USAAEFA PROJECT NO. 74-20

HOT BRICK III AIRWORTHINESS EVALUATIONOV-11D AIRPLANE

FINAL REPORT

JAMES S. REIDDONALD F. MACPHERSON JR CW4, AVNPROJECT OFFICER/ENGINEER US ARMY

PROJECT PILOT

JL14 W95

NOVEMBER 1974 JULW)

Approved for public !elease" distrizuimasmimitd

UNITED STATES ARMY AVIATION ENGINEERING FLIGHT ACTIVITYAEDWARDS AIR FORCE BASE, CALIFORNIA 93523

DISCLAIMER NOTICE

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1)ISCI.AIMFR NOTICEI

The4 fieeditigi. of ti~s report are not to be constnad as an official IDepartment ofthe4 %rrnm% pewrntmn unlew. to designated by other audtbonze documents.

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I "troyv 1bi% report when it is no longer needed. Do noet return it to the originator.

TRADE: NAMFS

Thme sa"W Of trade naines in thmis report does not constiluie an official endorsementOr approval of Ilse iase of the commercial hardware and software.

UINCLASSIFIEI)SECt.ATV CLASI.FI, - TNOf T*IS PAGE ~ fc*y,,t

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20. Abstract

on the low-speed high gross weitht regime. Structural tes.ing was limited to fluttertests of the wing store that c-ntained the 150-gallon fuel drop tatik modified withthe HOT BRICK Ill device, the wing at the HOT BRICK III store station, andthe right wing tip. Hlandling qualities tests included a still investigation,determination of control margins with high asymmetric lWads, single-enginemini mum trim and control Arspeeds, and static laterall-diectionall stability. Othertests included takeoff performance and an airspeed lysteF4 calibration. A largediscrepancy existed between the takeoff performance data presented in theoperator's manual and that obtained with the test aircraft. If the data from thisevaluation are representative of the OV-ID, then a defici -y exists, in that thetakeoff performance data presented in the operators. Ina is ex melyoptirnisti(. Four shortcomings were associated with operatin e OVID ai-planeat the heavy gross weight in the all-stores (E) configuration. e cnributinothe HOT BRICK Ill devike to these shorteominp is minimal. T'h hndling qualitiefsof the Y- I D HOT BRICK [il airplane are similar to the statit )% O ID airplanein the alk-!tores (E3) .,uifiguristi'.it. An adequate stall warning~f Ihidh rvided.Further testing should be accomplished to provide accurate take tf performancedata.

UNCLASSIFIEDSECURITY CL ASSIPICAlIOa OF T040S P&GUMM 04WO 5...e.

PREFACE

During the OV I D HOT BRICK III testing the aircraft was maintained by personnelfrom the United States Army Aviation Test Board, Fort Rucker, Alabama.Additionally, the following United States Army Aviation Engineering FlightActivity personnel provided significant 'contributions to the test.

CPT! Robert N. Ward, Aeronautical EngineerI LT Richard D. &ecker, Automatic Data Processing OfficerSP4 Paul R. Bonin, Aeronautical Engineering AssistantKathk;.en M. Dorris, Aeronautical Engineering TechnicianWalter S. Hall, Electronics TechnicianDean S. Smith, Aircraft Mechanic

DEDICATION

Thin report is dedicated to the memory of Maimr Frederick 0. Daniloff and CaptainKenneth F. Schrantz Jr, who were fatally injured on 22 February 1974 duringthe conduct of this evaluation.

-* 1_____________-.

OV-ID Arplane, US Army Scrial Number 69-17000.

TABLE OF CONTENTS

INTROD)UCTION

Ba&kgr;und . . . . . . . . . . . . . . . . . . . . . . . . 2TeF, Objective .. .. ............ ........ 2Description. .. ........... ............ 2Test Scope. .. ............ ........... 4Test Methodology. .. ............. ...... 4

RE'SULTS AND DISCUSSION

General. .... ............... .. .. .. ....Takeoff Performance........ ...... . .. .. .. .. .. .. ...[I indling Qualities. .. ........... .........

Control Margins. .. ...................Static Lateral-Directional Stability. .. ... ........ 9Dual-Engine StallJ .. .. .............. ... 9Singjc-Engine Control Margins. .. ............ 10Single-Engine Minimum Control Airspeed.. ..... .. .. 10

Flight Flutter Tests. .. .............. .... 11

CONCLUSIONS

Cenerl . .... .............. . ..... 13Deficiency and Shortcomings .. .......... ..... 13Specification Compliance .. ............ .... .. 14

RECOMMENDATIONS........ .... . .. .. .. .. .. .. . ..

