I-.

0 ~T1 F~E COPYCO USAAEFA PROJECT NO. 85-07

US ARMAIAIY

0 SYSTEMS COMMAND

FLIGHT CHARACTERISTICS TEST OF THE UH-60AWITH TAIL BOOM MOUNTED STRAKE

U ROY A. LOCKWOOD WILLIAM A. KELLYMAJ, AV PROJECT ENGINEER

S PROJECT PILOT

A RANDALL W. CASONA MAJ, AV

A PROJECT PILOT

"E

F OCTOBER 1986

A FINAL REPORT

ibTICELECTZI•* CMAY 2 919874S,

APPROVED FOR PUBLIC RELEASE, DISTRIBUTION UNLIMITED.

US ARMY AVIATION ENGINEERING FLIGHT ACTIVITY"EDWARDS AIR FORCE BASE, CALIFORNIA 93523. 5000

~ISCLAIMER NOTICE

"[he' findingo~ of lhiK report are not to he conotrited as an official D~epartment of

ltie Army position unless so designated by other authorized documents.

I~etrv hi reor wenDISPOSITION INSTRUCTIONS

Destoy hisreprt henit 6N no longer needed. Do not retmrn it to the originator.

'rRADE NAMES

Thie itse of trade namnes in this report does not constitute ant official endorsementor approval of !he iise of the corninercial hardware and software.

UNULe~bblr titoSECURITY CLA •1HIC kTION OF THIS PAGE ("an Dait PnCorodJ

REPORT DOCUMENTATION PAGE f REaD INsT ucTI-ONsi RFPORT NtJM8ER 2. GOVT ACCESSIOcN NO. 3. RECIPIENT", CATALOG NUMF3ER

___________ ______BEFORE COMPt, ETrNc, FORM

USAAEFA PROJECT NO. 85-074. TITLE (mid Subtitle) S. TYPE OF REPORT & PERIOD COVERED

FLIGIIT CHARACTERISTICS OF THE UHi-60A WITH FINALTA IL ROOM MOUNTEID STRAKE 11-19 AUGUST 19R6

6. PERFORMING ORG. REPORT NUMBER

"7. AUTHOR(a) ". CONTRACT OR GRANT NUMBER(@)

ROY A. LOCKWOOD, WILLIAM A. KELLYRANDALL W. CASON

S. PERFORMING ORGANIZATION NAME AND ADDOAES i0. PROGRAM ELEMENT, PROJECT. TASKAREA A WORK UNIT NUMUFRS

US ARMY AVN ENGINEERING FLIGHT ACTIVITYEDWARDS AIR FORCE BASE, CA 93523-5000 805-60108 (order #)

11, CONTROLLING OFFICE NAME AND AODRESS 12. REPORT OAE,US ARMY AVIATION SYSTEMS COMMAND OCTOQRR 19964300 GOODFELLOW BOULEVARD IS. NUMMER OF PAGESST. LOUIS, MO 63120-1798 i. CT..14, MONITORING AGENCY NAME 6 ADDRESS~ii dlifferent from Controlflnd 011100) IP SCRIYCAS.(f hareot

UNCLASSIFIED

,so OkASW IIC ATI ON/DO W-NG R-ý01N 0SCU OULE

IS. DISTRIBUTION STATIMINT (of this Repo-rf)

Approved for public release; ýiistrlbution unlimited.

Il. DISTRIBUTION STATEMENT (of the abltract •nlterd In Block 20, If dlfferent from Report)

16. SUPPLEMENTARY NOTES

It. KEY WOAOS (Continue on revfrat sId. It neces.sr aid identity by block number)

l'ortformt •ice and thlAd iing Qualit lserTaIil. Boom Mounted StrakceU1I-60A HIeltcopter

410. A&rqACT (Coorthlue a, o 1i .v &trb It nme•Wlt Idenltiy by block numberJ

Testing was conducted to evaluate the effect of a tail boom mounted strake onthe perfo rmance a1id hn1d dlin, gucinitties of the IJII-60A hel0'cop tr. A total of2.0 product tiv, r fight li hUki were flownduring the period 11 through -19 AuguL1986. The 1111-6OA hetlcopter's uverall performance in hover and sideward flight"to 45 Inotts true atrspeed (KTAS) is essentljaljly unchanged by the strake instal1-lat I on. There was nn overall effect of the ,trake instaltiation on the uIll-60A,heilcopter1 l tlight iontrol margins, except for a slight reductiLoil in right

