£USAAEFA PROJECT NO. 80-17-1

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€ AIRWORTHINESS AND FLIGHT CHARACTERISTICS TESTPART 1

YAH-64,ATTACK HELICOPTER

FINAL REPORT

PATRICK M. MORRIS GARY L. BENDERMAJ, EN GR .BNEMAJ, EN PROJECT ENGINEER "

PROJECT OFFICER P E

JOHN D. OTrOMEYER BARTHOLOMEW D. PICASSO,IIIPROJECT ENGINEER MAJ, IN

PROJECT PILOT

LARRY B. HIGGINS RICHARD SAVAGEMAJ, IN CPT, AR

PROJECT PILOT PROJECT ENGINEERDTIC

•AUG 2 3 1982'-"

0-x. SEPTEMBER 1981 :

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AIRWORTHINESS AND FLIGHT CHARACTERISTICS TEST FINAL REPORT.. PART I YAH-64 ADVANCED ATTACK HELICOPTER 30 May - 17 July 1981

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* GARY L. BENDER JOHN D. OTTOMEYERLARRY B. HIGGINS BARTHOLOMEW D. PICASSO, IIIRICHARD SAVAGE PATRICK M. MORRIS

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Airworthiness and Flight Characteristics Test Prototype YAH-64 HelicopterPerformance Qualities Handling Qualitie-Horizontal Stabilator Yaw SAS Hardovers

2L A=TRAC'r cm mw vee N new, and Identify by block niber)The Airworthiness and Flight Characteristics Test Part I of the prototype YAH-64 helicopter(S/N 77-23258) was conducted at three test sites: Palomar Airport, Carlsbad, California (elevation328 ft).(Bishop Airport. Bishop, California (elevation 4120 ft), and Coyote Flats, California telcuation9980 ft).) Approximately 32 productive hours were flown during 45 flights between 30 May and

4 '17 July 1981. Several design changes affecting performance and handling qualities were made to theYAH-64 since the last evaluation (EI)T 4) the most significant of which were the change in tail rotorrigging, the addition of a third l)( electrical bus and the incorporation of aNap-of-the-Earth/Approach mode of operation for the horizontal stabilator. Level flight perlormancc

Land vibration levels have been slightly degraded since EDT 4. while handling tiualitics and

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out-of-ground-effect hover performance remain essentially the same. Numerous anomalies in theoperation of the Digital Automatic Stabilization Equipment (DASE) were experienced which mayrender some of the handling qualities test results suspect. Two deficiencies (both DASE related) werefound during this evaluation: Yaw SAS hardovers, and disengagement of the DASE with failure of theNo. 2 eenerator. Nine previously unreported shortcomings (4 DASE related) were found.

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DEPARTMENT OF THE ARMYH%4 US ARMY AVIATION RESEARCH AND DEVELOPMENT COMMAND0430 GaooDELLOW BOULEVARD, ST. LOUIS. Mo 6313D

DRDAV-D

SUBJECT: Directorate for Development and Qualification Position on the FinalReport of USAAEFA Project No. 80-17-1, Airworthiness and FlightCharacteristics Test, Part 1, YAH-64 Advanced Attack Helicopter

* SEE DISTRIBUTION

1. The purpose of this letter is to establish the Directorate for Developmentand Qualification position on the subject report. The objectives of this testwere to assess the handling characteristics, performance and vibrationcharacteristics of the YAH-64 helicopter upon completion of full scaleengineering development. Due to schedule restraints, the test was divided intothree parts, of which this is the first.

2. This Directorate agrees with the report findings and conclusions, with theexceptions identified herein. Dispositions of redesigned subsystems/componentsaffecting the conclusions are also identified. Conclusions are discussed byparagraph as indicated.

a. Paragraph 52d. The No. 3 DC electrical buss was incorporated toprevent unacceptable voltage transients in flight critical systems following afailure in either the No. 1 or No. 2 electrical buss. This modification wasincorporated in all four prototypes, and was tested successfully in the threeother aircraft. It is believed the unsuccessful implementation is unique 1U'eto instrumentation, installation and peculiar electrical system for this test

-~ aircraft.

b. Paragraph 53a. Yaw SAS hardover failures were caused by failure of a* power supply internal to the DASE. The power supply output voltage was

monitored for total failure, but a "soft" failure where an incorrect voltage* . was output would not have been detected. The voltage monitor threshold will be

changed to detect off-voltage conditions as well as full failures.

c. Paragraph 53b. Disengagement of the DASE following failure of the No.2 generator is peculiar to this aircraft. The other prototype aircraft do notexhibit this characteristic with the addition of the No. 3 DC electrical buss.

d. Paragraph 54a. The extraneous illumination of the Master Caution light

was caused by the reset switching mechanism. A different design will be

incorporated for production.

e. Paragraph 54b. The APU configuration installed in the aircraft did nothave all changes defined during qualification testing. The APU start system,used for the APU qualification tests, successfully demonstrated APU starts

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DRDAV-DSUBJECT: Directorate for Development and Qualification Position on the Final

*Report of USAAEFA Project No. 80-17-1, Airworthiness and Flight

Characteristics Test, Part I, YAH-64 Advanced Attack Helicopter

* throughout all required environmental conditions. With the production design,the reported APU start failures should be eliminated.

f. Paragraph 54c. The environmental control system is currently underredesign to provide adequate cooling/heating to the crew and avionics.

g. Paragraph 54d and f. The erroneous activation of the engine out/mainrotor RPM and stabilator audio warning tone, with failure of the No. 1 or No. 2generator is unique to the test aircraft as discussed previously in thisletter.

h. Paragraph 54g. The persistent yaw oscillations were attributed to theDASE power supply failure previously mentioned. The phenomenon will beevaluated during A&FC Part 2.

i. Paragraph 54h. The stabilator system contains a feature to switch frommanual mode to automatic mode as the aircraft accelerates above 80 knots.There was an anomaly in the programming which resulted in inconsistentswitching to automatic mode when the stabilator was at 35 degrees LEU. Themonitor program was altered to correct the discrepancy and will be checked inA&FC Part 2.

j. Paragraph 55. The position on the shortcomings identified in thisparagraph is reported in the EDT 4, and is unchanged.

FOR THE COMMANDER:

CHALES C. CRAWFORD, JR.Director of Developmentand Qualification

IAq

TABLE OF CONTENTS

Page

INTRODUCTION

Background............................................... ITest Objectives............................................. I

m *Description ............................................ ITest Scope................................................ 2Test Methodology........................................... 4

RESULTS AND DISCUSSION

General................................................... 5Performance............................................... 5

General............................................... 5Hover Performance...................................... 5Vertical Climb Performance ................................ 5Level Flight Performance .................................. 6

Handling Qualities.......................................... 6General............................................... 6Control Positions in Trimmed Forward Flight ................... 6Static Longitudinal Stability ................................ 7Maneuvering Stability ..................................... 7Dynamic Stability....................................... 7Low-Speed Flight Characteristics ............................ 7

Forward and Rearward Flight ........................... 8Sideward Flight......................8Low-Speed Maneuvering ............................... 8

Stabilator Evaluation..................................... 8Control Positions in Trimmed Forward and

Rearward Flight.................11-lLevel Flight, Climbs, and Descents ....................... 11

Static Longitudinal Stability ............................ 12Level Flight Accelerations/Decelerations ................... 12Stabilator Malfunction During Acceleration ................ 12Contour Approach /Landing ............................ 13

Aircraft Systems Failures .................................. 13Stabilator Failures.................................. 13Auxiliary Power Unit Failures ........................... 13Stability Augmentation System Failures ................... 14Engine Compressor Stall ............................... 14Generator Failures.................................. 14

* -Digital Automatic Stabilization Equipment Evaluation ............. 1Vibration Characteristics.....................................1 7Cockpit Evaluation.......................................... 18Reliability and Maintainability .................................. 18Miscellaneous.............................................. 18Airspeed Calibration......................................... 20

CONCLUSIONS

General .................................................. 2Deficiencies ............................................... 21Shortcomings..............................................2Specification Compliance ..................................... 23

RECOMMENDATIONS.......................................... 24

APPENDIXES

A. Rfernces.....................................5B. Deeription................................................26C. Instrumntion.............................................756D. TestruTentaiusond Data.Analysis.Methods........................78D. Test DTa........and.at..Anly.is.ethods........................8F. Abev ations..............................................169G. Equipment Performance Reports ............................... 171

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INTRODUCTION

BACKGROUND

1 In June 1973, the United States Army Aviation Systems Command, sincerenamed the US Army Aviation Research and Development Command(AVRADCOM), awarded a Phase I Advanced Development Contract to HughesHelicopters (HH). The contract required HH to design, develop, fabricate, andinitiate a development/qualification effort on two Advanced Attack Helicopterprototypes and a ground test vehicle as part of a Government Competitive Test. TheUnited States Army Aviation Engineering Flight Activity (USAAEFA) conductedDevelopment Test I using two of these aircraft (ref 1, app A). In December 1976,AVRADCOM awarded a phase II Engineering Development Contract to HH forfurther development and qualification of the YAH-64 to include subsystems andmission essential equipment. During this program, Engineer Design Tests (EDT) Iand 2 were conducted by USAAEFA to evaluate developmental progress (refs 2 and3). In December 1980 USAAEFA conducted EDT 4 to evaluate the performanceand flight handling characteristics of a new empennage configuration (ref 4). InDecember 1980 and January 1981 USAAEFA, in conjunction with US ArmyAviation Development Test Activity, conducted EDT 5 to assess the integratedoperation of all YAH-64 subsystems (ref 5). AVRADCOM requested (ref 6)USAAEFA to conduct this airworthiness and flight characteristics (A&FC) testwhich follows incorporation of changes to improve flying qualities, structuralintegrity, flight performance, and vibration characteristics of the prototype aircraft.Additional changes are planned for the production aircraft. A test plan (ref 7) wassubmitted in March 1981, and an airworthiness release (ref 8) was issued inMay 1981 and revised in June 1981.

TEST OBJECTIVES

2. The objectives of this Part 1 A&FC test were to assess the handling charac-teristics, hover, vertical climb, and level flight performance, and to comparevibration characteristics with those from previous testing. A follow-on Part 2 A&FCtest, which will complete the A&FC phase of testing, is scheduled inDecember 1981.

DESCRIPTION

3. The YAH-64 is a two-place, tandem seat, twin-engine helicopter with fourbladed main and antitorque rotors, conventional wheel landing gear and a movablehorizontal stabilator. The helicopter is powered by two General ElectricYT700-GE-700R turboshaft engines rated at 1563 shaft horsepower (SHP) (sea levelstandard day, uninstalled). A 30mm chain gun is mounted on the underside of thefuselage below the front cockpit. The helicopter has wings with two store pylons oneach side to carry either HELLFIRE missiles, 2.75-inch folding fin aerial rockets,auxiliary fuel tanks or a combination thereof. An aerodynamic mockup of theMartin-Marietta Target Acquisition Designation System/Pilot Night Vision Systemwas installed. The test helicopter was HH air vehicle number 5 (US Army serialnumber 77-23258). Major changes to the helicopter since EDT-4 include:

a. A third mode (Nap-of-the-Earth(NOE)/Approach) for the horizontalstabilator.

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" ' "i . ' . ." .

b. Digital Automatic Sta hili/alijo I ( II pH en! (1)ASI) softwarc modifi-cations:

(1) Reduction in pitch ('on1mnud \ugmnhn-talttn Svstcm (AS) and StabilityAugmentation System (SAS) gains

(2) Reduction of pitch (AS a t hor ity \ ai rspe.ed

(3) Sideslip input to I)ASI piroided h% Air Data System (ADS)

c. Vibration absorber remoed

d. Full left tail rotor rigging at 2" degrees blade angle

e. Addition of a third direct cutrrent (DO electrical bus

f. External configuratioii changee affecting drag-:

(I) Stiffened wings

(2) External stringers for increased tail boom strength

(3) Removal of wing tip fairings

(4) Installation of rotating pulsc code modulation (PCM) instrumentationcanister and antenna brackets

(5) Instrumented main and tail rotor blades

Further description of the helicopter is contained in the operator's manual andappendix B.

TEST SCOPE

4. Flight testing for A&FC (Part I) was conducted at three test sites. A totalof 45 flights were conducted for a total of 32.4 productive flight hours. Twentyflights were conducted at Palomar Airport. Carlsbad. California (328 ft elevation)from 30 May through 7 June 1 98 1. and from 26 June through 17 July 1981.Twenty flights were conducted at Bishop Airport. Bishop, California (4120 ftelevation) and at Coyote Flats, California (Q980 ft elevation) from 10 June throuh25 June 1981. Three Army pilots performed the cValuation. However, duringlow-speed flights conducted at Bishop and Covote Flats (3 fliglhts). a HH pilot tlewas aircraft commander in the copilot gunner seat. HI installed, calibrated andmaintained the test instrumentation and perforined all aircraft maintenance duringthe test. Flight restrictions and operating limitations contained in the airworthinessrelease issued by AVRADCOM and the operator's manual (ref 9. app A) wereobserved during the evaluation. Where possible. flight test dath were compared tothe system specification and results obtained during FDT-4. The test conditions areshown in table 1.

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Table 1. Test Conditions1

*Tv Avg Avg Av* Tr'norrGss Longiidinal Densty CalibratIed Stahilator Wling

Test Weight CiG Alt iltsude Airspeed zlode Store%0lb) (FS) Ifo (KT)

HOVER 13720 204.5 (Aft) 6060 Auto lean

PROM CF 13440 206.2 (Aft) 11580 Auto lean

LE~VEL FLIGHT IHO 201.3 (FAwd) 8880 76-1412 Auto 16 IIELLFIRI

PERFORMANCE 152140 202.1 (Fwd) 8180 51-1492 Auto 8 ilIlLE1-IRI

* . STATICLONGITUDINAL 15560 204.4 (Aft) 6840 59 and 128 Auto 16 1IIELLI71RI

STABILITY

MAtNBUI G 15480 204.4 (Aft) "180 126 AoO 16 ILLEIRI

17860 20 1.3 1 Fwd) 8 20 35 RT Auto 16 1III LLFIRV

163-20 201.9 (FwdL) 5740 45 Fwd-R%&d Auto 16 tit LLI IRI4 5 LI'4RTLOW SPEED'

15060 202.3 (Fwd) 660 45 Iwd-RwT At l:-FR

14840 202.5 (Fwd) 10580 45 Fwd-Rwd Auto 8 ll[LL:IRIS

STABI LATOREV ALL'ATIO N

14960 202.5 (Fwd) 700 45 Fwd-Rwd 2 MAN' 8 IILLIFIRIControl 45 LT-RT'Positions __________

15180 202.1 (Fwd) 8000 52-120 MAN 811 LIIRI

ULevel. Climb, 15240 202.2 (Fwd) 6440 80t 10 NOE 4 X III LIFRIDescent

StaticLongitudinal 14800 202.1 (Fwdi 7940 73 Nol. 8 9 Ll~

Stabilitv

Acceleration,/ 142 '02.6 (Fwd) 640 0-125 Nt A lLIRDeceleration 190NE&NA' 8lF---R

Malfunctions 15100 202.6 (Fwd) 780 0-115 M AN 8 HIEIIFRI

Approach' 14920 202.6 (Fwd) 780 96-0 Auto & NOE 8 tIIELl-IRFLanding- & MAN'

EVAUAIO 14900 202.5 (Fwd) 580 0-98 Auto 8 IIIELEIE

NOTES:

Rotor speed 289 RPM (100%1), except hover performance2 Airspeed presented is true airspeed3Manual - Stabilator angle 25 & 35 degrees leading edge up (LEtJ)4 NOE: Nap-of-the-earth/Approach5Manual - Stabilator angle 35 degrees LEU

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TEST METHODOLOGY

5. Established flight test techniques and data reduction procedures were used(refs 10 and 11, app A). Test methods are briefly discussed in the Results andDiscussion section of this report. A vibration rating scale (VRS) (fig. 1, app D) wasused to augment crew comments relative to aircraft vibration levels. A handlingqualities rating scale (HQRS) (fig. 2) was used to supplement pilot comments on thehandling qualities. Flight test data were obtained from calibrated testinstrumentation and were recorded on magnetic tape. Real time telemetry was usedto monitor selected parameters throughout the flight test. A detailed listing of thetest instrumentation is contained in appendix C. Data analysis methods are describedin appendix D.

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RESULTS AND DISCUSSION

GENERAL

6. Several design changes affecting performance and handling qualities (para 3)were made to the YAH-64 since the last evaluation (EDT-4). Level flightperformance and vibration levels have been slightly degraded since EDT-4, whilehandling qualities and out-of-ground-effect (OGE) hover performance remainessentially the same. Numerous anomalies in the operation of the DASE wereexperienced during this evaluation. These DASE related problems may render somehandling qualities test results suspect. Two deficiencies (both DASE related) werefound during this evaluation: yaw SAS hardovers, and disengagement of the DASEwith failure of the No. 2 generator. Nine previously unreported shortcomings(four DASE related) were also found.

PERFORMANCE

General

7. In-ground-effect (IGE) and OGE hover and level flight performance weremeasured during this program. The drag of the aircraft has increased since EDT-4(para 3). Vertical climb performance tests were attempted, but inaccuracies in thelow airspeed system rendered the data useless. Engine installation losses necessary tocalculate installed power available were not determined during this test, therefore,the performance requirements of the system specification (ref 12, app A) could notbe evaluated. Performance tests were flown at the conditions listed in table 1.

Hover Performance

8. Hover performance was evaluated at 5- and 100-foot wheel heights usingtethered and freeflight techniques. Hover performance data, in nondimensionalform, are presented in figures 1 and 2, appendix E. OGE hover performance wasessentially unchanged from EDT-4 at low CT's and slightly better at high CT 's. TheYAH-64 requires 2142 SHP to hover OGE and 1831 SHP to hover IGE (5 ft) at agross weight of 14,525 pounds on a 4000-foot, 95-degree fahrenheit day.

Vertical Climb Performance

9. Vertical climb performance was evaluated using ADS information suppliedto a cockpit low airspeed indicator to maintain zero lateral and longitudinal airspeedduring the climb. Climbs were flown in surface winds of 3 knots or less. Althoughthe cockpit low airspeed indicator showed less than 3 knots, the aircraft appeared totranslate fairly rapidly to the rear during the climbs. This was assumed to be causedby an increase in wind velocity with altitude. Analysis of the data indicated theactual power required during vertical climbs was considerably less than indicated bymomentum theory analysis. Further investigation (comparison of low airspeed pace

: . vehicle data with ADS data at the same conditions) indicated that the aircraft couldbe translating as much as 15 knots rearward when the ADS system was indicatingless than 3 knots (figs. 83 and 84, app E). This suggests that during the verticalclimbs the aircraft was probably flying rearward and thus using less power thanrequired for vertical climb. Vertical climb performance should be re-evaluated duringPart 2 A&fC using an accurate low-airspeed syste.m.

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Level Flight Performance

10. Power required for level flight was measured in the 8- and 16-HELLFIREconfigurations at constant rotor speed in zero sideslip flight. Nondimensionaldata for the 8-HELLFIRE configuration are presented in figures 3 through 6,appendix E. Figure 7 presents a comparison of performance data from this testwith EDT-4 data. This comparison shows that the drag of the aircraft increasedbetween EDT-4 and this test. The change in drag is equivalent to an increase in flatplate area of approximately 3 square feet (assuming a propulsive efficiency of one).An increase in drag was caused by the changes in external aircraft configurationdescribed in appendix B (table 1). At the conditions of figure 7, the power requiredfor level flight at 146 knots true airspeed (KTAS) has increased from 2120 SHP inEDT-4 to 2215 SHP. Figures 8 through 10 present the 8-HELLFIRE data from thistest in dimensional form. Figures 11 and 12 present the limited amount of16-HELLFIRE data gathered during this test. These figures also include powerrequired information for the 8-HELLFIRE configuration at tVe same conditions(derived from figures 3 through 6).

HANDLING QUALITIES

General

11. The handling qualities of the YAH-64 helicopter were evaluated at theconditions listed in table 1. All tests were conducted using standard flight testtechniques (ref 11, app A) and maneuvers were flown at zero sideslip where rossible.Tests were performed with the copilot/gunner cyclic control stick in the retractedposition. Test results were compared, where possible, with EDT-4 data. Controlposition data was influenced by frequent off-center SAS actuator position.

*Control Positions in Trimmed Forward Flight

12. The control positions in trimmed forward flight were evaluated in level flightand dives for the 8-HELLFIRE configuration and in level flight only for the16-HELLFIRE configuration. Tests were conducted at zero sideslip with trim FeelON and Attitude Hold OFF. Data are presented in figures 13 and 14, appendix E.The longitudinal control position variation with airspeed was conventional in thatincreasing forward cyclic control was required with increasing airspeed. The shallowstick position gradient below 75 knots calibrated airspeed (KCAS) provided virtuallyno position or force cue for airspeed control but was not a significant problem sinceprecise airspeed control is not required during NOE and contour flight where thisairspeed range would be most often used. The shallow stick position gradient overthe airspeed range of 85 to 110 KCAS (less than 0.6-inch of longitudinal cyclic) is anairspeed range where precise airspeed control may be required, particularly duringflight in instrument meteorological conditions (IMC). The poor trim characteristicswithin this airspeed range were annoying to the pilot and frequent activation of trimrelease switch was required to establish a trim airspeed within ±5 knots indicatedairspeed (KIAS) of that desired (HQRS 4). The trim characteristics, thoughsatisfactory for flight under visual meterological conditions, should be furtherevaluated during IMC flight.

