369 kg(283 000 Ib),A= 65 °_ Figure 44. - SMCS performancewith relocatedsensor, PSD of lateral load factor at FS 515.6(203)- sensor location,case i.

Run 1 1.I vert gain = O.0, lat gain O.O ------Run 11.4 vert gain = 0.0, lat gain 1.5 .......... Run II.5 vert gain = 0.0, fat gain 2.2 SCAS on 0.28 xlO -6 O.28 xlO-5 _0.24 0.24 eege ._ e • 6 0 . 20 : .':.

2 (_ 2 V • ". _ .Q ..

--g 2 - "'"

(m /sec)

rad / sec 2 -!I : 0.16 _ 2 _I| " ' ,ft/sec) ,_.i_; I " rad / sec :| [ • o.12 : J I : O. 12 ;I I • l %..: e

\

O. 08 \_- '

o

_:: o. 04 .._\ 0.04 : Y •e O O I 0 10 20 30 40 50 60 Frequency (_) -rad / sec - L_ I, I . , .A " • -_ .' i I !

0 2 4 6 8 lO Frequency (f) - hz M = 0.85, alt B152 m(500 ft) AGL Wt = 128 369kg(283 000 Ib), ! %-- 65 ° Figure 45. - SMCS performance with relocated sensor, PSD of lateral load factor at FS 746.8(294) pilot station, case 2.

Th i s l arge response differencebetween two fuselage stationswhich are relativelyclose is not reasonable. Figure 45 data are believed to be the data in error. Subsequentto flight 1-56, the accelerometerat the pilot stationwas found to be unable to hold a calibrationand was replaced.

Because the pilot station response is of prime importancein ride quality evaluations,it was requested that the data of flight 1-55 be rerun. This could not be accomplishedin the schedule;however, the data of this flight do have limitedvalue and are presented here for completeness.

Figures 46 and 47 are similar to the data of figures44 and 45, respec- tively, but with the vertical SMCS on togetherwith the lateral SMCS.

All of the data in figures 41 through 47 were obtained at an aircraft weight of approximately128 369 kilograms (283 000 pounds). Data similarto figures 44 and 45 are shown in figures 48 and 49, respectively,for a weight of approximately120 204 kilograms (265 000 pounds). Comparisonsof these data indicatemore adverse high-frequency mode couplingwith increasing lateral gains for the lighterweight configuration.

It is concludedthat the frequencyresponse data show a significant improvementin pilot-stationresponse due to relocationof the lateral accelerometer. The PSD data of the lateral response at the pilot station, however, still show considerablehigh-frequency mode excitation. Taking all evidence into account, the lateral SMCS still appears to provide a net lateral response improvementfor the nominal lateral gain of I.5.

TRUNCATED ANALYTICAL MODELS As previouslymentioned,the flexible analyticalmodel of the B-I used in SMCS design analyses has been describedusing nomalized vibrationmodes of the structure. The model used I0 symmetricand 12antiso_ymetric modes in most analyses. In de_ng a_]xrti.cg_ l_modgls tO supportre?ring base simulator studies, however, it was found that this number of modes caused computer equT_m-e_iYYequirements to become excessive. The challengewas to retain the accurate aeroelasticimpact on short-period and Dutch-rollcharacteristics and the main essenceof the structuraldynamic motion as seen at the pilot stationand SCAS and SMCS sensors.

Run II.1, vert gain = O.0, fat gain = 0.O III Run II.6, vert gain = 1.5, lat gain = 1.5 ........ Run ll.7, vert gain = 1.9, fat gain = 2.2 -- SCAS on _6 ' " ._ 28- x I0 -5 x 10 2O 20 _n y g2 16 (m /sec) '2 ft / sec)2 , \ r.

4 4 . i

• ° I e

0 0 l0 20 30 40 50 60 ' . , . Frequency (0 J i - rad / sec I I I I " I I , I I I ' -' 0 2 4 6 8 |0 J Frequency (f) - Hz M = 0.85, alt _152 m (500 ft) AGL Wt _ 128 369 kg (283 000 Ib), A = 65°.

Figure 46. - SMCS performance with relocated sensor, PSD of lateral load factor at FS 515.6 (203) - sensor location, case 2.

Run II.I, vert gain = O.O, lat gain = O.O ..m_ Run 11.6, vert gain = 1.5, lat gain = 1.5_ .......... Run ll.7, vert gain = 1.9, fat gain = 2.2:: • . SCAS on ;' . : 0.28 , , :: 0 . 28 x 1 0 6 ." : :: .

-5 :i!! "

xlO : :, .

I i • 0.24 : ::, 0.24 '.

0.20 jI : 0 20

_ j •

o.16 I I: 2°'16 g _ . I : g 2

2 (ft /sec ) _, : li

(m lsec) rad l sec I : ! i rad / sec O . 12 I ; _.

0.12

\-

\'.

0.08 0.08 1 _ij_il:

I: I-

0.0 / 4 0.04 I .

I! ,, I: %,G " Z "" • e

o o

0 i-O 2-_ 30 ; 40 , 50 60 .Freque n c, ) (o J ) - rad2._ec I I I I ,I I : I" _" I I ' - • 0 2 4 6 8 lO Frequency (f) - Hz M = 0.85, alt _ 152 m (500 ft) AGL Wt = 128 369 kg (283 000 Ib), A = 65° Figure 47. SMCS performance with relocated sensor, PSD of lateral load factor at FS 746.8 (294) - pilot station, case 3.

8O Run 11.8, vert gain = 0.0, 1at gain = 0.0 Run 11.11, vert gain = 0.0, 1at gain = 1.5 ......... Run 11.12, vert gain = 0.0, 1at gain = 2.2 SCAS on 28 x .........

-6 "- 28 x lO ,.

-5 ": I0 " 24 !_ 24 _ '.

y • I i!

2 16 g (ft /sec) (m/sec)2 rad / sec rad / sec 12 t |.

8 \

8 :\

0 10 20 30 40 50 60 Frequency (m) - rad / sec I I I I I I I I I I !

0 2 4 6 8 lO Frequency (f) - Hz M = 0.85, alt _ 152 m (500 ft) AGL Wt _ 120 204 kg (265 000 Ib), A : 65° Figure 48. - SMCS performance with relocated sensor, PSD of lateral load factor at FS 515.6 (203) - sensor location, case 3.

Run 11.8, vert gain = O.0, lat gain = O.0 0.72 Run 11.11, vert gain = 0.O, lat gain = 1.5 x 10-5 0.64 _. .......... Run 11.12, vert gain = O.O, 1at gain = 2.2- x 10 -6 SCAS on ..

o e 0.64 "" e o e • 0 . m • o 56 "' 0.56 - • • O.48- :'.

: • 0.48 - • : • • : • 0.40 - ,. : 0.40 - _ny 2 0.32- g (m /sec) 2 Io

I

rad / sec 2 0.24 - 1: - !_ 0.24 - g 1: ::

(ft /sec) 2 I!

rad / seCo.16 'li :_ :

o. ] 6- I: I

j ;

I : k

0.08 - 0.08 \ : 0 _ 0 _ I I I 0 10 20 30 40 50 60 Frequency (_) - rad / sec I I I i I i I l I I I o . 2 4 6 8 l O Frequency (f) - Hz M = 0.85, alt _ i52 m (5O O ft) AGL jWt = 120 204 kg (265 000 lb),. / L= 65 ° Figure 49. - S_S performancewith relocatedsensor, PSD of lateral load factor at FS 746.8 (294) - pilot station, case 4.

l ) ynamic analyses Were conducted to identify the key modes contribut.[ng to dynamic motion at the pilot station (the SMCSsensor was nearby) and SCAS sensors located at the nominal center-of-gravity (CG) l o cation. The technique will be illustrated using the longitudinal-symnetric case. rn this instance , the normal acceleration frequency responses at the pilQt station and SCASsensors , due to excitation by the horizontal tail , were - employed as criteria to judge the degree of accuracy achieved with truncated models.

Thus, given a flexible aircraft describedby rigid-bodymodes plus a number of structuraldynamic modes, the structuraldynamicmode set is truncatedwhile none of the quasistaticaeroelasticeffectsof the eliminated modes are lost.

The data used to modify the aerodynamic derivatives are generated by the method described herein and are identified as F_ / Roratios , or [ IF / [ IN" The approach to generating the F / R ratios is as follows: (1) Select sufficient modes to represent accurately the dynamic characteristics of the real system.

(2) Assume those modes not selected for the dynamic simulationto be quasistatic.

(3) Excite the quasistaticmodes with the aerodynamicloadings associated with the dynamicmode, S control-surface displacements, and rigid-bodydisplace- ments. The loads picked up in each mode are" determinedby solving the simultaneousmodal equations.

(4) The solutionsto the precedingequationsprovide the information necessary to calculatethe F / R ratios used to correct the aerodynamicderiva- tives of the rigid-bodyand structuraldynamicmodes selected for the simul ation.

The equations to be used in the example are as follows: Rigid-BodyPlunge and Pitch Modes T

Symmetric Structural Mbdes, i = 1 to n

Before approaching the details of defining the specifics of the truncated simulation, a brief description: of the basic P/Rtechnique is discussed. If all of the n structural equations were eliminated but the aeroelastic impact on the short period retained, the equations would appear as follows: Rigid-Body Plunge and Pitch Mbdes The F / R rati o s st_wn, where all 10 structure m o des are i l wolved, are o bta ined as f o ll o ws. The set o f structural m o de equati o ns are set up as sh o wn bel o w and _1 thr o ugh _1 0 are s o lved f o r each o f the in d icate d unit l o adings o f _ , (_w / 2 V o ) & , (Vw /2V o )q , and _ , respecti v el y .

] i l i I i ! I I I I I. I , , ' i As an example o f how the F / R rati o s are developed,consider the derivative CN_. From the normal force equationwhere _ = 1.0 and all other rigid-body variablesare zero, the followingrelationshipis obtained.

nl through nl O were o btained from the simultaneoussolution o f the previous e_uation set-_or a loading Cni_ for a unit value of _.

The expressionis reformed to obtain the F / R ratio.

The F / R ratios are obtained in a similarmanner.

The extensionto the pitching-momentcoefficientsfollows the same line of logic.

Consider now the longitudinal-symmetric equationswhere three structural modes (I, 3, and 5) of the i0 are dynamic, but the quasistaticflexibility effects of the eliminatedmodes are retained. The bar over the coefficients indicatesa modificationdue to the quasistaticeffectsof structuralmodes 2, 4, and 6 through I0. TypicallyCN = ([CN] / [CN]) CN, where CN is for a rigid vehicle.

Rigid-BodyPlunge and Pitch Modes Symmetric structuralmodes, i = i, 3, 5 For mode 1 (typical o f modes 3 and 5 also) • The F / R ratio correctionsfor these equationsare obtained in the follow- _ng manner. Ass_ning that structuralmodes I, 3, .and5 will be dynamic, the simultaneousequationsto be solved for unit loadings appear as follows: The F / P , ratio data for CN and Cm coefficients, given the solutionof these equationsfor unit loadingsof _, (_w / 2Vo)&, (_w / 2Vo)q, and 6, are obtained as previouslyexplained. The F / R ratio corrections, to CNni, CN_i, Cmni, and Cm_. are obtained as follows: l where 72, 74 and 76 through 710 are obtained from the simultaneoussolution of the previous equation set for unit 71 loading.

c_,z J c_,<, a nd

c ._ 3 c _ s

are obtained for unit loadings of 73 and 75, respectively.

m m Ccn q' i C m_ s , a nd -----. C m _$

Cm_Z i C ro w I Cm_ / ,_

,r " are obtained using the pitching moment___elationships and the 7i solutionsfor the unit loading of 71, 73, and _5' respectively.

The data are obtained in a similarmanner for unit loadings of (_i / Vo), (_3 / Vo), and(_5 / Vo) .

' llm ratios for modifying the coefficients of the dynamic structuralmodes are obtained using the n2, n4, and n6 throughql0 responsesof the simultaneous equations for unit loadings of _I' _3' 775'(_i / VO)' (_3 / V°)'and (@5 / Vo).

Typ i ca 1] y :

C n,. _. c_, n2 4+C_,n .7?. C n, n,o_, O

-- %,_, = I 4-

__I c n , n , Where n2, n4, and n6 through nl0 are from unit nl loadings.

Where n2, n4, and n6 through nl0 are from unit (_l / VO) loadings.

It can be shown that all of the quasistatic aeroelastic information is in the truncated modal equations by using the following logic for a typical.

rig id-body aerodynamic co efficient.

A B f (F / R)3-mode modified X (F /R )7_mod e quasistatic = (F /R )10-mode quasi- system system static original system " The followingset of numbers for a test case of the technique illustrates the accuracyretained in the coefficients.

S F / R F / R Coefficient from A from B CN_ 0.752514 0.752519 CN& I.015199 1.015201 CNq 0.661091 0.661086 CN_ 0.620383 0.620382 CM_ 0.602100 0.602126 CM& 0.9484 20 0.948413 Q_ 0.732947 0.732943 CM6 0.588629 0.588754 Figure 50 illustratesthe degree of quasistaticand dynamic aeroelastic informationretained in the three structuralmodes plus correctionsset of equations. Shown is the frequencyresponse plot of the normal acceleration at pilot station and CG. Note that at zero frequency,the truncatedmodel response overlays exactly the original 10-modemodel response, indicatingthat all of the quasistaticinformationof the 10-modemodel has been retained in the truncatedmodel. Furthermore,the truncatedmodel dynamic response is a good representation of the 10-modemodel dynamic response.

Truncateddynamic equationsmay be developedfor the lateral-directional- antisymmetriccase in a manner similar to that shown here for the longit u dinal- sy_netriccase.

ANALYSISOF ,SMCSVANE AERODYNAMIC INTERFERENCE EFFECT During the developmentof the SMCS vane configuration, wind tunnel tests were conductedto determine the aerodynamiccharacteristics of the vane.

Fairly extensiveaerodynamicinterferenceeffectswere observed in the force and moment data for both the longitudinala n d lateral cases during component build u p tests. Reference 1 contains some of these wind tunnel data and analysisof their sources; refer to this reference as backgroundfor the material to be presented here. It is the purposeof this section to report the results of an analyticalstudy made to assess the importanceof the vane aerodynamicinterferenceeffectson the dynamicsof the aircraftresponse.

9O 0 10 20 30 40 50 Frequency (_) - rad / sec lO structural modes, 120 1 _H unsteady aero 1 _-_I 1 3 structural modes + corrections, pilot / _H quasistatic aero rad 4O 0 10 20 30 40 50 Frequency (_) - rad / sec M = 0.85 S.L.

Medium weight, A= 65° Figure 50. - Normal load factor due to horizontaltail frequencyresponse con_arisonsof full and truncateddynamic analyticalmodel, SCAS on.

The analyses conducted were made in the frequency domainwherenormal and late r al load-factor responses at the vane station were calculated due to SMCSvane oscillatory deflections at various frequencies. The main reason for doing this was that flight-test data existed for these frequency responses against which to check the reality of the interference effects modeled.

Another reason was that the basic analytical model was available to conduct this study; only the vane aerodynamic characteristics had to be reworked to inciude the aerodynamic interference effects estimated. The following para- graphs describe how this was done and discusses the res L tlts of the analyses made. Finally , an evaluation is made of these interference factors relative to the importance to future similar designs.

Discussions of reference 1 identified general areas where the forces caused by vane deflections were acting. As shown in figure 51 , they were on (1) vane itself, (2) forebody , (3) wing-aft body , and (4) empennage. As a simpi :i fication , it was assumed that these forces acted at point locations in each of the general areas identified. These point locations were determined from the wind tunnel test data of forces and moments for various stages of configurat ion buildup.

The SMCSvane alone force and moment coefficient data were determined analytically (reference 1) while the associated interference force and moment coefficients (CN , Cm , Cy , Cg , Cn) were determined directly from the wind-tunnel data. The structural mode generalized forces were calculated for each mode knowing the forces acting at the points defined previously and the structural mode deflections at these points (F¢i). However , before any of these data could be added , the transport time lag effect from the vane to the point of load impact had to be considered. In the frequency domain , this lag effect was included for each interference load by mutfiplying by -1

vo

e whereX is the distance(+ aft) fromthe vane stati o n to the pointin question, is the forcing frequency, and Vo the velocity.

TableVI showsthe rigid-body and structural generalized forcecoefficient fonmlationfor the SMCSvane,including intereference effectsfor the longitudinal-symmetric case. The SMCSvanelateral-directional-antisymmetric data,including interference effects, were assembled in an analogous manner.

The frequency response data of figures 22 through 24 and 27 through 29 are repeated here in figures52 through57. The normalloadfactorresponse data of figures 52 through54 showthatthe interference effects modeledproducea -- Horizontal tail Vane Wing-aftbody Forbody + a L , + a R (a) Normal forces Vertical tail Wing-aftbody

J

Forebody (b) Lateral forces vane + _R' - _L Figure 51. - Typical vane-inducedinterferenceforces.

TABLE r1. - S\ICS \":\.\E EFFECT I\ E\ESS 1\CLUD1\G 1\TERfERE\CE EFFECTS, LO\G1TUD1\.-li.-SY~I.'IETR1C C\SE, FRECUE\CY DO~l-U\ 150 f_,_\ - f %

, oo: . -7 , "" . ,\

angle 5G de g Phase i / / "..!_ / '_ /// " :_ _.

/ i I I I ° I I i • .._"-i • --

o L ? 2 : 4 ! 6 8 . "lo

• / _, ' : " . I -50 _'I_ l " "..°,i_ ' Frequency (f) - Hz "I Flight test data Analytical data, no vane aero interference .......... Analytical data, with vane aero interference 12 P, n z : :" o . a % •

I " "

8 " : g ": : •

:_\ _

:'" - ' :

/// -,_. /

•" JI "- _- - " -" -" "

,, ..' /" . . ._,:, € --- /

• • °" / , ' ° . - _. J :; - ,_ _.. _ # - . ..

0 ""_"_"I _'_i_'" I I I "" I I I I I ' 0 2, 4 6 8 10 Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = I19 296 kg (263 000 Ib), A= 65 ° Figure 52. - Effect of S_S vane aerodynamicinterferences, frequency response of normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS off, SMCS off.

15 ° Flight test data Analytical data, no vane aero interference .............. Analytical data, with vane aero interference 12 . .."

nz .. :: e • o • _CV ' " ' ' • , e 8 " " "i :' rad 4 • oe° • •

" , . . ' "

0 I I" " I I I I I i 0 2 4 6 8 lO Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = ll9 296 kg (263 000 Ib), A= 65 ° Figure 53. Effect of SMCS vane aerodynamicinterferences, frequency responseof normal load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS off.

•.C, .",x Phase 100 n _ ,, .-'" I °" -.

l .._. . "| ."" _,q_, : ".

e oo • • _ • • -. ang l e • " , , ' i ; - " °° ' _X " t _. , * " *

deg 50 : I _ ' _ ' .\. : J / _ .'IV'"

• e 2 ; 4 : 6 8 "',. 0 "'" Frequency (f) - Hz "'_ -50 Flight test data Analytical data, no vane aero interference .............. Analytical data, with vane aero interference :" .: ..

e e l • • ".

12 ' ' : .

# e nz , , 8 . . . . : " : "_ g rad o°° 0 2 4 6 8 O Frequency (f) - Hz M = 0.85, Air = 762 m (2500 ft) Wt = I19 296 kg (263 000 Ib), A_= 65° Figure 54. Effect of SMCS vane aerodynamic interferences, frequency response of normal load factor at FS 571.5 (225) due to SMCS vane deflection, SCAS on, SMCS on.

Flight test data -- Analytical data, no vane aero interference ................ Analytical data, with vane aero interference 4OO 300 !"_ " / ', : \ .

2 00 l _"" - _',r..--. , ,, , , _.

angle P hase I00 / " J/ "_ I _.

_; F requency (f) - Hz deg 0 " - ' ! I I • _ l a,. .. I I I 1 2 ; 4 6 ..... . . _ v L - W _ 0 -200- l O0 " % %_ ._ _ ..........j ow I i i l I lO j i o

I

rad k / Ii: I: t i I: _

;. . . ..;.- • "_, a- .- _ " • n I I I

00 - "-- . .

2 4 6 8 0 Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = 119 296 kg (263 000 lb), A = 65° Figure 55. - Effect of SMCS vane aerodynamic interferences, frequency response of lateral load factor at FS 571.5 (225) due to SMCS vane deflection, SCAS off, SMCS off.

Flight test data Analytical data, no vane interference ............... Analytical data, with vane aero interference g rad • ooS • •_ ' i _ . _ _ 0 2 4 6 8 0 Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = I1 9 296 kg (263 000 Ib), A = 65° Figure 56. - Effect of SMCS vane aerodynamic interferences, frequency response of lateral load factor at FS 571.5 (225) due to SMCS vane deflection, SCAS on, SMCS off.

Flight test data .... Analytical data, no vane interference ...... Analytical data, with vane aero interference ""____ - ,_ lOO "_--_-_ _ Phase _ / - I lb,_ _.

angle- 0 _'______._ , / . .__'_,., 0 deg .... 2""_" 4 6_ _ ' _ !Frequency (f) - Hz_._ -100 .

-200 % "- -300 .... " ' lO o, g rad _I _.

t 2 _ _ _o _ 0 2 4 6 8 I0 Frequency (f) - Hz M = 0.85, alt = 762 m (2500 ft) Wt = If9 296 kg (263 000 Ib), _A= 65 ° Figure 57. Effect of SMCS vane aerodynamicinterferences, frequency response of lateral load factor at FS 571.5 (225)due to SMCS vane deflection,SCAS on, SMCS on.

I00 significant degradationof the analyticalto flight-test-data matches in amplitudeand phase. A similar,but less severe degradationof the lateral load factor analyticalto flight-testdata matches due to inclusion o f the interferences is indicatedin figures 55 through 57.

Why better agreementswere not obtained of analyticalresults (with inter- "" ferenceeffects included)to flight-testdata is difficultto assess. First, however, it is to be noted that the interference data were obtained using staticmeasurementtechniques,whereas the flight-testdata were dynamic. The vane on the 0.036-scaleforce model from which most of the componentbuildup data were obtained produced small force readingsof questionableaccuracy for this type of study (momentdata were judged more satisfactory). The 0.1-scale forebodymodel was judged to produce acceptableforce and moment data. This, then, implies the need for more expensiveand larger scale models tested, using dynamic techniquesin order to support accurate analyticalvane aero- dynamic interferencestudies. The analyticalto flight-test-data comparisons without interferenceeffects displayedin the referencedfiguresare quite good, particularlyfor the normal load factors,and suggestthat such an expensiveapproach is not warranted to obtain satisfactory aerodynamicdesign data. A key conclusionreached as a result of this interferenceanalysis is that a static 0.l-scaleforebodymodel is adequate for obtainingvane aero- dynamics data that includethe significantlocal fuselage interferenceeffects; this model produced the basic data used in obtaining the best analytical to flight-testdata match displayedin this report.

IMPACT OF SMCS ON SELECTED LOADS BACKGROUNDOF USING SMCS IN THE B-I FATIGUEANALYSIS Because the SMCS system was designed to be fail-safe,not fail-operational, the original B-I structuraldesign conceptwas that the aircraftwould have full structuralintegritywith or without the SMCS. That is to say, the structuralloads for both design and fatigue analysesof the air£ramewere : to be done with the SCAS operative and the SMCS inoperative. The B-I SCAS was designed as triple-redundant, fail-operational, and (for the gust loads analysis) fully effectiveat all times.

During the B-I design development,studiesof the expected airframe loads were accomplishedusing the availableweight, stiffness,and control systems data as it evolved. Several years into the program, it was found that filter changes in the symmetricSCAS were having a considerableimpacton the forebody fatigue spectrum. Further analyses indicatedthat operatingthe SMCS mini- mized the effects of the SCAS changes and gave forebody loads of a more consis- tant magnitude. Also, the lower statisticalload le v els computedwith the I01 SM(_ ;let ire were thought to be more representative of those that would be see, i_ serviceusage. In view of these facts and with Air Force concur r ence, it wCJs decided that the B-I synmetricfatigueanalyses for the low-altitude imnetrationmission segmentswould be accomplished with the SMCS operating.

l : atigue analyses for all other flight conditionsand all design limit load analyses continuedto be performedwith only the SCAS activated.

GUST LOADS ANALYSIS DESCRIFFION As with the ride quality analysis describedearlier, the dynamic gust loads analysisperformedwas a generalizedmodal analysiswhere the equations of motion were solved in the frequencydomain. This analysis used 14 modes of motion: plunge and pitch rigid-bodymodes and 12 symmetricfree-freenormal structuralmodes. Also includedwere two active control surfaces: the hori- zontal stabilizerand the structuralmode control vane.

For the B-I aircraft fatigueanalysis,the missions for the expected serviceusage were each divided into segments. Flight conditionsrepresenting the mission segmentswere selected for analysis. For each flight condition, static aeroelastictrim loads (to be used as mean load levels) and gust dynamic response incranentalloads were computed. Both the trim loads and gust incrementalloads were issued for fatigueand fracturemechanics analyses as distributedgrid loads. Along with the gust loads were the estimatedtime to be spent at the flight conditionand a graph of the expected load exceedances per hour flight.

Aeroelasticloads for the steady-flighttrim conditionswere computed using wind tunnel-derived nonlinearaerodynamicdata for the rigid and theoreticallycomputed aeroelasticincrementloads. These static aeroelastic calculationswere performedwith structuralinfluencecoefficientsfixed at a point near the aircraft CG.

The gust response equationsof motion that were used are presented as followsin matrix notation (ref. 6). See the appendix for symbol definitions.