APPENDI XES

A. References. .. ............... ....... 16B. Description .. .... ................ ... 17C. Instrumentation. . . ...... *...........................190. Test Techniques and Data Analysis Methods. .. ........ 25E. Test Data. .... ................ .... 30

DISTRIBUTION

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J

INTRODUCTION

BACKGROUND

I. The HOT BRICK Ill is an active infrared countermeasure (IRCM) devicedeveloped by Sanders Associates (SA) and is installed on the OV-ID airplane ina modilicd Sargent-Fletcher I50-gallon external fuel tank. As a subcontractor toSA. Gnimman Aerospace Corporation (GAC) modified an OV-l D airplane to acceptthe device and conducted a limited airworthiness evaluation. The United StatesArmy Aviation Systems Command (AVSCOM) requested the United States ArmyAviation Engineering Flight Activity (USAAEFA) to conduct airworthinessverification tests on the OV-ID/HOT BRICK [II system (ref 1, app A). Theoriginal test airplane (SN 69-17018) crashed during conduct of the evaluation byUSAAEFA at Fort Rucker, Alabama, in February 1974. A second airplane(SN 69-17000) was modified and the tests were completed at Edwards Air ForceBase, California, in August 1974.

TEST OBJECTIVE

2. The objective of this evaluation was to identify any airworthiness problemsor flight characteristics changes in the aircraft caused by installation of theHOT BRICK ill system. The test data will serve as a basis for a safety-of-flightrelease for HOT BRICK IlI system testing.

DFSCRIPTION

3. The test airplanes were production OV-ID's (SN's 69-17018 and 69-17000),modified to accept the HOT BRICK il system. A detailed description of theOV21D airplane is contained in the operator's manual (ref 2, app A). Appendix Bgives a detailed description of the test aircraft external equipment.

4. The HOT BRICK IiI is an open loop IRCM device utilizing a mechanicallymounted IR source. The IR transmitter assembly is coupled with a modulatorassembly and is mounted on a modified 150-gallon external fuel tank. The IRsource consists of a ceramic radiating element heated by the combustion of JP-4fuel. ThL fuel for the equipment is drawn from a 15-gallon fuel tank mountedinside the modified 150-gallon fuel tank. The system requires 28 volts direct current(VDC) and is operated from the pilot control box (PCB) located in the cockpit.The HOT BRICK Ill device is further described in appendix B.

5. The OV-ID/HOT BRICK III airplane was tested in two external storesconfigurations which are presented-in table I. Table 2 ,Jefines the various airplaneconfigurations used during the HOT BRICK III tests.

2

Table 1. Exteri-al Stores Test Configurations.

Stores Loading

2

3 130-gallon drop tank

B, wiLh HOT BRICK III Fuselage

4 150-gallon drop tankwit~h HO2T BRICK III

5

6

1 ALQ-67 fuse Jammer

2

3 150-gallon drop tank

E, with HOT BRICK III Fuselage IAPS-94 SLA1O

4 150-gallon drop tank

with W~T BRICK III

5 LS-59A flasher pod

6 A14-80 radar jammer

ISLAR: Side-looking airborne, radar.

3

Table 2. Ai-plane Test Configurations.

Land ing Flap

Conf igurat ion Syrnbol Gear Position PowerjPosition (degi j_________Takeoff TO Down 15 1 Takeoff

Cruise CR I Up Zero For level flight

Power approach PA [ Down 45 For level flight j

TEST ;COPE;

6. '1tie OV-l D/110T BRICK III test program was conducted at Fort Rucker,Alabama. from I I to 22 February 1974 and at Edwards Air Force Base,California, from 17 July to 7 Augu:,* 1974. Nineteen test flights were conducted,with a total of 20 hours. Testing was conducted primarily in the: all-stores (E)configuration at a gross weight of 18,000 pounds, an aft center-of-gravity (cg)location (29 percent mean a~rodynamic chord) (MAC) and at pressure altitudesof 7500 and 14.000 feet. rt~e evaluation was performed within the limitationsof the operato; s n;ianuaI as modified by the safety-of-flight release (ref 3, app A).The results of the test were compared with the inforinL,. n contained in theappropriate sections of the operator's manual. In addition, compliance with theappropriate sections of military specification MIL-F-8785(ASG) (ref 4) wasdetermined.

TESTl METHIODOLOGY

7. Engineering flight test techniques used during this evaluation are d~iscussedbriefly in the Results and Discussion section of this report and in appendix D.Appendix C contains listings of the test instrumentation, the parameters that wererecorded on magneti.: tape, and those displayed on the pilot panel. An airspeedcalibration was accumplished using radar space positioning (figs. I and 2, app C).Dlata analysis mt thods are also presented in appendix D.