DD IJ AN 147 EDITION OF I1NO6S1 OBSOLETE UNCLASS[IFIEDv .7CUrlTY CL ASSI FICATION OF THIS PAr,. (W?,en bate E•nwol,,)

S sEcuRITY CLASSI•ICATION OF THIS PAOg•(", a KE•---•'d)

"cyclic control displacemenL (ilcLftdse in COrILrol margin) required in right

sideward flight. There was a minor improvement in longitudinal cyclic pilotworkload in ]eft and right sideward flight to 15 KTAS. The pilot did not

detect i significant difference in flying qualities with the strake installed.Th.I ; t t';t rf ltclt 'd Lit Io crllangL with tl !trie Ii l llt ;Ii II d. ,]lso pr v-i0 mo. f1•;Army Avialtun Engineering Flight Activity performance and flying qualitiesevaluations of the UII-60A 11elicopter7> without a tail boom mounted strake havetound no control margin or power required problems with sideward flight.Consequently, no further strake testing should be conducted on the UH-60A

helicopter. The UH-lH helicopter, which has a documented problem with tailrotor control margins during hover and sideward flight, should be considered acandidate for a tail boom strake evaluation.

•.•-•, IJ NCLAS S I F I )

" • ~ ~SEC URITY CL ASSIFIC ATION O F THIS I PAG E(Wher0 D atau A~n er eod)

I4.!* !I J

TABLE OF CONTENTS

Page

I~NTRODUCTION

Background ............. ................... .. *......... 1Test Objective........*.*.................... ..................... 1IDescription............. *00**.......... 1.Test Scope................................ 6 ......................... 2

T esat M'e thod o,1o gy .. 4a. .. .. ... . . . .. . . .. ............. 3

RESULTS AND DISCUSSION

General ................... ............................... 4Pcfracto .... **.*o 4

Handling Qualitie . . ............................ .. ..... 4

CONC USI NS ... ... .... ... ...*o~ess ooss... .s s.. sos 7

RECOM1I.ENDATIONS.............................................. . 8

APPENDIXES

A. References ......... ... ...................... to .............. 911. Descrip~tion . .... ............................ . ................. 10C. Tnstruimentnt ion ..................... o...................15D. Test- Techniq~ues and Dntn Analysis Metods........ 18E. Test D a ta...... .... ... ............... 21

1)I1ST itI BUT ION

I o-li

16 .. c A

INTRODUCTON

BACKGROUND

1. The Acrostructure Directorate, US Army Aviation Research andTechnology Activity, has shown that the installation of a strake(i.e., a spoiler attached to tie upper left shoulder of the tailboom) on a helicopter can result in improved sideward flightcapability and reduction of tail rotor power required in hoverand sideward flight. These results were obtained from testsconducted in NASA Langley's 4xY meter wind tunnel using two-dimensional tail boom cross sections. Preliminary flight testsof the strake mounted on an 01-58A helicopter have confirmedthat there may be an improvement in hover and sideward flightcapability due to the strake. the effects of a str,.re on theUII-60A was expected to be much more significant because the tailboom cross sectional area is larger and the main rotor diskloading is much higher than that of the OR-58A helicopter. As aresult, the US Army Aviation Systems Command (AVSCOM) requestedthe US Army Aviation Engineering Flight Activity (USAAEFA)to conduct a flight characteristics test of the UH-60A helicopterwith a tail boom mountod strake installed (ref 1, app A).

TEST OBJECTIVE

2. The objective of this test was to evaluate the effect of astrake mounted on the tail boon of the UII-60A helicopter.

DESCRIPTION

3. The UH-60A Black Hawk helicopter is a twin-turbine, singlemain rotor helicopter capable of transporting cargo, 11. combaL

troops, and weapons during day, night, visual meteorologicalconditions, and instrument meteorological conditions. Con-en-tional wheel-type landing gear are provided. The main and tailrotors are both four-bladed. Manual main rotor blade and tailpylon folding capnblittles are provided for air transportability.A moveable hurtzontal stabilator is located on the lower portionof the tail rotor pylon. The helicopter is powered by twoT700-(;E-700 turboshaft engines having an InsLalled Lhermodynamicrating (30 minnte) of 1553 shaft horsepower (shp) (power turbinespeed of 20,900 rpm) each at sea level, standard day staticconditions. Insthlied dual engine power is transmission limitedto 2828 shp. A more detailed description of the 1Il1-6OA helicopteris available in the Prime Item Development Specification (PIDS)(ref 2, app A), appendix B and the operator's manual (ref 3).

i N

V, 'i 1, 1 A

4. The test helicopter, JUH-60A US Army S/N 82-23748, is a sixthyear production Black Hawk which incorporates the Extern.-l StoresSupport System fixed provisions and the modified productiv,. stab-ilator schedule. The test aircraft also incorporated reŽorientedairspeed probes and engineering change proposal 0252 wt~fh up-graded the airspeed system with the production re' rict._.ýs, cnupiece spacer blocks, and fairings.