Frequencyresponse functionsof the generalizedcoordinates,q, were obtained by solving the equationsfor a unit sinusoidalgust inputat each of i00 equally spaced frequencies.

GUST L(I_DS STRUCTURALMDDEL A complete aircraft structuralmodel was derived from structuralflexibil- ity influencecoefficientsgeneratedby the B-1 InternalLoads Group using finite-element methods. These SIC's, which were used for both static and dynamic analyses,were for partiallybuckled skins so as to be representative of the aircraft stiffnessat the design-limitload level. On the lifting surfaces,the SIC points were arranged in streamwiserows to make it easy to compute the required slopes and deflectionsfor the aeroelasticanalyses.

Figure 58 shows the arrangementof SIC poin_ts, and table VII presents the structuraldegrees of freedom used in this model.

TABLE VII.- GUST LOADS MODEL STRUCTURALDEGREES OF FREEDOM Motion Type Component X Y Z @x @y @z Fuselage, fixed wing and fairings 57 Nacelle, includingengines 4 6 17 2 2 2 Moveable wing 4 4 58 llorizontal stabilizer 2 45 Total = 203 Z i0 I0 177 2 2 2 Normal elasticmodes, free from rigid-bodyplunge and pitch constraints, were computedby the method of reference 7, as shown in the followingequation.

where: -i T '. GUST LOADS AERODYNAMICS Aerodynamicgeneralizedforces were developed using the Doublet Lattice method with the w i ng-fuselage-horizontal-tail combinationrun at one time.

l : igu r e 58 presents a diagram of the aerodynamicpanel arrangement. Geometry adjustmentswere made to the theoreticalmodel to improve the correlation of the zero-frequency stabilityderivativesand pressure distributions with wind-tunneltest results. One particularadjustmentwas to leave a small gap at the root of the horizontalstabilizer. At each Mach number, the width of the gap was adjusted to get the best match of stabilizerCL_ and spanwise loading. For the SMCS vane, the aerodynamics were run separatelyusing the vane modeled with 30 boxes (5 chordwiseand 6 spanwise) in the form of a symmetricinvertedV. These data were correlatedagainst the averagewind- tunnel test data. Generalizedvane forces due to modal motion and gust were then added to those from the wing-fuselage-horizontal combination. No attempt was made to simulateany vortex wake produced by the vane.

Deflectionsused to define the structuralshape for aerodynamiccalcula- tions were selected from the SIC points available in the gust analysismodel. • It was found to be necessaryto review the shape of each mode to insure that the deflectionschosen defined a smooth shape with no sudden changes or rever- ses in slope. Points eliminatedrepresentedconcentratedmass it_ns and fairing-nacelle deflectionsthat were not representative of the true wing-body streamwiseshape.

Generalizedforces due to modal, vane, and horizontal-tail motions and a unit sinusoidalvertical gust were computedat zero frequency and eight frequencies between zero and approximatelyII Hz. In the process of computing • the aerodynamics,the downwash inductionmatrices generatedwere retained on magnetic files so that they could be used again with differentmode shapes, thus saving considerablecomputingcost.

At each of the i00 solution frequenciesrequired,generalizedaerodynamic forceswere obtained by spline-curvefits of real and imaginaryparts of the generalizedforces computed at the nine frequenciesas previouslydescribed.

To improvethe quality of the spline interpolations of the gust forces,the ° gust referencepoint was transferredto a point just forwardof the aircraft aerodynamiccenter. This transfer tends to flatten the curves; i.e., reduces the rate of oscillationof the functionswith frequency. After fitting, the gust reference point was transferredback to the nose of the aircraft. For convenience,the gust reference point is normallly placed at the nose of the aircraft so that when the solution frequencyresponse functionsare used to compute time historiesdue to a discrete gust input, the penetrationof the gust starts at time equals zero.

1.04 Fuselage station centimeters 500 lO00 1500 2000 2500 3000 3500 4000 4500 I I I I I I I I I I Inches 200 400 600 800 1000 1200 1400 1600 1800 Component Panel No. of boxes ____ _'._'_.. ' _ "_ *I* / Forebody 1 6 _I_ Body + wing stub 2 40 Outer wing 3 6 Outer wing 4 30 *SIC points used to define aero shape Wingtip 5 12 +SIC points not used to define aero shape Horizontal 6 6 Horizontal 7 36 Y motion SIC points -- FRL . t t , T . . + o Figure 58. Loads analysis SIC point locations and Doublet Lattice geometry.

l : inaladjustments were made by direct ratioing to bring the generalized forces in the rigid-body modes, due to rigid-body motions and gust, into agreement with the a v ailable wind tunnel measured values of CL_ and CM_.

I:acto r sde v eloped in the region of zero frequency were applied to the full range of solution frequencies.

GUST LOADS CONTROL SYSTI_ V _ Descriptionsof the active control systems (ref. 8) used in this gust analys[s are presented in figure 59. Also shown is the method used to form the transformation matrix, IT], for control-systems feedback,relating control- surface deflectionsto the generalizedcoordinates.

LOAD METHOD DISO J SSION Gust response loads were computed at selected structuralstations. (See figure 60.) The mode displacementmethod was used at each frequencyof solu- tion to compute distributedloads at the SIC points. Shears, moments, and torqueswere then computedby finite s L _mnations of the loads or the loads tilnes the appropriatemoment arms. Figure 61 presents the matrix equations used to perform the load computations. The mode displacementmethod finds a set of net external loads that are equivalentto the loads required to hold the structure in the deflectedshape attained (ref. 3). To obtain accurate results, a linear superpositionof the elasticmodes that are used in the solution must give a good descriptionof the structuraldeflections. This requ i resthat at least severalmodes of vibrationsthat are primary to each structuralcomponentbe included. Modes selected should not be restricted in frequencyto the maximum Fourier frequencyof the solution,but should be chosen to obtain all significantcontributionsto the structuraldeflections.

Although the final loads for structuralanalysiswere issued as grid or mass point loads, it was necessary to compute shears and moments in order to keep the ntmber of items used in the load calculationsto a manageablesize.

On each structuralcomponent,a referenceaxis and load stationswere estab- lished for computingthe shears,moments, and torques. The LOAD GEOM matrix containedone row for each load item. Basicallythe LOAD GEOM row elements _Jre: for a shear, ones in the columns required to select loads outboard of the load stationaxis, and for a moment or torque, the arms from the load station or referenceaxis to each required load point.

To improvethe accuracy of the computed shears,moments, and torques, each load acting on a SIC point was consideredto he a pressure evenly distributedover a load box around the point. For any load box cut by a load stationaxis, only the box area outboard of the axis was consideredin Symmetric SCAS at SL, Khp -- .36 Compensation Notch Low pass Servo Actuator Degrees Symmetric SMCS for M = .85 at SL, K_o= .565 Notch Shaping Washout Actuators Degrees The matrix representing the control systems in the solution equations was formed as: Gains x polynomials Mode defl.ectionsand slopes evaluated at S = j0_ at sen_or locations - n _ c _ ec _ _ , l 4 1 -n z c _ 0 . n _ .c c uz / Where: CG = FS 2649 cm (1043 in), CV = FS 516 cm (203 in), _ ' = o t:igure 59. Control systems for gust loads..

FS 2367 FS 2896 oc (932 (l 40) FS 922 FS 1377 FS 1872 F S 2520 FS 3366 FS 4018 (1582) (363) (542) (737) (992) (1325) Horizontal tail pivot

i I

' I I ' Pivot Dimensions: cm (in.)

+S Forebody Aftbody ZFB , Y t FB I I _ i I II I I Figure 60. - Structuralstations for gust loads analysis.

I c Point load at each mass point, by mode displacement method Calculation of loads ; shears, moments & torques N L I _I N L I j N _ I Acceleration at point i NMjI Pitch rate at point i

=

i " Control surface deflections Where: N = Number of mass points NM = Number of modes, rigid _u = T ( _J) ta ) NEM = Number of elastic modes _ V( _){ NLI = Number of S, M, & T load 2. aI Z _KI M NM_I iterns Figure 61. - Load calculationequations.

10 9 computing t:he load. The ._ment and to r sion a r ms for this-load were taken to tile centroid o1 the outboard area.

The partitioningof the load boxes was computermechanizedby subdividing a box cut by a load stationaxis into I00 small boxes (i0 equal division on a side). Area ratio and centroid coordinatesof each subbox were computed using as data the coordinatesof the load box corner points. The area ratio was the ratio of the subbox area to the load box area. Coordinatesof each subbox centro[dwere then checked, and all those outboard of the load stationwere retained. The retained subbox area ratios and area ratio moments were accunu- lated to obtain the total element values for the LOAD GEOM matrix.

Other load items computedwere load factors at selected stations,pitch F a te and acceleration, and control-surface deflections.

F r equencyresponse functionsresultingfrom these load item calculations we r e computer plotted for visual review and also saved on magnetic files for use in computingthe load exceedancecurves and in-phase component load cond i t ions.

GUST STATISTIC A LLOAD CALCULATIONS For each load item, the turbulenceresponse power spectrtm, _o(_), r esponsequantitiesA and No, and a frequencyof exceedanceof load curve were computed. The methods used were as presented in reference 9 and as follows: o 3600 V N - Ii0 The Von Karman continuousturbulencespectrumwas used with the gust scale L equal to 152.4 meters (500 feet). Also used were the sea, level vertical gust parametersfor low-levelcontour flying- Pl = 1.0, bI = 0.823 m / sec (2.70 ft / sec), P2 = 1 × 10-5, and b2 = 3.246 m / sec (10.65Tt / sec). In the cal- _.. culationsto computeA and No, the required integrations were performedfrom zero through the highest frequencyof the analysis,approximatelyii Hz.

• The load level that could be expected to be exceededone time per airplane life in each mission segmentwas determinedfor each load item. This was done to define a set of loads that could be used to develop distributedloading conditionsfor the fatigueanalysis. These load values were read from the e x ceedancecurve of each item at one over the total number of flight hours for the mission segment. By nature, all of these loads were positive numbers, and although they were consideredto representa load-cyclecondition,the phasing relationships between the loads were unknown.

LOAD PHASING The problem of load phasingwas solved by applicationof the correlation coefficientmethod to producewhat are referred to as in-phasecomponentload conditions. Reference i0 developedand used the correlationcoefficientto expressthe statisticalcorrelationbetween two gust response parameters.

o For this gust load analysis,correlationwas developedbetween all of the load items. Correlationcoefficients were not actually computed;but a matrix [B],the elementsof which were proportionalto Pij_i_j,was developed.

- Here i and j indicate load item numbers and, thus, the corresponding[B] matrix row and column numbers.

[B]= Real [H(_)] I_w(_)l [H_(_)I _] iii l'_ch column of [B] was then normalizedon the diagonal elementand weighted w_th the correspondingload item expectedvalue.

where: i=j Columns of the resulting [S] matrix represent load conditionswhere the diagonal elementsare the load item expectedvalues and the off-diagonalelements are the statisticallyin-phase componentsof the other load items.

CONDITIONMATCHING To develop distributedgrid loads that matched the in-phase component load conditionsthe mode displacement method was again used. As shown in figure 61, this method can be used to compute distributedloads for a given set of generalizedcoordinates. For this matching problem, the shears, moments, and torques at each load stationwere known, and the solutionhad to be made for the generalizedcoordinates. The load generationmatrix was defined as: Then the coordinatesand the shear, moment, and torque loads from the [S] matrix are related as: SLOADS = GEN LOAD] h _is equation is usually overdetermined, and the coordinatesare solved for in a least-squaressense. Due to the large differencesin magnitude of tile shears and moments involved,the solution results generallyprovide a poor load match. To improvethe solution quality, each load item was weighted so that it had equal magnitudeand, thus, equal significancein the solution.

. By multiplyingthroughby a diagonalmatrix of one over the loads, the following equationwas obtained.

1 = LOADS GEN

['°I h

Taking one column from the IS] matrix at a time, a general least-squares solution program was used to solve for the generalizedcoordinates. Distributed i_int loads for the in-phasematching conditionswere then formed by the mode d tsplacement method.

EXAMPLE CONDITION LOAD RESULTS Load results for a B-I aircraft low-altitudepenetrationconditionof M ' -- 0.85, altitude = SL, using sweep -- 67.5 degrees, and weight = 140 614 kilo- grams (310 000 pounds) are presented in table VIII. Also, presented in figures62 through i00 are plots of the load-itemfrequency-response functions, power spectrums,and load exceedancecurves for the items listed: (i) Wing, WS 985 (387.6), bending moment (2) Forebody,FS 1377 (542),bending moment (3) Forebody, FS 2367 (932), bending moment - (4) Normal load factor, nz, at CG - (5) Normal load factor, nz, at pilot (6) Delta (deflection) of horizontal stabilizer (7) Delta (deflection) of SMCS vane (SMCSon) 'I'AI_I,I! VIII. I,OAD COMPARISON, SMCSOFF VERSUS SMCS ON lixpected loads, increments from trim, one o_ccurrence per 1000 hours • M = 0.85 at SL, wing sweep = 67.5 °, wt = 140 616 kg (310 000 ib) SCAS on Loads l,oad :items SMCSoff SMCSon Shear WS 1341 (528) 38 330 (8 617) 39 149 (8 801) Bend. morn. 67 232 (49 588) 69 567 (51 310) Torsion ,' 3 243 (2 392) 3 246 (2 394) Shear WS 985 (387.6) 80 406 (18 076) 78 622 (17 675) Bend. morn. I 278 977 (205 764) 280 389 (206 805) Torsion } 11 977 (8 834) 11 413 (8 418) Shear WS566 (223) 123 958 (27 867) 117 842 (26 492) Bend. morn. 698 319 (515 056) 681 015 (502 293) Torsion _, 20 984 (15 477) 19.975 (14 733) Shear at wing pivot 171 710 (38 602) 155 567 (34 973) Roll morn. 649 463 (479 022 612 953 (452 093) Pitch morn. ' i 1 355 627 (999 864) 1 280 991 (944 815) Shear at IrFroot 58 681 (13 192) 60 189 (13 531) Roll morn. 237 236 (174 977) 238 159 (175 658) l'itchmorn. ,r 126 113 (93 017) 127 587 (94 104) I : B morn. FS 922 (363) 381 315 (281 245) 188 692 (139 173) FB mom. FS 1377 (542) 1 060 369 (782 092) 349 942 (258 105) I ; I_ morn. FS 1872 (737) 1 848 893 (i 363 680) 364 677 (268 973) I.B lllonl, tzS 2367 (932) 2 261 236 (1 667 810) 627 525 (462 841) AB morn. I-S2520 (992) 4 106 793 (3 029 030) 2 366 297 (I 745 300) AB mom. FS 2896 (1140) 2 438 061 (i 798 230) 1 I00 992 (812 054) AB mom. FS 3366 (1325) 983 160 (725 145) 603 152 (444 864) ~ Nacelle Sz 85 953 (19 323) 91 366 (20 540) - Anz at (:(; 0.956 0.949 An z at pilot 2.007 1.041 - horizontal , degrees 0.699 0.728 mode wine , degrees - 19.955 Stations, cm (in.)

Shears, N (lb) b loments and torques, N-m (lb-ft) 1.14 -. N-m Ib-ft m / sec ft / sec 14000 . , ! , 60 000 13oo o. I , L , ooo. !! ' ' 'II I I llO00. I I II ° I , 40 000 M 9 o00.

8000.

' 700o .

'0000. 6000, . ,

_ooo. , / \_

20 000 4ooo.3ooo . /f' _-j _' - \ I' 2000. I : 1000. _ . _ .

0 o 0 2 0 40 ' 6 0 180.00 140.00 1 120.0 0 ! 00.00

,oooo

80.00 K _ . oo \ \ P 40 . 0o \ _ . \ A s o ',, \ E -20. O0 \ " - 40.00 H 20.00 _ \ \ -60 . O0 - 8 0. O0 X -loo.oo \ -140 . 00 : : - 1 20.00 _ ! - 16 0 . 00 t - " - 18 0.00 20 q0 60 FREQUENCY, RAD IANS I SECOND Figure 62. SMCS off, wing bending moment frequencyresponse - WS 985 cm (387.6 in.).

(:_ ( ,b_ft )2

m / sec / ift /sec rad / m rad / ft I .OX lO + G 9 , \ ,o_ - ,.o x ,o. _ \ Ill ..........

........ _ I II ....... - 1 I I T r'- \ \ / I I zzz r _ _ z III ...........

P

V

,o , - , ! l fl ", I HII I .OXIO +05

I]H ..... '

I I I I 0 .02 .0_ . 0 6 S P A T |_ F REQUEN CYRAD / F T ' ' ' ' ' 2'0 0 . 04 . 08 . 12 .1 6 .

rad / m Figure 63 . SMCS off, wing bending moment response power spectrum - WS 985 cm (387.6 in.).

-. l.OXlO + 0 3 l .OXlO +02 l . OX l O.0 1 E X I. OXIO +00 C E E D A N C E S I .OXIO -0 1 / H U R l. OX l O -02 l. OXlO -o3 I. OXI O -04 z 40000 800 0 0 12000 0 160000 2 0 0000 240000 280000 1 3r r _NDING 14C)IHEN' T AT NING STATION 38 7 . 6 lb-ft • I ! i I " 1O0000 .200000• 300000 .400000 N-m Figure 64 . - SMCS off, wing bendingmoment exceedances WS 985 cm (387.6 inl).

N-m Ib-ft m / sec ft / sec 1 50000 140000 6 x 10 5 13oooo 120000 I10000 I00000 90000 4 x I05 i 8oooo TI V O 000 oU 60 000 E 50000 1" 2 × lO_ 4oooo 20000 , - ,

ooo , o ooo L / \

0 o_ " ----------_ _ _ _-------- 0 20 40 6 0 1 80 . 00 l_.oo 14o.oo "\ 1 1 20 . 00 I00.00 _.0o 6 0.00 i 40 . 00 \ H 2 0. 00 I A

_ o

-_o . oo I -20 . O0 t - 6 0.0 0 -8o.ooi -I00.00 - 1 20.00

- lqO.O0 \ I

-)oo.oo r _ k_ - 18 0.00

I

2 0 40 6 0 F R E QU E N CY, RA DI AN S I SE C O N D Figure 65. SMCS off, forebody bendingmoment frequencyresponse - FS 1377 cm _542 in.).

II_ Figure 66. - SMCS off, forebody bendingmoment response power spectrum - FS 1377 cm (542 in.).

l . O X l O +02 l. OX l O . 01 E X 1 . 0 X I 0 . 00 ¢ E E O A N ' C E S 1. 0X l 0 -0 1 / H U R 1. 0Xl0 -02 1. 0 Xl 0 -03 I. OX IO -04 4.0XI O +I 8.OXIO " F'0RE B DO Y BO _ [:)I N G _ A T FUS EL A GE STATI OXl_- t 2 I b-ft ' 06 ' 06 ' 6 0.4 x 1 0.8 x 1 1.2 x lO N-m "" Figure 67 . SMCS off, forebodybending moment exceedances - FS 1377 cm (542 in.).

N -m Ib-ft -. m / se e ft / sec _0000 300 000 _000 0 1 . 2 x l O6 e e oooo _0000 220000 M 2000O0 G 180000 0 6 A 0 . 8 x 1 _ ,6oooo 1 40000 1 20000 I 00000 20000 O- o_--_ " __ " _ - _ __ O 20 40 60 180.00 14 0.00 1 20.00 I 00.00 80.00 6o . oo \ 16 0 . 00 _ H 20.00 A S 0 E -20 . 00 _0.00 _ _ - I00 . 00

- 40.00 _ \

-1 20 . 00 .. - 14 0 . 00 -80.00 _ - 16 0.00 - 1 80.00 \ 20 4 0 6 0 FRE QU EN CY , RA DI A N S / SECOND Figure 68 . - SMCS off, forebody bendingmoment frequencyresponse - FS 2367 cm (932 in.).

Figure 69. - SMCS of, forebody bending moment response power spectrum - FS 2367 cm (932 in.).

- , I.OXIO +03 [1111 IItlllill iiii Jill i[iii ilil fill iilll IIitl ................... ,,, .....

I\1111111111111!111111111111111tll IIIII I_1111111111111!11111111111!1111 IIII!

1111111111111!111111111111111111 III11 . o× , o + O_ I lXl.l !ll !l! l! !1 l lllill !llil l l l Ill ,!!!!_iiiiiiiiiiiiiiiillliiiiiiiiii!!!!!

- lJlJ[_ illllll illlllllllllll i l i ill IIIII II1! IIIIIIIIIIIIIIIIIIIIIt1111 IIIII l iil J k li Illll llll lil l i l iil Ilillil lill t lllill N llllii l !iiliiilill l lil l lii t ll ll

. o x , o . o. ,, , , "I ]

IIII IIII I1\111 III! IIII IIII I IIit I jill t1111 I i IIIIIIIIIIklllllllllllllllllltlll t1111 X . OX l O +00 C )z() ( .............. II ...... II t, ' ' E lll : lkllllllll ; ]l ;::: l .... i D IIIIIIIIII_IIIIIIIIIIIIIIIIIII IIIII

< Il l A IIIIIIIt111_1111111!1111111111 IIIII

Il ii l i l i iliii i _' i i lli l iI ii 'li meD ' i ''' i _ . o× , o - O , II Il i l f l l lllt_ !ill I I I l l l lll lit :::::::i:i:::lill_llllll iiiiiiiii'iliii_ / ' IIIIIIII , IIIIII [ I_ ...... ,,,,, , ....

H ll : llllllllll li ll , \llllll li lll l l il IIIII 0 IIIIl! [lllllilllllil tll IIIII!11 _!111111111111IIIII

; I

i_iiii ii ii i ili ii; i IIIfi ,ill iilll liil llll I_III IIII IIII lllll .OXIO -07 !! ....

........................... iiii i ]lill]l ; lllll]llllllll]l]illll i _]l] '

,.ox,o-°, I!11111111111 I I! IIIIIIIIIII ilq

B.OX l O 1 . 6XlO' 2.4X 1 _ 0 6 F _ Y B__ J ' , ( 3INC_ I' I O H ENT AT F US_ A C _ . ,E S T A T ION g _ 1 b-ft . ° # !

1.2 x 106 2.4 x 106 N-m Figure 70. - SMCS off, forebody bendingmoment exceedances- FS 2367 cm (932 in.).

g g m / see ft / sec •0 7 000 . 06 _ 0 0 0.20 .06 o0 o .0550 0 •0 5000 M "04500 o . 15 _ . mooo N I .0_::)00 | T 03O00 Q O . 1 0 u E ._00 O . 05 . ol_oo _ .01000 W

ooo / . \

.00 5 00 0 .0 0 20 40 6 0 11 3 0.00 1 40 . 00 1 20.0 0 I00.00 80.00 60 . O0 16 0.00 H 2 0 .00 A S 0 E - 20. 0 0 - 4 0.00 P 4 0.00 _ - 6 0.O0 X,_ -80 . O0 - ioo . oo .!

- 12 0 . 00 \ - 16 0 . 00 - 180 .00 1"_

, ooo

" 0 20 40 6 0 - F REQU E NC Y, R A D I AN S / SEC O N D Figure 71. - SMCS off, normal load factor frequency response - aircraft CG.

[24 -3 lO ! .OX l O -04 -4 I0 0 . 02 .0 _ . 0 6 SPA T IA L FREQUENCY RAD / FT f I I I I I 0 0 . 04 0 . 08 O . 1 2 O . 16 0.20 rad / m Figure 72. - SMCS off, normal load factor response power spectrum - aircraft C G.

l.OX l O +03 i i i I l.OX lO + 0 2 \ \

,ox,o.°, I\

\ \

\

E X I.OX I O +00 \ C E - \ E o \ A \ N c \

E \

S 1. 0 X l O - 01 / \ \ H 0 \ u \ R l . OXlO -02

\

\ \ \ , 1. 0 Xl O - 0 3 \ I ,O Xl O -04 _ 0 .2 0 .4 0 . 6 0.8 1 . 0 1. 2 1 .4 LO A D F A C TO RA T AIRCRA F T CO Figure 73. SMCS off, normal load factor exceedances- aircraft CG.

1.20 g g --. m / sec ft / sec l . 2 . 35 ooo 1.0 . 3 oooo .25OOO 0.8 M A G .20000 N

0.6 I

DU . 15 000 E 0.4 • I 000 0 O. 2 . 05000 J L 0 .0 0 20 40 6 0 180.00 16 0 . 00 1 40 . 00 I 00.00 8 0.00 1 20.00 6 0.00 H 20.00 P 4o.oo!_ _.

s o "X A - 6 0. O0 _0.00 -80.O0 E -20.00 _ _ -mo.oo - 1 20.00 - 1 40.00 \ - 16 0.0 0 -180.00 " " 0 20 40 6 0 FREQUENCY, RAD IANS / SECOND Figure 74. - SMCS off, normal load factor frequencyresponse - pilot station.

rad_m rad / ft -_ I .OXlO+00 i'i i ........

!11 1I ........

t 1 111 I ,o° '_l "ll .....

,.o , ,o-o , i/lll

II i 11 I i| Ill i [I I l l l ___ I Ill t II]]l I , ! I II I -1 I\ . / : tll ....

1o ! 1 ! / i t11

o , !1/!

[ Ill : ..... "_- P

V _.... t-H_'_ --

- 2 " - 10 , _ I11 .... T-- , .............

° !!1 /! _____ - .... -

II I / ...........

ill _ -

=_ '3 III /

........ t-tl -- --------'---- ' -- 2 _---_ . _'-r_£ t I I

t ] I __z: ::__rz_: TI=_-,= -

o . 02 . 0,.t . o6 SP ATIAL F REQUEN(;Y RAD / FT I I I I I I 0 0.04 0.08 0.12 0.16 0.20 rad / m Figure 75. SMCS off, normal load factor responsepower spectrum - pilot station.