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Static Longitudinal Stability

13. The static longitudinal stability characteristics were evaluated at zero sideslip inthe 16-HELLFIRE configuration with Trim Feel.ON and attitude Hold OFF. Datafrom these tests are presented in figures 15 and 16, appendix E. The longitudinalcyclic control position gradient indicated approximately neutral stability at a trimairspeed of 59 KCAS and was essentially the same as that reported in EDT-4. Thestatic longitudinal stability at a trim airspeed of 127 KCAS, though still weak, was

* - slightly improved from that reported during EDT 4 for the 8-HELLFIRE* - configuration. Neutral or weakly positive static longitudinal stability, although

satisfactory and desirable for the low-speed attack mission, may contribute toK increased pilot workload during IMC flight where more precise airspeed control isrequired.

Maneuvering Stabiity

14. The maneuvering stability characteristics were evaluated in the 16-HELLFIREconfiguration using collective fixed steady left and right turns at a trim airspeed of126 KCAS. Data are presented in figures 17 and 18, appendix E. The stick-fixedstability (control position versus load factor) was weakly positive up to load factorsof 1.5 (fig. 17). Data points could not be stabilized above 1.5 g due to increasedpilot work load caused by saturation of the pitch SAS actuator and the resultant lossof pitch rate damping. The time history presented in figure 18 shows the typicalresults obtained during this test. Saturation of the pitch SAS actuator occurred mostfrequently between load factors of 1.5 and 1.7 and steady state pitch rates of 6 to9 degrees per second. When the SAS actuator reached its authority limit, the aircraftexhibited a pitch up or "dig in" tendency which significantly increased pilot workload for airspeed and load factor control (HQRS 6-7). The results of this test wereessentially the same as those reported during EDT 4 for the 8-HELLFIREconfiguration. The loss of pitch rate damping at load factors greater than 1.5 due tosaturation of the pitch SAS actuator is a previously reported shortcoming which stillexists.

Dynamic Stability

15. Tests were conducted in conjunction with level flight performiance in an effortto isolate the cause of a persistent yaw oscillation of ± 2 to 4 degrees. Time historieswere recorded with the DASE ON and OFF and with the ADS active and disabled(circuit breaker OUT). Test data are presented in figures 19 through 2 1, appendix E.Comparison of the three time history plots indicates that the persistent yawoscillation is driven by the SAS actuator in response to an erroneous sideslip signalprovided by the ADS. The yaw oscillations were noticeable at all airspeeds testedbut were particularly annoying below 90 KIAS. The persistent yaw oscillation of±2 to 4 degrees in level flight is a shortcoming.

Low-Speed Flight Characteristics

16. Low-speed flight characteristics of the YAH-64 were evaluated using acalibrated ground pace vehicle as a speed reference. The tests were flow IGE atapproximately 20 feet wheel height with surface winds less than 5 knots. Data arepresented in figures 22 through 29, appendix E.

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Forward and Rearward Flight:

17. The results of forward and rearward flight tests are shown in figures 22through 24. Figure 26 shows a discontinuity in lateral control position at a hovercaused by an unexplained shift in lateral SAS actuator position (para 35). Theseshifts did not occur abruptly, and were not noticed by the pilot during the tests.Figure 24 shows some minor discontinuities in lateral control position in rearwardflight at airspeeds greater than 25 KTAS. These minor discontinuities were causedby crosswind effects. Generally, the pilot was able to maintain precise aircraftcontrol with only minimal compensation (HQRS 3). The pilot work load increasedin rearward flight at both high altitude (fig. 23) and heavy gross weight (fig. 24)where the aircraft was operated near the power available limit (HQRS 4). Aircraftcontrol margins were satisfactory throughout and tail rotor torque requirementswere well below the maximum continuous limit. The forward and rearward flightcharacteristics of the YAH-64 are satisfactory.

Sideward Flight:

18. The results of sideward flight tests are shown in figures 25 through 28.Generally, precise aircraft control could be maintained with pilot work load higherin left sideward flight. Pilot work load increased at higher density altitudes andheavier aircraft gross weights (HQRS 4). Aircraft control margins were satisfactoryand tail rotor torque requirements were well below the maximum continuous limit.Figure 26 shows some unexplained discontinuities in lateral control position, similarto those experienced in forward and rearward flight (para 17). The sideward flightcharacteristics of the YAH-64 are satisfactory.

Low-Speed Maneuvering:

19. Directional control step inputs were used to evaluate low-speedmaneuverability in left and right sideward flight at 35 KTAS. A representative timehistory is shown in figure 29. The control inputs were made in the directionopposite to the direction of flight. These tests were conducted with a tail rotorblade angle limit of 27 degrees. In both cases a 15 degree per second yaw rate wasgenerated in 1.5 seconds, which meets the system specification requirement (ref 12,app A). In right sideward flight the maximum yaw rate obtainable was 15 degreesper second which occurred at 1.5 seconds using full left directional control. The testwas conducted with the yaw channel of the DASE both on and off with comparableresults. Adequate directional control exists with the tail rotor rigged to 27 degreeswith full left pedal.

Stabilator Evaluation

20. Tests were conducted to evaluate the characteristics of the stabilatorNOE/Approach mode. Additional data were obtained with the stabilator in theautomatic mode and in the manual mode at fixed 25 degrees and 35 degrees LEUangles (the maximum allowed stabilator angle). Flights were conducted at Bishopand Palomar airports and were flown at a forward center of gravity (cg) in the8-HELLFIRE configuration with DASE on. Data are presented in figures A and Band figures 30 through 39, appendix E. A stabilator system description is providedin appendix B.

."8

S..

FIGURE ASUMMARY OF CONTROL POSITIONS

IN TRIMMED FORWARD AND REARWARD FLIGHT

STABILATOR AUTO MODESTABILATOR FIXED 350 LEUSTABILATOR FIXED 250 LEU

~40Lu

30

20

uul

Lu 10

10

~-~- 10

~20

9 TOTAL LONGITUDINAL CONTROL TRAVEL =11.0 INCHES

8

7

6

5

4m~~

- 3 160 40 20 0 20 40 60

K-REARWARD FORWARDTRUE AIRSPEED

(KNOTS)

FIGURE BSUMMARY OF CONTROL POSITIONSIN TRIMMED FORWARD FLIGHT

STABILATOR AUTO MODEK-.- STABILATOR FIXED 350 LEU

- STABILATOR FIXED 250 LEU

~40u-I

30

20

'a 10C,,

0

~10

~10

0= -~ - 10=

: z t - -

0

220

u--

ce 6

LL

C.-,c

2: 4

u- 3

40 60 80 100 120 140 160

CALIBRATED AIRSPEED(KNOTS)

10

Control Positions in Trimmed Forward and Rearward Flight:

21. Control positions in trimmed forward and rearward flight with stabilator anglesfixed at 25 and 35 degrees LEU were evaluated from zero to 45 KTAS in 5-knotincrements. A ground pace vehicle was used as a speed reference and surface windswere less than 5 knots. The test aircraft wheel height was approximately 20 feet.iData are presented in figures 30 and 31. In general, as stabilator LEU angleincreased, nose down pitch attitude and aft cyclic control position requiredincreased. Other control positions were relatively unaffected by stabilator angle.Aircraft control margins were satisfactory. The longitudinal control positionvariation with airspeed is nonlinear but satisfactory. The small variations fromtrim in longitudinal, lateral, and directional control movements at all airspeedsindicates the low pilot work load (HQRS 3) in low-speed forward and rearwardflight. The discontinuity in lateral control position at a hover was due to anunexplained shift in lateral SAS actuator position (para 35).

22. Control positions in trimmed level flight were evaluated from 50 to 120 KCASwith the stabilator angle fixed at 25 degrees LEU and from 50 to 96 KCAS with thestabilator angle fixed at 35 degrees LEU. Data are presented in figures A and B, andfigure 32, appendix E. Results were similar to low-speed flight results (para 17), inthat as stabilator LEU angles increased, nose down pitch attitude and aft cyclicrequirements increased. Aircraft control margins were satisfactory . Increasing aftcyclic was required to fly increasing airspeeds (one-inch increase between 50 and95 KCAS at 35 degrees LEU, and 0.6-inch increase at 25 degrees LEU) however, thiswas not objectionable.

Level Flight, Climbs, and Descents:

23. Level flight, intermediate rated power (IRP) climbs, and minimum powerdescents were conducted with the stabilator in the NOE/Approach mode. Minimumpower descents were initiated from both level flight and IRP climbs by reducingcollective to a minimum setting which would maintain 100 percent rotor speed. IRPclimbs were initiated from level flight and minimum power descents. During thesemaneuvers airspeed was initially stabilized slightly below the stabilator switchingvelocity (80 KIAS) then varied ± 10 knots. Increasing airspeed through the switchingvelocity resulted in a nose up pitch attitude as the stabilator began automaticprogramming. Reducing airspeed through the switching velocity resulted in a nosedown pitch attitude as the stabilator began programming to the NOE/Approachposition (25 degrees LEU). The rate and magnitude of these pitch attitude changeswere essentially the same during level flight, climbs, and descents. The undesiredpitch attitude changes with slight airspeed variations will increase pilot work loadduring continuous flight at airspeeds close to the stabilator switching velocity. Theii following note should be placed in the YAH-64 operator's manual:

NOTE

Continuous flight at airspeeds close to the stabilatorswitching velocity (80 KIAS ±10 knots) with thestabilator in the NOE/Approach mode will result inaircraft pitch attitude variations caused by intermit-tent stabilator programming.

II

. . .

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;,,, ''"' '; 1 ;•f, ,' l:) li I !1;,· 'IV1Jili!'>.'llh'l11': (li J>.ilil"l:ljdl I(J 1.! n! 'i'i'.. \ 111 t!• •I ' 11• ;,_· ·.v:l'• lll':•::l · .<:t:fli,· .,_[:thdtl:, ;! !II I'· ,.,h

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the statJhti.ur s tnree modes of operation. IRP accelerations from a '}"

. 'I

were perfonned with the stabilator in manual mode (35 deg LEU), '.f\T( . t;rtode_ (25 de~ LEU), nnd in the automatic mode. Representative . bl . ..!~ -a?ccltpfffi~ht_hV-in f11,-tJfCs\~etlfi'Q~W,l iB mi8hPJ9.h~ mitfilNto¥ra e ai.fHMI rfroptffiyr1stfF' ffiWCffitrok~t18 eaH<fr R~J~~i aA\t>aH~il~}we-t~gran:_mtn£t8f pib~hlff1fuffigt1\9\tnedn0ifi~dfihai~Iw~~1 o tetitl;Hpt.rer itY iHiiliP,in:~Pitff>~ m&n!ooc~tt#OE1t9\p'f5¥01R~ Rt~.~~. Y~eEt;t o~Wife&~'ftu~(:! a'Wt!l ~dlffiffl f-61m~~fnJg 1':ic:±15lS'fat18h H.~lfi ~~ ·. ~tf · dffil!l t~ift'Hit\1PtW18SiiY·~~ ~l~'B~t inofHt\91 nffigu if ~i~ t\ffuviD ell et&:'th~Jert~ag\f' t<ffen~ h¥Mit.'tiE!-C~~~-<fffifflmg t}f :'Yta~mimr rtrrg.r{f~\!liMs f1·91tfegree~llu!.W:~\rHeMI\H~~P~WJ¥.g gt~~tt!~a:lYjilJgO\NIQ\'S ~I§eeHoo tllrf@!\lgtlccefuratrfigbilft{ftfugH-Vitfl\!ingwilelfifi~Y v~$oo1ty~h~Tllififumifle §taBililt§f s1mM:t:J\Wili\\1as ~11C~iinte:tetll~¥M~ 2tt>JfWPRip~c@@J~!Rffis ¥.Llt1Nff1ei!itam.fa~miw.s 3S>ttegrtlesst~JBtJ~cTheretiWift-ed~ngon'lr§iemt#gfu tj)tO!jtmns~r ftint1~91rabl~'-"iiff<hilli thl·n<a\!tedstics associated with either complete automatic stabilator programming or stabilator shutdown during IRP accelerations. The stabilator fail audio warning tone ~ttWilMara'd4.'(fi.~tioor\fflf~n~fAsl!M:J!Ui<ttff>l!hutdown; however, the shutdown served no useful purpose and increased pilot work load by necessitating manual stabilato\ pfog~tn m h~tll r Proi~.Uttn \Yf' cttt ed~ttioo a toi!e#gratioos ~ A-'t1Yrm!\!ev8 ~ni1'1ef\!1ati,fi:S'mlftlFC ph')igt-affilfiit.jgc\fkrffl:i(an!'ftfilom.BU21ibfuiliaih~ftlin~ (~6egttJBt9)owfit!i!1ca&el~Fhill\~ fifrdegtte~tteE tStaq;tU&fopetswi~lijf}Jr~w~e <fixSH.bh~olfflfu~ing:-Jiie,.\ §t~BHai8f p~ramtn\fllg ls1~'otttd <l!Je tmodHfi~s f6ll~rovutJc~t;nsfStml:ro.hrt~Nt '"Tff()~tafilln1fii f;ib:rll Q1trr~>t~tbilatt.Wti§mti~'tu\to,linf@lia:w.\lbtathm si~tffi-1 t~I!Sf§bitatl()r 's'Wltbltllig fuloc:ttyntative time histories arc presented in figur~s~37 and 38. Pik-t corrective action follovving the simulated failure required clecrleration to RVt)i·: 1

• :~c~'-"din£'. :1

~-Mrtt&W~~t~iDu~ A'~""~tlllttl:ished by applying aft cyc1:c . ; ,, ~itghrl:• lowering the collective. Corrective action was initiated two seconds i.t.\~ .:i-. •I! i:,:;;1

26 . . Stabi1ator malfunction during acceleration was simulated by conducting level flight IRP accelerations from a 20-foot hover at fixed stabilator angles of 25 and 35~ degrees LEU. Once set, stabilator angles were fixed by engaging the stabilator ., kill" switch located on the pilot's colleatlve lever. Control margins remaining and pi!ot corrective action required following the simulated failure were evaluated. Representative time histories are presented in figures 37 and 38. Pilot corrective action following the simulated failure required deceleration to avoid exceeding a stabilator airspeed limit. This was accomplished by applying aft cyclic and slightly lowering the collective. Corrective action was initiated two seconds after activation

.. , . 12 BEST AVAILABLE COPY

I. ·'

' \ '

of the stabilator aural warning. The aural warning which occurred at the stabilatorswitching velocity (80 KIAS) provided excellent pilot cue of stabilator failure. Theminimum control margins remaining during the recovery were 20 percentlongitudinal cyclic and 12 percent directional pedal. No undersirable aircraftcharacteristics or control margins were associated with stabilator failures during IRPaccelerations.

Contour Approach/Landing:

27. Contour approaches terminating with a quickstop maneuver and IGE hoverwere conducted from an initial airspeed of 90 KIAS (stabilator switching velocity+ 10 knots) with the stabilator in the NOE/Approach mode. A typical time historyis presented in figure 39. The stabilator programmed from 5 degrees leading edgedown to 25 degrees LEU in approximately 6 seconds as the aircraft deceleratedthrough the stabilator switching velocity. Maximum nose up pitch attitude duringthe quickstop was 25 degrees. This pitch attitude restricted pilots forward field ofview; however, the pilot could maintain desired heading using peripheral vision.The minimum control margin remaining during the maneuver was approximately10 percent directional control. Overall pilot workload during transition fromcontour flight to hover was minimal (HQRS 3). Contour approaches to an IGE hover

* . with the stabilator in the NOE/Approach mode are satisfactory.

Aircraft Systems Failures

Stabilator Failures:

28. Throughout the tests, low power descents at various airspeeds were made.Numerous failures of the stabilator automatic mode of operation were experiencedwhile in a low power descent at airspeeds of 60 to 80 KCAS. The stabilatorautomatic mode of operation could be regained by depressing the reset button;however, a subsequent automatic mode failure often occurred within a few seconds.The ADS processor was changed (Equipment Performance Report (EPR) No.80-17-1-5) in an attempt to correct this problem; however, the failures still occurred.The failure of the stabilator automatic mode of operation while in a low powerdescent is a previously reported shortcoming which still remains.

Auxiliary Power Unit Failures:

29. The auxiliary power unit (APU) was used to start the aircraft engines duringthis test piagram. The APU incorporates an automatic start sequence, which isinitiated by placing the APU switch from OFF to START and then allowing it toreturn to RUN. However, the APU failed to start on numerous occasions requiringthe utility hydraulic accumulator to be charged in order to attempt a second start.The start procedure was modified to place the APU switch in the RUN positionmomentarily, allowing the APU electric fuel pump to supply fuel to the APU priorto the start attempt. This procedure reduced the frequency of the starting problems;however, starting failures were not totally eliminated. All the start failures wereexperienced at the Palomar test site (328 ft elevation). None were experienced at theBishop test site (4120 ft elevation). The APU fuel control was replaced prior to the

13

btst thr~e test flights (EPR No. 80-17-l-3, app G) and no start failures were ~.XJl'rienced at the Palomar test site on six subsequent start attempts. The failure of the APU to start is a shortcoming. . ·

Stlbility Augmentation System Failures:

30. Throughout the test program multiple SAS hardover failures in the yaw axis ~ experienced. The hardovers occurred during taxi, hover, takeoff, low power des~t, and approach. Figures 40 through 42, appendix E are representative time histories of yaw hardovers. The hardovers were multiple cycle and full SAS actuator travcl. (20 percent of main flight control servoactuator). These hardovers produced ~w rates up to 23 degrees per second and sideslip angles of 40 degrees at 71 KTAS (ful, 40), which exceeds the sideslip envelope by 7 degrees. Additionally, the SAS frul\lre monitor circuitry did not prevent the hardovers. The ADS processor was cl\~d in an attempt to correct this problem; however, the hardovers were still ~perienced. Subsequently, the sideslip signal from the ADS to the DASE was romoved and no further yaw SAS hardovers were experienced. The yaw rate of :13 d•ees per second with a yaw SAS hardover failure failed to ·meet the require­m~nb of paragraph 1 0.3.2. 7.1 of reference 12, appendix A by 13 degrees per ~nd. The yaw SAS hard over failures (20 percent. of main.- flight control ~acr.tuator1travtl$13mlgi tdeficie,cy-whichllShouldrb41 amrehtedt,~·tt)vPafti 2iJOf thai.:.4&F.<i\ mte iSASr liati:lover.z monit&coitcuitry ;sbouJa b.eemodifiett to r~de ~It\Pl~tt~sSASJtardo~r.protection~ ,·,:J.Xlr•tu:n tur' dne gas t~mperatme (TGT) or 'ir-.1p~uxima.ta1):" >OOoC. was. aot..;:i T~x .:u!:.~cn-.·:: ·... lo;\·e~ed a1~d t.he ai~cr~~t la~ded. tmllflEL.Com~Stall:;s C'l'l h.;:>tr• fng:m~s ot:c•n·::-ro 1ollowm~ n left OirtlcUnnnl ,:n,\troiiitpht of le:;s than (llH!-.l:.~:r :l'::lJ tl' :1u·:·:.: 1 ,,,\ki right yaw ratC. Th..: collective ~~s 11\\lii\!tliQJidhQ~r Whl~prnfobnihgllo~bnt ~test.tat 9&'80:lfoitnp~tire gtl~i!lt4~,a tlJO~~~Jm~, sain tmftperattifct;''andT i4,920::J)o.unU&~ tgro51f;iweigh11 Hmging ~~91'~t~ls:: ~ere; 1me0untered. The. ~ft ·~oefit&-bperit.ec\;11\tat :engi!ne ~r lb!li~.o;I'he:: nrst •. engine contpmsor stall ;oOOUJTedcoo(tlt8Jnumlier l~gine ~.bilt'trini ,lm lGE · hover,Jollowing: a 1· smidll corrective:.·~ aftixtycilie. mplitDAttpoppmg n~~~·tlmatd··.-4nd: 1Ctransient: ioaximtim .birbirie ~'~turei;rfiMl)>iOf ~lU1~Xim~~ly: &QO!C:~as .rtdted: The ·collective was toweredUmdtthel aircniD landed . .Multiple compreSsor stalls on both engines occurred following a left directional Q)Qtro\ .ietru1HafJess than one-half inch to arrest a mild right yaw rate. The!collective w~ lowered and the helicopter landed. Popping noises were heard during the !l~scettlt!Md. teaglnl:l tt:ransient TG1t~·of·900°c:were o:oted1(ngbtMt~lighteiDlFbltowing

lh .. ~ tliah\rtktt·n. umben .. tw<thllgirie. • • .... :~.~d.t anged (~N.G. 80tlrl~l.l-~. ,,PP!~6no fw:tbM eompre8Sor1 stallsnvcre cencounteredn Ho\tewip'COndltJ.1ilis fwhloh · rf)dutdd th~erinillW.F,&tds ow~m· lnot2 (lup.HQ\ted~='·trrb'e de\1~nin•11tlte tMBC6 erDW&"i ~uation 1tmainesn The: pri>ductioo/~ngineE:'sltQ\lld tWaNalnat~at hi 1itertsity t\lt\\\~~nd.htl'-vrl &rOSS! \Weight to insure &aequate. engi1UrtSt;alhdltrgjntcrat~1 within a f>"~ngc that is near or on p:~~~ :: : ' d"! ·i:· ~~.~thv.Hy <pan! 38),, a No. 2 jlcnerator ~!~t~~: in an : :·::.. . · .... \~diH:~·der! control Input whtCh won'c occur as

32. Gonerator failures were M\ulated ir,. flight by turning off either the No. I or No. 2 go.nerator switch. A third DC electtlt: bus had been installed in the aircraft to p~Vimt s~tem failures and enoneous warning activation with the failure of a single aooorator. Failure of the No. 2 generator caused disengagement of the DASE. DASB di~m~t results in loss of SAS, CAS, hover augmentation system (HA.S), turn coordination and attitude hold. Since the actuators frequently operate· within a t't\1\ge that is near or on one limit of their authority (para 38), a No. 2 generator fuUm~ oould result in an abruptt uncommanded control input which would occur as

I ' 1. •' , ! li:If ,, t; f "' H I

. . 'I/ ·'·I"' : I .1, ,1.~

. ·. I,''

.. '

the actuator centered following DASE disengagement. This problem could occur inall three axes simultaneously and would be of the same magnitude as SAS actuatordisplacement from trim. Abrupt, uncommanded control inputs, which could beequivalent to 10 percent control authority, will increase the probability of resultingloss of aircraft control during night NOE operations. DASE disengagement withfailure of one generator failed to meet the requirement of paragraph 10.3.2.7.8of reference 12, appendix A. Disengagement of the DASE following failure of theNo. 2 generator is a deficiency that should be corrected prior to Part 2 of the A&FC.