5. The tailboom strake is comprised of four 0.0O3 inch (in.)6061-T6 aluminum alloy sections each 3.0 in. in height, 2.5 In.in width at the base, and 40 in., nominally, in length. Whenfastened end to end to the left side of the UH-60A tailboom, thefour sections run from fuselage station 485 to 643 and areoffset from the vertical axis of the aircraft by approximately40 degrees (deg). The base of the strake is curved to preventchaffing the rivet heads of the UH-60A tailboom surface and isfastent.d by means of aircraft structural screws anchored throughself locking nut plates. The 3.0 in. high aerodynamic surfaceof the strake is oriented 90 deg from the base so as to be normal.to the UH-60A tailboom skin at the point of installation. Thestrake was manufactured and installed by USAAEFA. Photographs ofthe test helicopter tailboom strake and tufting as well as a more

474 detailed description of the tailboom strake and its installationare included in appendix B.

TEST SCOPE

b. Flight testing was conducted at Edwards Air Force Base,California (2302 feet elevation). A total of 8 flights were con-ducted between 14 January L986 and 19 August 1986 for a total of6.9 flight hours of which 2.0 were productive (flown 11 through19 August 1986) Productive flights were limited to low speedflight characteristics in right (90 deg azimuth) and left (270)deg azimuth) sideward flight, USAAEFA calibrated and maintainedall the test instruimentatton and performed all required malnteii-ancte on t.hv helicopter. Flight restrictions and operating linif-tations observed during the evaluation are 1 italned in theoperator's manual (ref 3, app A) and the airworthiness release(ref 4). Testing was conducted at the conditions shown in table Iin accordance with the test plan (ref .'5)' approved by AVSCOM.

2

ft..

Table 1. Test Conditions 1

Average Average Average

Gross Longitudinal DensityWeight Center of Gravity Altitude

Flight Strake (ib) (FS) (ft)

1 OFF 20200 347.2 3730

22 ON 20180 347.1 3600

3 ON 20390 347.0 3420

NOTES:

INormal utility configuration, main rotor speed of 258 rpm,and mid lateral center of gravity location; Boost, FtabilityAugmentation System, Flight Path Stabilization and Trim - ON,Pitch Bias Actuator electrically disconnected at 2.7 in.

2 Fllght aborted when surface winds exceeded 5 knots.

TEST METHODOLOGY

7. A d('talled i1sting of the test Instrumentation is containedin appendix C. Established flight test techniques and dataredticttonl procedures were used (ref 6, app A) and aie describedIn appendtx D. The flight test data were obtained from testLnstrumentat ton displayed on the instrument panel and recordedby onboard magnetic tape recording equipment.

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RESULTS AND DISCUSSION

GENERAL

8. lests were conducted to evaluate low speed flight at theconditions presented in table I to determine the performance andhandling qualities of the UH-60A helicoptcr both with and withoutthe strake installed. Testing was performed at speeds up toapproyimately 45 knots true airspeed (KTAS) in left and rightsidewird flight. A calibrated pace vehicle was used as a speedrefeience. The belicopter was flown in-ground effect at a wheelheigt.t of approximately 15 feet. The low speed flight test dataare dresented in figurer 1 through 3, appendix E. The "I" barsdenjte the minimum and maximum excursions of control positions,engines and tail rotor torques, and tail rotor impress anglescuring the stabilized data points. The "I" bars are included toIndicate pilot workload. Testing was terminated after three

X productive flights when no significant improvement in performanceV or handiing qualites was noted with the strake installation.