I.OXIO +03 l .OX l O +02 I .OXI O +01 E X l.OX l O + 00 C E E D A N C E S l .OX l O -Oj / H U R l .OX l O - 0 2 l.OXlO - 03 . _ I .OXlO -04 0 . 4 O. B i .2 1.6 2.0 2.4 2.8 L OAD FAC T OR AT PILO T S TA T ION Figure 76. - SMCS off, normal load factor exceedances pilot station.

i deg deg m / sec ft / sec .0 6 500 O. 2 0 . 06000 .05500 . 0 5000 O.15 .msoo #4 .04000 A O.10 I . 0 30o0 U _ .0,2500 O . 05 . 0, .5o0 .0 1 000 • 0 0500 0 .o _---..__ __ 0 20 4 0 6 0 180.00 160 . O0 1 40.00 I 00. o 0 BO. O0 P 4 0.00 "_ ! 20 . 00 _ > H 20.00 A S 0 E -20.O0 -40.00 -80. 0 0 -80 . 00 6 0.00 - 1 20.00 - 1 40 . 00 - I _ :: _).00 _,

, oooo / \ /

- I BO.O0 0 20 40 60 F R E QU EN C Y. R AD IANS / SECOND Figure 77. S_S off, delta horizontal stabilizerfrequencyresponse.

1.30 Figure 78. - SMCS off, delta horizontal stabilizerresponsepower spectrum.

I . O XIO+0 3 I.OX lO +0 2 I .O XIO+ 0 1 E X I .OX l O +00 C E E . D A N C E S l .OX l O - 01 / H U R l .OX l O -02 l . O X l O- 0 3 F£gure 79. - SMCS off, delta horizontal stabilizer exceedances.

1.32 -. N-m Ib-ft m / sec ft / sec 1 0000.

8000.

GA 6 000.

40 000 9ooo. I 1 30 000 M 7000. i l IN 5t700. _ / _ I0 000 2ooo. _ . .

20 000 U 4000": //' _-' / 3000. !_ 1 000.

0 o 0 20 4 0 6 0 18 0.00 1 20.00 I 00.00 " X 80.00 _°'°° \..... ", H _.oo \ \ A ,,, -2o. oo _ -40.00 - 6 0.00 -80.00 __ - I 0O . 00 -140.00 -1 6 0.00

\ 1

. . - 1 80 . 00 , 0 2 0 40 FREQUENCY, RADIANS / SECOND Figure 80. - SMCS on, wing bending moment frequencyresponse WS 985 cm _387.6in.).

N-m )2 lb-ft )2

(

m/.sec ( ft/sec rad/m rad/ft J • OX 10+ \ --I - - \ \ I.OXIO+

1/

-f- II -, - I' II o

--\ -

U T - - - P 1\ U T I OXIO+ P .

- o - \ - --.- - \., J .OXIO+ ~ - _ .. _- - -- 1\ - \ .02 .04 .05 o SPATIM. FRElU:t-.CY R/lD/FT 0.04 0.08 0.16 0.20

o 0.12

rad/m rib~rc 81. - SMCS on, wing bending moment response power spectrum - WS 985 cm (387.6 in.).

I. OX l 0- °3 i - - I .OX l O - ° 4 100000 2 1 _0000 3 0 0 00 0 e EN O mO . OM e _ T A T UmO ST ATIO N _3 7 . 6 Ib-ft I I I I l O0 000 200 000 30 " 0 000 400 000 N-m Figure 82 . - SMCS on, . wing bending moment exceedances FS 985 cm _387.6in.).

N-m Ib-ft m / sec ft / sec 40000.

3 5000.

1.5 x I05 30000 .

M _00.

A G 1 . 0 x 1 05 _ 2 0000 .

U

, , . o. J r , t

_ooo. / "h 0 . 5 x 10 5 IOOOO. j _ o 0 0 20 40 6 0 1 8 0.00 1 40.00 _, 1 20 . 00 l OO.O0 130.00 _, 40.00 H 2 0 .00 i_o.oo _ ',., A

\

°

-20 . O0 - 6° .00 - BO .O0 - I 00 . 00 - leo . oo \ -40.00 J - 16 0.00 _, - 1 80 . 0 0 -1 4 0 .0 0 i _ k 0 2 0 40 6 0 F REQU E NCY , RAD I AN S / S EC O NO Figure 83 . - SMCSon, forebody bending moment frequency response - FS 1377 cm (542 in.).

N-m lb-ft)2 ( ( m/sec) ft/sec rad/m rad/ft - .

l- - - - III I.OXIO+ V\ l- - -I-- --1- - I-

,I

-I- --

- --

I\y.

II II

1\

I.OXIO+ 1/

-l-- -1-- - ,J..., - - - - - o 1\ - - U T P U T \ p 1\ o \ 1.0XIO+ .- - 1-

-

1\ - \ 1\ I.OXIO+ 1\ ..

- - - l- I-- ~\I- - ''1,.0 .06 o .02 .04 SPATlM... F"RECUN:V fW)/FT I I . I I 0.04 0.08 0.16 0.20

o 0.12

rad/m I Figure 84. - SMCS on, forebody bending moment response power spectrum - FS 1377 em (542 in.).

I. OXI0+03 ~ , ° I. 0X10+02 1.0Xl0 +01 E X 1.0XI0 + 00 C E E D A N c E \ • 5 1.0XlO -01 / H U R 1.0Xl0 -02 !

1.0Xl0 -03 i I.ox10 -°4 i 100000 200000 ,.300000 FOF_EY3OY BO',OINGMOMENT AT _ STATION 542 lb-ft II I I I I 100 000 200 000 300 000 400 000 500 000 N-m Figure 85. - SMCS on, forebodybending moment exceedances - FS 1377 cm (542 in.).

N-m Ib-ft m / sec ft / sec" I 4 5000.

!

2.0 x 105 40000.

,.s × '°s . _oooo._°°"1 _ N G k:_O0 .

1.0 x 105 u _ 2oooo.

D 0.5 x 105 loooo.

0 o _ 0 20 40 60 IBO.O0 1 60 . 00 1 20.00 I 00.00 BO .O 0 60.00 4 0.00 14 0 . 00 L H 20.00 A S 0 E -20.O0 _0.00 -_. O0 - I 00.00 - 1 40.00

. oo 1

, oo ,\ / _.\

" -,_o.oo _., "\ -, _ . oo \ 0 2O 4 0 60 -" F R E QU E NCY , RA D IAN S / S ECOND Figure 86. SMCS on, forebodybending moment frequencyresponse - FS 2367 cm (932 in.).

Ib-ft)2 N-m )2

(

( ft/sec m/sec rad/ft rad/m I.OXIO+ , - f-- - - f- 1\ \ J 1\ \ II

/

I.OXIO+ - ~ f- - - - - -- oa 1\ I.OXIO+

\

- .. .

- - - - f- 0- .

- - ._.- o u \

"K 07

~ I.OXIO+ - - - U -1- - T -f- P o 1.0XIO+

\

f- -.

-- - - f- ~ -- \ - \ -.

~

.

r-

--

- - 0 .02 .O't .06 5PAT1~ FR£Cll.EN::Y RAD/FT - , I I I I '- 0.04 0.12 0 0.08 0.16 0.20 rad/m Figure 87. - SMCS on, forebody bending moment response power spectrum - FS 2367 cm (932 in.).

lila I .OXIO+03 l . OXlO+02 1 .0X I0 +0 1 E X 1. 0XI0 + 0 0 C E E O A N C E S 1. 0X l 0 -0 1 I N U R 1 .0X l 0 - 02 I .OX I O -03 I .OXlO -04 1 00000 200000 300000 400000 ..5000 0 0 600000 F g_ _.E _ OY B O'- D ING _ A T _ S T ATION 932 I b-ft • i _- I ..... I !

" 200 000 400 000 600 000 800 000 N-m ~ n Figure 88 . - SMCS on, forebody bendingmoment exceedances- FS 2367 cm _932 in.).

, g g m / sec ft / sec .08 00 0 _ .

o . 25 .o _o / !

•0 700 0 . 06500 O. 20 . o6ooo . [ Y 5 5 00 H A "0 5 000 I . 04 0 00 T u . 0 35 00 O. 10 _ .o_oo . 025 0 0

o. 15 N o ._o o [ % /

.o l o ooI "' x J O . 05 .o1 5 oo \ _ X .00500_ _ _ _ - 0 .0 0 20 4 0 GO 1 8 0. 00 1 40.00 1 20.00 I 00.00 8 0 .0 0 60 . 00 4 0 .00 160 .00 1 H 20 . 00 A -20. O 0 -4 0 .0 0 • "-60.00 -I0 0 .00 -80. 0 0 _N - 1 2 0. 00 -N . _ . ,,.

- 16 0 . 00 - 1 8o. oo \ N

-1_ o .oo _ %

0 _ 4 0 GO F R E Q UE N CY. RA D IANS / SE CO N D Figure 89 . - SMCS o n, n o rmal l o ad fact o r frequencyresp o nse- aircraft CG.

1.42

(~2

rad/ft rad/m -l- f- 1\ - ~ - - - I- f- l- f- -1 1\ -- -- ~

1\

1\

2 O I.OXIO- - - i- f- - .- -

r-

1\ -2 II.

.- i'

I\U !\

IA -- ~ o 1\ U T 1\1\

II ~ I.OXIO-

i- T P 1\ i- 5 - iI o , ~ \

\

1.0XIO- . - i- -- 1-. .. - - - -~ -1\ - - - - -- - - 1;< 1\ ~ -~ ~.

i- 1\ ..

-t-.~ i- " .... 1...- .06 .04 .02 o SPATlf.L FREa..EN:Y RAD/FT I I

o 0.04 0.08 0.12 0.16 0.20

rad/m Figure 90. - SMCS on, normal load factor response power spectrum - aircraft CG.

i43 I.OXlO .03 I.O XIO + 02 l. OXlO +0 1 E X l. OX lO + 00 C E E D A N C E S I. OX I O - 0 1 I H U R l. OX l O - 02 l. O Xl O -03 1.0X l O -°q 0.2 0 .4 0 . 6 O .B 1.0 1.2 1.4 - .

L O A _ )F A CT _ A T AI F_I : _- T CG Figure 91. - SMCS on , normal load factor exceedances- aircraft C G .

g g m / sec ft / sec __ 0 0 O.

.08000 .0 7 00 0 O . 10 . o3ooo

ooo / %

. 01 00 0 0 . 0 0 20 40 6 O I _.00 I 16 0 . 00 I 1 40 . O0 1 _ ' 120.00 I00.00 J t i 80 .00 i i 6 0 . OO 40.O0 p t l , i H _ . 00 A t i s o \ / -,,, E _ _ . 00 __ _.

-4 0 .O0 i ' \ --60 .00 -80. O0 - 1 00.00 -120.00 - 140 . O0 - I _.00 -1 8 0 .0 0 " O 20 40 6 O F R E OU E NCY , RAD I A N S / S ECO N O Figure 92 . SMCS on, normal load factor frequencyresponse - pilot station.

g 2 g 2 m / sec ft / -sec F- 1 0-1_ ' _ l . O X lO -02 / - 2 _ __ i 0 o y i ,OX l O_O3 P ~ _ U T P O I -3- / lo i.O X lO - 0 4 m -4 o . 0 2 .o q . 06 SP ATI AL FREQU E N C Y RAD IF T | I I I . . . I I 0 0.04 0.08 0.12 0.16 0.20 rad i m Figure 93. - SMCS on, normal load factor response power spect_Jm - pilot station.

I .OXlO+03

\

I.OXIO+02 \ l.OXlO +01 \ \ i \ J E X l.OXlO+00

[ \ '

E o \ A \ N c k E \ S 1.0Xl0 -01 / \ \ H

o \ u

R |.OXlO - 02 l.OXlO-03 i I . OXl O -°h 0 .2 0 .4 0 . 6 0 .8 1 . 0 1 . 2 1 .4 LO A D F A C TORAT P ILOT ST A T ION Figure 94. - SMCS on, normal load factor exceedances- pilot station.

deg deg m / sec ft / sec 0.1 0 - .o3ooo

o_oo A J i

.02600 O.08 .o24oo .

•OP_2 OO

/ / "

.02000 J 0.06 _ .o1 6 oo , N O .o l r_o : .

i . 01 400 0.04 u .oI_o , E . 01 000 .O O SO0 _ { 0. 0 2 . oo6oo , I • .oo4oo I .0_0 0 .o "'_ _ , T_ .. - .L 20 40 60 1 80.00 140.00 1 2 0 . 00 I00 . 00 130.00 16 0.00 _ 60.0 0 1 p - 2 0 .O0 / H A 0

o . oo !

[-_1.00 -40.00 -_.OO -60.00 -I00.00 - 14 0.00

_, _ . oo \_ -I_.0 0 \

-IBO. 0 0 0 20 4 0 F R E Q UEN C Y, RAD I ANS I S E CO N D Figure 95. SMCS on, delta horizontal stabilizerfrequencyresponse.

1.48 / deg _2 f deg ,_2 m / sec _ft / sec /

\ra--_m / -r_--_T-

lo-1 _L_'X -

, x_

!

:-___\

,OXlO_ I

o \ - -- _

v ...... _--"........ _ ,.o× , o -O_ / I

p ..... / \ 1o-5 s -- \ \ / \ I

, . o x , o -°_ / - i " T

............. _ .___-_- ¥, =___--_ _-,-,_ ,_.-. -

, o x ,oO, ! r.j.k

-- ' -_--_-t : 10-7- _ ..... I- f .OXlO -°8 _=1==! : _--- - 0 .o 2 . 1_ . _ . o6 , _ PAT I . _L F RE Q UENC Y RAD / F'T L _ l i I t 0 0.04 0.'08 0.12 0.16' 0.20 rad / m Figure 96. - SMCS on, delta horizontalstabilizer responsepower spectrum.

E X I.OXlO+ C E E A N C E J .OXIO- 5 I H U R 1\ I.OX1O- oo...TA ~'ZCNT~ ST.6BILIZER. CE~S Figure 97. - SMCS on, delta horizontal stabilizer exceedances.

m / sec ft / sec 1 .7000 " - 1 . 6 000 1 . 5 000 1.4 000 A I . 3 000 / _ 1 .200 0 1 ' 4 I. I 000 3 _ o . 9ooo 0.8000 g o.7ooo / I 0 . 4000

1 . oo o 0

0._00 I I 0 . 3o00 0 .2 000 O . I 000 _ _ _ 0 o 4 0 6 0 1 80.00 160 .00 _,_ IL='O.O0 _ " -., 1oo.oo _ \ 80.00 4 0.00 P H c_. o b 1 40.00 _ A

o

-;2 0 . O0 N_ -4 O . 00 -60.00 "_ -8 0 .00 _ - I 00 . 00 - 1 40.00 - I BO.O0 - 180 .00 - 12 0.00 _ ' _ 0 2 O 4 0 6 O FREQUENCY , RA O IANS I SECOND Figure 98. - SMCS on, delta mode control vane frequencyresponse.

( deg _2 (deg _2

\m l sec / \ft l sec /

rad / m rad / ft _.

I .OXlO.01 " lO1 I.OXIO +Do I0 I .OX l O- 0 1 U T P -] u l O T P s O I .OXlO-02 - 2 l.OXlO -03 -3 lO 0 .02 . 0 '-I • .06 S PAT I A L F EQUE I ', _Y ' R , _ / F T I I I I I f 0 0.04 0.08 O. 12 O. 1 6 0.20 rad / m Figure 99. - SMCS on, delta mode controlvane responsepower spectrum.

I. OXIO+0 3 I_ Ill; ....... [::::: 11 I ] I : !':_l[ : ill ] [llj Ijjlljjj ijll ' _i ll t lj l IIII .................... IIII ] 111 1tll I111 lilt 1111 IIIIIIII IIII Itll I!ll _ llil lillilli Illi iili lili

,.ox,o + o_ _t ! t ! tl I I li lt l i ii

_ ikii : _ iiiiiiii _; iiii iiii__!_iiii ...................... IIII I I _ 1 111111 1 1 I I I I I I I I I t lM II111111 IIII ]111 Ill]

t i l_ lltl ll ll_ ll,l Iljl t111

illl' liltilil l t l i I l il IIII

,.=o + O , I ll i ' Il i l li II I l ! II

llll TIlL z 1 111 ::iiii : :,.iii i ii!! .................

.................. III ] II II IIIII III I l l l I I I I IIII \1111 1 111 I l l ! Ill ! IIII TITT i _ lllill Il ll li t l I I I t

!!!! .... ,, !!

] :l l e IIII IIIi_III : - iiii .... : III .... I111 Ill I1"1 IIII IIII ]Ill o IIII II I ll\lll I1t l I lll III1 A IIII Illlt_ll lilt ti l l IIit n Illi Illill _ l fill II I I l i l t C

_ ,- °× , ° - ° ' ,.. ! I lli 1, I il! .i t

# ll : i iiiiiii [ },i"ii i i .... !!ii _:' iiii I I II I11 1 111 1 IIII I l l[ IIII IIII111! IIII IIII illl H lll l I lllll I I I t I t I l l l _. IIII 0 IIII Iit11111 \ till Itll IIII u I I II iliiilli ,i l l! lill lili R

,. o× , o - °_ Jill I1 ! 1 N I J ! ! I J

......................... li l : t r , , I III , _ i i ' ' i , ,\ , I I II 111 1 " ilil ........

lit l I t l l tlJ t l l! i \ It ll II ' " ' II Illi iilll lt i I t li _ illl l i

,.ox , o - °_ I I I li lt i il "N. ! !!

til l [ I tz l [t l Tltl ._ t l l l 1111 IllllI l t IIll l I lt _ i ll / I1 11 IIIII!ll IIII II[I _ IIII IIII IIIIIIII IIII t111 nb , .]ll Jill Illillll I l tl I l il t f't.,.]

'. IIII IIIIIIII IIII IIII III1"

,. ox , o - _ Ilil IIIIIIII III Jill III 7-_

4 8 12 1 8 20 _ " " D O - T A _ _ V ,N _E , CE_:_ E _S Figure i00. SMCS o n, delta m o de controlvane exceedances.

The tabular load summary results and the plots are presented for both _ICS off (gains = 0), and SMCS on. In either case, the SCAS was considered to be operating normally.

As can be seen in the summary table, the effect of the active SMCS was to substantially lower the fuselage forebody expectedbending moments. Also lowered were the aft fuselagebending moments. The most obvious effect of the &ICS, as can be seen in the forebody frequencyresponseand PSD plots, was to considerablysuppress the response of the fuselage first bending mode. At the same time, the operationof the SMCS caused increasedresponses in some of the higher frequencystructuralmodes. The effect of this increasedmodal activity, as seen in the fuselage and wing frequencyresponse plots, was to cause slight increasesin the wing tip, nacelle, and horizontaltail expected loads.

Although the effect of :theSMCS activitydramaticallyreduced the magnitude of the forebody fatigueloads for low-altitude penetration,the total forebody fatigue spectrumswere not so drasticallyreduced. The SMCS was not used during the high-altitude cruise mission segments. The large nu m ber of flight hours spent cruising and the higher wing lift curve slope at a 25-degreewing sweep, cause the expected loads to be relativelyhigh. Thus, with the reduction in the low-altitudepenetrationloads, the cruise conditionsbecame dominant in the forebody fatigue load spectrum.

SMCS VANE EFFECT ON INLET /ENGINECHARACTERISTICS ]he objectiveof this section is to briefly describe and summarizeB-I flight-testresults that identify effects of ingestingvortices generated by the SMCS vanes into the inlet. Although vortices generated by the SMCS vanes were ingestedfrequently,no engine incidentsrelating to operation of the SMCS were identifiedduring the approximately1200 flight hours accu- mulated to date. This program includedmore than 200 flightswith three aircraft and 29 engines. Wind-tunnelresults with sub- and full-scalemodels were summarizedin referenceI.

Portionsof specificflights were dedicatedto demonstratingoperational suitabilityduring aircraftmaneuverswith the SMCS vanes deflected. Emphasis was placed on exploringcombinationsof SMCS vanes deflectionangles and air- craft maneuversduring operation at Mach 0.85. Effectson inlet total-pressure recovery and engine-facedistortionwere measured by a 40-probe instrumentation matrix at the inlet / engineaerodynamicinterfaceplane (AIP). Instrumentation, i ncludingan automatic in-flightcalibrationsystem,and flight-testprocedures are described. Resultsare summarizedto doct:nent this B-I experienceas an aid to futureprograms employingsimilar systems. Nomenclatureassociated with this sectionmay be found in the appendix.

1 5,1 TEST DESCRIFFION The B-I propulsionsystem is arranged in two nacelles under the fixed portionof the wing as shown in figure I01. Each nacelle contains two inde- pendent two-dimensional(2-D) external compressioninlets and two General Electric F-101 afterburningturbofan engines. Relative locationsbetween - the SMCS vanes and inlet nacelles are also shown.

Most flight tests investigatinginlet characteristics with the SMCS vanes deflectedwere conductedwith the simulated,fixed-inletconfigurationshown : infigure 102. Ramp configurationfor both inlets consists of the initial two ramps set at 7 degrees. On the inboard inlet, the third ramp is set at 5 degrees. In the outboard inlet, the third ramp is set at 9 degrees. Small differencesbetween inboardand outboard ramp configurations reflect an attempt to maintain good performancecharacteristics during both subsonic and supersonicoperation. The movable cowl lip is shown in its normal, takeoff, and landing positions. Duct flaw area distributionsare shown in figure 103. Maximum flow area is based on an average third ramp angle of 7 degrees. Design flow area is shown for referenceand representssupersonic operation with a variable geometry inlet.

Inlet boundary-layer air is removed through porous surfaces on the second movable ramp, throat panel, and small regions on the upper and lower end plates.

The bleed air is collected in two compartments. The air exits from the forward compartmentthrough fixed louvers and from the aft compartmentthrough two- position doors. The doors are open above Mach L _l.4 and closed at lower speeds.

Aft bleed doors were closed during all tests with the SMCS activated.

A bypass systan operates at supersonicspeeds above Mach 1.4 to match the inletsupply and engine demand. The bypass doors open to compensatefor reduced engine airflow such as occur on a hot day or during low-power settings.

The bypass doors remained closed during all tests with the SMCS activated.

The SMCS vane configurationis defined in reference I. Flight-testpro- visions includeda black box located in the crew compartmentto drive the SMCS vanes to a fixed deflectionangle. From this position,the vanes could also be driven at selected frequenciesto oscillate +i0 degrees. Using these -- provisionsto generate full-scalevane deflections, flight tests were conducted to identifywake ingestioneffectsduring combined aircraftmaneuversand engine throttle transientsas measured at the inlet / engine AIP.

Figure 101. - B-1 flight test aircraft, wings swept 65 degrees.

: Throat fixed link-- linkBypass ___ ' _- f ime d Ramp fixed link Movable lip__." _:::::`_:_:_i_i_i_i_iiiii_ii!:iiii_!iiii!!ii_iii!iii_i;iiiiiiiiiii_iiiiiii_iiiiii (takeoff / landing) __ _ . _____ "- - " _ _._ _ 2: 2 :i : i:fi 22 : 2 :i 2 )i:r: 2: i:i:i:5: 2 !iii 2 1 2 11:i[: 2 :i:r:i:iiii:i 2 1: 2 : 2 i : i: 2 :i: :-_ 2 :!:f:!:i: 2 :i:£_: 2 !i:ii1_ : _:J:!:_ : 2 :! : _ii!ii_:_:i:_ : _:J:_ : 1!i_ii_i_: y :i:i : f:i: 2 :5:_: : :::: 2 :i:i:[:_ : i:i:i!_:[£_i_ 2 i_: Fixed throatY_ I °v:_::r I .o., Fixed b _--BLC 1-[ two-position doors Adjustable link (ground adjustable) Figure 102. - B-I air inductionsystem.

18oo Duct flow area distribution ll 000 ECI-8 Aerodynamic interface _ P lane_ l Movable lip in lO000

I

10258 cm 2 cm2 9258 cm 2 9000 (1590 in.2)(( (1435 in.2) 8000i 4-1 U "0 <[ :, 7OOO 500O 600C J 4ooc I i I a I _ I I I _ I I I l I l I i I 80 lO0 12o 140 16o 180 200 220 240 260 in.

I I I I I |__ I I I I I I 200 300 cm 400 500 600 700 Nacelle station (at duct _) ° Figure 103. - B-I inlet subsonic diffuser flow area distribution.

FLIGHT-TEST INSTRUMENTATION B-1 flLght-testaircraftwere instrumented to measuremore than 1600 parametersof which approximately700 pertainedto the propulsionsystem.

InstrL_entation at the inlet / engine AIP included40 dual-purposeprobes in both engines in the left-handnacelle to measure both the steady-state and dynamic componentsof total pressure. Probes were installed integralwith eight engine-inletguide vanes, each with five probes locatedat the center of equal areas. (see figure 104.)

Computed distortionparametersare sensitiveto errors in individual total-pressure measur_aents. Several techniqueswere employedduring the flight-testprogram to minimize these errors. High-response, differential transducersmanufacturedby Kulite were referencedto a duct static pressure upstream of the inlet / engine AIP to maintain signal resolutionover a wide range of operating conditions. The referencepressurewas measured by an accurate,digital, absolutetransducerinstalledin a conditionedcompartment in the aircraft. The reference syst_iwas constructedwith sufficientvolume and orifices to restrict rapid changes during aircraft and / or inlet transients.

Referencepressurewas sampled four times per second.