S- 33. Simulated failure of the No. I generator produced erroneous activation of thestabilator automatic mode failure audio warning tone. The stabilator continued toprogram in accordance with the automatic schedule. The erroneous activation of thestabilator audio warning tone with failure of the No. 1 generator is a shortcoming.

34. In addition, failure of the No. 1 or No. 2 generator caused erroneous activationof the engine out/main rotor RPM audio warning tone. This could cause the pilot toreact to a false engine out situation and compromise mission accomplishment orcompletion. The erroneous activation of the engine out/main rotor RPM audiowarning tone with failure of the No. 1 or No. 2 generator is a shortcoming.

Digital Automatic Stabilization Equipment (DASE) Evaluation

35. During the course of this test program, several anomalies in the operation- . of the DASE were noted. These consisted of SAS hardovers in yaw (para 30),

persistent yaw oscillations in level flight (para 15), loss of SAS rate damping(para 38), and unexplained trim shifts in low-speed flight (paras 17, 18, 21). Otherareas of DASE malfunctions occurred in the operation of the HAS (para 36) and theattitude hold system (para 37). The DASE computer was removed following thecompletion of flight testing on 1 July 1981, and returned to the vendor. Amalfuction (undervoltage) in the computer internal power supply was found andcorrected. This undervoltage condition reportedly caused a deviation in all the

* DASE inputs, gains, and washouts. Additional tests were flown to re-evaluatethe problem areas related to the DASE after this malfunction was corrected. Amodified trim feel system was also evaluated during this DASE reevaluation. Thetrim feel modification consisted of removing the discrete OFF capability of theforce feel system. Due to a related problem with the SAS monitor circuits and theADS (para 15), the sideslip signal from the ADS to the DASE was removed. Thisprevented a re-evaluation of yaw SAS hardovers and persistant yaw oscillations inlevel flight.

36. The HAS was evaluated in a stabilized hover, during bob up, and whilerepositioning with and without retrimming. The evaluation was conducted todetermine if the DASE maintenance action had corrected the abrupt aircraft

* response when attempting to reposition with HAS ON. Tests were performed inwinds of approximately 10 knots. A time history of a gradual acceleration with HASON is provided in figure 43, appendix E. The HAS initially resisted the forwardlongitudinal control input with opposing SAS actuator movement up to full travel.As airspeed reached 35 KIAS, the HAS automatically disengaged and the SASactuator returned to the centered position. Pitch attitude excursions throughout themaneuvers were less than ±2 degrees. The abrupt aircraft response previouslyobserved appeared to have been corrected. Further evaluation of the HAS, toinclude operation in gusty wind conditions, should be conducted during operational

* testing.

15

I

37. Abrupt aircraft response following retriniming with attitude hold ON wasfrequently experienced prior to the DASE rnaintcnance corrective action. Followingthe maintenance action, the attitude hold feature was re-evaluated. Tests wereperformed in level flight by trimming at 1 10 KIAS, decelerating to 100 KIAS andretrimming. The test was repeated trimming at 110 KIAS, accelerating to 120 KIASand retrimming. A time history of retrimming at 100 KIAS is shown in figure 44.The attitude hold initially resisted the aft longitudinal control input in decelerationwith opposing movement of the pitch SAS actuator. While retrimming, the SASactuator centered, which resulted in the requirement for forward longitudinalcontrol to maintain the desired airspeed. Several attempts were required to achieve aprecise trim condition; however, once the desired trim airspeed was attained, thetrim condition was maintained within ± 3 KIAS in smooth air. No abrupt aircraftresponse was noted during this test. It appeared the abrupt aircraft response hadbeen corrected w i h the DASE maintenance action. Further evaluation of theattitude hold feature should be conducted in conjunction with instrument flightcapability testing.

38. Full SAS authority was not always available during this evaluation because theSAS actuators were displaced from their centered positions for long periods of time.Time to washout rate damping SAS inputs was greater than 10 seconds for all threeaxes. Since gust upsets are very short-term phenomena (i.e., less than one secondduration) these washout times are excessive. Any rate that persists for more thantwo seconds can reasonably be assumed to be intentionally caused by the pilotand should not be damped by the SAS (i.e., the rate damping should be washed outat least this rapidly). In turns, the excessive SAS washout time allows the pitchSAS actuator to reach the limit of its travel while trying to oppose the pitch ratewhich is necessary to maintain the turn. When this happens (at about 1.5 g) allrate damping is lost and pilot work load increases (para 13). In the absence ofCAS, aircraft rate response to step control inputs would be washed out to unac-ceptably low levels by the SAS. On the YAH-64 the solution was to increase theCAS washout time so that the CAS opposes the SAS following a control input.Figure 46 shows the roll CAS washout. These data were taken on the groundfollowing a lateral cyclic displacement from trim. These long CAS washout timeshave an undesirable side effect: Whenever the controls are displaced from the lasttrimmed position and there are no aircraft angular rates present, the SAS actuatorsare displaced from their centered positions for long periods of time. In fact, there isno CAS washout in flight. Figure 45 shows the roll SAS actuator displacementduring 30 KTAS rearward flight with the last trimmed control position being thatrequired for hover. Note that there is no tendency for the roll SAS actuator torecenter. This effectively reduces the rate damping authority of the SAS actuators(sometimes to zero). This effect occurs when lifting off to a hover, when changingairspeeds, during any transient maneuver, or at any stabilized flight condition if thecontrol forces are not trimmed to zero. This degradation of rate damping isparticularly detrimental during NOE flight where airspeed changes are made veryfrequently. It was suggested during this evaluation that the pilot depress themomentary trim release button prior to moving the controls from one trimmedflight condition to another and releasing the button when the new stabilizedcondition is achieved. This would ensure that the CAS would not cause the problemjust discussed. However, depressing the button effectively eliminates the CASfunction of the DASE. If the CAS is disabled when the pilot moves the controls,then it cannot augment the pilot control inputs, which it is designed to do. The

ecproblems caused by excessive SAS and CAS washout times were apparent

16

throughout the test program, although the severity was reduced following the DASEmaintenance action. The reduction of rate damping authority caused by excessiveSAS and CAS washout times is a shortcoming. The DASE software should bemodified to provide acceptable gust response, control response, and controlsensitivity independent of the force feel system reference.

39. A qualitative evaluation of a modified trim feel system (discrete off functioneliminated) was conducted during the DASE re-evaluation. The modificationrequired the pilot to either oppose the control force gradient or retrim whendisplacing the cyclic from a trimmed position. The location of the momentary trimrelease switch on the cyclic handgrip and the mechanical functioning of the switchmade it difficult to operate. Additionally, activation of the switch effectivelyeliminated the CAS function of the DASE (para 38). Absence of a discrete offcapability on the trim feel system was undesirable since many pilots, based onprevious experience in other aircraft, prefer to fly without a force gradient. Thediscrete off function of the trim feel system should not be eliminated.

VIBRATION CHARACTERISTICS

40. The vibration characteristics of the YAH-64 were evaluated with the verticalvibration absorber removed. Qualitative evaluations were made at the pilot's andcopilot's stations. Quantitative vibration data were gathered at the conditions intable 2, and are shown in figures 47 through 80, appendix E.

41. Vibration characteristics were evaluated in forward and rearward low-speedflight with the stabilator in the manual mode at 25 and 35 LEU and in theautomatic mode of operation (figs. 47 through 59). The 4/rev (19.3Hz) verticalvibration of 0.4 to 0.55 g at the pilot's seat at airspeeds greater than 25 KTASrearward flight are objectionable (VRS 6). The use of manual stabiliator full LEU(35*) did not change the vibration characteristics as compared to 25 degrees LEU.Although the vibrations at the copilot's seat were lower than at the pilot's seat forthe same flight regime (0.35 to 0.4 g) they were still objectionable (VRS 5).

42. The vibration characteristics in sideward low-speed flight were evaluated withthe stabilator in the automatic mode (figs. 63 through 69). The 4/rev (19.3 Hz)lateral vibrations in right sideward flight of 0.3 to 0.36 g at airspeeds above20 KTAS were objectionable to the pilot (VRS 5). The vibrations at the copilotstation were again lower than at the pilot's station and were not objectionable.

43. The vibration characteristics in level flight were evaluated with the stabilator inthe automatic mode (figs. 70 through 80). The 4/rev (19.3Hz) lateral vibrationgreater than 0.2 g at airspeeds below 54 KCAS were objectionable to the pilot(VRS 4). The 4/rev lateral and vertical vibration greater than 0.2 g at airspeedsgreater than 128 KCAS were objectionable to the pilot (VRS 4) (fig. 72). Thevibrations at the copilot seat were lower than at the pilot seat and were notobjectionable.

44. The vibration characteristics during the termination of the approach werequalitatively evaluated. These vibrations were similar to those encountered duringrearward flight at airspeeds greater than 25 KTAS and were objectionable to boththe pilot (VRS 6) and copilot (VRS 5).

17

................

45. The 4/rev vibration characteristics of the YAH-64 were generally degradedfrom EDT-4. The vibration characteristics of the other harmonics remainedessentially unchanged. The vibration characteristics are more severe at the pilot'sseat than at the copilot's seat and remain objectionable during several flight regimes.The 4/rev ( 19.3 Hz) vibration characteristics remain a shortcoming.

COCKPIT EVALUATION

46. The environmental control unit (ENCU) was evaluated on all flights. Ambienttemperature conditions ranged up to 95TF on clear, sunny days. The ENCU wasinadequate in terms of both air flow and air temperature and failed to keep eitherthe pilot or copilot crew station at an acceptable level of comfort. Failure of theENCU to provide adequate crew station cooling is a shortcoming and should becorrected prior to production. The ENCU should be further evaluated duringclimatic testing.

RELIABILITY AND MAINTAINABILITY

47. The reliability and maintainability features of the YAH--64 aircraft wereevaluated throughout the test. Twenty-two EPR's were prepared and submittedduring this test and are listed in appendix G. This section is intended to highlightthose undesirable reliability and maintainability features encountered.

48. The master caution light illuminated on at least three occasions without acorresponding caution, warning, or advisory panel segment light. The master cautionlight is designed to be a latched light while the segment lights are not. On at leasttwo occasions the pilot observed the caution, warning, and advisory panel when themaster caution light illuminated. None of the segment lights flickered indicating amomentary fault condition and the segment lights were then subsequently tested toinsure all lights would illuminate. These spurious illuminations of the master cautionlight occurred while taxiing into the parking area and prior to the connectionof external electrical power. Illumination of the master caution light without theillumination of a caution, warning or advisory panel segment light is a shortcoming.

49. Throughout the test, the Marconi engine and rotor speed indicator consistentlyindicated 102 percent (in the caution range) main rotor speed (N ) and No. 1 andNo. 2 engine power turbine speed (N.). The calibrated flight test instrume ntationindicated 100 percent N R and NP at the same conditions. The illumination of acaution range is distracting to the pilot since it indicates an overspeed condition. Theerroneous indication of 102 percent N R and NP is a shortcoming which waspreviously reported.

MISCELLANEOUS

50. Testing was accomplished to further evaluate deficiencies arid shortcomingsidentified during earlier tests. As a result of this evaluation, two previously reporteddeficiencies have been corrected and 1 3 shortcomings have either been corrected ordetermined to be satisfactory due to modifications of procedures or limitations.The following previously reported shortcomings (in order of importance) remainvalid shortcomings. A discussion of these shortcomings can be found in reference 4,

18

U. appendix A. The Program Manager and HH have identified corrective action for theshortcomings with asterisks (ref 15, app A).

a. The poor design of the pilot's fuel control panel.*

b. The absence of SAS pitch rate damping at load factors greater than1.6 due to saturation of the pitch SAS actuator.*

c. The restriction to the pilot's field of view caused by window edgedistortion, the overhead circuit breaker panel, canopy reflections, CPG helmet,and the pilot night vision system turret.

d. The excessive longitudinal breakout force (specification noncompliance).

e. The inadvertent directional control inputs during brake application.

f. Dissengagement of the stabilator automatic mode of operation while in alow power descent.*

g. The poor design of the collective pitch control friction mechanism.*

h. The lack of a reliable indication of parking brake st.tus *

i. The lack of an acceptable method for sampling the primary and utilityhydraulic fluid.

j. The high inherent friction of the engine power control levers.

k. The annoying tone present in the intercom system.*

1. The failure of the Marconi instruments to display the full green rangeduring normal operation.

m. The excessive accumulation of oil on the main transmission deck andin the upper fairing maintenance access area.*

n. The poor location of the pilot engine control quadrant.**

o. The washout of the rocket panel display, Marconi instrument indicationsand caution, warning, and advisory panel segment lights in direct sunlight.

p. The poor location of the tail wheel unlock light.

q. The constant illumination of the lower green segment light on the Marconivertical scale.

r. The illumination of the APU ON advisory light prior to the APUstabilizing at 100 percent RPM.*

s. The poor anthropometric design of the pilot's cyclic grip.*

t. The difficulty in accurately determining the engine oil levels without*I opening the engine cowlings.*

19

u. The difficulty in attaining a comfortable seating position with referenceto the cyclic and collective controls.

v. The illumination of the master caution light with green advisory segmentlights.*

w. The failure of the fire bottle discharge lights to illuminate with theactivation of the PRESS TO TEST switch.*

AIRSPEED CALIBRATION

51. The left and right pitot-static airspeed systems were calibrated in level flightand dives using both T-28 and AH-1S pace aircraft for the calibrated airspeedreference. Figures 81 and 82, appendix E present data from this calibration effort.The main-rotor-mast-mounted omnidirectional airspeed system was calibrated inlow-speed forward, rearward, left sideward and right sideward flight using a paceautomobile as a speed reference (figs. 83 and 84). A large dead band exists in the

" . longitudinal measurement of the low airspeed system. At actual airspeeds from-15 knots rearward to 3 knots forward, the longitudinal airspeed output wasindicating from 3 knots forward to 3 knots rearward. This affected the gathering ofvertical climb performance data during this test. It is not known if this error issignificant from an operational standpoint.

20

CONCLUSIONS

GENERAL

52. Based on the A&FC Part I flight test of the YAH-64 helicopter, the followingconclusions were reached:

a. Adequate directional control exists with the tail rotor rigged to 27 degreeswith full left pedal (para 19).

b. 4/rev vibration characteristics have degraded since EDT-4 (para 45).

c. Manual stabilator setting of 35 degrees LEU did not change the vibrationin rearward flight (para 41).

d. Incorporation of the No. 3 DC electrical bus did not prevent loss of DASEwith failure of the No. 2 generator (para 32).

e. Twenty-two equipment performance reports were submitted during this

test (app G).

f. Two deficiencies have been identified.

g. Ten previously unreported shortcomings have been identified.

h. Two previously reported deficiencies have been corrected.

i. Thirteen previously reported shortcomings have either been corrected ordetermined to be satisfactory due to modification of procedures or limitations.

DEFICIENCIES

53. The deficiencies reported herein are not necessarily intended to bar typeclassification as per AR 310-25 (see app D for definition of deficiency used in thisreport). The following deficiencies (in order of relative importance) were identified:

a. Yaw SAS hardover failures (20 percent of main flight control servoactuator travel) (para 30).

b. Disengagement of the DASE following failure of the No. 2 generator(para 32).

SHORTCOMINGS

54. The following previously unreported shortcomings (in order of relativeimportance) were identified (see app D for definition of shortcomings used in thisreport):

a. Illumination of the master caution light without the illumination ofa caution, warning or advisory panel segment light (para 48).

b. Failure of the APU to start (para 29).

. 21

4.

F ...-- - -

K

c. Failure of the ENCU to provide adequate crew station cooling (para 46).

d. Erroneous activation of the engine out/main rotor RPM audio warningtone with failure of the No. 1 or No. 2 generator (para 34).

e. Reduction of rate damping authority caused by excessive SAS and CASwashout times (para 38).

f. Erroneous activation of the stabilator audio warning tone with failure ofthe No. 1 generator (para 33).

g. Persistent yaw oscillations of ±2 to 4 degrees in level flight (para 15).

h. Failure of the stabilator to consistently achieve complete automaticprogramming from a full LEU manual setting (35 deg LEU) when acceleratingthrough the stabilator switching velocity (para 25).

i. The 4/rev (19.3 Hz) vertical vibration characteristics (para 45).

55. The following shortcomings (in order of relative importance) were identifiedduring EDT 4 and earlier tests. They were further evaluated during this test and"- remain valid shortcomings. The Program Manager and HH have identified correctiveaction for the shortcomings with asterisks (ref 15, app A).

.- a. The poor design of the pilot's fuel control panel.*

b. The absence of SAS pitch rate damping at load factors greater than1.6 due to saturation of the pitch SAS actuator.*

c. The restriction to the pilot's field of view caused by window edgedistortion, the overhead circuit breaker panel, canopy reflections, CPG helmet,and the PNVS turret.

d. The excessive longitudinal breakout force (specification noncompliance).

e. The inadvertent directional control inputs during brake application.

f. Failure of the stabilator automatic mode of operation while in a lowpower descent.*

* g. The poor design of the collective pitch control friction mechanism.*

h. The lack of a reliable indication of parking brake status.*i. The lack of an acceptable method for sampling the primary and utility

hydraulic fluid.

j. The high inherent friction of the engine power control levers.

k. The annoying tone present in the intercom system.*

1. The failure of the Marconi instruments to display the full green rangeduring normal operation.

,2 1

I

m. The excessive accumulation of oil on the main transmission deck andin the upper fairing maintenance access area.*

n. The erroneous indication of 102 percent NR and Np (para 49).

o. The poor location of the pilot engine control quadrant.

p. The washout of the rocket panel display, Marconi instrument indicationsand caution, warning, and advisory panel segment lights in direct sunlight.

q. The poor location of the tail wheel unlock light.

r. The constant illumination of the lower green segment light on the Marconivertical scale.

s. The illumination of the APU ON advisory light prior to the APUstabilizing at 100 percent RPM.*

t. The poor anthropometric design of the pilot's cyclic grip. *

u. The difficulty in accurately determining the engine oil levels withoutopening the engine cowlings.*

v. The difficulty in attaining a comfortable seating position with referenceto the cyclic and collective controls.

w. The illumination of the master caution light with green advisory segmentlights.*

x. The failure of the fire bottle discharge lights to illuminate with theactivation of the PRESS TO TEST switch.*

SPECIFICATION COMPLIANCE

56. The YAH-64 was found to be in noncompliance with the following para-graphs of the Phase 1I Advanced Attack Helicopter System SpecificationAMC-SS-AAH-H10000A during these tests.

a. 10.3.2.7.1 Yaw rate in excess of 10 degrees per second with a yaw SAShardover failure (para 30).

b. 10.3.2.7.8 Failure of the DASE with the loss of the No. 2 generator(para 32).

c. 10.3.4.1 Negative longitudinal static stability with the stabilator in theNOE/Approach made at airspeeds below the stabilator switching velocity (para 24).