PEPFORMANCE

9. The performance of the IMi-60A helicopter was compared withand without the strake during hover and sideward flight up to 45KTAS (fig. 1, app E) while maintaining approximately the sameratio ot weight over sigma. Engine torque required for left side-ward flight was the same with and without the strake installed.Engine torque required for hover and right sideward flight wasslightly higher (15 to 50 ft-lb) with the strake installed.Tail rotor torque required was the same (strake installed oruninntaltcd) during hover annd left cidpward flight up to 35KTAS. From 35 to 45 KTAS left sidward flight and from 5 to 20KTAS and 35 to 45 KTAS in right sideward tlight there was aslight reduction (0 to 30 ft-lb) of tail rotor torque requiredwith the strake installed. Power required variations as a resultof the strake installation were inconsistent and not significant.The UH-60A helicopter's overall peLformance durilng hover indsideward flight to 45 KTAS was essontially unchanged by thestrake installation.

HANDL, NG .qUALT IES

i0. liatidling qualities data with and without the strake installed

were compared during sidward flt ght tip to 45 KTAS (fig I, app E)Lo deterinino Chi, h ,,'ci -t.q of the i;traki( on flight. rontrol. margins.The directional control positlons were similar for hover andsideward flight j, Lo 35 KTAS left and right. From 35 to 45KTAS left, stdeward flight there was up Lo a 0.3 in. increase

I414

•,.

4 1 1

of right pedal required (decrease in control margin) with thestrake installed. At speeds of 35 to 45 KTAS right sidewardflight there was up to a 0.25 in. decrease of left pedal required

(increase in control margin). Longitudinal cyclic control posi-tions were similar during flights with the strake installed anduninstalled. Lateral cyclic positions were similar for hoverand left sideward flight up to 35 KTAS. From 35 to 45 KTAS leftsideward flight there was a requirement for up to 0.2 in. ofadditional left cyclic (reduction of control margin) with thestrake installed. Right sideward flight required approximately0.3 to 0.4 in. less right cyclic (increase in control margin)control with the strake installed. Flying qualities (as deter-mined by flight control excursions, annotated by "I" bars onfigs. 2 and 3, app E) were coi')ared for flight: with the strakeinstalled and uninstalled during sideward flight. Directionalcontrol excursions were similar up to 35 KTAS during left sidewardflight. At 35 to 45 KTAS left sideward flight, the directionalpedal excursions were reduced by +0.5 in. to +0.3 in. when thestrake was installed. Longitudinal cyclic excursiuns were reducedwhen the strake was installed during hover and left and rightsideward flight to 15 KTAS (by 40.5 to +0.2 in.). Longitudinalcyclic excursions were also decreased by +0.8 to +0.6 In. duringflight at 15 to 45 KTAS left sidoward flight with the strake

it installed. There was no change in longitudinal cyclin controlexcursions during right sideward flight above 15 KTAS. Therewere no changes in lateral or collective control excurtionsnoted with the strake on or off. The strake installation on theUIt-60A helicopter had no significant effect on flight ccntrolmargins. A slight improvement in longitudinal cyclic workloadduring left and right sidward flight to 15 KTAS was noted. Thepilot did not detect a tgni f iant difference in handlingqutaLitLes with or without tlhe strake installed.

11. The AerorLctrir ,rate, US Army Aviation Re-scarch andNAlTecnlmolgy Activity expected improvements in hovering performance

and flying qnnliti~es with r(-duced pednl requirement in sidewardflight: (as [ndiciat-.d in rof 7, app A) for a strake Instafllation

onop.or. StI ne, no sign fi'cant implrovetn( t: swere noted, Iitltng was terminated afte r three produc'ive [ tights.Since thL1 teftttle change with the sLrake instl-latton and since previous USAMAFA performance and lantdilig quali-Lien reportg on the aircraft have found no control margin orpower required problems during sideward flight, no father straketettnnyi should he considered on thle UII-60A. The 111-1i1- helficopterwhich has a docieiuntd problem wl h talt rotor control margi• •duringm hover ,ind siideward flight (rot H) might he n much bettercandidate for a tail boom Htrako evaluation.

12. t~arly termination of the test precluded compliance testlng tothe PIDS requirements (ref 2, app A) with the strake installed.

'I~

6

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CONCLUSIONS

13. The overall performance and handling qualities of the Uti-60Ahelicopter during hover and sideward flight to 45 KTAS was essen-tially unchanged by the InstallaLion of the st.rake.

a. Tile overall performance of tile Ull-60A helicopter wa8essentially unchanged by the installation of the strake (para 9).

b. There was no significant effect of the strake installationon the Ull-60A helicopter flight control margins, except for aslight reduction in right cyclic control d[splacement (increaseIn control margin) required In r•ght sideward flight (para 10).

c. There was a minor reduction In longitudinal cyclicworkload during left and right sideward flight to 15 KTAS.The pilot did not detect s significant difference in overallhandling qualities with the strake installed (parm 10).