An in-flightcalibrationsystem was developedto update individualprobe calibrationsonce per minute throughouteach flight. A schematicof this system,using a three-waypneumaticvalve for each total-pressureprobe, is shown in figure 105. The valve alternatelysequencesthe back side of the transducerfrom the normal referencepressure (operatemode) first, to a calibratepressure regulatedto a pressure approximately5 pounds per square inch above the referencepressure (calibrate mode), and then to the same pressure seen by the front side of the transducer (zeromode) (hencethe name ZOC valves).

The calibrateand zero positionswere each held for 2 seconds, and the operate positionwas held for the remaining 56 seconds. In the zero position, the total pressurebeing sensed is r o uted to the back side of the transducer : through an infinite coil, approximated by a coiled line 20 feet in length.

This provisionwas necessaryto prevent reflectingwaves affecting the frequency response of the probes and was determinedempirically. Both coils and pneu- -" matic valves, the latter grouped in gangs of five, are shown installedon the engine in figure 104.

.....

Q\ o l guide vanes Total-pressure probe O View looking aft O 359° a 359° a Kul ite AlP probes I I 06366 06374 ·06361 .06321 0 0 44 0 ·06362 ·06322 314 ~'O:\ 314° ~'\.'O:\ / / ·06363 ·06323 "~'o">:,>'\. ~ .

" ~'o ~ ~ .

• ~'o :,>'" • ~'O ~ !'> • 06' • 06' ·06324 ·06364 • ~'o ~ • ~'O ~ • 06':;>J: • 06';)/' 06' :;>s ~ • '0"> s:.

• 06'", ;}/,> 6' • ~'O !,>'I> ·0 :;>J::'> .~ :,>'" ·06325 ·06365 • '0" 71° • 06'", /& • 6'.;. (9 .~'O .~ o ;$'.9 06' /.9 ~"'6: QIi,ll "" '" ~ ~ ~ ;}.,o ...

80° '() ~ c c '" '"

. c c

'" .

• c:> .

o

·

o o e 269 ; ..

'" ~ '" ~'I> ,'I> a u ~'O", ".'" ~'O"> s:.'" ·0 ·0 • '0'" ~ 'i. ~'O"C s:.~ • 6'.;.

. ~.;.

QIi,1l • '0'" ~ QIi'~'i .~ 'o-.r- "."

o 9'0 o~ '()o • • ~ '0'" _,,'0 • 6'.; • u,.>.

• .~ ~'O", :,>'1>'0 .QIi:l~ • 06' ;}.9 • 06' :9.9 ~ ~'r, •• <:)"0 • 06' ;}"'&

• 06'''.9. · .~

• Q\i:l~3 06' :;>.; / "06'~ ".9..> (9 / ;}"'6' :.> • O\i:l~l 224 Y6' 224° 13&0

" 063)6

• o\i:l~' \ \ 197 0 197° Engine Engine 2 B-] j nlet/ engine aerOd)l1amic Fjgurc 104. tntcrfacc plane instrumentation.

Fuselage Engine compartment I nlet - - -- -- I i -I I I l I I I I l I

z ocI ' ' m ' ! Ical ibrate Pneumatic I

Ca l + 15 psi I ' ! I _ . 'I actuator l uperate _ I Cal I . II ZOC valve I

R2 + 5 psi I I I i • I

i I I I I Ii l Zer o "_ i[ I II I I _.Infinite coil I Engine I I R2 . I I I airflow I static I ! .AIP D robe j I , ,-

-- I _ ....... , ,_ _ __j

Inlet l ' I I t__ "i I! I Ii

_- ..... S i%n-a l--c o--n dT t i-o ne'-r ............ I I 7.5v supply I

I P° wer I

I Crew station I _. [ i I Cal, zero _s I l Tape II.__R2>10 ps!a TM _-- filter I

rec°rdeff I CBW Operate. _x2__._.j I l HP F _._j_ - I

16 ips I L • Normalize

' ; R2_ O.4Hz I

A / D I LP fi 1ter, J 16 sps i _ filter,

0.4 Hz II 12 H zI_ l

,.--I' ' ii I "_J ILP [- _rans_u_er ' ' I

I I I L I MV multiplexer I _ Figure 105. - B-I flight-test AIP data acquisitionand signal conditioningsystem.

lhlet_natic valves were actuatedby a pressure source regulated to 15 pounds per square inch above the calibratedpressure. Provisionswere includedto vent trapped pressures within the valve to either the engine compartment (ambient pressure) or an engine-face static pressure, depending on the operat- ing condition. Vent pressures were required to be the lowest pressure in the system at all times. Selection of vent pressure was accomplished by a crew compa r tment switch. A pressurized nitrogen bottle was us e d as a pressure source for the actuate and calibrate tanks.

Signal conditioningfor the transduceroutput is also shown in figure 105.

l.hcb signal is initiallypassed through a variable amplifierthat is used to establisha full-scalerange for the steady-statecomponentat approximately 24 pounds per square inch. Zero output levels can also be biased to favor normal operation. Subsequently,signals are split to record low- and high- frequencycomponentsseparately.

l_w-frequency or quasi steady-statecomponentsare generatedby low-pass filterstl_at effectivelyeliminatefrequencycomponentsabove 0.4 Hz (3 decibel level). Signalsare sampledby a sequentialanalog-to-digital converter onboard the aircraftand stored as digitalwords on a tape recorder located in the crew compartment.

lligh-frequency or dynamic componentsare generatedby high-pass filters tlmt effectivelyeliminatefrequencycomponentsbelow 0.4 Hz (3-decibellevel).

Signalsare furtheramplifiedbefore passing through voltage-controlled oscillators,multiplexed,and recorded on the same tape recorder (different tracks)used to record the low-frequencycomponents. With 21 continuous bandwidth (CBI_) signalsmultiplexedper track, four tracks were required to record the 80 high-frequencysignalsfrom both instrumentedinlets.

I)uring the zero and calibrate steps, all filters are bypassed, and the total signal (minus the referencepressure) is recorded both as pulse-code modulation (PCM)and CBW parameters. Calibrationsequence is operate- calibrate-zero-operate. Transitionbetween the zero and operate positions iml_sesa step change in pressure differentialacross the transducerdiaphragm.

This has no essentialimpact on the low-frequencycomponentbecause of the low-passfilter. However, the high-pass filter responds to thisstep change and res_ a lts in a damping characteristic with a period of approximatelyi0 sec- onds. As a consequence,dynamic data are invalidduring this time, and data r eco r ds for detailed dynamic analysis are selected to circum v entany problems.

Typical output signals, includingcalibrationcycles, are illustratedin figures106 and 107.

Parameter identification (scale in counts) PT = (CTSOP-CTSZ) CA L + 6002 CAL = (6002-6631) / (CTSOP-CTSCAL) RATE .CTSOP I , ,I CTSCAL

_ _3_6 _, ....i_- - -_I / t-l, I 633

CTSZF_ , - r I I ....

"' ' " 7 "' _oPE

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I,I I._. .ii iiiiiii i i _ ii I.I • Il l I ll li,O I '_ ll l o g IO l I ll l o ll i 00 I l l SA P .1PLES SAMPLES Figure 106. In-flight calibration cycle, total pressures at aerodynamic interface plane, flight 1-5, 64 sps.

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In stmm_ry, particularemphasis during the flight-testprogram was placed on obtaining accurate,total-pressuresurveysat the inlet / engine AIP.

As described,these features included the following: (1) Differentialtransducerswith a floating referencepressure (2) Accuratelymeasured referenceand calibratepressures (3) Automatic in-flightcalibrationprocedures (4) Signal conditioningto obtain good resolutionof both high- and Iow-frequencycomponentsfr o m the same transducer F1ight-testexperiencehas shown that, as one might expect with all the p n eumatic and electricalconnections,the system required considerable mainte- nance. However, it is believed that system accuracy approachedlevels availablewith a well-constructed wind-tunnelprogram, and this seems to be confirmedby c o mparisonsbetween wind-tunneland flight - testresults.

TEST RESULTS A total of 5 flight-testhours were dedicatedto exploringSMCS vane effectson inlet / enginecharacteristics. Flight conditionsand operating variablesare summarizedin table IX.

Several computedparametersfrom the 40 total-pressure measurementsat the inlet / engine AIP are used to present test results. Recovery (PTI / PT0)is the average of the 40 low-responsesignals referencedto free-stream,total pressure.

Se v eral distortionindexes are used and were computed by digital tech- niques. High-responsesignalswere filtered to 62.5 Hz and sampledat 360 samplesper second to representa one-per-revolution engine-frequency response • (signalsfully attenuatedat 125 Hz) . Circumferential(IDC)and radial (IDR) distortioncomponentswere computed for each ring (eachgroup of eightAIP pressures at the same radii) and combined mathematicallyto form a fan stall-marginindex (IDL). This latter index is normalizedto stall-margin allocations,and thus, values of unity computed from the high-response signals identifydistortionlevels approachingdesign limits. Inlet distor- tion, (I_MAX - PT_,IIN) / PTAVG, was also computed. Record lengthswere in the range between 5 and 20 seconds. Scans resulting in maximum values of stall- margin index are generallyused to identify trends.

'I'ABI , Ii I X. - SMCS VANEEFFECTSON INLET / ENGINE CHARACTERISTICS, FLIGHTTEST INVESTIGATION Dedicated AlP flt time a instrumentation Flt no. (min) Objectives / accomplishments status 2-19 65 Initial flight-testinves- Referencepressure tigationswith oscillating transducermal- van e s at 1 Hz, vane angle function,no con- ±i0 degrees, u = 3-8 degrees, version to = 0-3 degrees, initial engineeringunits throttle transients 2-33 45 Static vane deflectionangles ZOC valves mal- in 5 degree increments, functionedon ± full scale, u = i, 3, no. 1 eng, data 5 degrees, 8 = 0-4 degrees, reduced only for no. 2 throttle transients no. 2 engine.

2-36 60 Repeat of flight 2-33 Leaks in refer - - ence system pre- cluded AIP data reduction.

2-37 55 Repeat of flight 2-36 Satisfactory 2-38 60 Max rate throttletransients Satisfactory with vane deflected 20 degrees combined with nose left sideslip - 2-42 20 Conductpushovers and pullups Satisfactory with vane deflected 20 degrees aTotal _5 hours Engine-face,total-pressurecontoursare used to illustratevariations in distortionpatterns. High-pressureregions (pressureshigher than average) are shaded on the contour plots. Low-pressureregions (pressureslower than average) are unshaded. The magnitudeof the differenceabove or below the average pressure level is defined by the number on the contour. The annular region is formed by protrusionof the engine bullet nose at the AIP.

Envelopes encompassingall combinationsof circumferential and radial distortioncomponentsare used to identifythe magnitudeof dynamid activity and to compare them to design goals. Two types of time historieshave also been found useful in documentinginlet turbulencecharacteristics. Analog strip charts of the high-response AIP signals help to identifywake ingestion during transientmaneuvers. Digital time historieshelp to define transient conditionsbased on computed parameters.

OscillatingVanes Initial tests with the SMCS operativewere conductedwith the vanes deflectedsymmetricallyat nominal angles of +I0 degreesand then oscillated +i0 degreesabout that mean at a frequencyof 1 Hz. Vane deflectionangle thus oscillatedbetween 0 and _+20 degrees; the latter representingfull-scale deflection. Note that with these procedures,the vane is being used as an excitorand induces some discomfortto the crew members. With the vane oscillating,aircraftmaneuvers involvingcombinedangles of attack and side- slip were performed.

Time historiesof aircraft attitude and vane deflectionangles recorded during a 30-minutesegmentof flight 2-19 are shown schematicallyin figure 108 during operationat Mach 0.83. Angle of attack was varied between 3 and 8 degrees; sideslip angle was varied between 0 and 3 degrees. Tests were con- ducted at positive sideslipangles only to ingest the wake in the instrumented nacell e.

Flight times where increaseddynamic activity (identified from strip charts of the high-response AIP instrumentation) could be definitelyattributed - to SMCS vane deflectionare indicated. Increaseddynamic activitywas also noted at other times; howe v er, effectsof sideslipand vane deflectionangles could not be definitelyseparated. Ingestionis generally restrictedto positivevane deflectionangles (leadingedge up) greater than i0 degrees in combinationwith aircraft sideslip operation.

One major advantageof oscillatingthe vanes is to producerecognizable wake ingestioneffects in the data. The 1 Hz vane oscillationbetween 1.5 and 18.5 degrees during a sideslipmaneuver is shown in figure 109. Time histories Leading edge up L_.Vaneosc'illation at l Hz - -":':: ; ' _:(.':-_- : :_ : " . .:- T -7_ T " : " T-.- T --F . - " ":- : :--- : . _.- : T:..' : _'-" . 'T, 7 :" "_'7 • • , ;. _ ": : -" ::"= . -:i- 20 . : :. : ;.:r.:: : . :.".-,..:.:: .... :.,_--_-....:::.r:-!::===== = ===== = = ======= = ====l ": .... : -: : . ).:! = ",.: :'".... :..... t' u_ ":':;" ' , - ' : : i .... I'; ; ": * " : .... :" " : ' " " ' " : " , I : " • . : . . : :.: 16 "::: ::-, .:.-...:. v .l; ::- i.,:,.::4_:.. ; : . .s...-:. _ :t:.::::. . ::t:.-:: : ;.] - :l...,.- ..:.:.:..:-.!. :.._: ":' ::::::L :!> _: t -," _i__;:.;_-I; t ; ;i : ::,':Ti:_!: : - : i_--:i :::;_,;_ .:_ :::L!":!:: O " _ _ I . . . . . . . . . . . . . . . . . * . . . . . . I . . . . ' I ] • -u t.::' t .! 4 . . :: -i. 8-!::j > :: 12 .[ : i: r:'16.J':.i: > _.20 -..: ---:24_- 28 -i::::: .... : ":':'.. :.'. ':._::: . ".':.::.": I !':':.:.I:': I _:.:Z.:I..:::.!..'::_.'::.'.'.::.: . !'"L:"":"::'""'_ _' -12 .... --:.-_T.:-_T."--:i'.'--T_-'g: ::':"i :" :-:;i:!.::: : :r..l;:::':.',.:.: : : : =:.'. : ; . :':::: : : i" : :;::""'" I-_'" ......::" " ::':i _" ....... ;'. ........ : . .' :_ :1"::'" • :1:'.='_. ""- -:I: ' . ': : ! : : *, :: :: ,1:"1 ." 1 . : I: " : ...... : 1.-" • ............... • -; --G- ; - _. _._ ........... _, ................ ,.I...I .................. -._ - •"-:- = 'r ' --': : : = -_-.-::-.. .......... .--g .... F. .... , ................. |..l ......

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>'_ _I_. = ....... , _ ..... inle t :2. : :___ . _i: , i ::--i.-:: : ;:- : Increased A-IP dyr ) activity _:-:.-----'_ • . i ] . : i.i due SMCSvane deflection "_ u " No.

•- _ ' inlet ............................... "..... ; __ _ l:igure 108. - AIS / SMCS investigation - £ 1ight 2-19, Math 0.83.

0 ,,, -

x3029

RHSMC 0 _ ] Vane - _- l Hz- I - SMCS vane deflection angle, clegrees --20 LI I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I I !

i . | - i Sidesl ip angl - .. -.'T '.I . _,degrees.

MlOl2 0 Beta ' _ - 5 I l I If I If I I II II I II I I I III I II I II Ill IIII I II IIII II 2400 , . _ , - AIP PcMdata Counts 0 0:53"40 Flight time- seconds 2000 I I I I I I I I I I I I I I I I ! 11 I I I I I I I I I 11 I I I I I I I 11 I I I I I I O 2 4 6 8 lO "_ : "-:- ...... Representative AIP C -B I .......... ; .............. '.......

o !

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L ...... ._.. ' _ ' --1 ....

= D6338 : _ High pressure > " F.

,_ 4-J m D6339 ' • ' ......:i_ _ Low pressure : _ , -,, _ i ",._- Figure 109. - AIS / S_S investigation flight 2-19, Mach 0.83 / 16 000 feet, = 3 degrees,vortex ingestion in no. 2 inlet during sideslip operation, PLA = max.

of both low- and high-response AIP signals are shown during an increase in sideslip from 2.2 to 2.7 degrees. Wake ingestionis evident throughoutthe transient,and turbulencelevels increasesignificantlyat the higher sideslip angle. Although the individualplots are not preciselyaligned in time, low pressuresgenerallycoincide with positivevane deflectionangles. Similar results were obtained with the no. 1 inlet. (See figure ii0.) Smaller amplitudesare partiallyattributedto reduced engine airflow indicatedby lower fan correctedspeeds.

I_gine airflow demand has a significantinfluenceon peak-to-peakampli- tudes. Representative AIP pressuresare shown in figure iii during a throttle burst from IDLE to INTERMEDIATE(maximumairflow) power settings. Peak-to- peak amplitudesat INTERMEDIATEpower are greater by a factor of 1.7 than the amplitudesrecorded during IDLE power. Again, the lower pressures in the cycle appear to coincide with maximum, positive,vane-deflectionangles.

Static Vane Deflections Comparisonof results from tests describedpreviouslyindicatedthat resTJlts similarto those obtained with oscillatingvanes could be obtained with the vanes positionedat a constant deflectionangle. Crew members found these proceduresless objectionable, and their work load was slightlyreduced.

All subsequentflight tests exploring SMCS vane effectson inlet / enginecharac- teristicsduring maneuvers and engine throttle transientswere conducted in this manner.

Maneuvers.-Tests conditionsinvestigated during flight 2-33 (withSMCS vanes deflected+20 degrees) are summarizedin figure 112. Tests were con- ducted at Mach 0.85 and includedmaneuvers at combined angles of attack and sideslipwhere the maximum angle of attack was 6 degrees and maximum sideslip angle was 4 degrees. Representative AIP high-frequency,total-pressure signals are shown during maneuverswhere vane vortex / wake ingestionwas measured in the no. 2 inlet. Faulty ZOC operationprecluded analyses of data recorded in J the no. 1 inlet. A strip-charttrace,recordedduringnormal cruise attitudes, is shown for reference. Aircraft attitudesresulting in ingestionduring flight test generallyagree with those recordedpreviouslyduring wind-tunnel tests.

Steady-statetotal-pressure recovery and distortionparametersare shown in figures 113 and 114 as functionsof sideslip angle. Angle of attack is used as the independentvariable in figure 113 with the vanes deflected 20 degrees.

The largestdefect in recovery (approximately 0.06) was measured at 6 degrees angle of attack and approximately1.75 degrees of sideslip. Effects of the wake diminish as sideslip is further increased. At approximately3 degrees 100 ------..------r-----,.--------r------, Percent fan speed Mll12 NF OLLLLLLLLLLLL.L.L.L.L.L.L.L.L.L...L..L...L..L...L..L...L..J-l...J-l...J-l...J-l..~~~ .............................. ~ 20 -----.........-----.....,...----~-----r_---_, X3029 RHSMC Vane 0 I------~------+----..,---I-----_t_----___t - 20 L1..LLLLLLLLJ-L.L.J....l...1...J...L...l...1......L.J......L.J....l..J...l..J...l..J...l..J...J...L..J...L..J...L..J...L..J...L-1-L....I....l.....I....l....J.......J.:o Sidesl ip angle, degrees f- ,i-

-i

i=" , i- Ml012 Beta 0 i-

t·

i- .'

i- ,f-, I I I I I I r I r r r r I , , r , r , , , I I I I I I I I I I I I , , J I I I I I I I I -5 2800 ------,,------;,.------r-----,------; AlP PCM signal . I 40 CTS at 0.01 0.4 psi D6288 Counts Figure 110. - AIS/SMCS investigation, flight 2-19, O! = 3 degrees, vortex ingestion in no. 1 inlet during sideslip operation.

100 ,..-----.----r---"l:::::===---;t;r-----' INT Percent fan speed Mll13 NF

t

IDLE OLU-l..LLLLLl--L.LLL~__LL.L.LL.LL-L.L.L.L.JL.LL-L.L.L.L.J_.L.L~L..L.l_L.L.l.._l_LL..L..J 20,.------r-----,-----,------,-------, X3029 RHSMC 0 I-------+----:.---+------+-----+-------t Vane - 20 LU-l...l....u....l..L.J-I...L.L.IL..J....l-l-L..J...J....L...I..~L_1_l_l...J......L...L...L_L_ ............................... -'-'- .......... '__'_'_...I....I. .......

.- Sides 1 i pang 1 e, degrees::::J

I-- I-- Ml012 0 l- Beta l-

t

I-- I I I'll I I I II I I I I I I I I I I 1 II I I 111111111

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-5 2400 r------,-------.--,--..-----:T~--_;:.V""--r___, D6338 Counts ps i Flight time 2 4 6 8 10 • .• -- ... -'. - ·.A. jp -.c"sw- ;i~n.~.l·~. -- ~'. -:1.--- ... ~ f.-. ---.1.".;

; . . . . . 'I --- i ... JJI _ .• ~ ... ~

[ 063 38 -'\C~'(-t~~"'''li~ti"J~~r-~~\''r. ~I.· -- ft11;VWI.~I'~~~

----: .Low pores sure -" ---,----1 . 7X _. . -- '_'_'_1 .'--; ! D6339)IR~~~A~~~,:Ji~I~~~ .~~~_. !~\~I_'III~ ...

~ .. ~--_.::. ... ~.:.f.:" -.~_. __ LH, gh pressure . .' I Y: "; . . _l~\·

Figure 111. - AIS/SMCS invest;igation, night 2-19,

a = 3 degrees,vortex ingestion in no. 2 inlet

Juring IDLE-INT throttle transient.

Fit 2-33 test matrix Envelope of SMC vane 10 Angle of attack, wake ingestion from degrees • Normal turbulence Normal AIP dynamic activity wind tunnel data 8 O Increased turbulence (No vortex ingestion) Throttle transients -i_i_-- _ _:i '7-J _'2 -_ :_- " -- : - -i-"--_:--_ 6 _0.3 psi -" - -_ " ....

Mo re " _ 4 6347 -_- .... :_:-..:_-:_---:-:--._--_-_-_ 4 .----- • _ --- ....... :- : ...... •..: _: ..... - . - "!"i!i!iiii:_": .... 20 : 26 : 44 0 oz= 3 ° _= 0 ° -8 -6 -4 -2 2 4 6 8 -2" Sideslip angle, degrees t AIP dynamic activity with vortex ingestion _:- =-- '-- = 2.3 psi .... _ .... 2.4 psi ................ :----:'_........... 2.45 psi : _--_ A,P probe _-- ....... =-_=__.:_"_- _ _:_ _ L "_-=_.=_' :; _-.. _ ; --:_-_--_--: . : : : _ : " _.. ............ _: _-_ ,-_ ...... _ _- _--i_ii_. _____-_-i._-_.i L ----_i i:_ : _!--__._.

19:58:00 20:28:I0 20:43:00 e = 5"5 ° _ = 2° _ = 3° _= 4 ° OZ = 1° / _= 4° Figure 112• Flight 2-33 test matrix and AlP signals during 8ach 0.85 operationwith SMCS vanes deflected20 degrees, no 2 inlet, RB = 7 degrees, RC = 5 degrees INo.2 INLET I M= 0.85, alpha = variable, beta = variable RB = 7 deg, RC = 5 deg SMCV = 20'deg l.O0 Recove ry 0. ....... :! : 0.92 .....i i.1i.:..... .:i. /;!i _. :__..! : : ...... i.i:.i. : ..

•.. i.--i-: :!. ::i.-ii-:_!Sym Alpha WIR .i_i...]..i!ii-ii :i:i:.-:: • ' : :--: : "i_":i O I deg 297 pps ::' ._.: :. " : :: : :..---".. .. .......... - : .- ;. A 3 deg 323 pps .! " ::.-. !' .:::i: [.i::: • : :. :. : ':.:.-:: :; {'1 6 deg 348 pps .i. : .i ...i. i J .i.: 0.8 ............. ...... :: ; -:: _-:.Stal.. 1-margin indexl ._ - :-i!..::i :fi.::i.:: .'..: . . ; . : . .: : '.;::;.- . L'::: : ":'..:::-.;: L - ';'.:::_: ::: i::: : ".; ' ;.'.*._ . :; :_ " " .

IDL .... " .......... ' ...... :.........J.................... :...... :........ : '-.','.' :..': ....

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--- A '- " '_ ......... : _. . TT-,-_.. .... " ..... ;'-'+ .......... _, .... -t---_. _'-- A L-.. T ....... *....................... "-- - :::_ ; : _ - : : : : : _ . ;: . ; - ........ : A : : : ::: : :::- . ---:. + ; +--: --.: ---= _ i ._ + :-: : .: H-_ _ -: : _: _ : : __: : : _ ::: :. ? '_ :: . _ _ ? : + :__ :_ : _ _ ::: _ - 0 ! 2 3 4 5 Indicated angle of sides]ip, beta, degrees Figure i13. No. 2 inlet, effect of sideslip angle on steady-state inlet characteristics at SMCV = 20 degrees, flight 2-33.

[ 7 4 M= 0.85, alpha = 3 deg, beta = variable RB = 7 deg, RC = 5 deg l.O0 . Recover,' L ......... _...-. %. _Z_.:.:Z_:£1ZZ. Z 5 £Z.!'.-Z.:'-_-

o.96: !

• i. ' : : ".__i" " ' .:[7:_'.' : ::':x:. . :x:t'.:T_x::t_": . _ ':;.. ;. . ! " ' : .:i ..... , _ ......

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•i'..'.::. '. ....:.... ._. :... :. .....:: ..... SYM SMCV WIR .!::::i::.

..... _. • . .