RESULTS AND DISCUSSION

(ENE:RAL

8. An evaluation of the OV-ID HOT BRICK IlI airplane was performed todetermine the zarworthiness of tht OV-ID airplane when maodified with theHOT BRICK III device. Structural and handling qualities tests were conducted,with emphasis placed on the low-speed high groms weight regime. Structural testingwas limited to flutter tests of the wing store that contained tale 150-gallon fueldrop tank modified wah the HOT BRICK INl device, the wing at the HOTBRICK !Il store station, and the right wing tip. li-ndlin' qualities tests includeda stall investigation, determination of contro' margins with high asyrnmet~ic loads,singJe-er,. ne min0mim trim and control airspeeds, and stati: lateral-directionalstability. Other tests included takeoff performance and an airspeed systemcalibration. A large discrepancy exists between the takeoff performance datapresente,1 in 'he operator's manual and that obtained with the test aircraft. Ifthe data grum this evaluation are represevtative of the OV-ID, then a deficencyexists, ir that the takceoff performance data presented in the operator's manualis extremely optimistic. Four shortcomingi wer- associated with operating theairplane at heavy gross weights in the all-sto.cs (E) cenfiguration. The contributionof t-e HOT BRICK Ill device to these shortcomings is minimal. The handlingqualities of the OV-I D/IIOT BRICK Ill airplane are similar to the standard OV-IDin the all-stores (E) configuration. An adequate stall warning should be provided.Further testing should be accomplished to provide accurate takeoff performancedi ta.

TAKEOFF PERFORMANCE

() Takeoff performance testing %as not a part of the original test program. Duringiniiial takeoffs, poor performance was encountered with the test aircraft. For thisreason, takeoff pertormance was evaluated for the all-stores (E) configuration withHOT BRICK III and approximately 18,400 pounds gross weight. The distanceswere estimated by aligning the airplane opposite a runway-rei,-ining m.rker andobserving the closest marker at liftoff and when at 50 feet, as indicated by theradar altimeter. These markers were spaced at IO00-foot intervals along the runwayand distanes were estimated to the nearest 500 feet. The pilot technique andprocedtfre us,.d for takeoffs and climbs were those presented in chapters 3 and 14of the operator's manual. A large discrepancy between the takeoff performancedata presented in the opcrator's manual and that obtained during the conduct ofthis evaluation existed. During this evaluation, the test aircraft requiredapproximately twice as much takeoff distance than that presented in the operator'smanual. In addition, rotation to takeoff pitch attitude at the re,,ommended airspeedcould not be achieved. The minimum rctation airspeed was approximately 10 knotscalibrated airspetd (KCAS) grealer than recommended.

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10. The degrade'i takeoff performance of the test aircraft was initially attributedto subst,,4dard engine performance. An analysis of engine performance revealeddiscreprcie; between the torquemeters and engine test stand power available afteroverhaul ,'pp D). From this analysis, it was concluded that the torquemeters wereinaccurate ani the engines were developing spccification power. Other factors whichmay 1, ic.,;ontributed to the degraded takeoff performance are as follows:

Ah3vc-normal roughness of the propeller blades due to high operatingtime 101. hours) and being painted with low reflective lacquer(FSN 8010-083-6588).

t Above-normal roughness of the fuselage caused by application of lowreflectve lacquer (FSN 8010-083-6588).

c. Increased drag caused by wing stores (the contribution of the HOTBRICK III device to this increase is considered minimal).

d. The high gross weight requires higher takeoff airspeed and therefore alonger takeoff distance.

e. The right tire was deformed by high asymmetric weight distribution ofthe wing stores and this deformation increased rolling resistance (photos A and B).

f. Additional control surface and trim deflections required by the highasymmetric weight and drag of the wing stores.

Photo A. Right Main Tire.

C

Photo B. Left Main Tire.

11. The reason for the discrepancy between the takeoff performance of the testaircraft with that presented in the operator's manual could not be determined.If the takeoff performance obtained during this evaluation is representative of theOV-ID airplane, then the presentation of extremely optimistic high gross weightand ambient temperature takeoff performance data in the operator's manual isa deficiency and, if relied upor, could result in takeoff accidents. Further testingis required to verify/provide accurate takeoff performance data for inclusion inthe operator's manual. In addition. if the taktoff performance noted is verifiedthrough additional testing, it is a shortcoming, and takeoff performance shouldbe improved.

HANDLING QUALITIES

Control Margins

12. Lateral control margin tests were conducted in the CR anO. PA configurationsto determine the minimum trim airspeed and lateral control margin with anasymmetric wing loading. The normal loading in the all-stores (E) configurationwith the HOT BRICK Ill device installed results in 620 pounds more weight(140,800 in.-lb total aircraft moment) on the right wing than on the left wing.In the event of a right wing fuel transfer pump failure, a 1520-pound(307,300 in.-lb total aircraft moment) right-wing-heavy condition is possible withthe left drop tank empty (excet for trapped fuel) and the right drop tank full.