23

:2-

RECOMMENDATIONS

57. The following recommendations are made:

a. Correct the deficiencies prior to A&FC Part 2 testing.

b. Correct the shortcoming in paragraph 54g prior to A&FC Part 2 testing.

c. Correct the remaining shortcomings prior to production.

d. The DASE software should be modified to provide acceptable gustresponse, control response, and control sensitivity independent of the force feelsystem reference (para 38).

e. The SAS hardover monitor circuit should be modified to providecomplete SAS hardover protection (para 30).

f. The discrete off function of the trim feel system should not be eliminated(para 39).

g. Place the following note in the YAH-64 operator's manual (para 23):

NOTE

Continuous flight at airspeeds close to the stabilatorswitching velocity (1 10 knots) with the stabilator inthe NOE/Approach mode may result in aircraft pitchattitude variations caused by intermittent stabilatorprogramming.

h. Further evaluation of the HAS to include operation in gusty windconditions, should be conducted during operational testing (para 36).

i. Further evaluation of trim characteristics and the attitude hold featureshould be conducted in conjunction with instrument flight capability testing(paras 12 and 37;.

j. Production engines should be evaluated at high density altitude and heavygross weight to insure adequate engine stall margins (para 3 1).

k. The ENCU should be further evaluated during climatic testing (para 46).

1. Modify the stabilator programming to provide consistent automaticprogramming from all stabilator settings during acceleration through the stabilatorswitching velocity (para 25).

m. Vertical climb performance should be reevaluated during Part 2 A&FCusing an accurate low airspeed system (para 9).

24I

APPENDIX A. REFERENCES

I. Final Report, USAAEFA Project No. 74-07-2, Development Test 1, AdvancedAttack Helicopter Competitive Evaluation, Hughes YAH-64 Helicopter,December 1976.

2. Final Report, USAAEFA Project No. 77-36, Engineer Design Test 1, HughesYAH-64 Advanced Attack Helicopter, September 1978.

3. Final Report, USAAEFA Project No. 78-23, Engineer Design Test 2, HughesYAH-64, Advanced Attack Helicopter, June 1979.

4. Final Report, USAAEFA Project No. 80-03, Engineer Design Test 4, YAH-64Advanced Attack Helicopter, January 1980.

5. Letter, USAAEFA, DAVTE-TI, 9 March 1981, subject: Report, EngineerDesign Test, Government EDT-5 of the Advanced Attack Helicopter (AAH),USAAEFA Project No. 80-12.

6. Letter, AVRADCOM. DRDAV-DI, 1 April 1981, subject: Airworthiness aneFlight Characteristics (A&FC) Test of the YAH-64 Advanced Attack Helicopter,Prototype Qualification Test-Government (PQT-6).

7. Test Plan, USAAEFA Project No. 80-17-1, Airworthiness and FlightCharacteristics Test, Part I, YAH-64 Advanced Attack Helicopter, March 1981.

8. Letter, AVRADCOM, DRDAV-D, 28 May 1981, revised 18 June 1981, subject:Airworthiness Release for Airworthiness and Flight Characteristics Test Part Iof YAH-64 Helicopter, S/N 77-23258.

9. Technical Manual, TM 53-1520-238-10, Operator's Manual for Army YAH-64Helicopter, I May 1981.

10. Pamphlet, USAMC, AMCP 706-204, "Engineering Design Handbook HelicopterPerformance Testing," August 1974.

11. Flight Test Manual, Naval Air Test Center, FTM No. 101, Helicopter Stabilityand Control, 10 June 1968.

12. Specification, Hughes Helicopters, AMC-SS-AAH-H100000A, "YAH-64,Phase II Advanced Attack Helicopter Systems," 10 December 1976.

13. System Description, Hughes Helicopters, "YAH-64, Stabilator Control,Phase II," 4 August 1981 (preliminary).

14. System Description, Hughes Helicopters, "YAH-64 Digital AutomaticStabilization Equipment, Phase II," 1 July 1981 (preliminary).

15. Status Meeting (previously reported YAH-64 deficiencies and shortcomings),Hughes Helicopters, YAH-64 Program Manager, Development and QualificationDivision AVRADCOM, USAAEFA, Culver City, CA, 5 August 1981.

25

I-

APPENDIX B. DESCRIPTION

TABLE OF CONTENTS

General............................Dimensions and Genera! Data .................................. .

General ...............................................Main Rotor................................................7Tail Rotor ................................................ 4'Horizontal Stabilizer/ Stabilator ..............................Vertical Stabilizer........................................... 43Wing .................................................... 43Aircraft .................................................. 44

Flight Control Description ........................................ 44General .................................................. 44Cyclic Control System ....................................... 44Collective Control System.................................... 48Directional Control System .................................... 48Trim Feel System ........................................... 48Horizontal Stabilator ........................................ 48Flight Control Rigging....................................... 54Digital Automatic Stabilization Equipment ......................... 54

Hydraulic System............................................... 59General .................................................. 59Primary Hydraulic System.................................... 59Utility Hydraulic System..................................... 59

*.Servoactuators ............................................. 62Power Plant ................................................... 62Infrared (IR) Suppression System ................................... 64Fuel System................................................... 64

26

GENERAL

1. The YAH-64 Advanced Attack Helicopter (fig. 1) is a tandem, two-place twinturbine engine, single main rotor aircraft manufactured by Hughes Helicopters (HH),a division of Summa Corporation. The main rotor is a four-bladed fully articulatedsystem. It is supported by a stationary mast which transmits flight loads directly tothe fuselage. The tail rotor is a four-bladed semi-rigid, delta-hinged system incor-porating elastomeric teetering bearings. The rotors are driven by two GeneralElectric YT 700-GE-70OR engines through the power train shown in figure 2. Anauxiliary power unit (APU) is installed primarily for starting the engines and toprovide electrical and hydraulic power when the aircraft is on the ground and rotorsare turning. The aircraft is designed to carry various combinations of ordinancestores internally in the ammunition bay and externally on the four wing storepositions. The YAH-64 is designed to operate during day, night and marginalweather combat conditions using the Martin Marietta Target Acquisition Designa-tion System (TADS)/Pilot's Night Vision System (PNVS). The test aircraft, S/N77-23258, (photos I through 9) was configured with an aerodynamic mockupof the TADS/PNVS, 30mm chain gun and a HELLFIRE missile launcher loadedwith four dummy missiles on each of the two inboard wing pylons. An alternateconfiguration included a HELLFIRE missile launcher loaded with four dummymissiles on all four wing pylons. The major modifications since Engineering DesignTest 4 (EDT-4) consist of a 13 percent reduction in the pitch axis command

-. augmentation system (CAS) authority and a linear reduction of the pitch axis CASauthority with airspeed from 100 percent at 80 knots true airspeed (KTAS) to zeropercent at 280 KTAS. An additional change was the incorporation of anNOE/Approach mode of automatic stabilator functioning and the removal of thevertical vibration absorber. Various changes in external configuration which affectdrag were made since EDT-4 and are reflected in table 1, and photos 10 through 12.Neither the Back Up Control System (BUCS) nor the Electronic Attitude DirectionIndicator were operational during this test.

DIMENSIONS AND GENERAL DATA

A&FCPart I

Main Rotor

Diameter (ft) 48Blade chord (in.) 21.0*Main rotor total blade area (ft 2) 166.5Main rotor disc area (ft 2 ) 1809.56Main rotor solidity (thrust weighted, no tip loss) 0.092Airfoil HH-02**Twist -9 degNumber of blades 4Rotor speed at 100 percent N (RPM) 289.3Tip speed at 100 percent NR &t/sec) 727.09

* Includes tips**Outer 20 inches swept 20 degrees and transitioned to an NACA 006 airfoil

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Table 1. External Configuration Changes Since EDT-4

1. Removal of wing tip light fairing, Engineering Change Request and Record(ECRR) Number 39322, dated 7 August 1980

2. Installation of one instrumented main rotor blade and two instrumented tail: rotor blades

3. Installation of rotating PCM Canister (photo 10)

4. Installation of rotating PCM antenna brackets (photo 10)5. Installation of one instrumented main rotor pitch change link (photo 10)

6. [astallation of tail boom stiffeners (photo 11)

7. Installation of wing stiffening (photo 12).

39

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Photo 10. Main Rotor Rotating Instrumecntation

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Part 1

-' Tail Rotor

Diameter (ft) 9.17Chord constant (in.) 10Tail rotor total blade area (ft2 ) 14.89Tail rotor disc area (ft2 ) 66.0Tail rotor solidity 0.2256Airfoil NACA 632-414 (modified)Twist (deg) 8.8 washoutNumber of blades 4Rotor speed at 100 percent N (RPM) 1403.4Distance from main rotor mas

centerline (CL) (ft) 29.67Tip speed at 100 percent NR (ft/sec) 673Teetering angle (deg) 35

Horizontal StabilizerlStabilator

Weight Ob) 77.3Area (ft ) 33.36Span (ft) 10.67Tip chord (ft) 2.65Root chord (ft) 3.60Airfoil NACA 0018Geometric aspect ratio 3.41Incidence of chord line (deg) Variable (45 degree leading

edge up to 10 degreeleading edge down).

Sweepback of leading edge (deg) 2.89Sweepback of trailing edge (deg) -7.23 deg (swept forward)Dihedral (deg) 0

Vertical Stabilizer

Area (from boom C ) (ft2 ) 32.2Span (from boom C'j,) (in.) 113.0Root chord (at boom CL) (in.) 44.0Geometric aspect ratio 2.5Airfoil NACA 4415 (modified)Leading edge sweep (deg) 29.4Vertical stabilizer trailing edge deflection 16 deg left above W. L. 196.0

Span (ft) 16.33Mean aerodyiamic chord (in.) 45.9Total area (ft ) 61.59Flap area (ft2 ) 8.71 (fixed)

4 Airfoil at root NACA 4418

43

I.

A&FCPart 1

Aircraft

Fuel quantity (gals.) 369Design gross weight (ibs) 14525Maximum gross weight (Ibs) 17850

FLIGHT CONTROL DESCRIPTION

General

2. The YAH-64 helicopter employs a singe hydromechanical irreversible flightcontrol system. The hydromechanical system is mechanically activated withconventional cyclic, collective and directional pedal controls, through a series ofpush-pull tubes attached to four airframe-mounted hydraulic servoactuators. Thefour hydraulic servoactuators control longitudinal cyclic, lateral cyclic, main rotorcollective and tail rotor collective pitch. Cyclic and directional servoactuatorsincorporate integral SAS actuators. Hydraulic power is supplied by two inde-pendent 3000-psi hydraulic systems which are powered by hydraulic pumpsmounted on the accessory gearbox to allow full operation under a dual-engine

* " failure condition. A Digital Automatic Stabilization Equipment (DASE) system isinstalled to provide rtc damping. The DASE control authority is limited to10 percent of pilot control authority in pitch, roll, and yaw. The DASE alsoprovides attitude hold and a Hover Augmentation System. An electrically-actuatedhorizontal stabilator is attached to the lower aft side of the vertical stabilator.Movement of the stabilator can be controlled either manually or automatically. ATrim Feel System (TFS) is incorporated in the cyclic and pedal controls to provide acontrol force gradient with control displacement from a selected trim position. Atrim release switch, located on the cyclic grip, provides either a momentary orcontinuous interruption of the TFS in all axes simultaneously to allow the cyclic orpedal controls to be placed in a new trim position. Full control travel is 11.0 inchesin the longitudinal control, 9.0 inches in the lateral control, 12.8 inches in thecollective control, and 4.7 inches in the directional pedals.

Cyclic Control System

3. The cyclic control system (fig. 3) consists of dual-tandem cyclic controlsattached to individual support assemblies in each cockpit. The support assemblyhouses the primary longitudinal and lateral control stops, and two linear variabledisplacement tranducers (LVDT) designed to measure electrically the longitudinaland lateral motions of the cyclic for DASE computer inputs. A series of push-pulltubes and beilcranks transmits the motion of the cyclic control to the servoactuatorsand the mixer assembly. Motion of the mixer assembly positions the nonrotatingswashplate, which transmits the control inputs to the rotating swashplate to controlthe main rotor blades in cyclic pitch (fig. 4). The cyclic stick grips are shown infigure 5. A stick fold linkage is provided to allow the copilot/gunner (CPG) to lowerthe cyclic stick to prevent interference when operating the weapon systems.

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Collective Control System

4. The collective pitch control subsystem (fig. 6) consists of dual-tandem controlswhich transmit collective control inputs to the main rotor through a series ofpush-pull tubes and bellcranks attached to the collective servoactuator. Motion ofthe servoactuator is transmitted through the mixer assembly to the swashplate tocontrol the main rotor blades in collective pitch. Collective inputs are alsotransmitted to the load demand spindle of each engine hydromechanical unit(HMU). The HMU meters the fuel as appropriate to provide collective pitchcompensation. Located at each collective control base assembly are the primarycontrol stop, and LVDT, and a 1 g balance spring. The LVDT supplies electicalinputs to the stabilator control units.

5. The collective control stick (fig. 7) incorporates a switch box assembly, anengine chop collar, a stabilator control panel and an adjustable friction control. Theengine chop collar allows rapid deceleration of the engine to flight idle, primarily toallow immediate action in tile event of a tail rotor failure.

Directional Control System

6. The directional control system (fig. 8) consists of a series of push-pull tubesand bellcranks which transmits directional pedal inputs to the tail rotor hydraulicservoactuator located in the vertical stabilizer. Attached to each directional pedalassembly are the primary tail rotor control stops and one LVDT. A mechanicalhardstop was installed on the test aircraft directional pedal to limit tail rotor bladeangle to 27 degrees. However, this hardstop did not prevent actuator overtravel(beyond 27 degree blade angle) with DASE (command augmentation system) input.Two sets of wheel brake cyclinders are attached to the directional pedals and a360 degrees swiveling tail wheel is incorporated. The tail wheel may be locked in thetrailing position by means of a switch located on the pilot's instrument panel.

Trim Feel System

7. A TFS is incorporated in the longitudinal, lateral, and directional controlsystems. The TFS uses individual magnetic brake clutch assemblies in each of thecontrol linkages. Trim feel springs are incorporated to provide a control forcegradient and positive control centering. The electromagnetic brake clutch is poweredby 28 VDC and is protected by the trim circuit breaker. A complete DC electricalfailure will disable the TFS and allow the cyclic and directional pedals to movefreely without resistance from the trim feel springs. The trim release switch on thepilot's cyclic grip allows either momentary or discreet release of the TFS. The CPGhas a momentary release capability only.

Horizontal StabilatorAs 8. The horizontal stabilator is attached to the aft lower portion of the vertical

stabilizer. A dual, series 28 VDC electromechanical actuator allows incidencechanges of +45 to -10 degrees leading edge up (LEU) of travel. Safety featuresinclude an automatic shutdown capability which allows operation in the manualmode by means of a stabilator control panel located on each collective stick. Anaudio tone is associated with the failure of the automatic mode of operation. Astabilator kill switch, located on the pilot's collective stick, disables both theautomatic and manual operation to protect against a hardover failure. There are

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three modes of stabilator operation: the automatic mode, the NOE/Approach mode-.- and the manual mode. The stabilator is controlled in the automatic mode by two

stabilator control units (SCU's). Each SCU controls one side of the dual actuator.Both SCU's receive collective control position information from an LVDT. Twoindependent pitch rate gyros provide pitch rate information to the SCU's (one gyrofor each SCU). The Air Data System (ADS) provides airspeed to both SCU's.

* Additionally, the left-handed pitot-static system supplies airspeed to one SCU and* the right-hand system provides airspeed to the other SCU. Both SCU's receive

position information from both sides of the dual actuator. The maximum rate ofstabilator travel is 7 degrees per second.

9. The automatic mode is operational when the aircraft has normal AC and DCelectrical power applied. Automatic positioning of the stabilator during flight is

. primarily a function of airspeed and collective position as shown in figure 9. Thestabilator also responds with a low gain (0.2 deg/sec/sec) and limited authority(±5.0 deg) to pitch rate inputs to the SCU. Software limits in the SCU limit theincidence change in the automatic mode to those shown in figure 9.

10. The NOE/Approach Mode is selected through the NOE/APPR mode switchon the pilot's DASE panel and will stay engaged at any speed. The mode becomesoperational below 80 knots indicated airspeed (KIAS) and will bias the stabilator to25 degrees LEU at 3.6 deg/sec stabilator rate. The mode can be disengaged bymanual mode selection below 80 knots, or activation of the DASE release or AUTOSTAB reset switch. Acceleration through 80 KIAS will automatically engage thenormal automatic schedule and the stabilator will move at 3.6 deg/sec for the first10 seconds. Failure to revert to automatic schedule will result in system disengage-ment with visual and aural indications.

11. The manual mode can be selected below 80 KIAS through pilot's and CPG'smanual control switch on the collective stick. Manual control selection will result inSTAB FAIL caution and warning annunciation. Selection of automatic mode can beaccomplished by pressing the AUTO STAB reset switch on the pilot's or CPG'scollective stick. The stabilator will move at 7 deg/sec rate to the automatic modeschedule position. Acceleration through 80 KIAS in manual mode will automaticallyengage the automatic mode and the stabilator will move at 7 deg/sec.

12. The SCU's have a fault detection feature which will switch the stabilator modeof operation from automatic to manual if any of the followilng conditions aresensed:

a. A mismatch between the positions of the two halves of the actuatorequivalent to 10 degrees of stabilator travel (If there is a runaway failure of one sideof the actuator, this feature will disable the automatic mode after 10 degrees ofstabilator travel.)

b. A mismatch in the rates of actuator travel of more than 10 degrees persecond (This could only occur if one side of the actuator were extending while theother was retracting.)

c. The stabilator at a position of 20 degrees or greater with an airspeedgreater than 1I 0 KIAS

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53

d. A stabilator angle of 25 degrees or greater with airspeed less than 80 KIAS

. e. Improper AC voltage (less than 23 VAC for more than 0.15 seconds)or improper current (greater than 6.5 amps DC for more than 5.5 seconds).

SFlight Control Rigging

13. A flight controls rigging check was performed in accordance with proceduresoutlined in HH Experimental Test Procedure (ETP) 7-211500000, dated1 December 1980, (main and tail rotor controls) and ETP 7-211123600, dated21 April 1980 (horizontal stabilator). Horizontal stabilator schedule is shown infigure 9. Tables 2 and 3 present the collective and cyclic rigging.

14. Tail rotor rigging is shown below:

Full right pedal: 15.1 degrees thrust to leftFull left pedal: 27.0 degrees thrust to right

Digital Automatic Stabilization Equipment

15. The DASE provides rate damping (SAS), control augmentation (CAS), hoveraugmentation (HAS), attitude hold, and turn coordination. A preliminary DASEdescription is provided in reference 14, appendix A. The DASE is controlled by thedigital automatic stabilization equipment computer (DASEC). The DASEC receivesinformation from several sources. The heading and attitude reference system(HARS) provides the DASEC with aircraft angular velocities (3 axes), aircraftattitudes (pitch and roll), and inertial horizontal velocity (measured by the Dopplerradar). The Air Data System (ADS) provides lateral and longitudinal airspeed, andsideslip angle. The LVDT's provide longitudinal, lateral, and directional controlposition information. The electronic attitude direction indicator (EADI) providesturn rate. The DASEC processes this information and commands control inputsthrough the electro-hydraulic servo valves on the longitudinal, lateral, anddirectional servoactuators. The DASE authority is limited in each axis to-10 percent of the pilot's control authority. Figure 10 presents a block diagram of

the DASE.

16. The SAS function of the DASE system provides rate damping in pitch, roll,and yaw axes. Each axis is separately engageable through a magnetically held toggleswitch on the DASE panel shown in figure 11. The CAS is used to augment the pilotcontrol inputs. CAS is an automatic function of the DASE whenever pitch and rollSAS are selected and yaw SAS is selected below 60 KTAS.

17. A limited authority HAS mode is provided through pitch and roll SCAS usingrates, attitudes and Doppler corrected inertial velocities from the HARS. HAS isused to reduce pilot workload in gusty conditions by assisting the pilot inmaintaining a desired hover position. HAS is engageable through the DASE panelo .using the same switch designated for attitude hold below 25 knots ground speed and50 KTAS whenever pitch and roll SCAS are engaged. Additionally, a heading holdmode is provided through the yaw SCAS using aircraft heading information from theHARS. This function is engageable whenever yaw SCAS is engaged and the HASswitch is selected.

54

."

M. L

6.