;-N_t.; + .:i:> !:.:_;_.._:::i?_::>: .i :-: ,: • '-:- .:: ! ::.i_"_-_:! _ 0 316 pps-:---i_ i" •" : : i'. ":".'t:...!::.:_'::::'."::-.. ' . i l d :: _.... : : " ___-.._[. :.!_::{__.:k_:____.-.._._ _ 3.3 eg 316 pps ._-:_ - . ::.i! _ :- -:-: -" ::. :-__'_-_'_!::! g ,20 deg 323 Dps :'- :": :: .- :-. ::::::::::::::::::::: : ::::::::::::: . :_:l:.:_:::.!_k ":::: :.: ::.: :i ,. ..'._ ............... I ..... _.... _ "" _....... _........................... _..... _.......

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-.:;::_.--:- -t:::!::-it!N--!:iStall-margin index :::q.;::":i.:;::-i -.--.- IDL ::_.L. .. :i:-'r-.:.i-'-.i:::'::::'{::: ::]': :::: :i.: :::. ].:'.

..... l ...... I : .............. 1 :7 ::............. J......... , ....................... t-..........

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.............. _........ ' ........ _........ I ........ _" ....................... _........ l .............

0.2 :: ::.:.i}::}} :. .:.::_.:::ib:::::_:!.:::!:: Distortion .:-!:. !: :_: ;:::-!-::._::-!:.:: . L -::-" ::::.Z :::::::::::::::::::::::::: _:: Z Z---L--!-: Z [ LL :::Z-_F£f.:- . : .-:..:]: . : .:..:::::::: : :::::: : ::::::::: : ::: [PTMAX-PTMIN I -..:........:.-. ii::.-:!.::.i:!:qi:.!

I DT -:iii ::i i_i [ ii!i:.|!!i_-i : _-i::i: i:_i!ii:i[_iiii L PTAVG /::.::!i:::i:::: i -:ki:i;: i -! i :;".::.i::.J .._£._i....:.: J _!=:_: :: :i: ::i.k.:::!_Z.-_. : ::L::_::_:_:..... _i£_i..::_i--..-.-:_.:_:_ :: ::-L :. _. . ] ;:.:_._IL:_:_ ::_b:.'_i:_:.k-_::.!-_:. i L_:_;_:_N: : .:-_!:._:..-!_-...:::_ : _;II:.-L-: i:._i:- -.- O. 1 . :. :i ::: :..1:.i_:.:.: - :.!.ii:_ : !:!:.!!:::!::::!-.iiN-{ii::__._.'.::, iii. " :: (_Z_L .:i.:: :.

.-._.::.: : : - .::- ._ :.h:::L-::,: :: _i-"Zx:_:": 7 _:::. ] ..,,n : .:_ •. _ k i: 8;.:-i:_:.t__:._._ - 0 l 2 3 4 5 Indicated angle of sideslip, beta, degrees F]gure 114. No. 2 inlet, effect of sideslip angle on steady-stateinlet characteristics at various SMC vane angles, flight 2-33.

sideslipwake effectsare no longer present in the inboard inlet. At lower angles of attack, sideslip angles greater than 3 degrees are requiredbefore wake ingestionbecomes evident.

Effect of vane deflectionangle is shown in figure 114 during operationat 3 degrees angle of attack. Pressure recovery values from this figure indicate a similar loss during sideslip operation for both the 13- and 20-degreevane deflections. No losses in recovery are evidentwith the vane held in a neutral pos it ion.

No definitivetrends are evident in the steady-statedistortionindexes.

floweret, with the measured increase in turbulencepreviouslyshown in figure 112, this is sometimesindicativeof an in-phasecomponentpresent in the flow as shown in the time historiesof total-pressure recovery (figure115) where SMCS wake ingestionwas most evident. Time historiesof stall-marginindex are also shown. Similar traces during operationat normal attitudeswith no vane deflec- tion are shown for comparison.

Resultingenvelopesof circumferential and radial distortioncomponents are s}_wn in figure 116. Total pressure contours representing maximum values of stall-marginindex are shown in figure 117. Maximum dynamic effects corre- spond to operating conditionsresultingin lowest total-pressure recovery at 6 degreesangle-of-attack and 2 degreesof sideslip. Maximum stall-margin indexwas 0.74, and the associatedtotal-pressure contour shows a well- developed,low-pressureregion in the hub.

An extensive series of tests with the SMCS vanes was also conducted during flight 2-37. Performance and distortion characteristics are summarized for the 11o.2 inlet at normal aircraft attitudes (_ = 2.7 degrees, _ = 0 degrees) in figure I18 for the complete range of SMCS vane deflection angles from 20 to -20 degrees. No indications of wake ingestion are evident.

A series of tests investigatingSMCS vane effects during sideslip opera- tLon were conductedover a range of angle of attacks. Results are generally shown for both inlets,and in each case, resultsare shown with the vane at a neutral position to help separate vane deflectioneffects from sideslipeffects.

Re_Its at a nominal 1 degree angle-of-attack are shown in figures 119 through 123, 2.5 degrees angle-of-attackin figures 124 through 126, and 6.0 degrees _- angle-of-attackin figures 127 through 132.

Generally,these figures illustratethat the higher the angle-of-attack, the less sideslip required for the wake to enter the inlet and affect the distortionpatterns. Data recorded during tests at 6 degrees angle of attack are also used to illustrateadditionaldata analysis techniquesbesides the digital time historiesof quasi steady-state measured and computed parameters.

, .

, .

Fi Itered to 62.5 Hz sampled at 360 sps a= 5.5° {3 = 2° SMC vane =+20° a= 3° {3= 0° SMC vane = 0°

PT1/PT~ ,0 613:4

'111

0.8 11 ~ I PTi IPT:,: l... 0 0 0 0 0 [J"""", I, " , '" "I

1. 0 ...----,.-- ---r-----r----,

I DL ~J\'1',~t~':~:\V!f~, \'i/\>-'~tJ.l.I, \Wt,\I,r'lf\'

0.0 20: I 9: 57 I .0 2.0 19: 57: 49 I .0 2.0 Time - seconds' Time - seconds

a= 3° f3 = 4° SMC vane =+20°

PTl/PT:':ttt=u

1. 0 r----,---;r:- d

IDL o. 0 ..........-W-L-Iu....L.J.-L~U-LLLLl..I....L :: L1-1..L1-1.

2.0 1.0 20:43:03 Time - seconds Time - seconds Figure 115. - Time histories of total-pressure recovery and stall-margin index, Mach 0.85, with-~~S vane deflected 0 and 20 degrees; no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.

o.og.oo 0.02 0.0" 0.06 O.OB 0.10 0.12 D ....

IDC IDC IDC Figure 116. - Dynamic circumferential and radial distortion components, Mach 0.85, with SMCS vane deflected 0 and 20 dep,rcc-~~no. 2 inlet, RB = 7 degrees, flight 2-33.

, , Pressure above PTI Filtered to 62.5 Hz in shaded area sampled at 360 sps Pressure below PTI max IDL scans in unshaded area ,number on PTI - PTI [I+ _ contour ) _ ] Use + if contour in shaded area Use - if contour in u_shaded area = 3 ° SHC vane = 20 ° Figure 117 . - Dynamic total pressure contours during Mach 0 . 85 operation with SMCS vane deflected deflected 0 and 20 de_rees, no. 2 inlet, RB = 7 degrees, RC = 5 degrees, flight 2-33.

Recovery Flight 2-37 1 . 00 ? . _i_i::i::.: : . - .: : : .: : _-i : : . :!::!-ii ,!:: . " -: - : . : • "......

............................ : • : T.. ::...: ,, • i ....

0.98 ?: ]; ::?i:i-_"-i!i_i_i_iiii;:i!iT-; I NO. 2 INLET I i!71:!]!:. !i=:::7:_i:ii=:i;i:!ii_i: ; :=LJ--.L__:-_ RB = 7 deg_ RC = 5 deg _- ...........

.' _ . "I. : : i_.;..;...

PT1 _:-:.. -- -.

PTO .... .-! _ . M = 0.85, a lpha = 2.7 deg, beta =Odeg. -.;: : '.. .

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96 ............................................

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• ; : U: . .q.::;: ; : ; t ' - ;: ; ";;_ -: t:-,: ' _;U: ; :;; : [::;::-::I:,:-;. : .=:I;:;:;:::I:::: : t :I : ': :;:: ":.; ; ' : .... : ,--_--,---..--_ --iTi_ii:::J:::f=:[:li::[:!:=:ii_P:i:O Steady-state i--_i:_:::_-i.:.;.:!::_!!::i-i!:=iif_!::_! ] i!i::_ Dynamic (from max IDL scan filtered ' , ..' , _ '. | ".'...: . . ' . • .... ._' ,=,._ ..... :"_-'--::-:::::':::_: ::' : -" to 62.5 Hz sampled at 360 sps) •"--_'"-_-':' : -":_'-: : ::.=d, =_ _ . .. r ..................... ' - -:"=- -: _.:i:i'::i:i F ::i! :iiI::::i 'h.:., :". -'_ .., ............... ::':iii]4['iiq=_=_ Di stort ion .--:it:: :':i_:i,71:.:! i:!: .: : ! .i - ._::i!:i:!!::_:i:.:::h::ii : _::-=._ .:_: : -ti : rPTMAX-PTMIN1 • _I _::i! T-t!::.i:i: : : ::i 0.2 '_ ' ''.: ; .......... 1 .... , ..3 F. ............. t ......................................... _. .

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::::h :.:ii:.]i::::!i."_ ...... : -:;:::.='::' v :: : :!!::' : : : ::l':::]:7"h::'!:!:!:: .':i !": i' _ . :." : ............................. ' LLI:-_:-- 2 . L Z I . ..L.L: "_-i2:=J%__.L." " ' -- • " L =.= " __.-'_ .... ;_. " ......... _: -F : J.............,.... _ ....... T ...... -.--_.._ ...... - - _4: ::: __::!_:=:f_ = : :..;m " _..::.A.: !:,,_::::_!.:G).I_.: _::.::.-._ :: " ...... =- _ ..... l ........ _v ........... ----_,, i . _ . _ -- " ....

: " ':"_:" : :-_ :: : _:--::=i : i::::i.i : !!i- - . : i:" i"i:t - : : _': := i: : !-!.:? -:::i!-' : i-::::f:!'iir-:;':-.-.!:::!-::T:_:_---, ....................... _....... t: - _-- : ::: ..... : i ................... ! ................................. " ......

-20 -10 0 10 20 Structural m o de c o ntrol vane defl e cti o n angle , deg-rees Figure 118. - Effects of SMCvane position on no. 2 inlet steady-state and dynamic characteristics , Mach= 0.85 and alpha = 2.7 degrees , flight 2-37.

[8O M o ( P E G) I 0- -" 0.5 JI I LI[1111 I_ 1 1 1 1 1 L .L L LL -2 I _I_ .L . L.L.J A L TITUDE ( D E G ) ( K FE E T ) o II LAI''' , , ,_,_i,, , , llJl ll l k_, ,, ,, , J ,_ , , i ,1 .... ,, ,, . 5 I IItllll _J ll J l l ll lJqll lJ L_ '% I I t I I 25 ' E - - I ST R U C ' _ J R AL H OOE CO N T R OL V A N E _ SI _ ....

- _ 0 J. tl t l [ll ll_ 1 _ U!l[ l ll l t ll ll l ltl ! I L L L 2LJ. -2 _ 1 J JJ . J . _L LLI l J 0 _. 0 80 120 0 _ 0 80 120 _ 2 0: k9:32 SE CO NOS SECONDS

I No .l inlet I INo.2inlet I

RB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg _ i __ _ LO -- _ ..... , ST A LL _R GIN I ND EX "-- 7 0 ]1 f t 11 1 [ I I I I 0 I II I , . _ .... _ , 0 I tl [ ] 1 11 1 I I II Ii iii iii I I I I I 1 FAN DI S C H ARGE H A CH N O . SI G N AL G N A L ( A FT C i N TRO L )

°" r - i ]

/ iitiil[i itll_l:_ ::ii_111 I.i]ll_ll II!ll;llI_[ illlll L 0 F _ l _ ' l _ l l / t : _ll_ ll lr illll : :1 , _:.: t _ --.--. _._ - L .... _ L _I 0 gO 80 120 0 _ . 0 SO 120 SE CONDS S EC O ND S Figure I19. - Effects of sideslip angle on steady-state inlet characteristics with SMCV= 0 degrees and alpha = 1 degree, flight 2-37.

MO (DE G ) 0 . _ k L l I CI_: [ _a ,EL _ - _ I II I lllllll l llllllll I l t l llFtr 111111 11 1 II II l ll l ll l ) [ l[ll[l l ,,,, _, , 11 ,, .c I I IJ _l l _ [ _l_ l _ L L Ll IT I ll ll _ * l l _ l Jllrl_J ll l l ll l l Jl !

(_ F ) 0 - S,_'V O - C _ G ) o k.0 80 !20 0 _ .0 80 1 2 o 2 0 : W ]:02 i_ SEC ON DS SECONDS

INo. 1io,e_ I I mo._ io,e_ I

RB = 7 deg, RC = 9 deg RB - - 7 deg, RC = 5 deg 1.0 ...... % T _ LI ._ qGtN INDE.X _ - _ 1 . 0 - STALLMA R GI N I, N DE:X _ 60. 360 COR REC T E D MEIGHTFL. O_ GOR R ECT1ED V _ EIG , HT FLO _ W I R .

(P PS ) " (PPS ) - o._., i i , _ o_ _ No. s _r . _.. I I o .s t _ A _ O_s, : _,_r_ _ c , No. s _ r , _ . " " _ AN FAN p _ - F -, - - I : 0 JJJ i l Ji il _ Ji IL I JlJlL I I t lllll l f ItJi l J l it ILIIJlilf O_ {{{IJ[lll ili _ lli:llli[lil L_[I_II[I IliJ _ ll il? _ li,_ 0 W O 80 120 0 40 80 1 2 0 SECONDS SEC O NDS Figure 120 . Effects of sideslip angle on steady-stateinlet characteristics with SMCV = 20 degrees and alpha = 1 degree, flight = 2-37.

1.82 ..

" Wi ng Pressure above PTI \ in shaded area

I No. 2 i n 1e t I

Pressure below PTJ Ramp

~ Cowl

in unshaded area 5 deg M = 0.84, RB = 7 deg, RC number on Looking aft ~] PTI - PT I [1 + ( can tau r ) SMCV = 20 deg

- ~

, f Use + if contour in shaded area Use.- if contour in unshaded area Beta !Beta = 1.6 degl = 0 degJ

I Beta = 4.0 deg I

1.4 deg Alpha = Alpha = 1.4 deg 1.3 deg Alpha W1R W1R 311 pps W1R = 312 pps 313 pps

IDL = IDL = 0.296 0.304 IDL = 0.420

= lOT lOT lOT = 0.056 0.055 0.076 Dynamic total-pressure contours (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 121. - Effects of S~~l1 wake on no. 2 inlet dynamic total-pressure contours during sideslip, SMCV = 20 degrees and alpha = 1.4 degrees, flight 2-37.

_ 0 ( DEG) a Lr I T ' J _ I _ E _ TA (v FEET ) ( DE C , ) 0 - i_ AL T I T_ JDE INDICA TE DAi _ ,LE O F S IDES L IP , _ TT _ MCV , .oF ) 0 C_ E[ G ) L _ L _ . -- _ J J- U J - _LLU ,-- U . _U -I 0 20 _ 0 60 80 1 00 0 20 _ ,0 60 80 100 _ 20:5].:[ 1 5 SEC C_ NDS " SECONDS

i.o., 'n'o! ' I"o. 'nlo l

RB = 7 deg , RC = 9 deg RB = 7 deg , R C = 5 deg PT -"_ PTO 0. 9-L , L , U . .I _ . LI L. L . L .L ; L _ .L L L _ :L. L L . L - L . i i J J i i _ i i 1 ID I. I _.

0 I J ,_ LL . L L _ . _ & _ I 0 I I t I I I I I I I I I I I I I I ] I I _ I I I 11 I ] ] I J I ] I I 1 I D t _ _ J ' _ ID T Wl R __ WI R l_ O - _J [_ . I I I I I I I _ I t I _ I 162 I I I I I I I 11 i I I I I I I I I I [ I I I I I I I o._ - - -- F AN D l SC H ARGE MACHNO. SIGtW _ L J / 0. _ t FAN DI S _ HI _q GIE I' _ C _ NO. S l G / ' _ / ..

1 _ 'T c o _o_) P _" c ,u ,T C O m _O L ) 0 20 ' _ 0 €'0 80 I00 0 20 kO 60 gO tOO SECO t _ S S _ CONDS Figure 1 22 . Effects of sideslip angle on steady-state inlet characteristics with SMCV-- -8 degrees and alpha = 1 degree, flight 2-37.

[84 Pressure above PTI in shaded area No. 2 inlet

I I Wing

Pressure below PTI in unshaded area Cowl _ I_ _ Ramp , number on, M= 0 85, RB = 7 deg RC = 5 deg PTI - PTI [l+ [ contour ) _ ] " ' - ^ SMCV = -8 deg Looking Aft Use + if contour in shaded area Use - if contour in unshaded area _J _:_ : _ : i : _ : _ : _ : i : _ : _ : _:! : ! : i Alpha = 0.8 deg Alpha = 1.0 deg Alpha = 1.0 deg WIR = 307 pps WIR = 305 pps WIR = 312 pps IDL = 0.351 IDL = 0.291 IDL = 0.338 IDT = 0.055 IDT = 0.054 IDT = 0.059 Dynamic total-pressure contours (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 123. - Effects of SMCVwake on no. 2 inlet dynamic total-pressure contours during sideslip,SMCV = -8 degrees and alpha = 0.9 degrees, flight 2-37.

(O r: ( ;.)

( K F[ET ) l BETA ALTITUD E i I 0 Ill|llll, I''lill,I Illll,l,r lllll l ll Illlltrl Illlll ; I 50- 2.5 . .

T E4 PfEP . A T U R F t i I _ [L ON[T R OL VAN E_P OS I T1 0_ N T J ( TTO- T TSTO ) ST R UCT W. . _ E C COF ) O- ( : W EG ) I -50- Iltllllll ZlILfI11_ Zlllllll IIlllltt ll,l ( lll IllJllll_ "25 _ _!_ { _. 0 8 0 | 2 0 _+0 80 120 SECONDS S ECONDS _ 21: ) .7:52

Imo. ' io'e l

RB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg PT O P TO , I I I I 0 S-- IIIfl IIIIII iL l I ) lll ll ll I lll lllll

..... o 0 O "

0 . 2 g I I o,s_'r,o_ I I I 0 .2 o I s toRT,o.

0 I I IIII I I I III II II t lllllllll ll lllllll lt fll l ll l l l llllll l ll l ll { I I iiiiii t lllJf l t f l l1 1 1111 1 1 11llll lllllllllll l lJ ll ll l lf ll ll _ lll j C_t R _ CT _ _I G H TFLO _ 1_ 60 ' I , C O.C O TED W EI G H T F LO _ ; • _.

WI R WI R CPP$) - (PPS ) - " 1 _ 0 _l lll,llll Illllllll IIIlltll IIIIlllll Ill*tmll _l£1Iltll _ 60 Illflllll Iltllllll (llllllll {tlffllt tltfl]l] fflllllll 0. 5 F_ _l_*,_ ; e A_ NO . s (c.,,_, o.s- , , F AN ms c _ _c HN O . s(c . _ O_ IIIIIlll III!llllll Imlllllll I1{111111 IIIJtllll (lllllll 0 _ IlZ*lllll iltflrl] i!llllll llfIIIlll flllllll IlmmIItll 0 k O 8 0 120 0 40 80 120 SE CON Ds SECO N DS Figure 124. - Effects of sideslip angle on steady-state inlet characteristics with SMCV = 0 degrees and alpha = 2.6 degrees, flight 2-37.

.- HO ALPHA ( D r : 6 . ) 0- INDI_ q' _ A,' _E _ 'A T TAC K J 0 .5 J-J -- LLL J I I lll lll l l I U J_ . L L I I I _1 I I -2 I t[ AI_T I TUD E ALTI ' PJDE BL = TA ( K FEE T) - ( D I EG. ) _5-1 _ i rl llt ,l il,r ji itl i ,i ! l j 0 IlJllllll ililtllll tltlllllf tl[llllll Illllflll

, o _ ,, , ,._ "-_----T ..... q---- - -r:_ - _:-- -1

(T T O - T r S_) I - S _UCT_ U U . H O0_ CON T RO L V _ E I I ( OF ) 0 - (_ G) 0 -- ._0 f f I I I I I I I I I t I I I I I I ! I / 11 11 llZltle : l lillllllll .2 _IJ--L L _ L _- LL .L L . L _ L J - J _ J _ J _ J - L ._ J I J- L I '-I LL J - . _ J .

0 20 _, 0 60 80 100 0 20 _ . 0 60 80 1 00 21 : 12:55 _ , . L . _ ( ,ND_ < _ E_. . t l DS

i o ., i0 , eq

RB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg , .,: ......

I _c, , ;) L_ t v:, , , ' L R; pT _ PT O ,,, , ,,"_ . l i 4 _ L _ . L J -:_ j , _ i i_i, ,i , [ , J l_ ;,, I OT I I _ IOT -r

I

" 0 iiiiiiiii l 0 I[ ii • 360- .

! . / - - WIR W IR - - ( P Ps) CPPS ) " I _ 0 I I _ I I I I I I I I I I I I l I [ I I , I t i_ 0 _llll f lll _ I I I I I I I I I I I t [ I l J ] ] i I I I I I I I I : I [ I I I I I 0.5_ _ _ FAN D I SCHARGE HACH NO. S I G N AL (AF T C O NT R O L) (AFT CON TR O L) i II iI II II_; lllI_IllIlllIll!ll lll''_J'l_l _I_ 01 t I I I I I I I _ 011 I I I I I I I 1 4 _II I I I I I 116 (} I I I' I I I I I _ (} I I I I I I I I 00 _ I I ' I I I I f 12 0 _ O _ O SO I_ 0 SE C OND S S EC ON DS Figure 125. - Effects of sideslip angle on steady-state inlet characteristics with SMCV = 20 degrees and alpha = 2.6 degrees, flight 2-37.

_- Pressure above PTI @ o co in shaded area I I

Pressure below PTI I No. 2 inlet I Wing

I I in unshaded area , number on, PTI - PTI [l +- L contour ) A ] M = 0.85, RB = 7 deg, RC = 5 deg Cowl Ramp A r . .

• SMCV = 20 deg use + if contour in shaded area Looking aft Use - if contour in unshaded area • . * ... ,:.. , .........:..... . ..

_,, _-!... : .!_!

m_-!!

t

_ , "_.i:_ : _ "/ . , ,. , , ............ : _........... ..............._...... >. ............ : .. : . : ..:.

\ - _-- . ...- / I Beta = 0 deg Beta = 1.7 deg l I Beta = 3.8 deg I_ Alpha = 2.6 deg Alpha = 2.6 deg Alpha = 2.7 deg WIR = 317 deg WIR = 317 pps WIR = 318 pps IDL = 0.391 IDL = O.397 IDL = 0.449 IDT = 0.064 IDT = 0.064 IDT = 0.I04 Dynamic total-pressure c o ntours (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 126. Effectsof SMCVwake on no. 2 inlet dynamic total-pressure contours during sideslip,SMCV = 20 degreesand alpha = 2.6 degrees,flight 2-37.

i. o i-. - - -i i ......... ......

0 _ U J. / _IItltLJ . J. / .J .L / .I . U. / _- I J -L / J I .2 _ t ll l l ltr lllllllllllllllllll ALTI ' / %ID E B E TA .

(K F EET ) ( .DE G . ) i tt llfll iiiii111 _ fi l l 1 1111 i - 5 I_ Lllllll1 _I ti _ lllllllltlllll 25.... S TRUCT ' JRAL MOOEC O N TR OL VANE _S I T I _ -- - C T TO- T rS _ ) A T U -- S / 4CV 0- -- -- ( OF) 0- CDEG) ,

+° - ?1

-$0. ] e tllrll iil11111 tlllllltlll_tlel tl I_1111_ t J llllllsl _2 5 _ 1 LL . L .U .LL _ L J I J L.L . LZ.L . L . I -LJ _ J J . Ai I I I I I _J2I I _ , J I I t LL . U. L-U- I -LJ " 0 _ 0 80 120 0 k.O 80 1 2 0 . g [ C OI , IDS Ok: _ . "_ 1$: 5 7

I.o.1 inletl I"o.2 +nlot l

RB = 7 deg, RC = 9 deg RB = 7 deg, RC = 5 deg . 1 I I ]j l lll l]l 1 I , * , r, ,ll, , 11 _,, , g . _ I I I II I 1 1 ] [ll l j] ll Il lll Illl G I N IND E X STA L L t' A R GIN IND(D( IDL I Dt..

- (PPS) Wl_

++T+ ' '

" 1 6 0- ( PPS) 0 . .5-- FAN D[ _ E f _ A CH N O. $ I G t _ 4 . 0 . 5- F A N D[SC _ GE _4 ACH N O. S ' |Gt, l a , L ( A F T CONTROL) I ( A F T c Ontro L .) !

51_t_t ) _,I_S S _. O NDS Figure 127. - Effects of sideslip angle on stead y -state inlet characteristics with SM Cg = 0 degrees and alpha = 5.8 degrees , flight 2-57.

1.0 - - II'{)ICATED An;Lf OF ATTACK V I--~I -"'" ~ ~ ___ --.r ~ AL_ (DEG.)

I

" ~'l 0.5

-2 - _._-~. __ .

- ALTln.oE L ALTlTl.OE

~I'{)TCATED AI<;~ OF SIDELIP BETA

(K FEET) ~ (DEC.)