I

RECOMMENDATIONS

14. The foilc'wtng recommendations are presented regarding thefl.ight characteristics of the UH-60A helicopter with a tail boommounted strake:

a. No futher strake testing should be conducted on the UH-60A(para 11).

b. The± U1I-lil helict)pter has a documented problem with tailrotor control margins during hovc.r and sideward flight andshould be considered for a tail boom strake evaluation (para. 11).

..-

A PPENDIX A. REFERENCES

1. 1Ltter, AVSCOM, AMSAV'-ED, undated, Subject: Flight Character-isLics Test of t-he Uli-60A with Tail. Boom Mounted Strake, USAAEFAProject 85-07.

2. Prime Item Develonr•ent Specification, Sikorsky AircraftDivision, DARCOM CP-2222-SIOOOF, 18 December 1981.

3. Technical Manual, TM 55-1520-237-10, Operator's Manual, IR-60.AlieZL!opter, Headquarters Department of the Army, 21 May 1979 withchange 37 dated 17 July 1986.

4. Letter, AVSCOM, ANSAV-E, 31 July 1986, subject: Release ofUII-60A BLACKHAWK Helicopter S/N 82-23748 to Conduct a Flight Eval-uation of the UH1-60A Helicopter with Tail Boom Strake Installedand Two (2) Light Weight Video Cameras Installed on the Stabil-ator, USAAEFA Project No. 85-07.

5. Test Plan, USAAEFA Project No. 85-07, )P'ligh•ia ,•,%:ter'ti•'.Teet of the U(I-10A wi'ith Tait Room Mo,:anted St'ake, June 1985.

6. Flight Test Manual, Naval Air Test Center, FTM No. 101,He<licopter ,9tabiLity amd Control., 1; June 1968.

7. Letter, National Aeronautics and Spare Administration, LangleyResearch Center, 14 December 1984, subject: Request for Flight-Test Investigation of Tail Boom Strake on UII-60 BLACKHAWK.

8. Final Report, USAASTA Project No. 71-18, Vail Rotor Pekr'ofwm-.v1t!' (t'll '½"a. nil la t F!. i 1 !f~1'id.i, Qd t' Teoto 1111-i//!lr'!'•,'ont)r., .1anuary 1972.

9 Final RH porl , ItSAAFIFA P ro ct No. 84-28 !,',(Ih 't a . i.t. .uit+, .'-• 'I *I" , . .•t " ,,.,*i t hI .l+ pI' A+[,+ 4, A.',z ,+, (P .A o l'(

u~I "..'rt'r,'c , to be pub lished.

I, . tFinal RcporL, IISAAEFA Projecct No. 86-01, ,'r''. r"'.'Vb r-,

rt,"'r , .:U r",' ,•::,gt; 4/:., " .:.rbl ,',r'r'I i",,nkia r t.L'.!.z,', to bepuhi ished.

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APPENDIX B. DESCRIPTION

GENERAL

1.- The Sikorsky UH-60A (Black Hawk) is a twin turbine engine,single main rotor helicopter capable of transporting 11 combattroops plus a crew of three. It Is equipped with three nonretract-able conventional wheel-type landing gear. A movable horizontalstabilator is located on the lower portion of the tail totorpylon. The main and tail rotors are both Lour-bladed with acapability of manual main roilo blade and tail pylon folding.The cr088-beam tail rotor with composite blades is attached tothe right side of tji pylon and is canted 20 deg upward from thehorizontal. The primary power plants for the UII-6Oa helicopterare two General Electric T700-CE-700 front drive turboshaftengines, each rated at 1553 shaft horsepower (slip) (30 minutelimit) at a power turbine speed of 20,900 revolutions per minute(rpm) (sea level, standard day, installed). A complete descrIptionof the aircraft is contained in the operator's manual (ref 3,app A).

BASIC AIRCRAFT INFORMATION

2. General data of the Ull-60A helicopter are as follows:

Gross Weight

Maximum gross weight 20,250 lbEmpty weighe Approximately 10,620 lbPrimnary mission gross weight 16,324 lbFuel capacity (inealuud 360 gal

Mair Rotor

Number of blades 4Diameter 53 ft, 8 in.Blade chord 1.73/1.75 ftBlade twist -18 deg (equivalent)Btade tip sweep 20 deg aftBlado area (one blade) 46.7 sq ftAirfoil

sect I oti ( root to tip de,- ignat tiOl ) SCI095/SCI 095R18thlIvkilq! ( perc t cho rd ) 9 .5 pe rc,.Ž

Main roto mast titL (forward) 3 deg

Tail Rotor

Number of blades 41)i1aR nv Lu r 11 Ft

V 10

A. %

I Jil %. "41*Ct .