:4 z0 ne -2c - :

L) -f u b t4 tij -11 :4 tl W JjWitj

o- - - - -- -

I.55

Table 3. Computation of Blade Angle TravelPilots Collective and Cyclic Controls

Computation Travel 'Tolerance(deg) (deg)

LONGITUDINAL CYCLIC

I. Forward = 1/2 (Item* 9 - Item 2) =(If Item 2 is leading edge down 20.9 20 (minimum)add Item 2)

2. Aft = 1/2 (Itcm 3 - Item 10)(If Item 10 is leading edge down 11.2 100 (minimmn)add Item 10)

LATERAL CYCLIC

*3. Left =1/2 ([temn 13 - [tern 17)*'.. (If Item 17 is leading edge down 11.4 10.50 (mniln)tim)

add item 17)

4. Right 1/2(Item 18 -Item 14)(If Item 14 is leading edge down 7.6 7.0 (minimum)add Item 14)

COLLECTIVE

5. Full pitch travel (Item 5 - Item 6)(If Item 6 is leading edge down 19.2 1 8.0, (11111ltn)add Item 6)

6. Collective pitch full downMeasured 6 3/4 radius -1° to +20(Theo. chord line)

Measured (a' pitch housing(Bolt pad mchined surface 2.4 inches -9.? -100 to -7'inoard of lead-lag hinge)

*Iteinl numbers obtained from table 2

56

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S-j -Ij __7

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-A-4---- 4 C_-

DASE

BUCS/ST YAW

P L Tc C P G

OFF

NOE/APP ROLL

OFF OFF

ATTD/HOVE RHOLD PITCH

OFF OFF

Figure I I DAS Control Panel

58

18. A limited authority attitude hold mode is provided through pitch and rollSAS. Attitude Hold is engageable through the DASE panel above 60 KTASwhenever pitch and roll SAS are engaged. Attitude Hold will automatically disengagewhenever the airspeed is decreased to 50 KTAS.

19. A limited authority turn coordination function is provided through yaw SASusing sideslip information from the ADS. This function is automatically providedabove 60 KTAS whenever yaw SAS is engaged.

HYDRAULIC SYSTEM

General

20. The hydraulic system consists of four hydraulic servoactuators powered* simultaneously by two independent 3000-psi hydraulic systems. Each servoactuator

7. simultaneously receives pressure from the primary and utility systems to drive thedual-tandem actuators. This design allows the remaining system to automaticallycontinue powering the servos in the event of a single hydraulic system failure. Thetwo systems (primary and utility) are driven by the accessory gearbox utilizingvariable displacement pumps, independent reservoirs and accumulators. The APUdrives all accessories, including the hydraulic pumps, when the aircraft is on tileground and the rotor is not running. The accessory gearbox is driven by the maintransmission during flight and provides for normal operation of both hydraulicsystems during autorotation. An emergency hydraulic system is provided to allowemergency operation of the flight controls in the event of a dual system failure.

Primary Hydraulic System

2 1. The primary hydraulic system (fig. 12) consists of a one-pint capacity reservoir,which is pressurized to 30 psi using air from the shaft-driven compressor- anaccumulator, which has a nitrogen precharge of 1600 psi, designed to reduce surgesin the hydraulic system; and a primary manifold that directs the fluid to the lowerside of the four servoactuators. The primary system also provides the hydraulicpressure for operation of the DASE and BUCS functions.

Utility Hydraulic System

22. The utility hydraulic system (fig. 13) consists of an air pressurized 1.3 gallonreservoir and a 3000- psi accumulator which drives the APU starting motor. Theutility manifold direct- fluid to the upper side of the servoactuators, the storespylon system, tail wheei lock mechanism, area weapon turret drive, and rotor brake.Other manifold functions include an auxiliary isolation check valve which isolatesthe area weapon turret drive and external stores actuators when either a low pressureor low fluid condition exists: a low pressure sensor isolates the accumulator as anemergency hydraulic source for the servoactuators in the event of a dual hydraulicsystem failure. The accumulator assembly stores enough fluid for emergencyoperation of the flight controls through four full strokes of the collective stick andone 180 degrees heading change. The emergency system may be activated by eitherthe pilot's or CPG's emergency switch. An electrically activated emergency shutoffvalve is designed to isolate the utility side of the directional servoactuator and thetail wheel lock mechanism when a low fluid condition exists. However, the electricalconnections were not installed during this test.

5 96'

HEAT LATERAL

EXCHANGER/ SERVO ACTUATOR(

DIRECTIONALSERVO ACTUATOR

COI LECT IV[

SERVO ACTUATOR

PR IMARY SIGHTlciPUMP CGE ii

AIR-REL IEF

HAN D l LONGITUOINAI

PUMP [KI SERVO ACTUAT[OR

%-J)) k J

PPIWARYMAAN 110 Di

PR IMARYt10N

Z FLUIIF~

Figure 12. Primary Hydraulic System

60

LONGITUDINAL

DIRECT IONAL LOIrI:AIISERVO ACTUATOR SEROILTv

WEAP~ONJURRLT

"*AU

Fiur I3ItliyIdrui Sse

HEEaLOCK LTI6!

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-

Servoactuators

23. Individual hydraulic servoactuators are provided for longitudinal, lateral,collective, and directional controls. Each servoactuator (fig. 14) consists of aballistically tolerant housing, a single actuator rod and dual frangible pistons, aLAP assembly, BUCS plunger, and various parts for routing of both primary andutility hydraulic fluid. The system is designed to accomodate all flight loads with afailure of either system. However, DASE and BUCS functions would be lost with

-"failure of the primary system. The BUCS plunger assemblies were installed duringthis test, however, electrical connections were omitted.

POWER PLANT

24. The power plant for the YAH-64 helicopter is the General ElectricYT700-GE-700R front drive turboshaft engine, rated at 1563 shp (sea level,standard day, uninstalled). The engines are mounted in nacelles on either side of the

" - main transmission. The basic engine consists of four modules: A cold section, a hotsection, a power turbine, and an accessory section. Design features of each engineinclude an axial-centrifugal flow compressor, a through-flow combustor, a two-stageair-cooled high-pressure gas generator turbine, a two-stage uncooled power turbine,and self-contained lubrication and electrical systems. In order to reduce sand anddust erosion, and foreign object damage, an integral particle separator operateswhen the engine is running. The YT700-GE-700R engine also incorporates a historyrecorder which records total engine events. Engines S/N 207-239R and 207-258Rwere installed in the left and right positions, respectively. Pertinent engine dataare shown below:

Model YT700-GE-700R

Type Turboshaft

Rated power (intermediate) 1563 shp sea level,standard day, uninstalled

Output speed (at 100 percent NR) 20,952 RPM

Compressor 5 axial staies,I centrifugal stage

Variable geometry Inlet guide vanes, stagesI and 2 Stator vanes

Combustion chamber Single annular chamber

with axial flow

Gas generator turbine stages

Power turbine stages 2

Direction of rotation (aft looking forward) Clockwise

62

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-jUL4 L &AJ = C

5 z am~ -&

(J~4 UU AJ

4a.ccA

ca

r.

Ic!633

°h . .

Weight (dry) 415 lb

Length 46.5 in.

Maximum diameter 25 in.

Fuel MIL-T-5624 (JP-4 or JP-5)

Lubricating oil MIL-L-7808 or MIL-L-23699

Electrical power requirements for history 40W, 115 VAC, 400 Hzrecorder and NP overspeed protection

Electrical power requirements for anti-ice I amp, 28 VDCvalve, filter bypass indication, oil filterbypass indication, and magnetic chipdetector

INFRARED (IR) SUPPRESSION SYSTEM

25. The IR suppression consists of finned exhaust pipes attached to the engineoutlet and bent outboard to mask hot engine parts. The finned pipes radiate heatwhich is cooled by rotor downwash in hover and turbulent air flow in forwardflight . The engine exhaust plume is cooled by mixing it with engine cooling air andbay cooling air (fig. 15). The exhaust acts as an eductor, creating air flow over thecombustion section of the engine providing engine cooling. Fixed louvers on the topand bottom of the aft cowl and a door on the bottom forward cowling provideconvective cooling to the engine during shutdown. The movable bottom door isclosed by engine bleed air during engine operation.

FUEL SYSTEM

26. The YAH-64 fuel system has two fuel cells located fore and aft of theammunition bay. The system includes a fuel boost pump in the aft cell for startingand for high-altitude operation, a fuel transfer pump for transferring fuel betweencells, a fuel crossfeed/shutoff valve, and provisions for pressure and gravity fuelingand defueling. Additionally, provisions exist for external, wing-mounted fuel tanks.Figure 16 is a schematic of the fuel system. Figure 17 shows the locations andcapacities of the two internal fuel cells.

,- . 27. By using the tank select switch on his fuel control panel (fig. 18), the pilot can!- . select either or both tanks from which the engines will draw fuel. With the tank

select switch in the NRML position, the left (No. 1) engine will draw fuel from the* forward fuel cell and the right (No. 2) engine will draw from the aft cell (fig. 19).

When FROM FWD is selected on the tank select switch, the two fuel crossfeed/shutoff valves are positioned so that both engines draw fuel from the forward tank(fig. 20). The FROM AFT position allows the engines to draw fuel from the afttank only (fig. 21). The tank select switch is disabled whenever the boost pump ison. When the boost pump is on, the fuel crossfeed/shutoff valves are positioned toallow only fuel from the aft cell to feed both engines (fig. 22). The air-driven boostpump operates automatically during engine start and may be activated by the switchon the pilot or CPG fuel control panel (fig. 18).

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28. The pilot and CPG also have the capability to transfer fuel between tanks usingthe transfer switch on the fuel control panels (fig. 18). Moving the fuel transferswitch out of the OFF position closes the refuel valve and starts the air-driven pumptransferring fuel in the selected direction (fig. 23).

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APPENDIX C. INSTRUMENTATION

The airborne data acquisition system was installed, calibrated, and maintained byHughes Helicopters. The system used pulse code modulation (PCM) encoding, and

-magnetic tape was used to record parameters on board the aircraft. A boom wasmounted on the left side of the aircraft, extending 52 inches forward of the nose. Apitot-static tube. an angle-of-attack sensor, and an angle-of-sideslip sensor weremounted on the boom. Calibration of the boom airspeed system is presented infigure 1. Instrumentation and related special equipment used for this test follows:

Pilot Station (Aft Cockpit)

Pressure altitude (ship)Airspeed (boom) (sensitive and digital)Tether cable tensionMain rotor speed (digital)Engine torque (both engines)* Engine turbine gas temperature (both engines)*Engine power turbine speed (both engines)*Engine gas producer speed (both engines)*Angle of sideslipEvent switchTether cable angles (longitudinal and lateral)Longitudinal control positionLateral control positionDirectional control positionCollective control positionStabilator angle*Normal acceleration

Copilot/Engineer Station

Airspeed (left pitot)Altitude (ship)Main rotor speedEngine torque (both engines)*Engine turbine gas temperature (both engines)*Engine gas producer speed (both engines)*Total air temperatureTime code displayEvent switchData system controlsFuel used (both engines)

PCM Parameters

Time codeEventMain rotor speedFuel temperature (both engines)Fuel used (both engines)Engine fuel flow rate (both engines)Engine gas producer speed (both engines)Engine power turbine speed (both engines)Engine torque (both engines)

75

ilk ...... ... ..

tUL*-..

-i W~.j~j -P& I Ae...

t.

EQ 776

Engine turbine gas temperature (both engines)Airspeed (boom)Airspeed (ship, right and left)Altitude (boom)Altitude (ship)Total air temperatureAngle of attackAngle of sideslipTether cable tensionTether cable angle (longitudinal and lateral)Control positions:

Longitudinal cyclicLateral cyclicDirectionalCollectiveStabilator

Aircraft attitudes:PitchRollYaw

Aircraft angular velocities:PitchRollYaw

Stability augmentation system actuator positions:LongitudinalLateralDirectional

Air data system:Longitudinal airspeed

*' Lateral airspeedResultant airspeedPressure altitudeAir Temperature

Control actuator positions:Tail rotorCollective pitchLongitudinal cyclicLateral cyclic

Main rotor shaft torqueTail rotor shaft torqueRadar altitudeCenter of gravity normal accelerationVibration accelerometers:

Pilot station ( axes)(pilot seat)Pilot floor (3 axes)Copilot station (3 axes)(copilot seat)Center of gravity (3 axes)

Ii7

• 77I

APPENDIX D. TEST TECHNIQUESAND DATA ANALYSIS METHODS

GENERAL

1. The helicopter performance test data were generalized by use ofnondimensional coefficients and were such that the effects of compressibility andblade stall were not separated and defined. The following nondimensionalcoefficients were used to generalize the hover and level flight test results obtainedduring this flight test program.

. a. Coefficient of power (Cp):

Cp -SHP (550)SPA(nR) 3 (1)

b. Coefficient of thrust (CT):

CT ThrustT pA(62R)2

(2)

c. Advance ratio (p):

1.6878 VT

12R (3)

d. Advancing tip Mach number (Mtip )"

Mtip =1.6878 VT + (92R)

a (4)

Where:

SHP = Engine output shaft horsepower (both engines)550 = Conversion factor (ft-lb/sec)/SHPp = Air density (slug/ft 3 )A Main rotor disc area (ft2 ) = (1809.56)= Main rotor angular velocity (radian/sec) = 30.26 (at 289 RPM)

'-- R = Main rotor radius (ft) = 24Thrust = Gross weight (lb) during free flight in which there is no acceleration or

velocity component in the vertical direction. Tether load must beadded in the case of tethered hover.

* 1.6878 Conversion factor (ft/sec)/ktVT = True airspeed (kt)a = Speed of sound (ft/sec) = 1116.45 .VI0 = (T + 273.15)/288.15T = Ambient air temperature

78

i .

7 7- 7-: -6 7

For a rotor speed of 289 RPM, the following constants were used:

A= 1809.56 ft2naR = 726.34 ft/secA(&2R) 2 = 954657879 ft4 /sec 2

A(92R) 3 = 6.934025959 x 10'ift/sec3

SHAFT HORSEPOWER REQUIRED

2. Engine output shaft torque was determined by the use of the enginetorquemeter. The torquemeter was calibrated in a test cell by the enginemanufacturer. The outputs from the engine torquemeters were recorded on the onboard data recording system. The output shp was determined from the engineoutput shaft torque and rotational speed by the following equation:

21r x Np x QSHP=

33,000 (5)

Where:

NP Engine output shaft rotational speed (RPM)Q Calibrated engine output shaft torque (ft-lb)33,000 = Conversion factor (ft-lb/min)/shp

HOVER PERFORMANCE

3. Hover performance data were gathered during 5-foot and 100-foot tetheredhovering flight. Power was varied between data points from the minimum requiredto maintain tension in the tether cable, to the maximum power available. Cabletension was measured and added to the aircraft gross weight to determine thrust(required in equation 2). To further increase the range of CT and C., main rotorspeed was varied from approximately 97 to 100 percent.

LEVEL FLIGHT PERFORMANCE

4. Level flight performance data were reduced to nondimensional form usingequations 1, 2, and 3. Each speed-power was flown at a predetermined CT withrotor speed held constant. To maintain the ratio of gross weight to air density ratio(W/a) constant, altitude was increased as fuel was consumed.

4 5. Test-day (measured) level flight power was corrected to standard-dayconditions (average for the flight) by assuming that the test-day dimensionlessparameters CPt, CTt, and up are identical to CP , CT s and u,, respectively.

79

a o l u ' " -m m k " ' !I ' I ... m- i ' " t , ' . .-

From equation 1, the following relationship can be derived:

pSHPS =SHP1 L

tPt (6)

Where:

subscript t = Test day

subscripts= Standard day

6. Test specific range was calculated using level flight performance curves andthe measuerd fuel flow.

SR VT

Wf (7)

Where:

SR =Specific range (nautical air miles per pound of fuel)VT= True airspeed (kt)Wf=F uel flow (lb/hr)

HANDLING QUALITIES

7. Stability and control data were collected and evaluated using standard testmethods as described in reference 11, appendix A. Definitions of deficiencies andshortcomings used during this test are shown below.

Deficiency - A defect or malfunction discovered during the life cycle of an itemof equipment that constitutes a safety hazard to personnel; will result in seriousdamage to the equipment if operation is continued; or indicates improper designor other cause of failure of an item or part, which seriously impairs the equipment'soperational capability.

Shortcoming - An imperfection or malfunction occurring during the life cycleof equipment which must be reported and which should be corrected to increaseefficiency and to render the equipment completely serviceable. It will not cause an

* immediate breakdown, jeopardize safe operation, or materially reduce the useabilityof the material or end product.

_ VIBRATIONS

8. The PCM vibration data were reduced by means of a fast Fourier transformfrom the analog flight tape. Vibration levels, representing peak amplitudes, wereextracted from this analysis at selected harmonics of the main rotor frequency. TheVibration Rating Scale, presented in figure 1, was used to augment crew comments

* on aircraft levels.

80

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Z .0LL aLJ0

Uj 0

60. 81

AIRSPEED CALIBRATION

9. The boom pitot-static system and both ship's systems were calibrated by usingthe pace aircraft method to determine the airspeed position error. Calibratedairspeed (Vcai) was obtained by correcting indicated airpseed (Vi) using instrument(AVi) and plosition (AVp ) error corrections.

Vcal = Vi + AVic + AVpc (8)

10. True airspeed (V,) was calculated from the calibrated airpseed and densityratio.

Vt

(9)

Where:

a Density ratio (--) where p is the density at sea level on a standard day.p00

WEIGHT AND BALANCE

11. Prior to testing the aircraft gross weight and cg were determined by usingcalibrated scales. The aircraft was weighed with no fuel, in the clean configuration,with instrumentation on board. The aircraft weight was 11, 61 pounds with alongitudinal cg location at FS 214.1.

HANDLING QUALITITES RATING SCALE

12. The Handling Qualities Rating Scale (HQRS) presented in figure 2 was used toaugment pilot comments relative to handling qualities and work load.

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APPENDIX E. TEST DATA

INDEX

Figure Figure Number

Nondimensional Hover Performance I and 2Level Flight Performance 3 through 12Control Positions in Trimmed Forward Flight 13 and 14Collective Fixed Static Longitudinal Stability 15 and 16Maneuvering Stability 17 and 18Dynamic Stability 19 through 21Low-Speed Flight Characteristics 22 through 29Stabilator Evaluation 30 through 39Stability Augmentation System Failures 40 through 42DASE Evaluation 43 through 46Vibrations 47 through 80Ships's Airspeed Calibration 81 anC 82ADS Airspeed Calibration 83 and 84

84

~FIGURE 1-0O NONDIMENSIONAL HOVER PERFORMANCE

YAH-64 USA S/N 77-23258

WHEEL HEIGHT = 5 FEET . . ...

AVG AVG AVGSYMBOL ROTOR DENSITY

SPEED ALTITUDEOA(RPM) (FT) (0C)

O 289 6320 25.0

281 6340 25.5o 290 11440 13.5

U 282 11440 13.5

280 11440 13.5

NOTES: 1. WINDS 3 KNOTS OR LESS.

2. TETHERED HOVER TECHNIQUE. -

3. SHADED SYMBOLS DENOTE FREE FLIGHT TECHNIQUE.4. CLEAN CONFIGURATION.

112

108

104

100

XU

" . 92

00

AU

800U

96

01

684

00

2 76 0 84 8 92 9 0 10 10LiS T

800

Lo

68472 7 80 4 8 92 6 10 10 10

MANRTRTRS OFIINT 0. TPN.X11T A~R~

NONDIMENSIONAL HOVER PERFORNANKtYAH-64 USA S/N 77-23259

WHEEL HEIGH4T = 100 FEET

AVG AVG AVGSYMBOL ROTOR DENSITY

SPEED ALTITUDE OAT(RPM) (FT) (C

282 5800 19.0

289 5820 -19.0

289 11600 13.0U282 11700 13.5

C>285 11800 14.0

286 11860 14.0

NOTES: 1. TETHERED HOVER TECHNIQUE

2. SHADED SYMBOLS DENOTE FREE FLIGHT TECHNIQUE.

3. CLEAN CONFIGURATION.

4. WINDS 3 KNOTS DR LESS.

4 108

104

86-

CD

iA46

AA

P, 72 7 s 4 889A9

- ------ - --

NONDIMENSIONAL LEVELI. FLGHT PERFORMANCEYAH-64 ,S2 K.fl 77-23258

NOTES: I. 8-HELFIF CONF!KUJRTION2. MAIN OTf? -r-' - 290 RPM3. AVG LCGhVUDINL CO LOCATION

FS ?'O2J (2"4. ZEPC 'S I s1P5. GUFVLS OEQTVFD FROM FIGURES 8 THROUGH 10

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NONOMENSIONAL LEVEL FL~IT PERRMCYAH441 USA-S/N 77-23258 -

NOTES: 1.. -8-4IELLFIRE CONFIGURATIOlt--------2. MAIN ROTOR SPEED =290 RPM3. AVG LONGITUDINAL CG LOCATION FS 2M2A '(FWD)4.- ZERO) SIDESLIP5. CURVES DERIVED FROM FIGURES 8 THROUGH 10

76

72

68

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88

FIGURE 5.*NONDIMNISIONAL _ LEVEL FLIGHT PERFORMANCE_

YAH-64 USA S/N 77-23258

-NOTES-. 1. 9;-HELt.FIRE CONFTIURATIONZ.- MAIN ROTOR SPEED = 290 RPM-

3~AVG~ IONGITUD4#AL- CG* LOCATION--FS--202,L- (NOD)4. ZERO SIDESLIP'5: CURVES DERIVED:FROM FIGURES 8 THROUGH 10

-92

88 --

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4 56.76 80. 84 88 92 00 104

- - tWRUS COEFFCIENT C -X 104W'

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NOND74ENSI-ONAL LEVEL FLIGHT PERFORMAI1E-YAH-64 USA S/N 77-23258

iUtES 'I. B-HELLFIRE-T0NFGURATTOWU2. MAIN ROTOR SPEED = 290 RPM~3 3. AVG L0?NGITU114IAL CG LOCATI0t4-FS- 202.0 (F-WD)4. ZERO SIDESLIP5. CURVES DERIVED FROM FIGURE 11 THROUMh 13

104--- . -- ~

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80

7676 80 84 8s 92- 96 100 104

THRUST COEFFICIENT, C- X 10' Ahk

*9~0

FIGURE 7LEVEL FLIGHT PERFORMANCEYAH-64 USA S/N 77-23258

AVG AVG AVG AVG AVG AVGGROSS LONG.. CG PRESSURE ROTOR CT CONFIGURATION

WEIGHT LOCATION ALTITUDE OT SPEED*.(LB) (FS) (FT) (00) (RPM)

14525 202.0 (FWD) 4000 35.0 289 0.007924 8-HELLFIRE.-

3400

3000

2600

S 2200 AF-cc ~E DT -4

o 1800

1400

1000

60040 60 80 100 120 140 160 18D.