I ,1, ,I

TEMPERATURE (TTO- TTS TO) ~TT Sl'CV (OF) (DEC) - \0 ~o L.LUl.:..llJ.l..U.J.l..U.J.l..U.~.lJ0 Ci.J..L'-ll1-L.LU.:..L..O..L:. .L1 J.l..U...LC1.JLU.LL.LU.L:. lJ s o 1 20 SECOI'{)S

~IS:00:52

0.2 _.

- -- -:I;~lJ~ J~j

lOT Pn-tAX-PfMIN] [ PTA\IG O_ULLLLll-l.11J.JL11.lJLLl-LJLLl-LJ..1.LL: .. l Lt LLl ., - - ~..kIGHTFL~ WIR (PPS) -

]

,1J'Uli

0.5 FAN D~SCHAAGE ..lcH I-Cl. Slkw.

(AFT CONTROL) - u_ .

_.1.l.1UJ' 120 so ~O SECOI'{)S SEClNlS Figure 128. - Effects of sideslip'angleon steady-state inlet characteristics with SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.

SY M TEST p _L_ . T / PT 1 4 ,N . .I A I _ P _ L _ E TA P T O PCNF R -I PC N F R- 2 P _ NFR- 3 PO 'I rR - 4 I: _tr) UCTION TI M E O 20 3 " / 6.01 0, 8 7 2 3 059 5. 8 0 . 1 9.' / 9 0 . 04 9 8 .69 82 . . 5 3 9 7.06 1 / 31 / ' 7 8 2 2 3 1 8: 0:31 13 _ 03 7 6 .0 2 D. E_ 230 1 5 5 .6 1 . 8 9.6 99 .0 9 99. 42 f . / _ . 24 9 7 . 35 1 / 31 / 78 2 3 2 1 8 :0: 5 1 • 2 03 " / 6.03 0. 85 22 1 85 3 6. 2 2 .9 9.6 99 .13 99.0 4 6 [ : 1 .' / 6 9"7.61 1 / 31 / ' t B 2 . 2 3 18: 1 : 1 8 .... AIP pressures / PT0 at max I DL - engine l 1 .040 . i i _ i - ......................... i.................................................... i ......................... - _ .......................... -........................... t.......................... f .....................................................................

......................... !................................................ 4......................... _. ........................... "........................... =: ......................... i ....................................................................... ; : :'"" : : : . .

........................ i....................................... i................. i......... -- ,.ooo... ................... " ........... : ..... ............ ........... ........... N o . 96o". :i:i. L .....:.:..::::.:.._::]._...i....]..._!.::::&..::ii ........... i.:_:i:iLLiiiiii i_ii_if ..... L ....:._......_. i _I LL ... :... ::::::::::::::::::::::::::::::::::::::::::: ili i O . 840 ....................................................... _ .................... _ ..... . .................................................................................................. _ ........................

: : s :. . - :- : ........................... :............................ •.......................... .. ; .......................... _.................................................................................. : ...........................,......................... . . ................

_. l " : .

...................................................... ;........................... ' _ ........................ • ..................................................... T ........................... '[ ......................... T ................. T .......

.................................................... _ ........................... ._ ........................._..................................................... i .......................... . , . ........................... " .......................... " ...........................

o . 8oo ........................................................ i........................ _...........]......................................... ,:. ..................... ,; ...................... _ ......................... ;_ ............................

• : * | . : :" l : : I ....

...................................................... .. ........................ .: ........................... _........................... ._ ........................... ._ ........................... . * ........................... . ...........................- .........................

O . 760 ; ! _ ! ! * : : 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 a 2 . 2 2 . a 2 2 _ 2 2 2 2 e 2"_l-- - Probe aa e a e 21111 11 1 J ! I 00 000 0 0 000 9999999999 88 8 8 _ ' 4 3_. I 09 B 765 4 3 e I 098"765 4 3 _ I 098 " / 65 4 3 a I 098 - / 6 number (see figure 104) AlP pressures / PTO at max IDL - engine 2 I .040 _ _ .; i ............................................................................................................ • r ........................... " ........................... f ..................................................... - ...................................................

• ........................ : ......................... ; .......................... ,, 1 ........................... , i . ........................... , _ ........................._ ........................ _, ...................... _.....................

,.ooo :: .......... %'_ ;_:_ : _F_ .............. _f_ ....._i ........... i_:_:J_:_ : _:_i -_:_-_ - _=_:_:

iiii iii

-. _'TZ?IIZZf?Z_L .......... & L Z 22 iXl ...................... j. 2 ZIZZ L ZZZ:iZZZZZTZ£ • , . : . : . I .

O.800 .......................... _ ............................ _ ............................ : ............................ _ ......................... ' _ ................... ' _ ...................................................... '; ............ _" .............

: ; i : : : : " .......................... |............................ , .......................... ,.......................... . , ........................... ,_ ........................... .i, ........................... . ; ,.-_ ........................ , ........................... , ...........................

; : ; { : : : : = ............................ i ........................... ; ............................ ; ........................... • ........................... _........................... ; , ........................... _ " ........................... , i . .......................... ,,7 , , ..........................

. ! : : [ " : " " : • : : 0 . 760 : _ ' " _ _ _ ! ' " 1 56666666666666666666666666666666666666 6 6 3 "_'3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3"3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 3 8666665555 5555 5 54444 _ _ 4444 3 33 3 3 33333 _22_ = 3 4 3 _ , I 0 9 8 " / 6 5 4 3 2. I 0 9 8 'v 6 5 4 3 _ 1.0 9 8 " / 6 5 4 3 2 I 0 9 8 3' 6 Figure ] 29. - IndividualAIP probes, total-pressure recovery array, SMCS vane deflected20 degrees,Mach 0.85, _ -- 6 degrees, flight 2-37.

Filtered to 62.5 Hz sampled at 360 sps No. inlet No.2 inlet 0.10 C 0.08 +-J L +-J .

0.06 l/l ICJl "'0 0.lJ't rc "'0 0.02 rc 0::: , Circumferential distortion Figure 130.

Variations in circumferential and radial distortion components during sideslip operation with the SMCS vane deflected+20 degrees, Mach 0.85, a= 6 degrees, flight 2-37.

, ,

P res,ure, nsha o ab O Vearea .... eti

I No. 2 inl Wing i i Pressure below PTI in unshaded area M = 0.86, RB = 7 deg, RC = 5 deg Cowl :_Ramp , number on, SMCV = 20 deg PTI- PTI [l _+ [ contour ) A ] Looking aft Use + if contour in shaded area Use - if contour in unshaded area J ":::i: :: _ _ " I J _ _ .1 _ 1 _ ' : " I_ _ 1 _ _, I ,t.:: " :i_iii : _:i_ - ?i!ii_!??! . " .o .I . .j I _ . j , i : :: i:i: : : : :::: ::::::::: : : :: : : : : :iii!!_i : :: :: : :_ : _ik ;. , . ,_ V _ : j J _1 :" " : ' :" . ._ : ' 1 _ ' " " 1': ': ": '1: : : :: : " ": ': :: :: :: : : : :" ' : ': : : :: ."" : : : " " 11 _ ' -" , -. - . ...... _ .:.:.: ....... t- . : . : : ....... _ 4 : • Alpha = 5.8 de 9 Alpha = 5.6 deg Alpha ; 6.0 deg WIR = 346 pps WIR = 349 pps WIR = 348 pps IDL = 0.506 IDL = 0.688 IDL = 0.547 IDT = 0.091 IDT = 0.185 IDT = 0.099 Dynamic total-pressure contours (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 131. Effects of SMCVwake on no . 2 inlet dynamic total-pressure contours during sideslip,SMCV = 20 degrees and alpha = 5 8 degrees flight 2-37 ° y • Pressure above PTI Wing in shaded area A Pressure below PTI T in unshaded area M = 0.86, RB = 7 deg, RC = 9 deg Ramp _ _- _ Cowl ,number on SMCV = 20 deg PTI - PTI [l± t contour ) _ ] Looking aft Use + if contour in shaded area Use _ if contour in unshaded area Beta = 0 deg [Beta = 1.8 deg I iBeta = 3.6 deg[ Alpha = 5.8 deg Alpha = 5.6 deg Alpha = 6.0 deg WIR = 347 pps WIR = 348 pps WIR = 350 pps IDL = 0.580 IDL = 0.607 IDL = 0.706 IDT = 0.099 IDT = 0.120 IDT = O.163 Dynamic total-pressure contours (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 132. Effects of SMCVwake on no. 1 inlet dynamic total-pressure contours during sideslip, SMCV = 20 degrees and alpha = 5.8 degrees, flight 2-37.

For example, local recoverieson a probe-by-probe basis are illustratedin figure 129. Probes within a particularrake are connectedby straight lines and are arranged in order from hub to tip. This particularplot uses sideslip as the independentparameter. SMCS vane wake ingestionis evident in the inboard inlet at 2 degreesof sideslipand is becoming evident in the outboard inlet as sideslip is increasedto 3 degrees. Distortioncharacteristicsresult- -- Lng from scans that produced the maximtm value of stall-marginindex are illu- strated in figure 130. Combinationsof circumferential and radial distortion componentscomputed at 360 samples per second are shown for selected aircraft attitudeswith the SMCS vanes deflected 20 degrees. Maximum values of stall- margin index are identifiedand compared against inlet design goals (IDL = 1.0).

Correspondingtotal-pressure contours are illustratedin figures 131 and 132.

Maneuverswith the _CS vane deflected 20 degrees (figure133) illustrate the sensitivityof vortex ingestionto angle of attack. At zero sideslip, vortex ingestion is evident in the no. 2 inlet between angles of attack of 7 and 8 degrees, consistentwith the maneuver envelopespresented earlier.

(See figure 112.) Similar angle-of-attack excursionswith the vanes held in a neutral position are shown in figures 134 through 136 for comparison.

Steady-stateinlet characteristics during sideslipmaneuverswith the SMCS vanes deflected13 degrees are shown in figure 137. Total-pressure recovery decreasesapproximately5 percent when sideslipangle increasesfrom 1.7 to 4.0 degrees. Recovery losses are accompaniedby significantincreasesin distortionlevels. Total-pressure contours representing maximtm dynamic values of stall-marginindex as a functionof sideslip angle are shown in fig- ure 138, and a well-developed,low-pressure region is evident in the hub.

Stall-marginindex increasesfrom 0.5 to 0.75 as sideslip is increasedfrom 1.7 to 4.0 degrees.

Engine Throttle Transients. Engine throttle transientswere conductedat selected conditionsto demonstrateengine stall-margincapabilitybeyond that requiredby the inlet during SMCS vane wake ingestion. Initialtests were con- ducted during flight 2-19 and 2-33. Qualitativedata and operatingconditions are summarizedin figures III and 112.

During flight 2-38, maximum rate throttle transientsbetween IDLE and -- IhFFEPJ_DIATE power-leversettingswere conductedon the no. 2 engine during sideslipmaneuvers at differentaltitudeswith the SMCS vanes deflected 20 and -8 degrees. Steady-stateresults at nominal angles of attack between 1 and 5 degreeswith sideslip out to 4 degrees are shown in figures 139 through 141.

- ~'"'~~ ,

I. rmI- - D

~ , _111'1!, ':'LTlTLlJE .J.LTIT1JDE (if. r~ET) UlJJJJ " JJUl' .1.Lllll;LJ .20 40 80 SECC1'VS , No. 2

I No. in 1 et

in 1 et I

RC RB = 7 deg, RB = 7 deg, RC = 9 deg 1.0 :.o~ V'I I

pn.E I 1

RECO'/ERY ~ PrJ PTO o• 8 ..LI-LJ...LLJ...LU-LLJ...LL..LLJ...LLLLl.LLl.LL.LJ..l..L.

0.3 :.0 .,..-----r-----,-----,- \.0 STALL ",I1'VEX I~L IOL O...L.Lu.l.Ju.l.J..l.1..1..l.1..1..l.1..1.Lll.l..Lllll.llLLll 0.2,------,----,------.---, ,- __ -, Q .2 lOT o ..LU..l.1..1.l.l J.l.l..J..LUlllJlllJLl..LLi..LLllLllll.l.ll.l..LllllllllJ..l.l"JJ1.l..l..l..lll..

HIt) 3&0 F====1'==--==r==

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T L '~lP (PP5) (PPS)

wlR Ll

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I =----- v

I

JJJJl J." Q '" ?LA (CE~) Figure 133. - Effects of angle-of-attack maneuvers on steady-state inlet characteristics with SMC vane at 20 degrees, flight 2-42.

L CDEO /

O, _1 11111 !1 1 i ii iiii i i illlflTll |Tlllltll [llltllll IIit1,111 _ I1 1 S l I _1! 1 11_ • ( K FEET) BETA (D E G) 0 il tllllll Iffllllll I lJtll l l! I J l l llll I _1 111[1111 1_I [ 11_ 11 1 11 Il ll[ t l t! I l rlr tlt ll L_ 0 120 _ _ ' 0 80 120 SE C ONDS S EC O N D S RB -- 7 deg , RC -- 9 deg RB = 7 deg , RC -- 5 deg L.O , E E R E CCV E RY p 'r _l p ' r_ PTO PT0 0, 8 Irlll l l ll t lllllll I _ 11111 1i l I 1 1 r I I I l l lllllIl L I I I III1 Ill It It i1 0, 8 li l J _l ll rlltl ll ll I lllfl l ll 0 I l llll l f l l l ll f fill II I II III II I 1 11111111 t II III I I II IIII I I1?111 I | Q llJlt II t I l ll II l l l I f III III I III II 1 111f l l[I I1111111111111 1 1 1 1 ll| PTAVG 0 i I l l 0 | lll t ll l l ) IIIIlllll l l l ll f IIIIll _ 111 1 flflllljljlflllt II IIIII I / C_ RE CT_D WE IGHT FL OW , COR R ECTED WEI GHT F L OW W I R W 1 R _PPS) ( PPS ) 0. 5 0. _ _ _ F AN D I SCH / _ G E ,- _ , N O . STGN _L (_ ,FTC0N TI_(_,.) (A F T C O N T I _ OL ) 0 Ill|Ill !|l t |f t l|llll! !1|11111 I|11111|1 |1111|1 0 _ I IllJl[I IIIIIIIIJ till|Ill P l..A . P L_

" I (DE G ) (D_ G )

0 o Illllrtll IIIItllll Illlltlll Illllll Jllllllll ItllltlllJ 0 _l'_ll l l l: l l llllllllll l _ o _ . 0 0 %20 SE CONDS sEc Ofl DS Figure 134. - Effects of an angle-of-attack maneuver on steady-stateinlet characteristics with no. 1 engine at intermediate,flight 2-42, SMCS vane deflectionangle = 0 degrees.

II I~DICAiE:J .G.l'GL:: 'JF ~TT~CK

1<0 :.r:[un[~_==::-RU'eER

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~11:l2:lO

INo.2 inletl ~No. inletl 7 deg, RC = 5 deg RB = 7 deg, RC = 9 deg RB 1.0 RECOVERY I ~ ~ RECOVERY PTl PT~ 0.8 7.1 1.0 1.0 STALL /'o4ARGIN It-VEx STALL "1ARGIN [N:lEX lDL ICL I~

U

o liJ.1.l.lll.1.tl.u..J.l..lll.1.Li.J.Jl.UJl..U.Li.J.JLL!.ll..U.J..!..ull.1.UUlll.l.ll.U.l.-'-L llll.l.llllu.lll U 2 O. n .2 DISTORTlCJ'.l [PTKAX-PT><INj .-..

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rTKAX-pT><IN] P~iVG , , o o l60 loO I I C:)RRECTED wE IGHT FLOW

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tJ

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F"'N OISCJ-'"Cl.RGE MACH ~. SIGf'.4A.L FAA! OISCHARGE KOoCH NJ. StGN4L (AFT CeNTROL) (AFT CONTROL)

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PLA (OEG)

I

..

20 GO 20 SECC/-DS SECOf,()S Figure 135. - Effects of an angle-of-attack maneuver on steady-state inlet characteristics with no. 2 engine at intermediate, flight 2-42, SMCS vane deflection angle = 0 degrees.

\.0 1===+==::[::3""~C~H ~'U'~eE3:'=+===F==l

fAD ').; ALTlruoE

(K FEET)

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lu

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a 0.2 LA ~ .• o ~ " lOT

- "'"

DISTORTION [PT'-'A)(-~""Wll PT;'VG j a 36 WIR ~PPS) W IC:JRRE>::::~ rWETSI-" .:'L1 ~.

a ~ 6 \ a.

FAN 'JISCI"¢.RGE "'ACH f.IJ. SIGNAL (AFT CCt,TROL) FAN ff !

~ " 'J POtiER L::~E~ ..:.tGLE -< col,> "C" 1---1 --,

I

f\ f---+---"-----+-- -, !

,.

" L,,~ ,0 SECCtJ)S Figure 136. - Effects of an angle-of-attack maneuver on steady-state inlet characteristics with no. 2 engine at intermediate, flight 2-42, SMCS vane deflection angle = 0 degrees.

Figure 137. Effects of sideslip angle on steady-stateinlet characteristics with SMCV = 13 degrees and alpha = 3.3 degrees, flight 2-38.

in shaded area No. 2 inlet Pressure above PT I I I _Wing Pressure below PTI in unshaded area M = 0.85, RB = 7 deg, RC = 5 deg Cowl -_4_>- Ramp , number on SMCV = 13 deg PTI - PTI [I+_ t contour ) A ] Looking aft i Use + if contour in s'haded area Use - if contour in unshaded area Beta = 0 deg Beta = 1.7 deg Beta = 4.0 deg

I l

Alpha = 3.4 deg Alpha = 3.2 deg Alpha = 3.2 deg WIR = 348 pps WIR = 347 pps WIR = 347 pps IDL = 0.514 IDL = 0.504 IDL = 0.750 IDT = 0.090 IDT = 0.099 IDT = 0.163 Dynamic total-pressure contours i (max IDL scans, filtered to 62.5 Hz, sampled at 360 sps) Figure 138. -Effects of SMCVwake on no. 2 inlet dynamic total-pressure contours during •o '_ sides-lip, _I C v'= 13 degrees and alpha = 3.3 degrees, flight 2-38.

- 1.0 aDlCATfD At-&L~ OF ATTACK MAC.H N..f"8ER ~ -_.

'- ~ "- V'-- -- -v-v ALPHA (OEG.)

MO

[1

0.1 \0 - ALTlruoE BETA :"I.TlflKJf - (r FEET) (OEG.)

_1-ll.LLJ..ll.I.JL.U.1.LU.LLJ..WJCLl..l.LU.LLJ..ULWJ.LU.LUULWJl.U..1.LU.ll.LJ LLLU &0 20 &0 o 20 SECCN>S SECCNlS

~20:00;4b

in 1 etl

I No. 2 in 1e tl

7 deg, RC = 9 deg RB RB 7 deg, RC 5 deg 1.0 RECOVERY

r

-v l/\-vv

'\

PTI ~ PTO PTe

:r

0.8 1.0 STALL MARG IN IIllEX STALL MARGIN I/{)f)(

)

V rv- IOL -"'y IOL

1"4

- U - o r-- 0.2 n..,.~

IV' I

J

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lOT lOT

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-~ [PTl<AX-PTMIN] VG PT o - 1&0 or I--- ~RECTED 11GKT FLJ WJ" WI' - (PPS) (PPSI

,,"~LJ,~ _ JJUJ1..l.l-l-

1&0 0.1 FAN DISCt-'ARGE r-oACH tD. SIGNAL (AFT CONT"OL) - Iv-.

FAN ¥ - 0 -.--- 200 PCWER LEVER AAGLE PLA PLA (;lEG. ) (DEG.)

.r-- t-------.

0- &0 40 &0 20 o SE(CIllS SECCN>S Fib~re 139. - Steady-state lnlet characteristics during no. 2 engine transients, sMcv ~ 20 degrees, alpha = 5.5 degrees, and beta = +4 degrees, flight 2-38.

!>Il!CATED Al'GLE OF ATIACK .oLPHA (DEG.)

", ]~--'~t[l

a- ."

2- S-r----,-----r----,-----,----, I 1>IliCATED Al'GLE OF SIDESLIP BETA (DEG.)

I

~::~>]""""g,~','~,blJ

a 20 40 60 80 100 20 80 lOa SECOI'()S I SECOI'()S f-----21:SI:S8

I No. lin I e tl INo.2 inletl

IDT '~IR WIR (PPS) (PPS) FAN ~p F/>I< -¥ o ' . , "-Ll.LU..LU..LU..LU..u..L.u..L.u..L.w

200_ :- """Ell LEYOl AN::LE

PeA PLA (DEG. ) (DEG.)

- r-r-'-Li ~- - --\.-, r'- _ ll1.-t ' 'I

l

60 80 lOa SEem,DS SECCKJ5 Fi~lre 140. - Steady-state inlet characteristics during no. 2 engine

transients, SMCV = 20 degrees, alpha = 3 degrees, and

beta = +4 degrees, flight 2-38.

1.0- .. _- u_ r----.---~--..._--__,___--__, MO AL_ (oEG.)

BETA AL TIT\X:lE (OEG.) 0 (<FEET) -5 240 o SECQr-()S SECDt<lS . ~.

l---21: 40: 48 20 deg -8 deg ISMCV

•

INo. 2 in 1 etl 1 in 1 etl

'No.

deg, RC = deg RB 7 deg, RC =-5 deg RB 9 1.0 1.0 .. ~.

.J "\. rv REJVERY PTI ~

- PTO

?TO 0.'

1.0 101.

101.

0.2 lOT lOT WI" (PPS) FAN DISCtAAGE r-tACH N>. SIGNAL (pET CONTROL) I.

- o - '00 - H""tE.,;: Lt.VF.R MoGLE Pl"MER LEVER A1'C.LE

t

PLA PLA I CCEG.)

(oEG.)

-lrJ lfL-.-.-

I~

--l rulm~DiATE

lJU ~

\LJ

IDLE , _Wlll.llll o 0 240 240 160 160 80 SEcoms Figure 141. Steady-state inlet characteristics during no. 2 engine transients, ~1CV = 20 degrees and -8 degrees, alpha = 1 degree with positive sideslip, flight 2-38.

SU_V_RY B-I experiencefrom air-induction system flight tests with vortices generatedby the SMCS vanes is summarizedwith the followingobservations: _ (i) Agreement between flight-testand wind-tunneltest resultswas generallygood. Small differenceswere noted in maneuvers (combinations of angles of attack and sideslip)that resulted in wake ingestion.

(2) Inlet total-pressure recovery and distortioncharacteristics can be adverselyaffected by wake ingestion,and considerationof these factorsplays an importantrole in establishinggood inlet / enginecompatibility.

(3) Tests were conductedwith the vanes deflectedand held at a static conditionand also with the vanes driven plus and minus full-scaledeflection at a frequencyof 1 Hz. Tests with oscillatingvanes verified SMCS vane wake / vortex ingestionin flight. Tests with static vane deflectionsproduced resuits consistentwith oscillatingvanes and were less objectionableto the crew.

(4) SMCS operation has not resulted in any flight restrictions. Dynamic values of stall-marginindex remainedwell within allocationsfor all conditions investigated,and there were no flight incidentsrelated to the conduct of the SMCS tests.

(5) Effects on total-pressure recoveryand engine-facedistortionwere generallyrestrictedto windward sideslipoperation combinedwith leading- edge-upvane deflections. The 13-degreevane deflectionangle produced effects similar to the 20-degreedeflectionangle on inlet performance.

(6) Sideslip angles resultingin wake ingestionare smaller during opera- tion at higher angles of attack due to the increasedoutwash from the fuselage forebody. Also, the range of sideslip angles resultingin wake ingestion decreasesas angle of attack increases. No wake ingestionwas encountered above 7.5 degrees angle of attack.

S_Y OF SMCS FLIGHT TEST.RESULTS " The objectivesof the SMCS flight-testprogram were fourfold: (i) obtain specific dynamicresponse data to validate analyticalmodels of the aircraft and control systems, (2) determinedetail SMCS performancecharacteristics, (3) determine impact of SMCS on handling qualities,and (4) determine impact of SMCS on operationalcapabilitiesof the aircraft.

l'our aircraft were built under the B-I contract;A / C-I and A / C-2 were used for detailedengineeringand developmenttests while A / C-3 and A / C-4 were used for operationaltests. As of 16 March 1979, B-I SMCS-relatedtest t i me was as Follows: A / C-I 13 . 6 hr A / C-2 3.3 A / C-3 116.5 A / C-4 0.3 Total 133.7 hr A considerableamount of the flight-testdata has already been discussed in associationwith the topics of previous paragraphsof thi s report. It is tileintent of this section to present flight-testdata not touched upon by the specific topics covered earlier.

_S PERFORMANCE IN TIME-HISTORY-DATA FORMAT Accelerationtime histories are perhaps the most dramaticmeans of demon- strating the ride quality problem solved by the SMCS. In the vertical axis, the sy n _etricfirst-fuselage mode at approximately3 Hz is the big motion producerat the nose of the aircraft due to turbulence. In the lateral axis, tile antisymmetric(lateral)first-fuselage mode at approximately5 Hz pr o duces the largestresponse to turbulence. Figure 142 illustratesthese motions with typical time-historyplots of vertical and lateral accelerationat the nose of tileaircraftat the SMCS vane location. The aircraft was flying through low-altitudeturbulenceat M = 0.75 with the SMCS off at the time that these recordswere taken.

Figure 143 shows data similarto that of figure 142; only during this flighttest, the SMCS was operated to determine its effectiveness. The data sl_cn were recorded on a-flight at M = 0.70 where the B-I was flying at altitudesof 305 to 610 meters (i,000to 2,000 feet) above the terrain in the local EdwardsAir Force Base area. Considerable light to moderate turbulence was present nearly continuously. The SMCS was turned on with both the vertical - and lateral cockpit gains setat 1.5. To demonstratecomparativeaircraft per- formancewith and without SMCS operating,severaltime periodswith the SMCS on and off were recorded. Figure 143 is typical of these data. The pilot station verticalaccelerometerwas not operativeduring this flight, so the readingof the vertical accelerometeron the radome is shown.