Blade chord 0.81 ftBlade twist (equivalent linear) -18 degBlade area (one blade) 4.46 eq ftAirfoilsection (root to tip designation) SC1095/SC1095R8thickness (percent chord) 9.5 percentCant angle 20 deg

Gear Ratios

Input Output

S-rpm Ratio

Engine to main 20,900.0 257.9 81.0419rotor

Engine to tail 20,900.0 1189.8 17.5658rotor

Tail rotor to 1189.8 257.9 4.6136main rotor

STRAKE INSTALLATION

3. The UH-60A tailboom strake (photos I and 2) is comprised offour 0.063 in. AL ALY 6061-TG sections, FED SPEC QQ-A-250/11stock, each 3.0 in. in height, 2.5 in. in width at the base,and 40 in., nominally, in length. When fastened end to end tothe left side of the UH-60A tailboom the four sections run fromfuselage station 435 to 643 and are offset from the verticalaxis of the aircraft by approximately 40 deg. The base of thestrake sections which fasten to the tailboom skin employ a curv-ature across their width which is more severe than that of thetailboom in order that the strake sections may ride above existingtailboom rivet heads and prevent rivet chaffing. The radius ofcurvature for the sections are 2.5 in. and 3.0 in. for the tworear sections and two forward sections, respectively. Both endsof each strake segment are chamfered at 0.50 in. x 45 deg. Thefree edge along the base of the strake (uppermost edge whenmounted onto the tailboom) has been modified with a slight 0.50in. radius reflex to prevent chaffing of the aircraft skin alongthe length of the strake. The transition from the base of thestrake to the 90 deg extension also utilizes a 0.50 in. radiusof curvAture. Additionally, slight cutaways exist in the forward-no~t and two aft segments to allow for clearance of existingrivet patterns along bulkheads within the tailboom. The radiusof each corner in all cutaways is 0.125 in. Blind rivets alongthe tailboom at the points of attachment of the strake were

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periodicially replaced by stainless steel self-locking nut plates,(D.A. 500-83-V-N248), which are fastened with AN426 AD3-6 solidaluminum rivets where work space permitted, and OLA 500-85-M-Y734blind aluminum rivets elsewhere. The strake in secured withMS-27039 1/7 10/32 in. airc'raft structural screws. The longestspan between successive nut plates is 5 in. The ends of eachstrake segment extend no more than 2.25 in. from the closest nutplate, with the only exception being the forward end of theforward segment, where the closest nut plate is 4.75 In. aft.The strake was manufactured and installed by the US Army AviationEngineering Flight Activity.

14

W -I 7 ~ -

APPENDIX C. INSTRUMENTATION

GF N RAL

1. The test Instrumentatlon was installed, calibrated and main-tained by US Army Aviation Englneering Flight Activity person-nel. A swiveling pitot-static tube and angle of attack and side-sllp vanes were Installed on a test boom 162 In. forward of thenose of the aircraft and 29 in. to the right of centerline. Alsoinstalled were two video cameras, one mounted at each stabilatortip, for the purpose of monitoring the aorodynamic behavior ofthe tailboom tuftingi during testing (photo I, app B). Each cameraassembly weighed approximately 2.7 lb. Images wore recorded ontape adjacent to the engineer's station and observed in realtime using a video monitor nearby. Equipment required only forspecific tests was installed when needed and is discussed in thesection on special equipment (parA 2). Data was obtained fromcalibrated instrumentation and displayed or recorded as indicatedbelow.

Pilot Position

Altitude (radar-dual rangto)*Rotor speed (sensittve)%nginv torque* **Turbine gas temperature (T4.5)* **

Engine gas generator speed * **

Control positionsLongitudinalLateralPedalCol locttve

S ta i ln i po M I t. ltronEvoiet switch

~opj lo~t/J.... ,er Station

Rot or A I)Ued*Ing lnu LorquL* **

lot ;I I r t., vnpt ra tIro

Fue I us ed ( Lota I ZVr)lb I [ tL cnarl pos I onTi'ne rode displayRun nitimh:rI'vent HW[LcIe