TRUE AIRSPEED(KNOTS).

9'

*FIGUREB8LEVEL--FLIGHT--PEWOWANCE..._TAH-64 LISA, S/N 77-23258

* -AVG -7AVG- tG-- '-~ -AN -AM-- AGROSS LOCATION4 DENSITY ROTOR -CT CONFIGURATION~'--WE IGHT J-LONG OAT -LITD SPEED_ ---

149 20.0 -f~ .25 --- 62- -95- - .073- -fL-I

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FIGURE 9LEVEL FLIGHT PERFORMANCEYA~t-64 USA S/N 77-;23258

AVG AVG CG AVG AVG AVG- AVGGROSS LOCATION DENSITY ROTOR CT CONFIGURATIONWEIGHT LONG LAT ALTITUDE OT SPEED(LB) (FS).) (FT) (0C) (RPM)15500 202.2 (FWD) 0.0 (MID). 9460 15.0 290 0.009031 8-HELLFIRE

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ulL_2400

oc 2000

CURVE DERIVED FROM1600 -- FIGURES-3TH$ROUGH 6

1200

80040 60 80 100 120 140 160 180

- TRUE AIRSPEED (KNOTS)

* 93

= - ..-. tEYUFUQITPERFORMW~E

VG. V.QAVG.'-- AVG -AVGL AVGGROSS LOCATION DENSITY GAT ROTOR C T CONFIGURATION-

WEIGHT LONG LAT ALTITUDE SPEED(La) (FS) (BL (T) (0C) (RPM) X0

*15 .400 201.9 (FWD) 0.0 (MID) 13220. 5.5 289 0.010178 B-HEL LFl RE._

.25

- --.20 - ... _

o 0I o 00 0-0010 90,

u j~ ---- - - - - - -

I--

K --- ~-- 1600

(f7t

-No

40 6U 120.

80'740 6 80IDO 1~- 140 160 -18

FIG)RE I I

AVG - - - 5AG :.AVG AVG 'V

GROS4 LOC ATI 14---DENS IOOP' C. OFGURATIONWEIGHT LONG. IA - ATUU:-.-OT s~

(LB) (FS) K (AL) (FT) (C) (QRPM) X1o'-157W' -6V~.-o fi) 7-'300 20.5 16-HELLFIRE

.20 - ~1.00

15- c5k0~ T5o 00

.10 - 00.90

0 0.80

.2800 .

Cl0.70

S 2400

wU I

2000

1200-

40 DAHE CURV DERVE FR10M60 18..-TFRURE 3ISPE (KROOTH)

LA95

FIGURE 12'LEVEL FLIGHT PERFORMANCEYAH-64 USA S/N 77-23258

___-AVG--- -- AV6-CGv- AV6- -AVG AV- AM~.*GROSS LOCATI DENSITY OT ROTOR CT CONFIGURATION

WEIGHT- -LONG t____AT ALTITUDE OT SPEED-(LB) (FS) (BL) (FT) (0C) (RPM) X10 4

.16420--i- 201-4 (FWD) .0_.. (KID) 10440 -11.0 290 _0.009866 .16-HELLFIRE

1 . 5- -

- U -- -- - - -- -1 .00-

0A0

-- .... DAHD.CREDlVE- FRO1

40 60 80 10/.0 10168

TRUE AI(8PHELL(KIRE)

* 96

t**L FidsiT TR IMED FORWARD FIH

L777L~An AVG--'~~- TRIPMmu DINET h. rol~ FLIGHT4- -t~1A~ (F)- (C) (RPM)

b~ ~ u.LJr ..- OG~ .- O.5- 2S LEVEL

-. ~~2. j-j--K--r Siydetsu devate desceniding conditfon-

L.7:

7:~ ., I

7T717-...

ow:~ bRLC~It TRAVEtI 4 1~4E

:34 -

IIII W7WvlML~AtT~JT~NS-

7 .7 - - -- ~ -.---.-- --

A* 140 -

40 0- 12 6 8

A-4- *A"M-_ _ __ __ __

x. -_Q' .77 m~ Il.232n8

A7I7T777,~~~ ~ OAT.-~ D ODIIE-s NJTTON ROTOR FLRIT

- -- ______.-___________________________ ------- 7 -- 7

~~ t**_ _____

...... ......- ~ -- -

zoA

I -M

It-o 14- 8

L1 -l d -E

w-ll 98 4_ _ 4_

____________ - FIGURE 15COLLECTIVE-FIXED STATIC LONGITUDINAL STABILITY

YAH-64 USA S/N 77-23258VG A VCGAVG AVG AVG _ TRIM DASE

AVGS CGDENSITY RTR FIHGROAS CONDITIONOWIEIGHT lOG LAT ALTITUDE SPEE ROTR NFIGTO'L t (fS- -(BtY (FT) ("C) (RPM)

*-15940 :--204.3 (AFT) 0.0 (MID), 6600 21 289 LEVEL ONNOTES: 1. 16-HELLFIRE CONFIGURATION

2. STABILATOR AUTQ MODE3. TRIM FEEL ON4. ATTITUDE HOLD OFF

* -20 5. SHADED SYMBOLS DENO0TE TRIM

10

10

I-.--

~ - ~ '10-

- -. - -TOTAL DIRECTIONAL.CONTROL TRAVEL =4.7 INCHES

3[

~-~--2 ----

TOTAL LATERAL CONTROL TRAVEL 90INCHES

---- 7 TOTAL LONGITUDINAL CONTROL -TRAVEL --110INCHES- -

LL 4

3040 50 60 70 80 90CALIBRATED AIRSPEED (KNOTS)

f Q99

r .... YAH647FIGURE 16

COLLECTIVE-FIXED STATIC LONGITUDINAL STABILITYYAH-64 USA S/N 77-23258

AVG AVG CG AVG AVG AVG TRIM DASE

GROSS LOCATION DENSITY ROTOR FLIGHT CONDITIONWEIGHT LONG LAT ALTITUDE SPEED CONDITION

(LB) (FS) (BL) (FT) (°C) (RPM)

15560 204.4 (AFT) 0.0 6840 20 289 LEVEL ON

NOTES: 1. 16-HELLFIRE CONFIGURATION2. STABILATOR AUTO MODE3. TRIM FEEL ON4. ATTITUDE HOLD OFF5. SHADED SYMBOLS DENOTE TRIM

cz 10

g 0

S10

-...- TOTAL DIRECTIONAL CONTROL TRAVEL = 4.7 INCHES

4

4 010

- 0 *- 3p

TOTAL LATERAL CONTROL TRAVEL 9.0 INCHES

ti 5

• 4

TOALLNGTDIA CNRO TVL 1. NCE

z Iw .)- 99

TOTAL LONTUDAL CONTROL TRAVEL 9 1.0 INCHES

V) 7* =4

0z -c )- 0e e

U- 3

TT100 I0 120 130 140 150 160

CALIBRATED AIRSPEED (KNOTS)I00

7

RIMMUR 17MANEUVERING STABILITYYAH-64 USA S/N 77-23258

- AVG WC6 - - - -AVG AVG- AVG- FLIGHT DASESY14 GROSS- 1 OCATION DENSITY OAT ROTORCODTNCNIIN

--WE IGHT.. LONG- LAT ALTITUDE SPEED -CODTN CNIIN[LB) CFS)- (BL) (FT) (0C) (RPM)

.___0 _J.370D____264_4_AFT)_- -0 -7440 18-5 289. RT TURN ON*0 15260 204.4 (AFT) 0.0 6940 20.0 289 LT TURN ON

- -- *-N~T -:1. --16. HELUtIRE CONFIGURATION-2. 126 KCAS

-- ---. FORCE TRIM ON _

4. STABILATOR AUTO MODE5. ATTITUDE nOLD OFF

Lu

-10

Op

0-20

CE OLL.

~ -TOTAL LONGITUDINAL CONTROL. TRAVEL. .11.0 -INCHES

C) 0 U

-Iz. 3*~~~~~~ -*._ _ _ _ __ _ _ _ _ _

2- 1.0 1.2 1.4 1.6 1.8 2.0 2.2

C. G. NORMAL ACCELERATION (G)

* 101

4 ... ... . . . . .

IDU'I-

CDC

,r C

C~

C6V

I4/f -A .11 I.~d69 jjI- i ad 79S

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

. . . . . . . . . .

C3,

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

0 0 M ..

. .. . . . . . . .. . . . . .01

-w ) C6

C- - C-C

C;CD

3c C CDCC

.. . ~ .. . . .-. . .

CD ~

00 . .o ... .. . . . .-.. . .

- Fm 0 0

I~ ~ ~ .A E ,'S I-0d41.' OZM=o

0 -103

l . ~ ~ ~ ~ ~ ~ . .. ..~w ... ..... .' ." , • • • ° . .. ... . . ..-. . .. . .

. .. .. .. ... . .. . ... .... ...

Ins

ca

............. ....... ..... .. .....• i

0 . ..

... . ..... . ...

. . . . . . ..C, 4 . . . . . . . . . . . . . . ... . . . . .. . . .. •.. .. . . .. . . .

g, 0 . ..... . . . . . . .... .. ... . .. ... .a t

ia

S a '-- 22

CZ-. ....... " ' . M -'

:c, CD " .. ".x"' : r

ar Q

CD . .. .. .. .. ......... !.. ............. ...>. .......... ....

" -' . . . .... ....... . . .. .. . i .' . . ... . . . .. . . .

(5.- L a. . . . . . .' /

. .. . . . . . . . . . . . . . . . .. . ..... .. .

La a

. . . ." " . ... .3 ......... ... .. ....... ... .. t3 .... ....

""'(SIONX) (93G1 (IJ37 linA wOm " NI"-" 33dSHIV 3n i1V'd3Il IISNV dIIS30IS . . .NOII1SOd 10HIN03 IVNOI133dIO

(SN ONN) (SIONA) (93.) (-N33H3d)

"a 33dSHIV 3nu1 IVNIOAnISNOI a33asMi WOOS 31-AV dIIS30IS HOWV NOIIISOd boivnijv SVS 14VA

t......... . C.(.7

IN ~. I EI0V

2-tI 1- t . * . .- .104.

... ... ... . .

. ... ...

0 0. .. . . . .. 0.. . . . . . .

.. .. . .. . .. .. .. ... .....

m CD C3 La 0

La

... .L... .....a .... ...

O CJZ~.... .....n - - - - - - -

.. ... . . .. . . . .

C-ZN I- 2c

Q WC-.r 0 0 . :C : La:NA .... .. ... . .

00-......... ... .. . . . . . . . . . . . . . . . . . .a. . . .. . .

..0 .C C .. . . . . . . . . . .

... ... ....

0D Zn

0 o .. . . ..1. . . . . . . . . . . . . . . . .. . . .

..... .. . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . .*~~J .j.X .. . . . .. . . . .. . . . . . . . . . . . . . .. . . .

.. . . . .. . . . . . . . . .. . .

La cod

.. .. . ..

~ - La

CD=- .0

A..~ N ... .. .. qC.... .

.. ... ... .. . . .~~- .. ..... . I qN-

.1 *NL £ 1. 1JOl . .. ........ I .... .3~ ....... ~..... ... ... .. .. .. ..... .. ..... ...... ... ..s

LOWSPED FI"JE 22LO PDFORWARD-AND REARWARD.-FLIGUTYAH-64 USA S/N 77-23258

_MG V-C-- AVG -- AVG AVG - -WHEEL BASEGROSS LOCATION. DENSITY OAT ROTOR HEIGHT CONDITIONWEIGHT LONG LAT ALTITUDE SPEED'

*(L.B) (CFS)_ -bT6Q, (fT) - ( C) (RPM)- (FT)15240 2 62.4(Fwd)- -0.0 ( d) 6580 27.0 289 __ 20 _ON

- NOEST I; -8-Hs1 tftr,-ronftgul atm.

2. _denotes extreme travel from trim.3. Test done in winds of 5-knots or lesm.

- -4O000,__ -- ---. 4. Trim feel Wff

j HXIMN CONTINUOUS

'-0-- 0

30I-> UJ

C>

00

I--6

-TOTAL COLLETfIVELCONTROL TRAVEL 4 1.8 INCH4ES

-00

R L) .

zt I- 3 TOTAL DIRECTUINAL CONTROL TRAVEL - 4. INCHES

CI.- .U

00 40 0 0 40 60--RERAD OWR

TONOAL

0I106

FIGURE 23.LM-SPEED FORkARD AND REARWARD FLIGHT.

YAH-64 USA S/N 77-23258

LCAVG-- AVG W AVG AVG AVG WHEEL DASEGROSS LOCATION DENSIY OAT ROTOR HEIGHT CONDITIONWEIGHT LONG LAT ALTITUDE SPEED-(LB) (FS) (e ) (FT) (C) (PH) (FT)

15240 202. (FWD) 0.0 (RID) - 0S40 4.5 289 20 ON

NOTES: . 8-ellftre configuration.

* 2. 1 denotes extreme travel from trim.

3. Test done in winds of 5-knots or less.

4. Trim feel Off

4 0,000

bL . .

M AI|NI CONTIINUS TAIL ROTOR• -1 TORQUE LIMIT

30,

20

U. 10

P. 1. 0-

01

TOTAL COLLECTIVE CONTROL TRAVEL = 12.8 INCHES

, r. U_

", , --TOTAL- DIRECTIONAL CONTROL TRAVEL =4.7 INCHES. .- . .

U 01

TOTAL LATERAL CONTROL TRAVEL 2 9.0 INCHES

2• 3

TOTAL LAT AL CONTROL TRAVEL 9. 0 INCHE

IIO

5

TOTA LOGTUIA CONRO TRVE 20 11. INHERERWR FOR-A9

TREARPE

(KNOTS

70-

.." . . .- - - . ..- -_ . :: - . - . - . - . ,--,, .- _ - . ,. . .. .." • -. -,_. ,. .' _ ---.. . . .. . . . . .. .

FIGURE 24LOW.SPEED-FORWARD AND REARWARD FLIGHT.

YAH-64 USA S/N 77-23258

-- VG. AVG-M,- -AVG AVG AVG WIIEEL DASEGROSS LOCATION DENSITY OAT -ROTOR KEIGHT CONDITIONWEIGHT LONG LAT ALTITUDE SPEED

6LB1 (FS) (6L- . (FT) (°C) (RPM) (FT)

..65-- - 2.1-.9(-Fwd)-.O.O(M4. - 5680 - - -2110'- 290 - - 20 ON

. ..... .... ..... ......... .............. --- T'---+. 4- fHe. -I ire configuration.

.2. I denotes extreme travel from trim3. Test done in winds of 5-knots or less

.... ..-- ---- ----- ,---- .. . ... .................. . .... . . .. . ..-- 4;---Trin feel on-.

-... ..---- -- ---- ....... 5.- Ta rotor torque not available

- -30• -_ .. .. .C- - e- e0 4 eG

-- 00

0

i+ ~-0..- U. .. S..

TOTAL COLLECTIVE CONTROL TRAVEL = 12.8 INCHES

US

:. - TOTAL DIRETIOML CONTROL TRAVEL = 4.7 INCHES---

bb" .171- "-*0 - I. -..... .. . ... . .

* 0.

,,.TOTAL LATERAL CONTROL TRAVEL 9.0 INCHES

W • -

i 9 TOTAL LONGITUDINAL CONTROL TRAVEL =11.0 INCHES-.. - -

8

60 40 20 0 20 40 60REARWARD FORWARD

TRUE AIRSPEEDa (KNOTS)

Jo

'F~R25

SIDEWARD FLIG~tTAW-64 iA S/l9 77-23258'

MG- - AVGC--- -AG--- - . P.Y AVG- WHE DASEGROSS LOCAJIONt DENITY OT ROTOR HEIGHT CONDITION

--- -- EIGHT: LONG _ LAT _ ALTITUDE OT SPEED(LB8) (FS) (BL) (Ft) (0 C) -(A*)- (FT)

14960-.. 202.3_(MN)0.O. (MM-_ .6660. 27.5 289 . 20 ON

NOTES: I.- 8-Helli re configuration.

3. rest donre in winds of 5-knots or less.

~~ ROTOR TORQUE LIMIT

S30

I,-,

t- 0

0.

C . I.-

--TOTAL LNIUIA.CNRLTAE-1-.ICE

6TOTAL_ LOATACNTL COTROVL TRAVEL= 11.0 ECHE

4-p TTAL LAERCTONTRO -MTRAL L x-. INCHES

60n 40 2 2 0 sLE RIGH

TRE ISPE- -(ARM)

K . ~ 4 OTAIRECLONA ~OTRQLTRAVL 47 10 I4)

- FIGUE26SIDEWARD FLIGHT74 YA -64 USA S/N 77-23258

AVG AVG GG AVG AVG AVG_ WHEEL DASEGROSS LOCATION DENSITY OAT ROTOR HEIGHT CONDITIONWEIGHT LONG LAT ALTITUDE SPEED(LB) (FS) (BL) (FT) (00) (RPM4) (FT)

14440 202.5 (FWD) 0-0 (MID) .10600 4.5 -W9 20 ON

.. NOTES-- 1. 8-Hellftre-configuration.

2. ' denotes extreme travel from trim.3. Test done in winds of 5-knots or less.

cc 40, .4. Trim feel Off

oo )TAIL ROTOR TORQUE-a.- cii LIMIT

I~l .--I-- . -

ppp338OO.......O

-~ 10

0

10

8 TOTAL LO2TDNA.OTO9IAE=11OICE

6 . .~ a- 6 TOTAL LONGITUDINAL -TR AVEL, lH ..I.. ..S

!' * i'J '_z . ,_ 61TO~l L TE AL ~ e 'CONTROL RA Ez = 9:0 (N HES ---------

Cc - .- i

7

44

60 40 ' 0 20 40 6

C, I.-

0- 3

... 4 TOTAL DIRECTIONAL CONTROL TRAVEL 4.7 INCHES

320

i2 2

C3)

60 40 20 0 20 40 60LEFT RIGHT

TRUE AIRSPEED(KNOTS)

110

FIGURE 27SIDEWARD FLIGHT

TAH-64 USA S/N 77-23258

AVG CG AVG AVG AVG WHEEL DASELOCATION DENSITY OAr ROTOR HEIGHT CONDITION

LONG AT ALTITUDE SPEEDIFS) k L) (FT) (°) (RPM) (FT)

- C UOl.8(Fwd) O.O(Mid) 5800 21.5 290 20 ON

NOTES: 1. 16-Hellfire configuration.

2. denotes extreme travel from trim.

3. Test done in winds of 5-knots or less.

4. -Tri-feel on.

5. Tall rotor torque not available

.'. ! O 0CC 3. 0003 O.3 30

-TL COLLECTIVE CONTROL TRAVEL = 12.8 INCHES

.,TaL LONGITUDINAL CONTROL TRAVEL 11.0 INCHES

,OTAL LATERAL CONTROL TRAVEL 9.0 INCHES

TOTAL DIRECTIONAL CONTROL TRAVEL 4.7 INCHES-I

60 40 20 0 20 40 60LEFT RIGHT

TRUE AIRSPEEDd(KNOTS)

V T

P, 4 PI -

ip j ..........

......... ....

... .........

112

W2

....................... ....0 - ... ....

000

Oca~

000

0So- 0 ~~L

.. . . ....~

..........c> -

t~. .... 0 La.

C ~~ .. ... .. ... .... .

~t;

o0 C

0 .. .. ... ..