2O0 A2008 Vertical acce] 3 Hz first fuselage symmetric at FS 572 (225) (vertical) mode +0.129 g •1.0 g-- -0.129 g _k +0. 132 g 0 g-- -0.132 g - A20I I Lateral acce] 1 5 Hz first fuselage antisymmetric (lateral) mode i i sec SCAS on, SMCS off M = 0.75, alt = 152.4 m (500 ft) Figure 142. - Typical dynamic response near crew station due o to turbulenceduring low-altitude,high-speedflight.

F,; C Oo SMCS off SMCS on Radome u E "- _ Radome E ._ vertical ._ vertical _ _ velocity _ velocity "_. _ A2027 r-,_

[

Fs56(22) _ _

Pilot- us Pilot- u m lateral _ station m c_.

station • lateral °i_

velocity _JL _ +_

A2011 FS 747 (296) Right ° I _ ^ . . ^ A vane F-, dell ,_i Left ° vane defl Figure 143. SMCS performancein turbulence M = 0.70, _t = 305 m (I,000 ft) AGL A= 65°• !

Vertical and lateralmotion at the front end of the aircraft with the SMCS off are shown by the first two time historiesof figure 143. As indi- cated, the primary motion in the vertical axis was the first fuselage bending mode at approximately3 Hz. The lateralmotion was composed of whole-vehicle motion near 1 Hz and the first-fuselageside-bending mode motion of approxi- _ nmtely 5 Ilzsuperimposed. When the SMCS was operated,as shown in the next two plots, considerableattenuationof the 3 Hz motion was achieved. Very little motion of the aircraft at lower frequenciesappears to be present. The effect of the SMCS on the lateral axis motion was not as dramatic as on the verticalmotion, but the 5 Hz motion was partiallysuppressed. The whole- vehicle lateralmotion was not attenuated. It is to be recalled,however, that the SMCS is designed to attenuatestructuralmode response without adversely affectingw|mle-vehiclemotion (handlingqualities). The last two plots in figure 143 show the SMCS right and left vane motion during the time that the SMCS was operating. As shown, the maximum vane deflectionsseldom exceeded 16 degrees, whereas, +20 degrees were available. Both the 3 Hz vertical and 5 Hz lateral structuralmotion can be seen to drive the vane deflections;the largest component is due to the verticalmotion.

Another measure of the SMCS effectivenessis the amount of structural damping the system is able to provide to the key fuselage responsemodes. The upper left-handplot in figure 144 shows that it was possible to excite the s_unetric first-fuselage bendingmode with a sharp horizontal-tailinput pulse.

; : tom the time history of the vertical load factor at the SMCS vane location (afterthe horizontal-tail pulse was removed), it was possible to extract the structuralmode damping ratio, _. Figure 144 shows a plot of _ obtained in this manner v ersus SMCS vertical gain setting _as set in the cockpit). A typicnlnominal gain settingof 1.5 is indicated. The upper right_handplot shows the time-historyresponseof the normal load factor with the vertical gain at this setting. Figure 145 demonstratesthe vertical SMCS performance over a wide range of vehicle weights.

Attempts were made to excite the antisymmetricside-bending modes with sharp lower rudder pulses and to extract structuralmode damping ratios.

- This did not prove to be a successfultechnique. A different excitation techniquewas used. As has been mentioned in an earlier section,A / C-I and A / C-2 of the test B-I aircraft have syst_ns installedallowing the SMCS vanes to be oscillatedat various amplitudes (A) across a range of frequencies up to i0 Hz. This capabilitywas used in an attempt to extract structural mode damping for lateral side-bending modes. The techniquewas to select a resonant frequencyand amplitudesufficientlylarge so asto provide transient data when the forcingmotion was cut off. The data of figure 146 shows that it was possible to excite the 5 Hz first-fuselage lateralbending mode rather cleanly. .Except for SMCS off response,the decay response could not be used to extract the structuralmode-dampingratio because of high-frequency mode SMCS off SMCS on, vert gain 1.5 ! I t I, I ! _ I 0 ] 2 3 0 l 2 3 nz Time- sec nz / '_ Time- sec FS 572 FS 572 (225) (225) n z nz FS 2649 .._ FS 2649 (IO43) (I043) LH 6 cv J _ RH _ cv / __

j-

0.15 Structural I mode O.lO time history damping 0.05 ratio plots 0 , , , , I . • _ , • I 0 l 2 SMCS vert gain setting A = 650 , M : 0.85, alt = 914 m (3000 ft) Figure 144. - First fuselage symmetricstructuralmode damping from horizontaltail pulse excitations.

oM = 0.85, alt = 18i9 m(6000 ft), wt = 131 544 kg (290 000 Ib) , 0 M = 0.85, alt = 914 m (3000 ft), wt = I16,575 kg (257 000 Ib) O M = 0.85, alt = 914 m (3000 ft), wt = 123 969 kg (273 300 Ib) OM = 0.85, alt = 914 m (3000 ft), wt = 144 698 kg (319 000 Ib) M = 0.85, alt = 152 m (500 ft), wt = 127 008 kg (280 000 Ib) A= 65 ° Structural mode frequency,3 Hz 0.20 I I O. 15 f structural I mode 0.l0 I damping I ratio i l Nominal 0.05 Setting g I I " 0 I I I I I 0 0.5 1.0 1.5 2.0 2.5 SMCS vertical gain setting Figure 145. Effect of SMCS vertical gain setting on first fuselage symmetric structuralmode damping at various aircraftweights.

21 1 Lateral acceleration Exciter - - at SMCS vane FS 572 (225) off SCAS off, SMCS off I ±0.J8 I

'

I A=150 I f=5 Hz I SCAS on SMCS off - I i +-0.20 A=150 i I SCAS on, SMCS i I Lat gain 1.0 A=I50 I I SCAS on, SMCS I I Lat gain 1.5 +0.15 gl I A=200 I I I SCA_S o n, SM_C-S I ! Lat gain 2.75 +- O. 18 g_ I A=300 I _ l sec-_ A= 65 ° , M = 0.85, Alt. = 762 m (2500 ft) Figure 146. First fuselage antisymnetricstructuralmode damping from forced SMCS vane oscillations.

contamination. The data do indicate that the SNCSis effective in eliminating structural response motions at the lateral system gains tested.

SMCS PERFORMANCEIN PSD-DATA FORMAT in addition to time-history data, anotherconventionalway of lookingat ride qualityperformanceand the effect of the SMCS on performance is in the form of PSD plots of load factorsat the pilot station. Figures 147, 148, and 149 show typical data of this type for the vertical and lateral load factorswith SMCS off and on. Because of other test requirements,none of the B-I flight-testaircraft had a gust boom installed. Approximatevertical and lateral gust intensities were estimatedusing nose-boomangles of attack and sideslip angles in order to provide normalizingfactors for the PSD data.

Gust intensities for the data shown were estimatedat 1.22 to 1.52 meters per second RMS (4 to 5 feet per second RMS).

l : igure 147 shows that the approximately3 Hz first-fuselage vertical bending mode, previouslyshown in the time-historyplots of figures 142 and 143, ks the main contributorto the verticalmotion at the pilot station. As in figure 143, the data of figure 147 demonstratethat the SMCS is very effective in reducing the pilot vertical load factor.

PSD plots of the pilot station lateral load factor are shown in figures 148 and 149 for the SMCS off and on conditions. Insteadof a consistentsingle- peak responseas in the vertical case, the lateralresponse exhibits two d ifferenttypes of responses,dependingon fuel loading. Figure 148 illustrates a s ingle-peak response as seen in the time-historyplot of figure 142. Fig- gure 149 shows the other common responsewith two peaks between 4.5 to 6 Hz.

In this latter instance,the time, history data appear more random than the lateral accelerationtrace of figure 142. These data were taken with the SMCS forwardsensor package in its original location. The SMCS is seen to signifi- cantly reduce the main response peaks but tends t o excite some of the higher frequencymodes in the immediatevicinity. It was this couplingthat led to . the sensor relocation study discussed in detail earlier.

-- SMCS AND HANDLING QUALITIES One of the design goals for the SMCS was not to interferesignificantly with basic handling qualities. The impact of SMCS operationon the B-I handling qualitieswas determinedfrom horizontal-tail and rudder-doublet transient responses. The results of these tests are shown in figure 150. The wings were at the 65-degreesweep position,Mach 0.85, and altitude 1524 meters (5000 feet). During these tests, the CG locationwas varied.

SCAS only SCAS + SMCS (cockpit gain se t 1.5) --4 0 . 00 0 25 0 .0 025 0.00020 0 . 0020 _nz 0.00015 g2 0.0015' g2 (m-_sec) 2 (ft lsec) 2 tad / see rad / sec 0.00010 0 . 0010 0.0005 0.00005 0 - 0 _ 0 lO 20 30 40 50 60 Frequency - rad / sec I, I I I I I I I I I I 0 l 2 3 4 5 6 7 8 9 lO Frequency - H z Wt = 125 194 kg (276 O00 Ib), CG at .35 _w = 65 ° , M = 0.85, alt _152 m (500 ft)AGL Figure 147. - Vertical SMCS performance in turbulence as shown by power spectral density of vertical load factor at pilot station, FS 747 (294).

SCAS only .... SCAS + SMCS (cockpit gain set 1.5) 0.40 xlO-- 3 "- -- L 0.35 xlO 0.35- 0.30 _ny

o.30- -2J

g2 (m /sec)2 O.25 rad / sec g2 0.25 -(ft /sec)2 rad / sec 0.20 0.20 - 0.15 0.15- 0.10 0. I0 - i ¢ '_\ I \ € '\ 0.05 0.05 - / \ \ ¢/ _ ., , \\ O- 0 _ -_ _-. -_- 0 10 20 30 40 50 60 f Frequency _ rad / sec I I I I I I I I I I I -. 0 1 2 3 4 5 6 7 8 9 10 Frequency-, Hz Wt = 125 194 kg (276 000 Ib), CG at .35 _-w A= 65°, M = 0.85,alt_ 152 m (500 ft)AGI Figure 148. - Lateral SMCS performance in turbulence as shown by power spectral density of lateral load factor at pilot station , FS 747 (294) , single-peak response.

SCAS only SCAS + SMCS (cockpit gain set 1.5) 0.12 0.12 m xl0 -4 xl0 -3 0.I0 0. I0 0.08 0.08

g2

rad / sec (ft /sec)2 II rad / sec I 1t_ (m /sec) 2 g2 !I 13 0.04 0.04 \ / _ / 1 O. 02 O. 02 I I r _ \ / "x .. / / I ,","x 0 _ O J 0 l0 20 30 40 50 60 F r equency- rad / sec i I I I I 1 I I I I I 0 1 2 3 4 5 6 7 8 9 I0 Frequency - Hz Wt = 128 369 kg (283 000 Ib) , C .G. at 0 . 46 Ew A= 65° ' M= 0.70, alt _305 m (1000 ft)AGL Figure 149. - Lateral SMCS performance in turbulence as shown by power spectral density of lateral load factor at pilot station, S 747 (294), double-peak response.

Dutch roll mode 2.0

25" °

frequency - 1.5 rad / sec l.O .5 0 I I I I I 0.2 0.3 0.4 0.5 0.6 C.G - MAC 5- Q Q Short-period mode --_--B frequency - 3 rad / sec 21 _SCAS on, SMCS off --O--SCAS on, SMCS on !

0 I I I I I 0.2 0.3 0.4 0.5 0.6 C.G.-MAC Wt = I 1 7 482 kg(259 0001bs) to 128 822 kg(284 O001bs) A = 65 ° , M = 0.85, alt = 1524 m (5000 ft) Figure 150 . - SMCS impact on short-periodand dutch roll frequencies.

In the longitudinalmode, short-period mode frequencyand damping ratio were extractedfrom the transient followingthe pitch doublets. A slight reductionin the short-periodmode frequencywas observed when operatingwith the SMCS on (figure150) for all CG positions. No significantchange in the short-period mode damping ratio was observed.

In the lateral-directional mode, the Dutch-rollmode frequency and damping -- ratio were extractedfrom the transientfollowingrudder doublets. Results obtained for the lateral-directional handling qualitieswere similar to those for the longitudinalhandling qualities. The Dutch-rollfrequencyis slightly reducedby the SMCS at all CG _ositions. (See figure 150.) Again, no signifi- cant effect of SMCS on Dutch-rolldamping could be detected.

In order to determinewhether SMCS would interferewith aircraftmaneuver- _ng, a roller-coaster maneuver was executed with the SMCS off and on. Evalua- tions of recorded data of the SMCS vane deflectionsshowed no significant motion; maxin_m vane deflectionsrecorded were less than 2 degrees. This was determinedto have a negligibleimpact on the maneuveringof the aircraft.

Subsequentto the previouslydescribedtests, pilots have reported a slight increase in stick force required in terrain followingwith the SMCS on over that with the SMCS off.

SMCS HICH-GAINTESTS B-I A / C-I and A / C-2 were used in the high-gaintests to demonstratesyst_ gain margins over the expectednominal gains. It had been initiallyplanned to do all of the testingof the relocatedforward SMCS sensor package on A / C-I only, includingthe high-gaintests under discussion. However, after only a few of the tests associatedwith the relocated forward SMCS sensor package had been completed,A / C-I went into layup for modifications. In order to continuewith testing, the forward SMCS sensor packagewas relocatedon A / C-2.

A summaryof the high-gaintests on A / C-I and A / C-2 is presented in table X. At nominal gains, there appears to be no differencein the SMCS performanceon A / C,1 or A / C-2. At high gains, however, A / C-2 wan limitedby a 35 Hz limit cycle which was not evident in similarA / C-I data. These data _- show that A / C-I and A / C-2 have nearly a factor of two gain margins over expected nominal settings for the ride quality design point of Mach 0.85 at low altitudes. It is at Mach 0.55 at low altitudeson A / C-2 in a lightweight configurationthat the 35 Hz limit cycle prevents obtainingthe expected nominalgains. From the lightweightconfiguration data of flight 1-41 obtained at Mach 0.85 at high altitude where the dynamic pressure schedulegain in the • o i I TABLE X. - SMCS HIGH-GAINTEST_ SUMMARY i Wing Heavy weight Light weight sweep M = .85, low alt M = .85 low alt (degree s ) 65 Vertical Lateral Flt V ertical Lateral Flt gain gain no. gain gain no.

1.9 2.2 1.9 2.2 expected expected expected expected nominal nominal nominal nominal 4. 8 3.9 2-23 4.5 6. 0 1-48 35 Hz limit cycle no 35 Hz 3.0 6.0 a2-25 35 Hz evident Heavy weight Light weight M = .55, low alt M = .55, low alt 55 Vertical Lateral Flt Vertical Lateral Flt g ain g a in n o. gain gain no.

3.0 3.7 3.0 3.7 expected expected expected expected nominal nominal nominal nominal 4.5 4.7 2-23 3.0 3.1 2-23 35 Hz limit cycle 35 Hz limit cycle S.0 5.5 bl-50 no 35 Hz i a ,Repeat of 1-48 runs b R epeat of 2-23 runs SMCS is about the same as at the low-altitude mach @.55 flight condition,it was expectedthat a vertical gain of six would be obtained. However, as the table shows, this value was not obtained.

These inconsistencies in A / C-I and A / C-2 data led to repeatingwith A / C-2 a set of previouslyrun A / C-I tests and vice-versa. These tests were completed on flights 2-25 and 1-50. Analyses of these data indicate that the 35 Hz gain limitationsare unique to A / C-2. Ground tests were conductedon both aircraft in an attempt to identify the causes of the 35 Hz limitationsof A / C-2; these tests were unable to isolate the causes of the 35 Hz.

CREW EVALUATIONSOF SMCS EFFECTS The bulk of the SMCS operationalsuitabilitytests were conductedon B-I A / C-3. Ten B-I flight crewmen participatedin the tests: four pilots (PLT), three flight-testengineers (FTE),and three Offensive System Operat o rs (OSO).

The crews flew regularlyscheduledmanual terrain-following(MFF)and automatic terrain-following(ATF)flight-testmissions with the SMCS on and off. Crew- members were instructedto maintain awarenessof comfort and personal per- formanceduring the various TF missions. Each crewmemberwas instructedto complete a test questionnaireto document his evaluationof the B-I ride and effectsof this ride on his performance.

The curves in figure 151 through 158 summarizethe subjectiveresponses to questionson overall ride quality and the effect of turbulenceon: flight path / nonflight path control tasks; readabilityof instrumentsand displays; reaching / usingcontrols;crew fatigue,motion sickness,and physical discom- fort. These data are in the form of compositeresponses from all crewmembers.

Overall Ride Quality Subjectswere asked to rate ride quality during TF flight as a function of SMCS on and off for the followingconditions o f turbulence: smooth air and light, moderate, and heavy turbulence. The combined ratings for all L subjectsare shown in figure 151. The data reveal the following: (i) Ride quality ratings decrease (ride qualityworsens) as turbulence increases.

(2) In smooth air, use of the SMCS results in little improvementin ride quality.

(3) Use of the SMCS improves ride quality in light, moderate, and heavy turbulence.

N=7 Poor- \ _ SMCS off a " a , c_ N = lO s N = lO '_ Fair- SMCS on 4 PLT CD 3 FTE _ . 1 , / 30SO 4 J _ p m Good- , , $ II N = 7 N = 7 • _ •_ S S • cr ." 4 PLT N= I0 -o 2 FTE Excellent -- l OSO N = number of test I I I I participants Smooth Light Moderate Heavy Turbulence rating Figure 151. Ride quality ratings for varying degrees of turbulence.

Turbulen c e effect rating N = 4 PLT _orkload Performance effect effect Unable to perform task- Max effort Major-- _ . SMCS off Moderate-- _ Large-- i _ • Incr effort Minor -- None -- Moderate-- __ SMCS on - None [ None-- t I I I Light Moderate Heavy Turbulence rating Figure 152. - Effects of turbulenceon flight path control tasks.

Turbulence effect rating N = lO Workload Performance 4 PLT effect effect 3 FTE .

Unable to perform task- 30SO Max effort Major-- N = 7 4 PLT I Large - N = 7 2 FTE Moderate- N = I0 _-" SMCS off 10SO Incr effort Minor-- f None- = 7 {Moderate--_ _" N_ SMCS on _s.. _ ' None I None - N = lO I.

I I I Light Moderate Heavy Turbulence rating Figure 153. Effects of turbulence on tasks other than flight path control.

Figure 154. Effects of turbulenceon readabilityof instrumentsand displays.

N=IO Extreme -- 4 PLT 3 FTE c N=7 *_ Moderate- _ , ,_ SMCSoff 3 050

-

11 N = 7 N = lO 11.1_ SMCS on

_ . // / . . _ _ / _ _ . 4 P L T \

._u Slight-- X_ // / // - \ 2 FTE _ ./ / " % ..= 7 I OSO None -- N = 10

I I I I

Smooth Light Moderate Heavy Turbulence rating Figure 155. - Effects of turbulence on reaching / using controls.

Extreme -- N = 10 N=7 c \ 4 PLT m _-" 3 FTE "_ -_ ... - SMCS off Moderate-- ._ SMCS on 30SO 4- J u N= I0 N=7 _- 4 PLT SIight -- . o N = 7 2 FTE 10SO "_ N= 10 None -- I I I I Smooth Sight Moderate Heavy Turbulence rating Figure 156. - Effects of turbulence on crew fatigue.

c_ N= I0 c Extreme- SMCS off 4 PLT rO SMCS on 3 FTE 4-' PIoderate- 30SO tl..,, _- N=7 iz1 N = lO 4 PLT Slight- 2 FTE u 10SO t / 3 r- .o None - 4J I I I l Smooth Light Moderate Heavy Turbulence rating Figure 157. Effects of turbulence on tendency for motion sickness.

_- N = lO ._ Extreme -- fD L SMCS off 4 P LT SMCS on 3 FTE U _- Moderate- 30SO N=7 4--1 L.

o 4 PLT OE Slight -- _ _ _k_1, 2 FTE '_ "f _'_ I oso " :g None- N = 7 " U U 3 -- >..

¢- " I I I I Smooth Light Moderate Heavy Turbulence rating Figu r e t58. El;fectsof turbulence on physical discomfort.

One subject added the followingcomment to his ratings: "In smooth air, SMCS is not needed. Standard of comparisonis the F-Ill which is excellent."

(OSOA) Flight-PathControlTasks Pilots were asked to rate the effects of turbulenceon flight-pathcontrol tasks during T F with SMCS off and on. Ratingswere to be made with reference to a turbul_ence-effect rating scale. The combined ratings for all subjects are shown in figure 152. The data reveal the following: (i) The effort to perform flight-pathcontrol tasks and the negative effect on task performanceincreaseas turbulenceincreases.

(2) Greater effort is required , and the negative effect on subject performanceis greater with SMCS off than with SMCS on; i.e.,workload is less When SMCS is being used.

Non-Flight-Path ControlTasks All subjectswere asked to rate the effectsof turbulenceon non-flight- path control tasks during TF with SMCS off and on. Ratingswere made to a turbulence-effect rating scale. The combinedratings are shown in figure 153.

The data reveal the following: (I) The effort to perform non-flight-path control tasks and the negative effect on task performanceincrease as turbulence increases.

(2) Greater effort is required,and the negative effect on subject performanceis greaterwith SMCS off than with SMCS on; i.e., workload is less when SMCS is being used.

One subjectadded the followingcogent, "Operationof equipmentdifficult- particularlyCITS." (FFE C) Readabilityof Instrumentsand Displays ° - All subjectswere asked to rate the degree of difficultythey experienced in reading instrunentsand displays during TF flight as a func t ion of SMCS off and on for four conditionsof turbulence: smooth air and light, moderate,and heavy turbulence. The combined ratingsfor all subjectsare shown in figure 154. The data reveal the following: (1) The difficultyto read instrumentsand displays increasesas turbulenceincreases.

(2) Difficultyratings are higher (readability less difficult)with SMCS "" on than with SMCS off.

One subject added the followingcon_nent: "'E' scope always extremely difficultto read because of location - turbulenceno factor."

Reaching /Using Controls All subjectswere asked to rate the difficultyin reaching controls or in performingcontrol actions during TF flight as a function of SMCS off and on for four conditionsof turbulence: smooth air and light, moderate, and heavy turbulence. The combined ratings for all subjectsare shown in fig- ure 155. The data reveal the following: (i) The difficultyto reach and use controls increasesas turbulence increases.

(2) As turbulenceincreases,it is easier to reach and use controlswith SMCS on than with SMCS off.

Five of the subjectsadded a conlnent, as follows: (i) "In manual TF, pitch stick force too heavy." (PLT C) (2) "Rating is based primarilyon use of central integratedtest system (CITS)." (FTEA) (3) "CITS controls somewhat difficultin turbulence." (FTE B) (4) "Operationof CITS is difficultat best. SMCS helps." (FTE C) (5) "Some of the ratings are due to the locationsof the controls."

(OSO C) Crew Fatigue All subjects were asked to rate the degree to which TF flight introduced any specialtendenciestoward fatigue. Ratingswere requestedfor the SMCS off and on modes for four conditionsof turbulence: smooth air, and light, _ m o derate,and heavy turbulence. The combined ratings for all subjects are shown in figure 156. The data reveal the following: " (I) Fatigue effects increase as turbulenceincreases.

(2) There is little differencein fatigue-effect ratingswith SMCS off versus SMCS on except for the moderate turbulencecondition. For moderate turbulence,the fatigue-effects rating for SMCS on is higher (less fatigue effect) than for SMCS off.

Three of the subjectsadded a con_nent, as follows: (i) "Ratingsmade for AUFO TF." (PLTA) (2) 'Turbulenceis not the primary factor in producingfatigue - TF itself produces a high level of fatigue." (FTE B) (3) "Turbulencedefinitelyincreasesfatigue." ( F TE C) Motion Sickness All subjectswere asked to rate the degree to which TF flight introduced any specialtendency for motion sickness. Ratings were requestedfor the conditionsof SMCS off and on for four levels of turbulence: smooth air and light, moderate,and heavy turbulence. The combined ratings for all cre_nen are shown in figure 157. The data reveal the following: (I) The composite data show little tendency for motion sickness for all turbulenceconditionsfor both SMCS off and on modes. F TE were more affected than pilots.

: One subjectcommentedthat: 'Turbulenceis not a major influence. The rough- ness of terrain has more impactwhen you are stuffed in that 'blackhole' with no windows." (FTE C) PhysicalDiscomfort All subjectswere asked to rate the degree to which TF flight introduced any specialtendenciestoward physical discomfort. Ratingswere requested :for the conditionsof SMCS off and on for four levels of turbulence: smooth air and light,moderate, and heavy turbulence. The combined ratings for all crewmen are shown in figure 158. The data 'revealthe following: -" (I) Physicaldiscomfortincreasesas turbulenceincreases;although at heavy turbulence,the effectsare rated only as slight.

(2) Less physical discomfortis reported for light, moderate, and heavy turbulencewith SMCS on than with SMCS off. For smooth air, there is no differencein physical discomfortratings for the SMCS off versus SMCS on modes.

One subject commentedthat: "Ph y sicaldiscomfortresults from feelingsof irritation,aggravation,and anxiety produced by the rough ride and high workload,hard to pin down further." (PLT D) AdditionalRide Quality Observations All subjectswere asked to add any additionalobservations(not covered by the ride quality questions)concerningride quality characteristics or SMCS effectsduring TF flight or any other characteristics realting to TF which have a bearing on crew comfort and efficiency. Comments includedthe following: (i) "SMCS is very effectiveand required in the B-I to aid the flight crew in performanceof the TF task." (PLTA) (2) "Essentialfor effectiveB-I MrF, desired for effectiveB-I ATF."