• Ship sy,,-. m/in tL r llhraLted* *BoLh vilg ino,4

I*I I~l) co '%

Digital Pulse Code Modulation Data Parameters

Altitude (radar)Total air temperatureRotor speedEngine torque**Turbine gas temperatures (T4.5)**Engine gas generator speed**Engine power turbine speed**Engine fuel flow**Engine fuel used**Engine fuel temperature (at fuel used transducer)**Auxiliary Power Unit (APU) fuel usedAPU fuel temperature (at fuel used transducer)Tail rotor shaft torqueTail rotor impress pitchStabilator positionBallast cart positionControl positions

LongitudinalLateralPedalCollective

Stability augmentation system output positionLongitudtnalLateralDirectional

Control mixer input positionLongitud fnAlLateral.Directional

Aircraft attitudePitchRollYaw

Afreraft nngtilnr ratePitch

YawLinear acceleration

Center of gravity normalCenter of gravity lateralCenter of gravity longitudtinal

Time of dayRon number

**JBoth L'ng inea

16

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Data status wordsPilot eventEngineer eve~nt

SPECIAL EQUIPMENT

Ground Pace Vehicle

2. A vehicle utilizing a calibrated fifth ;.rheel to determilneaccurate ground speed was used in conjunction with wind specd anddirection to provide a precise true alropood reference for thetest aircraft during low speed tests.

Wen thc'r Station

3. A portable wentlivr station, consisting of an anelmometer,sensitive ternpernture gage, and barometer, was used to recordwind speed, wind direction, ambient temperature, and pressurealtitude durinS low speed tests.

17

APPENDIX D. TEST TECHNIQUES AND DATA ANALYSIS METHODS

AIRCRAFT RLIGGING

1.A 1`I tp,10 voolirolti onpiI niq'r a rigging rhuick wtgim pi rrifirvil onOv1 mai n tind I nt] rotors durltig 6 previotia test program conductedby the US Army Aviation Engineuring Flight Activity (USAAEFA)(ref 9, App A). The rigging data are presented in table 1.

AIRCRAFT WE~IGHT AND BALANCE

2. A weighing accomplished during a previous test program con-ducted by USAAEFA wasn used an a reference for all $Lrake tooting(ref 10, app A). The Aircraft was weighed in the instrumentedcondition With full fuel and oil onboard. In addituion, themovable ballast systemn, the OxLernal stores support mystem (EBBS)wings with fivings and racks, four external fuel tnnks, and loandbAllant at various fixed locations was installed during theweighing. With the weight of fuel, the weight of the EBBS, theweight of the external tanks, and the weight of fixed bellastsubtracted from the total weighti, the basic weight was 11918.8pound. with the longitudinal center of gravity (ell) at fuselngePi~atton (FS) 356.5 (ballast cart eg FS 301). To this weight wasaidded 5.5 pounds to account foe romote video cameras located atFS 698.0 and 60.0 pounds to account for the video monitoring/recording equipment with support rack located at FS 278.0. Thebasic aircraft weight warn therefore modified to 11,984.3 lb witha longitudinal cg at FS 356.3. This basic weight and eg wasused in all subsequent strike tooting.

3. The fuel weight for each flight wen determinod prior to enginestart and following engine shutdown by using external sight gagesto evaltiate fool volume and by measuring the opecifir weight orthe fuel, Lead weights were secured incide tho Aircraft at fixedlocatlons iind within the movable ballast cart to tadjunt theweight and cg for tust purposes. The ballast cart had A maximumcapacity of 2664 lb and a tutal travel of 72.7 in. (VS 301 toFS 373.7). To minntntin the desired weight during low speedteatit:ng, hallhat: wan ~j ddecd periodically tic FS 124.9) na ruvi wapNhdlrtieci at tI 1111 tnt i rv, In ,'oivn ivt' t onI I 11w order lv uiiipl too i iiLvnt: pot1LH Ton * l a Intalt vi l ,, t.11 hibiitt (. I'Ar.~a Willi I1iOVt'd 1111 IlIilrequ L rud. Viat desired I eat weight. was varied as re~quired ilt tHihegintifng of' eaci, flighit to maInftnin n consistent gross weightover sigma ratio.