In VS

- InKi li .i *2&I.126 g

(T1s/9i30)l3) (ii31 linU wouAN (±N3383d)31"1 MVA NOIIISOd lloivflov MIVA NOIIISOd lOlLINOD 1 VNOUi.R3I1 NOIIISOd 8oivfliov svs MvA

. . . .. . . . . . ]JNt L *-Su lllmd . .. 1'ddQ . -96S1 =dQl113

FItGURE 3OLOW-SPEED FORWARD AND REARWARD FLIGHT

* TAH-64 USA S/N 77-23258

AVG AVG CG AVG AVG - bAa WHEEL DASEGROSS LOCATION DENSITY ROTORWEIGHT LONG LAT ALTITUDE OAT SPEED HEIGHT CONDITION

(LB) (FS) (BL) (FT) (0C) (RPM) (FT)15180 202.5 (FWD) O.(MID) 720 16.5 289 20 ON

NOTES, 1. Stabilator fixed 250 leading edge up.

2. 8-Hellfire configuration.

.3- 1 denotes extreme travel from trim.

4. Test done In winds of 5-knots or less.. "401J5 Trim feel off.

CD 2 0,50

ll 0" 2o. -

U> Mi AXIMUM CONTIIUOm TAIL~ 0 ROTOR TORQUE LIMIT

~0La 40-

301

W ~ 2

S 101

V 0 - ~ - -----0. ~010

TOTAL COLLECTIVE, CONTROL TRAVEL = .18 IN-CIES

I-16>) 0. U..

TOTAL DIRETUIONAL CONTROL TRAVEL - 4. INCHESA 0 :

-. 4

C>1 - -------

TRUE-~ -JP-E

TOTAL LATERAL CONTROL TRAVEL 190 I.NCHES

5U.,

_j U-

11

- - LdM-SPtED FOMM AND EAMRO FLIGHT______ -- YR-64 US SIN 77-2325

AV -cr - AN- AVG Am - iIEL- GASEM#033 1~XOIIGAT ROTOHUT CONDITION~ WEIGH____ - LONG .ALjTUDEJ SPEE--

(IB~ (S (k)(FT) (00). (RPM4) (FT)- 262 5 tM) OQIID) __~ 700 16Z- -..29 20 ON

- - - ----- ----- .NO=S_ L- StA llator Meda-35P lesdlaq edge -up-

-~ -2. 8-iellfire conftguratioir.

.- - denotes riteme travel from trim.* .ge . .4., _Test done in .winds--of 5-knots or less.-

-- - ---- -~-/ - ~ . Triw-feel off.+TARA -

-* -TAVL -4=-1-7A__ _ _ _ _ _

TOM- ~ ~ ~ ~ ~ _ _ LkERAL -OTO RvL -6ICE

TM- OGTU DINALaE TVECtU ROL TVEL =11.0 INCES

7~ . . . . .

. ---- - - - - __ - . -- ---

115 j

FIGURE 32CONTROL POSITIONS IN TRIMMED FORWARD FLIGHT

YAH-64 USA S/N 77-23258

AVG AVG CG AVG AVG AVG TRIMGROSS LOCATION DENSITY OAT ROTOR FLIGHT

..WEIGHT LONG LAT ALTITUDE SPEED CONDITION(LB) (FS) (BL) (FT) (0C) (RPM)15180 202.1((FWD) 0.0 (MID) 8000 16.0 290 LEVEL

NOTE: 8 HELLFIRE CONFIGURATION0 STABILATOR FIXED 35 DEGREES LEADING EDGE UP0 STABILATOR FIXED 25 DEGREES LEADING EDGE UP

U40

o 3 0J I----- ----- B--- -

I 20

20

10

"-iTOTAL COLLECTIVE CONTROL TRAVEL =12.8 INCHES

k,'.- "11

'J Z 10

- 07

6

4-- TOTAL DIRECTIONAL CONTROL TRAVEL = 4.7 INCHES

_1

oD I

,. ..- -.

TOTAL LATERAL CONTROL TRAVEL = 9.0 INCHES

9 TON6TALDATEAL CONTROL TRAVEL 11.0 INCHES

I--I

M,,

* -

S 9 LONGITUDINAL CONTROL TRAVEL 11I.0 INCHES

" - a 7

6

0 20 40 60 80 100 120 140 160 180

CALIBRATED AIRSPEED (KNOTS)

I",

-.ECTIVt-rLJXED.STAT-I1CLO1h1$ISA trY_____

W T. ot -__DENITYI ~ ~ D CNDI~ _

__-- f 7T) 0YY

I :: -_~lAW-mAQA-m.. .... .... .

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

TOTAL UIR[CiGA W RETRVEtL~ h~

-b7- --7:

1 - - - gi

TO A. L XrU1RA Z TRAV

rk7 -

_ aL

L.4C0 3060 U T.- - .*cAL I 8ArE D SP Yi (P jKOTSt

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oc

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100

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doi0l~vig.S 3A~IID31103

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I *iNI S *-IL i -icna -zs~ls =41.124

. .. . . . 1

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CDC

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GldS 3I V~anmo 03dbi woa31NVdIIS301S SGV NOIIISOd KPOJf1V SVS MVA

.. . . . . . .. . . .

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. . . .. . . . . . . . . . .. . ... ...

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u . . . . . . .

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

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.. .. . .. .. . .. . . .. .. . .

.... . . . ..-

... ... .... C .0 .C..- . .... ..

Cc .. ... ... ...

... .. .. ...

Lo 39 .. ... .. C

o CL 0

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-~~~~. ...- -. . . .... ... . - - .-

AC

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< C-

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- CC>

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(930 (S(A (90

31ON -)31N -0ISI GLA~I W00 31N II3

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co S . .

om 0 0

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. -cJ

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(0S/30 (90 N33~)'I-OII~ HovivssioHU 3 4jws

31 IOM3ai ivilN 31l~l3 OIISdIO NO ld0 ~

(AS93- 000 i38d i i ijAID d31Vb0 MV oi i v O I~ ovivsv v OISd1dN)W0i

*ZIZ -Q

F' .................... .......

.. .. . .....

to j. . 13. 'S S. . i

I,, Z<

a -m QI

izi

. . . . . . . . o

IS tc- c3

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318CO aiivlm NII~ ino v i~ OISdIUO V3V

m; ci

C (30

L-J, 1GURE 41-*~ ~ CHRA~dLARACTERISTICS ____

- ~ YAH6-64 :USA SIN' 77-23258 ' .

7 GROSS ___ 16CAtrON _ ESTY OAT ROTOR FLIGHT* .~~ -IGHT' L LAT -ALTITUDE -SMED C~DTO

(FS -- (

15180 202.5 NOD rA!16). 720 .1.6 289 :LOWSPEED.

3, 8 HELLFIRE CONFIGURATION

* ~WHEL HEIHT 20-PEET0- VERTICAL!

-~C OLATERAL _ /EV =38.5_Hz_

~0A A* - A

- ~~~~~~~ ~ o 0- ,. . - -4't .3 H1-

LUI

0 -.a A t-

-. -/ A ----------

i---- ~~~------ ------ _ _ _ -

M/EV 4 48 R*Z:

0 - m MR- a-__ _ &A - kmaf kaam

* 60 46 20 0 20 -40 60*REARWARD' FORWARD

TRUE AIRSPEED (KNOTS)., -

* 131

- ______- VBRATION CHARACTERISTICS _

YPJ4-64 'USA S/N 77-23258

AYGI_.CG - -A_ YG Y _ _

__ GSSLCAIO DENITY OAT - ROTOR FLIGHTWE_ I __ L.O LAT ALTITUDE SPEED . CONDITION'

1.--- J58h 22,5 FWD) 0. (MrD)- 720 16 .5 289 LOW SPEED-

-NOTES-. 1._ P.LO.T_.LGOR . .___ ,~TMLATqOR FIXED 250 LEADING EDGE UP

3. 8HELLFIRE CONFI TI AJN- -- - --- ~ -:.---~-.~~ 4.- WHEEL HEIGHT- 20 FEET

0 VERTICALo LATERAL

U AU111IUiT~~L - -8/REV 38.5 Hz

0

-a -- --c 1.- -, -- - -

- - - - - -4/REV =19.3 -HZ

0-

-- a---- ------- ------- A

LII

10-4 .. a

1/REV =4.8 Hz

60 40 20 0 20 40REARWARD FORWARD

TRUE AIRSPEED (KNOTS) .-

* 1321

__VIBRATION CHARAcTtRISTICS

-~AVG AWGCG AVG. AVG AVGC~i~IT Y OAV -T FLIGHT

WEIS~$T -:LOKG. LAt, ALTITDE 'SPEED CONDITION- -- tBV -f~t~--f~h--~--(Fh--eY---RP4)----

-,-~0 202.5- CFO)o0.0iv4i0 720 16.s,----N LOW SPEEDNOTES: .1. COPIL T SEAT

2. STABILATOR FIXED 25~ L.EADING EDGEWU 0 VERTICAL3---~~~ ~ ~ ~ LATERAL CN IGR IO

4.- WHEEL HEIGHT 20 f EET -- A LONGITUDINAL

U 8,LREV -.38-5 Hz

0

0 -0

0

Li.2 0

0 _

-------- - --- -2--E 9.6 -Hz------

.. .. .. . .. .. . .. .. . .. .. . 11REV 4.8 Hz

*REARMA D FORWARD _

TRUE AIRStEI KNOTS),

6 133

................ ........-n~*~*l

* ~ ~~~~~~~~~~ ~~~~~ :7---~- .....:: 3 i 2 :r: ~ -'~'. . ____

2 -~~~PUACy~ U'g~7 h

I~~~T_ :_-LM ~T A OL i~

- -~ 4-vIPT':t -

A OM4- ..,. .......

-- 4- -: 4.7~

r _7 U -

EF* -l_

... 2 . __:.. 7 :- j4L

*~~_i __I -F______

1REY -: 4: 8Hz.t -.-.

60L 40 20 _ 020 _ .40 6

d 134

YAH-z34 i 5

AVG- AYGCG AVG AVGGROSS LOCATION 2.OAT ROTOR FLIGHTWEIGHT LONG LAT P.SPEED CONDITION

15180 202.5 (-FWD). 0.0 (iK J. 6.5 289 LOW-SPEEDNOTES: 1. AIRCRAFT C'I

_2._.STABILATOR F 1XK i'FADING EDGE UP-----3. 8 HELLF!'r"E IT.7 c (I7N4. WHEEL HEIGH]" 20 F!_';

u-- 0.3 [ AL1 8/REV =38.5 Hz

0-.2 00

0 000

0 00 0

0 4/REV =19.3 HZ_ 0

U10 t3'

-- 2/REV = 9*. Hz:

aU g2 l) o oa-....

-4E 4;--z

la J

REA/WARD4FORWARD

TR.UE A1F 7iTU (K:*"CTS)

FIGURE 52VIBRATION CHARACTERISTICS

.. YA9-64 USA SIN 77-23258

AVG AVG CG AVG AVG AVGGROSS LOCATION DENSITY OAT ROTOR " FLIGHTWEIGHT LONG LAT ALTITUDE SPEED CONDITION

(LBY "FS)- (BL) (FT) (0c) (RPM)

14720 202.5 (FWD) 0.0 (MID) 790 16.5 289 LOW SPEED

NOTES: 1. PILOT SEAT2. -STABILATOR. FIXED 35c LEADING EDGE UP3. 8 HELLFIRE CONFIGURATION4. WHEEL HEIGHT 20 FEET

0 VERTICAL

0. 3 0 LATERALA LONGITUDIIAL 8/REV = 38.5 Hz

0.2 0.0.8.o °088-0.1 O0

0 0 4/REV =19.3 Hz- -....

0 0

00II

cc __ * c

E, 3 0 c . ...

I/REV 4. 8 Hz.

S60 40 20 3 040 60-- REARWARD FORWARD

TRUE AIRSr'E'Oi ""NOTS)

.. .... . ... . ... . . . .. .. . . ... . . . ..

FIGURE 53VIBRATIONCHARACTERISTICS

YAH-64 USA S/N 77-23258AVGA___ G- AVGL-.---

GRO$SLOC~tONDENSITY OAT ROTOR FIHWEIGHT LO0NG :LAT ALJTUDE SPEED 'CONDITION-(L83) -F~ T L F P T ~(P414720 202., MI.:)j 0.0_ .. Ii 16.5 289 _ LOW SPEED _

NOTES-: 1-:~ -PILGT--fLOORL---- L ---STAB ILAIR.XD!352 LEADJ{G- -EDGE -UP

1i .8 HELLFPIRE CO?'FIGURATION4. WHEELHEIGHT_20 PFEET

0. VERT ICALO LATERAL

8/REV =38.5 Hz

..G0 00

- .. .4/REv = 19,3-Hz

0 - -- - ---

a00

0 oAAcm A A. '

2/REV 9.6 Hz

-

----- ----

60 40 20 _0 20 40 60*REARWARD_ _OWR

TRUE AIRSPEED (kNOtS)

* 137

FIGURE 54VIBRATION CHARACTERISTICSYAH-64 USA S/N 77-23253

AVG AVG CG AVG 4VG AVGGROSS LOCATION DENSITY OAT ROTOR FLIGHT-WEIGHT LONG LAT ALTITUDE SPEED CONDITION(LB) (FS) (BL) (FT) (°C) (RPM)

14720 202.5 (FWD) 0.0 (MID) 700 16.5 289 LOW SPEED

NOTES: 1. COPILOT SEAT2. STABILATOR FIXED 35": LEADING EDGE UP3. 8 HELLFIRE CONFIGURATION4. WHEEL. HEIGHT 20 FEET

0 VERTICAL

01 LATERALI A LONGITUDIRAL 8/REV 38.5 Hz

.1 0 0 0 0 0 0 0

.. . ... C3 ... . .. 0 0 -

_- o 0.64/REV = 19.3 Hz

0.4 ...00

0 0

.... - 0 . .... . . . ----_ _- - . ._

I2/REV =9.6 Hz0. 3n

03

1/REV 4.8 Hz0. 1

40 20 0 20 40 60EARWARD..... FORWARD

TRUE AIRSPEED (KNOTS)

..- 13

FIGURE 55VIBRATION CHARACTERISTICSYAH-64 USA S/N 77-23258

- A.VG. -AVG- CG_ AVG AVG _AVG

GROSS LOCATION DENSITY OAT ROTOR - FPLIGHTWEIGHT LONG LAT ALTITUDE SPEED CONDITION

K(LB)- _(ES) (BL) (FT) (--C) (-RPM)14720 202.5 (FWD) 0.0 (MID) 700 16.5 289 LOW SPEED

NOTES: 1. COPILOT FLOOR- 2. STABILATOR FIXED 35 olLEADING EDGE UP-3 "_ 8 HEaLFIRE CONFIGURATION4. WHEEL HEIGHT 20 FEET

- 0 VERTICAL0 LATERAL

0.3- -LNIUINL-H 8/REV 38.5 H

0.2

- 4/REV =19.3 Hz

0.4La

0

a~ l-. 1o a taI

0.3

- O~ - -REV 9.6 Hz

-O is16a a Q A it~ 9 SI ea a-a

60 40 20 0 20G 40. 60* - RLAMWRD -FORWARD_

TRUE AIRSPEED (KNOTS)

FIGURE 56VIBRATION CHARACTERIS)TICSYAH-64 USA Sj/N 77-23,258

AVG _ AVGCG __ VG AVG AVG_GRS OAINDENSI TY OAT ROTOR FLIGHT

WE1f LONG LAT ALTITUDE SPEED CONDITION(tV(PS) (BL) - (PT) (Yc

14720 _ 202.5 (FWDb)._ 0.0 (MID) 700 16.5 289 LOW SPEEDNOTES:. 1. AIRCRAFT CG

2. -STABILATOR FIXED 35' LEADING EDGE U- P3. 8 HELLFIRE CONFIGURATION

4. WHEEL HEIGHT 20 PEET

o VERTICALO LATERAL

---A LOINGITUT N A L-8/REV 38-5 Hz

0 0

0 0 0 00,90

41REV 19.3 Hz-J-

00-~. 0 o 0 tOa 0 10 0 a

- - 2/IREV =9.6 Hz

---- g. a aO~ -1ao9 2 19 9

I/REV =-4.8 Hz

60 40 20 L0 40 60...REARWARD FORWARD

iRU

4

FIGURE' 57VI RAT IONJHARACIERI STICS.YAH-64 USA S/N 77-23258

AVG -- AVG.G&- -AV - AVG - AVGGROSS LOCATION DENSITY OAT ROTOR FLIGHTWEIGHT LONG LAT ALTITUDE SPEED CONDITION'LST) PT-T.- (..-y ( (CRPM)

15240 202.4 (FWD) 0.0 (MID) 6580 27.0 289 LOW SPEED

NOTES: 1. PILOT SEAT.S2.8HELLFIRE CONFIGURATION3. WHEEL HEIGHT 20 FEET

0. VERTICAL -

o LATERAL

03 A LONGITUDINAL 8/REV = 38.5 Hz

.2 .E_ 3 E... . ..0 -0- O

00

. . . . . . .4/REV = 19.3 Hz

- 0.-

0 0 0

-.J

CD -A

m 2/REV = 9.6 Hz

Lu.

0.2

,- *1!

0 ift- a g 1a c claaga a a a a eI/REV = 4.8 Hz

0,3

0 0.2

60 40- 20 -0 - - 60* REARWARD FORWARD

TRUE AIRSPEED (KNOTS)

S 141

VIBRATI -ON CHARACTERISTICSLOAHN US S/ ROTOR3LIGH

- AVG AVG-CG- .AVG. AVG ~AVG

WEIGHT LONG LAT ALTITUDE SPEED CONDITION(z)(F5 (BL) (FT) (CY- (RPM)

1524 224(FWD)_ 0.0 (MID). 6580 127.0 289 -LOW SPEED

NOTE: 1.COPILOT SEAT4 28 HELLFIRE CONFIGURATION

. HEEL HEIGHT 20 FEET

-0 VERTICAL13 LATERAL.-EILONGITUDINAL 8/REV TU38.5 Hz

S0.20.2 0

C30 0

• ~ ~ I04 20" FDO ID._650 270 28 aO PE

0

4/REV = 19.3 Hz

z.4

Q0

0000

00

D Z

ca 2/REV = 9.6 Hz- 0.3

- 0.2

I-

._ -01 - -

0.3 2/REV =4.8 Hz0.3 -- ----

-0.2 ..

60 40 20 0 20 40 60*REARWARD FORWARD

TRUE AIRSPEED (KNOTS)

O 14--

S. 1 ~~~~FIGUR~E 59 .rT:h 1 Tf7l

K ~ 2.Ji7~ VIBRTION CPARACTEPISTCS j717. .... YA1464 USik S/N 77-23258 A

AV9 ~ A&& AVG A qAiA-..-i IA~Q~ , DENSITY :OA

U_ WEIGH AtTETUD _______

7 -------- PCEAT- C

NEIGH 20 FEET :

............................ .. ..........................................

- ~ LO~41TtDNL _______________

K --

.. .. .......

*A r;6"

m+4

.--4 . ~ .. w

.. . . . .......... . . . . . . . . . . I-1i1

. . . . . . . . . . . _ _. .... ......__ _ _ _

___-7 t--_ _7. 4 REARPE KNT)______

A 7~~i~fUSA fi 7 8 L~E

I.----.IWAT

~II~I:Zi3L-u 7 TT&DI?4AL - ,

VilEV 38-5- 1i

. ... .... - If

- -A _ __ -a-IV___

---- - -- - -- -

.F:..T. . ........a.17 ti*.

. 9r .... .

: -77-----:'

i& w4 lb

- -- E ------

* 144

0i

- VIBRATION CHARACTERISTICSYAH-64 USA S/N 77-23258

- AVG CG . AVG AVG AVGLOCATION DENSITY 'AT ROTOR FLIGHT

LONG LAT ALTITUDE SPEED CONDITION(BE ) FT)- C (RPM)

210.9 (FWD) 0.0 (MID) 5680 21.0 290 LOW SPEED

.TES: 1. PILOT SEAT2. 16-HELLFIRE CONFIGURATION3. WHEEL HEIGHT 20 FEET. VERTICAL- 8/REV 33.7 HZ

E0 LATERALA LONGITUDINAL. [

•0

4/REV = 19.3 Hz

2.41 0~ e 0 o(

0

I--I

2/REV 9.7 Hz

w2

* 1/REV 4.P Hz

40 20 0 2 0 40 60SARWARD FORWARD

TRUE AIRSPEED (KNOTS)

-I-

FIGURE 62.- - YI8RA130N~CIAgACERillC -- -

YAN-64 USA S/-N .77-23258

AVG AV~t - M.G- - AVG FLIGHT-GROSS, LOCATIOt* DENSITY OT ROTOR CONDITION

-EGT- -LON- ---LAT---- --AL-TI-TJOE- __O-T -__SPEED(LB) (FS) (BL . (FT) (00) (RPM)

2 1656 lOW o~tW - 558001. 29UJ LOW -SPEEDNOTES! 1. COPILOT SEAT

-a.-1--IftFIRECOHtATN3. WHEEL HEIGHT 20 FEET

0 LATERAL 8/E 87Hz

0.3

0.2 C3- -~0

K ---- -- -- 4/REV-7 19.3 Hz

0.2-- --- -Y

o -- Al

2/REV 9.7 Hz0.3 - - - - _ _- - -

0. - - - - - - -

-J-

60 40 20 0 20 40 60

RRA--- TRUE AiRPSPEED (KNOTS)-' f"R

* 146

YAH-6 1

_ _AVG _AVG CG Av AVGGROS'S LOCATION ROTOR FIHWEIGHT LONG LAT SPEED ODTO

(Lay7 0"S BL n) R ~14960 202.3 (FWD) 0.0 I'M! ~ ?. 289 LOW SPEED

NOTES: 1. PILOT SEA2. 8 HELLFIR[E : '

3. WHEEL HET

- -- 0 VERTIC-ALO LATERAL.

03A. L0NGlTUL),rlitAL 8/REV- 3&5 Hz

U 0.3

0 0.2300

- ~ ~A ~ A A L A

-4/-R.EV = 19.3_H-z-

0

2/REV 9.6 Hz

Lu

0.2 -

0.3

0.2

0.3

60 40 20 P040 60*LEFT RIGHT

FIUR 64

___ VIBRATION CHARACTERISTICS___ ~-*-- YAH--64 USA S/N 77-23258

AYQ ~ _ ~ AVG.: G AVG AG AG__ _

GROSS LOcATI ON DENSITY OAT ROTOR FLIGHTWEIGHT LING LAT ALTITUDE SPEED CODTO

____14960___ 202.3 '(FWD)- _0.0_(MID_) 6660 2_97_.5. 289_ _. -LOW .SPEED_NOTES -- . COPILOT SEAT

2. 8_HELLFI-RE, .CONFi1GURATION-3. WHEEL HEIGHT 20 FEET

SLATERALALONGITUDINAL 8/REV =38.5 Hz

00

±~~~4 -.- -.- /REV = 19.1 Hz

LjJ

C) '

~~~ -- a_

TRU AISPE (KNOTS)

-4-----14S

-- ,:Vtb tkikk ISTICS:

- t AV. AVIG AGROIS ___ jLcAI*W; JfltISIT' OAT -ROTOR FLIGHT- WEORT ~N. L.. .... ~AtTU S-ECONDITION

~jtF W~ LO' SPEED-4-

*~~... ...4...LKE~W LIT. 8/REV_ 38.5 Hz

mi -4/RIE V. -19. 3, -

*. . . . . .