(PLT C) L (3) "Turbulencemakes this aircraft hard to stabilizeon a bank angle, -- adds to an already high workload." (PLT D) (4) "Ride qualitieswith SMCS are definitelybetter than without - the differenceI don't think is really able to be seen in the layout of this questionnaire." (FTEA) (5) "Lack of outside visual reference contributessignificantly to disorientation and motion sickness. Mo v ement about the crew compartmentis hampered by TF, especiallyin turbulence. Any movement in the crew compartment also contributesto disorientation and motion sickness. There is a significant improvementin the ride with SMCS on." (FTE B) (6) "I consider SMCS essentialfor long term TF flight." (OSOA) (7) "Due to part of the controls and displays location, it is difficult to accomplishsome weapon-orientedtasks under turbulentconditions. The sensitivityof both the navigation (NAV)panel and Stores ManagementSystem (SMS)FWD / REV switch make it difficultto use under turbulentflight condi- tions." (OSO B) (8) 'The discomfortsin the OSO stationare are severe in light or moderate turbulencewith SMCS off due to the fact that we have no outside reference." (OSO C) Handling Qualities Pilots were asked to describe the effects, if any, of SMCS activationon aircraft handling qualitiesduring MTF. Commentsincluded the following: Pitch Control (I) "A / C ride is smootherwith SMCS on and thereforecontrol is easier."

(PLT A) (2) "Assists by damping." (PLT B) (5) "COUld have some detrimental effect because rapid control inputs are apparently countered b y SMCS -- not considered a problem to date , more evaluation required." (PLT C) (4) "Increases pitch forces , se_ns to slow aircraft response." (P L T D) L ateral / Direct ional Control (1) "A / C ride is moother with SMCSon and therefore control is easier."

(PLT A) (2) "Assists by damping." (PLT B) APPENDIX NONENCIATURE This report was the result of the contributions of a number of authors and each has used nomenclatureunique to his particulardiscipline. In order -_ to help the reader to quickly locate a given symbol, this sectionhas been organized so that a general section, is presented first followedby nomenclature associatedwith three sectionsof this report which are especiallyheavy in specializednomenclature. These sections are: "FlexibleAircraft Equations o[ Mot i on," "Impact of SMCS on Selected Loads," and "_S Vane Effect on Inlet / Engine Characteristics."Under this system, similar symbols often have differ- ent meanings; the reader is cautionedto identify symbolswithin the context of their use.

GENERAL A amplitudesetting of B-I oscillatingsystem for the SMCS vanes A / C aircraft alt altitude ATF automatic terrain following AUTO automatic BP butt plane b [ _F referencelength mean aerodynamicchord CITS central integratedtest system CG center of gravity cm centimeters deg degrees El bending stiffness [ frequency,cycles per second F, Flex subscriptdenoting flexible F1t flight ft feet ];l ' Ji flight test engineer [ :/ R flexible-to-rigid ratio

[

[

]R flexible-to-rigid ratio of bracketedparameter FS fuselage station g Accelerationof gravity CJ torsionalstiffness GVT ground vibrationtest crew sensitivityindex; subscriptZ denotesvertical axis, li() Y denotes lateral axis Hz hertz (cyclesper second) hr hours in. inches _K _bREF k reduced frequency,_ , V o o K SMCS gain Khp SCAS gain scheduledwith altitude kg kilogram - KMN n yaw SCAS lateralaccelerometergain Y kq pitch SCAS gym gain k_ SMCS gain scheduledwith dynamicpressure q K_ r yaw SCAS gyro gain Knz pitch SCAS normal accelerationgain ib pounds L-gain lateral SMCS gain Lh dimensional, force along Z-axis due to plunging motion h, + down Le , dimensionalforce along Z-axis due to pitching motion o, + down J [ALO]Real [L@I Real V° [Lh]Imag Lni dimensionalforce along Z-axis due to structuralmode generalizedcoordinatemotion hi, + down Lwg dimensionalforce along Z-axis due to vertical gust velocity V C g,+ down L_ dimensionalforce along Z-axis due to control surfacemotion _, + down m meter M Mach number MrF manual terrain following Mh dimensionalmoment about Y-axis .dueto plungingmotion h, + nose up Mo dimensionalmoment about Y-axis due to pitchingmotion 6, + nose up Mni dimensionalmoment about Y-axis due to structuralmode gen- eralizedcoordinatemotion hi, + nose up Mwg dimensionalmoment about Y-axis due to vertical gust velocity w , + down g N newton N ntmber of test participants Qih dimensionalgeneralizedforce in structuralmode i due to plungingmotion h, + for ni increased Qi@ dimensionalgeneralizedforce in structuralmode i due to pitching motion @, + for Ni increased [AQi_IRoal IQi0]Real-(_) [Qih]imag Qini dimensionalgeneralizedforce in structuralmode i due to structuralmode generalizedcoordinatemotion hi, + for ni increased Qiwg dimensionalgeneralizedforce in structuralmode i due to vertical gust velocityWg, + for Ni increased Qi6 dimensionalgeneralizedforce in structuralmode i due to control-surface motion _, + for ni increased SL sea level V-gain vertical SMCS gain WL waterline damping ratio @ pitch angle about elastic axis A sweep angle of leading edge of lifting surface a Vertical gust intensityderived from angle-of-attack vane _V measurements OBv lateral gust intensityderived from sideslip angle vane measurements elasticaxis-bendingslope •( ) power spectral density of subscriptparameter frequency,radiansper second RIDE QUALITY EQUATIONSOF MOTION RELATED qo 1 / 2 pV2, dynamic pressure p density of air AP _o pressure coefficient Vo resultantvelocity of the CG componentof resultantvelocity; subscript denotes axis V r ) along which component acts Wg vertical component of gust velocity Vg lateral componentof gust velocity Sw wing area bw wing span _w wing mean aerodynamicchord Z vertical deflection Y side deflection Zx' Zy' _z distance along the x, y, and z-axis,respectively - - distance from vehicle CG to control surfacek hingeline - _' z( ) (+ aft), subscriptidentifiessurface distance from control surface hingelineto surface CG _' _( ) (+ aft), subscriptidentifiessurface distance from X-axis to surfaceCG in-planeperpendicularto ) plane of symmetry (always+), subscriptidentifiessurface.

_°F perpendiculardistance from CG to thrust axis; + down W airplane weight m airplane mass mass of control surface, subscript identifiessurface m, m() M. the ith mode generalizedmass, f f f dX®dz y,z) Am(x,y, Ix, Iy, Iz moment of inertia about the X-, Y-, and Z-body axis, respectively I product of inertia; positive when the principal X-axis is xz below the body axis at the nose of the vehicle IR engine rotor moment of inertia I( )[L moment of inertiaabout hingeline; ( ) subscriptidentifies surface n( ) load factor; subscriptdenotes axis along whichcomponent acts Euier azimuth angle @ Euler pitch angle Euler roll angle see figure 6 X, Y, Z body-axis coordinates Xe' Ye' Ze earth-axiscoordinates 11 altitude (+ up from sea level) p rolling rate about X-body axis • p rolling accelerationabout X-body axis q pitching rate about Y-body axis see figure 2 q pitching accelerationabout Y-body axis r yawing rate about Z-body axis r yawing accelerationabout Z-body axis _R rotationalrate of engine rotor relative to airframe _. natural frequencyof ith mode i forcing frequency angle of attack; angle between the projectionof the resultantvelocity vector on the XZ-plane and the X-body (referenceaxis) _ see figure 5

|

sideslip angle; angle between the I resultantvelocity vector and the

J plane of symmetry XZ

I" control-vane dihedral angle 6 con.trol-surface deflection;positive deflectionproduces ( ) positive force (+CN,+Cy)(see figure 3); subscript identifies surface control-surface acceleration;positive in the sense that 6( ) , .

6 ( ) is positive; subscript identifiessurface 6 H rolling tail control differentialdeflection,+ deflection produces +Cz g accelerationof gravity gsi structuraldamping constant,mode i n. deflectionof the ith normalizedstructural i mode at normalization point n. rate of change of the ith mode at pQint of i normalization _. accelerationof the ith mode at point of i normalization > see figure 4 ¢.( ) the i normalizedmode shape; i.e., ratio of i local deflectionto deflectionat normalizing point (nondimensional) ; ( ) superscript denotes location _! )' slope of the ith normalizedmode; ( ) i superscriptdenotes location T _! ) fuselage torsionalangle, ( ) superscriptdenotes location F force Z aerodynamicforce in Z-direction N aerodynamicnormal force (N = -Z) L aerodynamiclift force (L_N for small _) C aerodynamicchord force (C = -X) D aerodynamlcdrag force (D _ C for small _) X aerod y namicforce in X-direction _ see figure 3 Y aerodynamicforce in Y-direction T thrust (T = X) L aerodynamicrollingmoment about X-axis M aerodynamicpitching moment about Y-axis N aerodynamicyawing moment about Z-axis Nwg normal force due to unit vertical gust velocity Mwg pitchingmoment due to unit vertical gust velocity Q[ generalized force in ith mode fff F(x , y ,, ) i(x , y,z)dxdydz ..

generalized force in structural mode i due unit vertical Qiwg gust velocity Yvg side force due to a unit lateral gust velocity rollingmoment due to a unit lateral gust velocity Lvg Nvg yawing moment due to a unit lateral gust velocity Nni yawing moment due to mode i deflection Yni side force due to mode i deflection M_i pitching moment due to mode i deflection Nni yawing moment due to mode i deflection Lni rolling moment due to mode i deflection Qij generalizedforce in mode i due to mode deflectionj C chord-forcecoefficient CC = Swqo DCC = --chord-force coefficientdue to control-surface deflection, CC6( ) a6( )subscriptidentifiessurface N . normal-forcecoefficient

CN = %%

OC N CNf O ff normal-force curve slope 8CN normal-forcecoefficientdue to downwash lag and CN_ = _Cw vertical acceleration 8CN CNq = - / q Cw_ normal-forcecoefficientdue to pitch rate 8CN CN__ = a / Q_.2\l wl normal-forcecoefficientdue to pitch acceleration \4V2 / 8CN CNni = 8Ni normal-forCedeflection coefficientdue to structuralmode 8CN normal-force coefficient due to structral mode CNni = 8_, deflection rate 8CN normal-forcecoefficientdue to control surface CN6 = 86( deflection,subscript identifiessurface

() )

8CN normal-forcecoefficientdue to control surface CN6( = ) O n deflectionrate, subscriptidentifiessurface

• ()

M Cm = ° w_w_=-o pitching-moment coefficient 8Cm Cm_ = 8_ pitching;moment curve slope _Cm pitching-momentcoefficientdue to downwash lag Cm t _ = &I \ and vertical acceleration 3Cln __ Cmq = __qC-w_ pitching-momentcoefficientdue to pitch rate H I

- \2Vo !

aCm Cmq = /_F . ,2\ pitching-moment coefficient due to pitch acceleration _Cm pitching-moment coefficient due to structural Cm_i = 3ni mode deflection 3Cm pitching-momentcoefficientdue to structuralmode i _ deflectionrate Cm pitching-moment coefficientdue to control surface

Cm6( = ) 36( ) deflection,subscript identifiessurface

8Cm pitching moment due to control surface deflection rate, subscript identifiessurface

) )

° = a subscript_, B, etc, as shown indicatesthat ()(_) bracketedparameter is a nonlinear functionof the subscriptvariable Qi generalized-force coefficientin the ith Cni = Swq ° mode C . = __8C_i_ generalized-force coefficientdue to angle of attack _i_ 8_ in ith mode 8Cni generalized-force coefficientdue to downwash lag C_i& = 81_Cw_1-----'_ and vertical acceleration

VV° I

aCni generalized-force coefficientdue to pitch rate C_iq = / qC-w_ in ith mode 3Cni generalized-force coefficientdue to pitch Cni_l -tqc w 21 acceleration in the ith mode _CNi generalized-force coefficientdue to the jth mode C_I_j" = _n]. shape in the ith mode 3CNi generalized-force coefficientdue to rate of change "" fli of jth mode in ith mode C_i_J -- (-_o) aCni generalized-force coefficientdue to control-surface deflection in ith mode, subscript identifies control C r Ji_( ) a 6 ( ) surface aCNi generalized-force coefficientdue to control-surface

= ) a 6 ( ) k deflectionrate in ith mode, subscript identifies

control force Y Cy -- - side-force coefficient qoSw aCy Cy_ = _s side-forcecoefficientdue to angle of sideslip _Cy side-forcecoefficientdue to rate of change of

Cy_i = O / _bw\ angle

12---_° _ sideslip (lateralacceleration) by symmetricstructuralbending (dihedral) CyB q i = 3CyB side-forcec o efficientdue to sideslip angle caused - _ i _Cy Cyr = side-forcecoefficientdue to yaw rate

01 l

\2Vo !

2%

CYr = a / _b,.2 \,___I side-forcecoefficientdue to yaw acceleration \4Vo2 / _Cy Cyp = / pb w \ side-forcecoefficientdue to roll rate _Cy Cy_ = a / _bw2 \| | side-forcecoefficientdue to roll acceleration \4Vo2 / _Cy side-forcecoefficientdue to control-surface Cy 6 = ( ) _6( ) deflection,subscript identifiessurface _Cy side-forcecoefficientdue to control-surface

C y_( = _( ) ) deflectionrate, subscriptidentifiessurface _

J N Cn = Swbwqo yawing-momentcoefficient _Cn C_ fl = _ yawing-momentcoefficientdue to sideslip angle _Cn yawi n g-momentcoefficientdue to rate of change Cn fi = _bw) of sideslipangle (lateralacceleration) ;CnB y awing-moment coefficient due to sideslip ang l e Cn_ni _i caused by symmetricstructuralbending (dihedral) _Cn Cnr = / rbw \ yawing-momentcoefficientdue to yaw rate I I _Cn Cn. = yawing-momentcoefficientdue to yaw acceleration r i_bw2_ _Cn Cnp = 'Jpbw\l I yawing-momentcoefficientdue to roll rate

\2Vo /

_Cn Cn_ = D /pbw2 _ yawing-momentcoefficientdue to roll acceleration \4Vo J.

_Cn = yawi n g-moment coefficientdue to control surface Cns( ) 2 6 ( ) deflection,subscriptidentifiessurface aCn yawing-momentcoefficient due to control surface (:n_ = ( ) a_( ) deflection rate , subscript identifies surface L CI = rolling moment Swbwq o -- aCl Cz_ = aB rolling-moment coefficientdue to sideslipangle aC_ rolling-moment coefficientdue to rate of change -- C_ = 81_bwl of sideslipangle (lateralacceleration)

\ -vo !

aCz8 rolling-moment coefficientdue to sideslip angle aNi caused by symmetric structuralbending (dihedral) C_Bni = ____ aC_ C_r = / rbw\ rolling-moment coefficientdue to yaw rate _C_ C_r = / +h2\ rolling-momentcoefficientdue to yaw acceleration aC Cgp = o / pb\|___Z_.l rolling-moment coefficientdue to roll rate ac = _ Cg_ / nh2\ rolling-momentcoefficientdue to roll acceleration

i wj

8\4V02 / aC£ rolling-moment coefficientdue to control surface C_6( ) a6( ) deflection,subscript identifiessurface aC_ rolling-moment coefficientdue to control surface _ C_( ) an( ) deflectionrate, subscriptidentifiessurface aCNi generalized-force coefficientdue to sideslip angle C_i = a6 in ith mode _CNi generalized-force coefficientdue to sideslip angle Cni" = 6 l_bw_ rate of change in ith mode aC.ni = 6 generalized-force coefficientdue to sideslip in the Cni6 an. ith antisymmetricmode due to bending in the jth nj J symmetricmode aCni generalized-force coefficientdue to yaw rate in Cnir = Irbw_ ith mode aC_i generalized-force coefficientdue to yaw acceleration Cni" = • 2_ in the ith mode r rbw _ aCni generalized-force coefficientdue to roll rate in CHip = pbwl ith mode

\2Vo !

_C_i generalized-force coefficientdue to roll acceleration Cni_ = i _bw2 _ in the ith mode a t[ 1 4Vo 2 I qoSw K s = mVo- qoSw_w Kq = ly qoSw K# = mVo qoSwbw Kr = Iz qoSwhw Kp = I x qoSw Kni = Mi i / 1- i, j used as subscriptto identify structuralmode k subscriptidentifyingcontrol surface II subscript identifying horizontal-tail control surface r subscript identifyingrudder cv subscript identifyingstructuraln_de controlvane r R subscript indicatingreal part I subscript indicatingimaginarypart o subscript indicating trim value tL hingeline I;R L fuselage reference line MAC mean aerodynamicchord ( )TF aerodynamictransfer function - LOAD EQUATIONSOF MOTION RELATED a acceleration l A gust response factor Oo /O w b turbulencefield parameterdenoting gust intensity I B ] phased loading conditions,one column per condition,each elementproportionalto P i- 3 °" _" i 3 I E J expected load values for each load it_n {F} forces at each SIC point: real or in the frequency domain, complex g structuraldamping parameter Ih I real generalizedcoordinates H(_) frequencyresponse function for a load item II*(_) denotes complex conjugateof H(_) Ill(w)] frequencyresponse functionsof load items, one row per item for i01 frequencies

j #T

- [LOADS ] loads; shears,moments, and torques; real or in the frequency domain, complex LOAD geometry to compute shears,moments, and torques,one row GEOM per load item, one column for each SIC load point I D AD I load generationmatrix, one row per load item, one column GEN 3 for each normal elasticmode EM_ mass matrix [_M] generalizedmass matrix for rigid-bodymotions [¢]T [M_ generalizedmass matrix = [M][_], rigid and elastic modes -_ \ N numbe r of c r ossings of zero per hou r with a positive slope o N(y) average number of level crossingswith a positive slope equally or exceedingy per hour P turbulencefield parameterdenoting proportionof time in turbulence q I complex generalizedcoordinates QM] generalizedaerodynamicforces due to modal motion Qc] genezalizedaerodynamicforces due to unit deflectionsof control surface, column 1 for horizontaltail and c61umn 2 for the mode control vane _iI generalizedforces due to a unit sinusoidalgust R] rigid-bodymode-modifying matrix S Laplace operator S_ phased loading conditions,one column per condition,each column scaled so that the diagonal element equals the corre- .-'- sponding expected load value [ Sl C] structural flexibility influence coefficients ISLOADS I phased loading column from [S], shears,moments, and torques IT .] transformation matrix, for control system feedback,relating control-surface deflectionsto the generalizedcoordinates V veloci ty _H' _cv control-surfacedeflections;subscriptH for horizontal stabilizer,subscriptV for mode control vane mode deflection L [ ¢] J mode shapes by columns, superscriptdenotes type of modes; RBM for rigid-body,and E for normal elastic _w' _o root mean-squarevalues of gust velocity and output item response,respectively oi, o. root mean square value of response items i and j J _(_) gust power spectrum,normalizedon unit gust intensity, o2 W (_) load item output response power spectrum o pitch rate o .. correlationcoefficientexpressingthe degree of statistical ij linear dependencebetween load item i and load item j frequency structuralnormal elastic-modefrequency ] square or rectangularmatrix; where i = number of rows, and i,j j = number of colunns E _ diagonal matrix cQlumn matrix , _. L _] row matrix [_T matrix transpose [] matrix inverse ENGINE / INLET RELATED A / l) analog to digital AIP inlet / engine aerodynamic interface plane AIS air induction system Aduct duct area, psi / count CAL calibration CBW constant bandwidth CTS counts CTSCAL counts output during calibration step CTS_4EAN mean value in counts during operate step CTSOP counts output during operate step CTSZ counts output during zero step db decibels I!CI-8 inlet configuration identification FS full scale HP high pass IDC circumferential distortion component _.

IDL engine stall-margin ratio (function of IDC and IDR) IDR radial distortion component ips inches per second K i000 feet LP low pass 25O NF engine fan speed N2 nitrogen P pressure _P pressure difference PCM pulse-codemodulation PDYN Instantaneous pressure (dynamic) PLA engine power lever angle PPCM time-averaged pressure (steadystate) pps pounds per second psi pound per square inch PT total pressure PTI average total pressure at inlet / engine AIP V rI local total pressure at inlet / engine AIP PTO free-streamtotal pressure FIIv_X maximum total pressure at inlet / engine AIP V FMIN minimum total pressure at inlet / engine AIP RB first movable inlet ramp angle RC second movable inlet ramp angle R2 referencepressure SMCV structuralmode control system vane sps sample per second VCO voltage controlledoscillator WIR correctedengine airflow X2 amplifier (times 2) X6 amplifier (times 6) ZOC a three-way value (refer to page 161) t_ Z_ fractional increment on total pressure contour, AIP 6002 reference pressure, absolute units 6631 calibration pressure, absolute units -_ REFERENCES I. Wykes, John H.; Borland, ChristopherJ.; Klepl, Martin J.' ; and MacMiller, _ Cary, J.- Design and Developmentof a StructuralMode Control System, NASA CR-143846,October 1977.

2. Etkin, Bernard: Dynamicsof Flight. John Wiley and Sons, Inc., New - York.

3. Bisplinghoff,Raymond L.; Ashley, Holt; Halfman, Robert L." Aeroelasticity.

Addison-WesleyPublishingCompany, Inc., Reading,Massachusetts.

4. Landahl,M. T.: GraphicalTechniquesfor Analyzing MarginallyStable Systems. Journal of Aircraft, September-October 1964.

5. Rustenburg,John W.: Developmentof Tracking Error FrequencyResponse Functionsand Aircraft Ride Quality Design Criteria for Vertical and Lateral Vibration. ASD-TR-70-18,January 1971.

6. Stenton, Thomas E.: TheoreticalFrequencyResponse Functionsand Power Spectra of the XB-70 Response to AtmosphericTurbulence. NASA CR-1621, August 1970.

7. Dugundji, John: On the Calculationof Natural Modes of Free Free Structure. Journalof the AeronauticalSciences,February 1961.

8. Flight Control System Description, Rockwell International, North American Aircraft Division,E1 Segundo,TFD-71-807_June 1971 (Revision August 19753.

9. Austin, William H., Jr.: Developmentof ImprovedGust Load Criteria for United States Air Force Aircraft. SEG-TR-67-28,September1967.

i0. Fuller,J. R.; Richmond,L. D.; Larkins, C. D.; and Russell, S, W.: Contributionsto the Developmentof a Power Spectral Gust Design Procedure for Civil Aircraft. FAA-ADS-54,January 1966.

J 1. ReportNo . 2. Government Accession No. 3. Rec i pient ' s Cata l og No .

NASA C R- 1 44 8 8 7 4. Titl e and S u b title 5 . Re po rt Date January . 19 8 0 ANAL Y S E S AN D TE S T S O F THE B- 1 A I R C RA FT 6. Performing Organization Cod e STRU C TURAL MOD E C ONTROL SYSTEM 7. Author(s) 8 . Pe rf orming Or gan iz a tion Repo rt No .

John H . Wy k es , Thomas R . Byar , Cary J . M ac M il l er, and NA- 79 -4 05 David C. Greek 10 . WorkUnitNo.

9 . Performing Organization Name andAddress Rockw e l l International N ort h A merican A i r c r aft Division 11 . Co ntr a ct or Gra n tNo .

815 Lapha m Street NAS 4 -2519 El Se gu ndo, CA 902 4 5 " 13. Typeof Report andPeriod Co vered 12 . S po n s oring Agency Name andAddress Contracto r Report - Final N ationa l A eronautics and S pace A dministration 14 . Spon s oring A ge ncyCode Washington, D.C. 20546 H- 11 09 15. Supplementary Notes NA SA Technical M onitors: Jim M e K a y a n d L arry Felt , Dryden F l ight R esearch Center NASA Progra m Manager: Jack Nugent, Dryden Flight R esearch Center 16. Abstract An 18 -mon th p rogr am w a s cond u cted t o c o mpi l e a n d doc um ent for p ubl icati on i nformation perta i n in g to analy s e s and flight tests of the B- 1 St ructura l M ode Contro l Sy s tem (S M CS). Thi s i s the second pha s e of a c _ ntinuing effort; re su lts f ro m the fir s t p h ase study are docu m en te d in De _;i gn and Development o f a St ructural M ode Contro l Sy s te m , by Jo hn H. Wykes , Chr is topher J. B or l and, M artin J. Klepl , a nd Cary J. M ae M iller (NASA C R - 1 43 8 46 , Octo b er 1 977). Thi s repo rt covers the _ fo ll owin g topics: _ ' _ (1 ) F l exib l e aircraft equations of motion (2) Description of flexib l e aircraft ana l yses mode l (3) Comparison of analyses and flight-test performance results of the S M CS ( 4 ) A summary of the study of the forward S M CS sensor package relocation ( 5 ) Truncated ana l ytica l mode l s us ed in simulation effo r t (6 ) An analysis of the S M CS vane in terference effects (7) I m pact of S M CS on selected l oads (8) Flight - test resu l ts of the S M CS vane effects on in l et / engine c ha racteristics r (9) S ummary of SM C S f l i g ht - test results 17. Key Words(Suggested byAuthor(s)) 18 . Distribution Statement S tructural m ode control system Large flexible aircra ft Analyses and tests S T AR catego ry : 05 19. SecurityCla s sif. (ofthisre po rt) 2 0. Security Cla s sif. (ofthispa ge } 21. No.of Pa ges 22. Price * Unclassified Unclassified 268 * Fo r s al e b y th e Na ti o n al Te ch ni cal I nf orma t io n Se rv i c e , Sp r i n gfield , Vi rgi n ia 2 2151 i" ,t D