NV

Td b Ic 1 .IlCl 11 m tid "Ikd i l O L U L ' K i• . ji ; Ii Lu t'l i L L u o l

Mann Rotor RtggLig

Flight Control Po sttton Sw ,N hpl ate Tilt Coll ictive

(Duroee) |Bblade Pitch

Co L Ivct I vc 1,onjvi tud inni Loteard PedaLI Longitudin~t1 Lateral~ (Degraee")

H-igh * *-42 -3.3 24.3

1,o w AFT Ur -9.4 -7.4 8.81 g1h AVT I,T * -9.2 -Y.6 24 . ilow FWI) RT 1.1 .0 7.2 93.1lhFWD)T * 17.3 6.5 23.4

.AFTLT 1"T -1.1 .1 -7.7 23.6d AVT • i.r1 * -11.7 -7.5 16•6

Mid FWD RIIT -15.6 6.2 15.5Mid * * * -7.4 ]-2.6 17.0

T[ ll Rot~or ULSgA I1

FIP ght Cointrol. Pomition Collt , ttvc Blade Pitchlt the Root(Vogruum)

M I d i(LT L ,

M'd RTr 16.2

bllM 1ll ) 0 .9

LAow HI 11 8.2

1.7.

I,ow ]:T 7..

I o~d 1.'(mt on alppropr-hritt, cmi. r'olt. waO i W pl i [ mluld

TEST TrCIINlQUES4. Conventional test techniques were used during the conduct of

the handling qualities tests. Detailed descriptions of the testtech•niques are contained in reference 6, appendix A.

I0

2()

APPENDIX E. TEST DATA

INDEX

Figure Figure No.

Low Speed Flight CharacturisicsStrake tnstalled/Unlnstalled ComparisonStrake Uninstalled Parameter Excursions 2Strake Installed Parameter Excursions 3

4121

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flu p Ilw

J;. I PARA.ET. . ýý-ýN 5;440 V.:

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LE T. TUE ýIRSkEDý (KNOTS)* 179 -Tr-ý-r

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DISTRIBUTION

1lQDA (DALO-AV, DALO-FDQ, DAMO-HRS, DAMA-PPM-T, 6

DAMA-RA, DAMA-WSA)

US Army Materiel Command (AMCDE-SA, AMCDE-P, AMCQA-SA, 4

AMCQA-ST)

_US Army Training and Doctrine Command (ATCD-T, ATCD-B) 2

"US Army Aviation Systems Command (AMSAV-8, AMSAV-ED, 15

AMSAV-Q, AMSAV-MC, AMSAV-ME, AMSAV-L, AMSAV-N,

AMSAV-GTD)

US Army Test and Evaluation Command (AMSTE-TE-V, 2

AMSTE-TE-0)

LUS Army Logistics Evaluation Agency ()ALO-LEI) I

US Army Materiel Systems Analysis Agency (AMXSY-RV, AMXSY-MP) 8

US Army Operational Test and Evaluation Agency (CSTE-.AVSD-E) 2

US Army Armor School (ATSB-CD-TE) 1.US Army Avation Center (ATZQ-D-T, AT';Q-CDC-C, ATZQ-TSM-A, 5

ATZQ-TSM-S, ATZQ-TSM-LTH)

US Army Combined Arms Center (ATZL-TTE) 1

UIS Army Safety Center (PFSC-SPA, PESC-SF) 2

US Army ost and Economic Analysis Center (CACC-AM) I

LUS Army Avlatloii Rosearch and TechnLology Activity (AVSCOM) 'I

NASA/Ames Res•earch Center (SAVRT-R, SAVRT-M (Library)

US Army Avlation Research and Tuchnology Activity (AVSCOM) 2

AviaLion Applied Technology lDirectorate (SAVRT-TY-LIRI)

!_SAVRr-.TY-rSc (Tech Library)

A

%~~

US Army Aviation Research and Technology Activity (AVSCOM) 1

Aeroflightdynamica Directorate (SAVRT-AF-D)

US Army Aviation Research and Technology Activity (AVSCOM) I

Propulsion Directorate (SAVRT-PN-D)

Defense Technical information Center (FDAC) 2

US Military Academy, Department of Mechanics 1

(Aero Croup Director)

ASD/AFXT. ASD/EN1F 2

US Army Aviation Development Test Activity (STEBO-CT) 2

Assistant Technical Director for Projects, Code: CT-24

(Mr. Josoph Dunn) 2

6520 Test Croup (ENML) 1

Commande.r, Naval Air Systems Command (AIR 51158, AIR 5301) 3

Defe.nae tntelligence Agency (DtA-DT-2D) I.

115 Army Aviation Sy~temH Command (AMCPM-BiH-T) I

US \ýrmy Aviation SyHtems Commnnd (AMSAV-1:AA) 1

Nnsn L~nnplvy He~ienrch Center (Henry Kelly) 4

N N

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