7-g t _ _ __ _ _ _ __ _ _ _ __ I 1 . z

LU -- ~

~~~ ...H ..... .I.

* 4-

... ... : ~:t ...........-............. .... ... . . . . .

S::.1:~ Z 2Z :::.Zr: . ..... .

0~~.... .... -2Z 2 :- ..

....... J ......0~~~. ....Vt~ RIiLL.

o 149

- FIGURE 66

vzkOA ROTOkC- - WISHT~ - --_OATl COND IT IOn_

-- -0~5--*--~*-----i06O -- ~- 4-SPEE&-z - -Wft -

3- - - MITE~R EGKIM-AT- -

-AT -EPA;,~HTZ FE- e - - * - - : O A T E A L ~ . . .8 / R E V 3 8 .5 H z

4LI V 19.3 Hz_ _

- IL. ZZ~13 7717IIXI_7~~

-~ -- -I E - --------- ---------

--------- -- --

-- ~O~~G-E4a

LEFT__ _ _ _ _ _0 O.L-60-AtX~EEDte TT

* 150

7.7

7 . . ..~~ ..7 ..

............ ... .... ~ ..... .2 L ~ L E

CAT IM-TTOR:.-.I.FSET

q-. .7V'L EA --- ----- -17. 777--

-AUiEGHT. 20 FEET .>j...

'17. 7 .~t W ~ i K ..... . ... .

-- ~~~~~~1 -- - -- - . . .. .. .

--- ------- _

:7'. 7~

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

7:

... ... - 71

FIGURE 68ITIONLCARACTERISTICS-

YAH-64 USA S/M 77-23258

-AVG-- ---- AVG- AV -AVG --n-1-HTGROSS -LOCATION DENSITY OT ROTOR CODTN

--- E-1T------O*4------- LA-- ALTITUDE OAT- SPEED--_ CNDIIO(LB) (:)(BL) (FT) (0 0) (RPM)- ~ ~ . (-68--2~ FWO)G (f) - 5800- ---- 21.5- - -210- -- LOW--SPEED-

NOTE717PILOT SEAT-------- 1±FI--ftffiGRTO M -- O -1-M

- -Etl HEIGHT 20 FEET

0-LAERL /REV 13 Hz

o ___ 4REV 9.3 Hz

4b04~-9 4 -W &..--- L--

a g

-40- 0 4o06

TRUE AIRSPEED (KNOTS)

FIGURE 69VIBRATION.CHARACTERlSTICS

- ~Y~f-E4~ SAN 77-293258-

AVG AVG CG AVG AVG AVG_-~GROSS LOCAkTiON___ DENSITY OAT ROTOR FLIGHTWEIGHT LONG LAT ALTITUDE SPEED CONDITION

16083 210.8 (FWD) 0.0 (MID) 80 2. 9 LOW SPEED[OTES: -1. -COPILOT SEAT

2, __16-HELLFIRECNFIGURATION~3. WHEEL HEIGHT 20 FEET

0.EI1A 1 I- 8/REV-= 38.7 Hz-03 LATERAL

G.2~~~ LONGITUDINAL- -.--

----. 0. _4/REV = 19.3 HZ _

Q _

w ~ ~___ 2RV=97H0,2

0.1~~~1_E _9_,- 7_ Hz__ _ -- - - -__-

) 0.3 AI_,jREV_= _4.8 Hz_

0.2--------- - - - - --

0 0.1

60 40 20 0 20 40 60LE.FT -- TU ISED-- - --- RIGHT

TREARPE KOS

I-.

-- 7--: -7ln 7,l

Fw AVU~IA1 Ai

G~s j~A.NUi77TY-DA ROTOR . FLINHTrir J~LL4 AtTrTUDE: SPEED CONDITION

-5 -7 ----- LEE

S-RffZ9 I.0 9TAt 1AhO

-fOYR~-AL__ffiLLIL ti LATERW -.---..

- _ -38. -H---~- - -

. . . . . . . .. . . . .IA A -,

................L .: .a

t:..... .......

.. .... .~i .....-.

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

..................... . ........ .7:

..................

L -I

154 ...-I-7

7 1- VIBPATION Cr A -R~PI'TIGS

-YAR-64 USA S;-N 77-23258

AVG - __ _AVGCG AVG~ AVG AVGGROSS LOCATION DENS ' TV OAT ROTOR FLIGHTWEIGHT L ~~LAT ALTITWi E SPEED CONDITION

(L151~ -- B (FT) (0,CY (RPMY--- ---

14960 2 02.0 (-FWD) 0.0 (MID) 5620 29.5 290 __ LEVEL

NOTES:. 1. COPILOT SEAT

lw .2. 8-HELLFIRE CONFIGURA\TION

0 VERTICAL13 LATERALA LONGITUDINAL

0w3 - 81M--=- 38.7 Hz

0b0

AI

-J-

LU

m:. -- - 074

00

_ 0.30o ' I

204 0lo 2 4 6

CAIRTL,"iSEE KOS

FIGURE 72VIBRATION~ CRCERISTICS7YAH-64 USA S/N 77-23258

AV --_AVG_: CG AVC- AVG AGIvGROSS - LOCATION 5EN',S T Y OAT ROTOR FLIGHT-

WE16HT LONG LAT ALTITD SPEED CONDI'TION.(UBY is :__) ~ F) ',-C) -(RPn)~-

150201 _202..5_(FWD)_ 0.0, (MID) 5020 20.5 286 LEVL__NOTES: 1 .1 PILOT SEAT.

2--. S HLLLFIRE CONFIGURATION0 VERTICAL

.- ~ LATERAL,A LONG IrlUDINAL

0*3 - S~/-REV S. 5-z

0 00 .1

a-_-

4/3t 9n 3o H

0 0.4

-03 -2/REV =9.6 44z

0.3-

0.3 ---~1/REV =4.8 Hz

0;

20 40 60 61 10C 120 140 160

6ISUD KOS

FrGURE 73 -

VIBRATION CHARACTERISTICS

--AVG _AY;_CG .ffiAYGi_ _.._Y _AYG.GROSS - LOCATION DgNSITY OAT RTRFLIGHWEIGHT LONG LAT ALTITUDE SPEED CONDITION(LB) CL F

15020,2- 'Efl ~ (M1DY)_ 5020::- 2O5- __289 -- -LEVEL

NOTES: -I. COPI-LOT SEATS------- -- 2. - & HELL FIRE CON FI-GURATI104 - --- -*-

0 VERTICAL

A LONGI TUQINAL

-- -- -8/REV =38.5 Hz

G.,2 ----- - - - -

- 0

-~ .& - ~4/REV =19.3 Hz

-0-.2 z ~ - ~

0.3 --- 2R _A~.6 JR_

L-

0.1-

0.3-------- - - - - I/E 4.8 Hz

0 ~0.2- - -

204o 60 80 100 120 140 160

CALIBRATED AIRSPEED (KNOTS)

FIGURE 74

YAH 64 USA S/N' 77-23258

GROS OCAIO -DENS.ITY _OAT _ROTOR.. FLIGHT-WEIGHT - LONG LAT ALTITUDE SPEED CONDITION

(LB) (FS)0CTR M

15020(FWDQI N14- 502 _210_5 --289 . LEVEL

NOTSK4 PLOT FLOMR -

-2-.- 8-HELLFIRE_ CONFIGURATION __

- -0 VE RTI CAL.

- -A3NGITUlINAL

WV 38. 5

... 77~~~ _7 --.-4 R Z

____ A-:- -4 &A

--.- 04 -7 - --- 7--

4- .' 111- A

---------- ------

20 0. --60 80--+ A_.2. 14 6

.!CALIBRATED AIIkSPEII (KNOTS)

* 158

71 .~W ~ ./ .l .t . ... . .

SN _7m : ' T7

:iww ~ ~ ~ ~ ~ L f.. ...... .j'84EL~~E OF$R~~Qi

Ty::---

r 11

. 0:1:

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

Ct-&

* .* : t

f. .4 ....

.. ..... ::... : ...:: : .~ I.. ...

CALIBRATED -AMSEEtN

_ _T * 7.-1 ...-...........- I

.. .. ... ....... ----.0 Jig

608 10 j 4

I-ON C HAA CTEC 7-.W~f USA S/l 7-

DENS ITY Ftz FI GHT -

ALTITrJD rO431

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8,/ REV 38.5 Az

t4:I

2/REV =96$

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CALIBlRAT E~ i

I~~~~~ ~~ ........./N ~:I...!.~AVGt .& ....... .

GROSS LOCATION i~II4 DT * (RFIHWOW~H L( NG LM' ~AM;1TI T7CODTJ

15400 _ 201.9 _WMD~ 1'O 2M- 32C . 89 LEVEL_NOTES: 1. COPILOT SEAT - - ~ LL...

.LATERAL. -- ..-. JL~77 __

03LONGITUDINAL... --

0.2 0 _ _ _ _ _ _ _

-. -- -& -0 - 4j

_ A -&A_ I

U coA*

Lo +o . ____ ____4 RE = 1.3+H

- . . . . ..:

(-)-

J: ------- j~- - ------..

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S. ~ I -... i:t:.::::j..t.wFAL t.-ew

1611

~j~l4c P Vol1914

Mm -"rt icfff

11

4 ';LONGITUDINAI-- i J .' .. 5/l :385SHLz-

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4 __Y

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:, r.w...lW CTERIS TICS - -

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4-

7IiT 1REEO..(KNO)..! ....

163

FIGURE 80VIBRATION CHARACTERISTICS

AVG AVG.CG _ AVG AVG AVGRS LOCATION DENSITY OAT -ROTOR FLIGHT

WEIGHT LONG LAT ALTITUDE SPEED CONDITION- fs) ff-tjL (-iT) -{' Cy -IRPt)-

15780 201.3 (FWD) 0.0 (MID) 7300 20.5 290 LEVEL

NOTES: 1. COPILOT SEAT- -

2. 16- E ILLFIRE CONFIGURATIONO VERTICALO LATERAL

03 8/REV _38.7. iz

0T-~ AL-GTUJA -

o4/REV 19.3 Hz

- -

-0-3_ _2/REV 9,7 _Hzlxi-

Or 2-

- -1/REV =4.8 Hz

co~ 0

K.--20 40 60 80 100 120 140 160

K-CALIBRATED AIRSPEED (KNOTS)

77 ~1P T SAWEED C~IJ1T$ X-ThH. WING1 I .

AVGAVGUG ~ AY __AV~ ME:- FL IGHT

- --US -t~------ - -

*- ~ ~ ~ a HELL.- --...- FI -X ...... TV--.-_- fI $

-7 - - 7 -7

77717: i

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MOTK 6

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..... AVG-CG Vu_ AV v . _ FLIGHTtAT TON D14TVROTOR

-:WE I Git- LOAG ALTTThD DPE

L52) ?2. 5 1 <40J 0. "s ~ 1?3 2, 2 9 LEVEL

*i -

'V)E S: TV. --8,LKLEC1E Cf~,'IJA T .1 ON- YAC![ MET:if U 1'.TZE.Z

77777-7---3 Fb-AD3EiD SYMBOLS '' KEI" Th.Zt PAC

-20

ig4

69 iR-ERO

'NSj T 1)tjl E

FIGURE 83

LONGITUDINAL AR~SPEDAIBATOYAH-64 USA S/N 77-23258

NOTES: 1. ADS -AIR DATA SYSTEM2. TESTS CONDUCTED AT 20-POOT WHEEL HEIGHT

USING A.N AUTOMOBILE FOR PACE3. 8-HELLFIRE CONFIGURATION

0

u- 30

20 -LIMt Of ZERO EfRRO_

0.

-- 0

0

~30 0 00

~40

~20

S 10 LINE OF ZERO ERROR7

oO~0 000 000 o

10

604 0020 40 60REARWARD 0OWARD

PACE VEHICLE TRUE SPEED(KNOT$)

167

NOTES:~~~~ IAEA URE CAIBATOL - AIR DATA -SYSTEM

YAH-64 USA S/N 77-23258

NOE:1. ADS = AIR DATA SYSTEM2. TESTS CONDUCTED AT 20-FOOT WHEEL HEIGHT

USING AN AUTOMOBILE FOR PACE3. 8-HELLFIRE CONFIGURATION

"_40 - --

30

20 LINE OF ZERO ERROR0

ow10 0-

200

F4 0

3000 o-r

0

60 .40 20 0 20 40 60...LEFT *.-RIGHT

PACE VEHICLE TRUE SPEED(KNOTS)

4 168

' --'' -' -- - -

APPENDIX F. ABBREVIATIONS

a Speed of soundA Main rotor disc area (ft 2)

AAH Advance Attack HelicopterAC Alternating CurrentADS Air Data Systemapp AppendixAPU Auxiliary Power UnitAVRADCOM US Army Aviation Research and Development CommandA&FC Airworthiness and Flight CharacteristicsBL Butt LineBUCS Back-up Control SystemC CelsiusCAS Control Augmentation Systemcg Center of GravityCL CenterlineC Coefficient of PowerC[k Copilot/gunnerCT Coefficient of ThrustDASE Digital Automatic Stabilization EquipmentDC Direct Currentdeg DegreeECRR Engineering Change Request and RecordECU Electrical Control UnitEDT Engineer Design TestENCU Environmental Control UnitEPR Equipment Performance ReportETP Experimental Test ProcedureFABS Forward Avionics Baysfig. Figurefs, FS Fuselage Stationft FeetHAS Hover Augmentation Systemg Acceleration of GravityGW Gross WeightHH Hughes HelicoptersHMU Hydromechanical UnitHQRS Handling Qualities Rating ScaleHz HertzIGE In Ground EffectIMC Instrument Meteorological Conditionsin. InchesIR InfraredIRP Intermediate Rated PowerKCAS Knots Calibrated AirspeedKIAS Knots Indicated AirspeedKTAS Knots True AirspeedLEU Leading Edge Uplb PoundLVDT Linear Variable Displacement Transducer

169

NMMtP Advancing tip mach numberNAIPP Nautical Air Miles Per Pound of FuelNOE Nap of the EarthN Power turbine speedN R Main rotor speedOAT Outside Air TemperatureOGE Out of Ground EffectPCM Pulse Code ModulationPIO Pilot Induced OscillationPNVS Pilot Night Vision Systempsi Pounds per Square InchQ Engine output shaft torqueR Radius (ft)ref ReferenceRPM Revolutions Per MinuteSAS Stability Augmentation SystemSCAS Stability and Control Augmentation SystemSCU Stabilator Control Unitsec SecondsSHP, shp Shaft HorsepowerS/N Serial NumberTADS Target Acquisition and Designation SystemTFS Trim Feel SystemTGT Turbine Gas TemperatureUSAAEFA US Army Aviation Engineering Flight ActivityVDC Volts Direct CurrentV Maximum Horizontal VelocityVkC Visual Meteorological ConditionsVNE Never Exceed AirspeedV True airspeedVRS Vibration Rating ScaleWL Water LineWf Fuel flow rate

Greek and Miscellaneous Symbols

A Incremental changep Advance ratiop Air density (slugs/ft 3 )a Air density ratio92 Main rotor angular velocity (radians/sec)- Approximately

4/rev 4th harmonic of the main rotor

170

APPENDIX G. EQUIPMENT PERFORMANCE REPORTS

The following EPR's were submitted:

Number Subject

80-17-1 Yaw SAS Hardovers

80-17-2 Engine Compressor Stall

80-17-3 AUP Operation (sea level)

80-17-4 ENCU Performance (inadequate)80-17-5 Stabilator (automatic mode failure)80-17-6 Pressurized Air System (engine start)

80-17-7 Tailwheel Assembly80-17-8 Engine TGT Fluctuations80-17-9 Engine Power Levers (rotor locked start)80-17-10 Engine Cowling Support80-17-11 Marconi Scale (Np + NR indications)80-17-12 Engine Nose Gearbox Seals80-17-13 IR Suppressor Assembly80-17-14 Intermittent Master Cuation Light80-17-15 Flight Control Rod Ends

80-17-16 SDC Manifold80-17-17 Collective Friction

80-17-18 Stabilator Bearings (bushings)

80-17-19 Thermistor Installation (tail rotor gearbox)

80-17-20 Engine noise80-17-21 Nitrogen Bottle Color80-17-22 Tube and Line ID Tape

.

171

| n a • . . l p I- . r r •

DISTRIBUTION

Deputy Chief of Staff for Logistics (DALO-SMM) 1

Deputy Chief of Staff for Operations (DAMO-RQ) 1

Deputy Chief of Staff for Personnel (DAPE-HRS) 1

N Deputy Chief of Staff for Research Development and Acquisition

(DAMA-PPM-T, DAMA-RA, DAMA-WSA) 3

Comptroller of the Army (DACA-EA) 1

US Army Materiel Development and Readiness Command

(DRCDE-SA, DRCQA-E, DRCRE-I, DRCDE-P) 4

US Army Training and Doctrine Command (ATfG-U, ATCD-T,

ATCD-ET, ATCD-B) 4

"- US Army Aviation Research and Development Command

" (DRDAV-DI, DRDAV-EE, DRDAV-EG) 10

US Army Test and Evaluation Command (DRSTE-CT-A, DRSTE-TO-O) 2

US Army Troop Support and Aviation Materiel Readiness Command

(DRSTS-Q) I

US Army Logistics Evaluation Agency (DALO-LEI) I

US Army Materiel Systems Analysis Agency (DRXSY-R, DRXSY-MP) 2

US Army Operational Test and Evaluation Agency (CSTE-POD) I

US Army Armor Center (ATZK-CD-TE) I

US Army Aviation Center (ATZQ-D-T, ATZQ-TSM-A,

ATZQ-TSM-S, ATZQ-TSM-U) 3

US Army Combined Arms Center (ATZLCA-DM) I

US Army Safety Center (IGAR-TA, IGAR-Library) 2

US Army Research and Technology Laboratories/Applied Technology

. Laboratory (DAVDL-AS/DAVDL-POM) 2

US Army Re-,earch and Technology Laboratories/Applied Technology

f - Laboratory (DAVDL-ATL-D, DAVDL-Library) 2

US Army Research and Technology Laboratories/Aeromechanics

.Z. Laboratory (DAVDL-ATL-D)

"l US Army Research and Technology Laboratories/Propulsion

Laboratory (DAVDL-PL-D) 1

US Army Materiel Development and Research Command (DALO-AV) I

US Military Academy (MADN-F) 1

Defense Technical Information Center (DDR) 12

Program Manager (DRCPM-AAH-SE, DRCPM-AAH-APM-TE) 6

US Army Operational Test and Evaluation Agency (CSTE-TM-AV) I

General Electric - AEG (Mr. Koon) 1

US Army Materiel System Analysis Agency (DRXSY-AAS) 1Hughes Helicopter (Mr. Gerry Ryan) 3

US Army Missile Command (DRCPM-HDT-T) I

a.

I.

a