7

The variation of' minimum trim airspeed with an asymmetric load is shown infigure A and in figure 1, appendix E. The control margins at various airspeedslor symmetrical and maximum asymmetrical fuel loads are shown in figures 2through S.

FIGIUE AwJINTIMM AIRSPEED WITH ASYETC LOAD

OV-b 0us& 3/N 1700STORE~ COMF 16 UXAM OM4 19 IT1 IHdTICK=

IA ~t 00L EL A

~ SYMPICTRICAL FUEL LAMO

40 to so IM IM No ho

ASYMM4ETIC WAD -1.1(910ff 64MG INeAVY)

13. After determining the minimumn trim airspeed to be 155 KCAS for the CRconfiguration and 136 KCAS for the PA configuration at the 1520-poundasymmetric load condition, airspeed was decreased to a target airspeed of 97 KCASin the PA configuration. Approximately 30 percent of aileron control remainedat this airspeed. A left lateral force of only 5 pounds was required to maintainwings level at 97 KCAS.

14. Landings Were easily accomplished witht a 1200-pound (247,545 in.-lb totalaircraft moment) right-wing-heavy condition using an approach airspeed of1 20 knots indicated airspeed (KIAS) and approximately 100 KIAS touchdownairspeed. The discussion in the operator's manual on operations with highasymmetric wing loadings is satisfactory for the OV-ID/HOT BRICK III airplane.The lateral control margins and lateral trim capability of the OV-ID/HOTBRICK IlI airplane were satisfactory with asymmetric wing loads of up to1520 pounds.

Static Lateral-)irre/ional Stabdity

15. The stati- lat:ral-directional stability of the OV-ID/HOT BRICK III airplanewas evlhfatCd in the *To, CR and PA configurations at airspeeds from 86 to138 KCAS and the conditions listed in paragraph 6. The test results are presentedin figures 6 through 12, appendix E. The static lateral-directional stability wasessentially unchanged from previous results presented in the Army Preliminary17valuations (ref- 5 and 6, app A). Although the pedal position gradient wasapproximately linear, lightening of the pedal forces was apparent at low airspeedsin the PA configuration. This slightly increased the pilot effort required to establishand miaintain a stcady-heading sideslip. Within the scope of this test, the staticlateral-directional stability is satisfactory.

I)ual.Fngine Stalls

l6. Stall characteristics were evaluated in the all-stores (E) configuration with HOTBRICK Ill at 18,000 pounds gross weight in the TO, CR. and PA configurationsat an alt cg. Altitude effects on the stall airspeed (VS) were evaluated by performingthe stall series at two altitudes: 7500 and 14,000 feet pressure altitude. The testtechnique was to trim for level flight at approximately 1.2VS, obtained from theoperator's manual for the test configuration. Then airspeed was slowly decreasedat a rate of I knot per second or less until achieving a stall. Stall was definedby a mild uncontrollable nose-down pitching motion. A comparison of the testdata with the stall airspeeds from the operator's manual is presented in table 3and in figure 13, appendix E. Time histories of the stalls are presented infigures 14 through 16.

Table 3. Dual-Engine Stall Airspeed.

Calibrated Stall AirspeedGross Pressure (kt)

Configuration Weight Altitude -"

(lb) (ft) Test Operator'sData Manual

17,930 7600 77.5 77TO

17,860 14,020 77.5 76.5

18,050 8180 85.5 85CR . . ..

18,110 14,860 87.5 85.5

17,860 7740 71.0 72PA . . . . . . .. -

17,591) 15,000 72.5 71

17. Control effectiveness about all three axes during the approach to the stallwas excellent. The stall was characterized by a mild nose-down pitching with notendency to roll. Stall recovery was easily accomplished by releasing the backpressure on te control stick. The stalls occurred without warning.

18. The lack of stall warning on the OV-ID/HOT BRICK III airplane would behazardous, especially during a short field landing approach and obstruction takeoff,where a stall could result. The lack of stall warning is a shortcoming and failsto meet the requirements of paragraph 3.6.3 of MIL-F-8785(ASG). Stall warningshould be incorporated to provide the crew with an adequate cue to approachingthe stall angle of attack.

Single-Enine Control Margins

1). The single-engine control margins were evaluated in the TO, CR, and PAconfigurations at the conditions listed in paragraph 6. The variation of trim andcontrol position with airspeed is presented in figures 17 through 22, appendix E.Te critical trim control for all test conditions was the rudder trim. The airspeedat which full trim was required in the CR configuration for either propellerfeathered was approximately 145 KCAS. For the TO and PA configurations, thisairspeed was afproximately 140 KCAS for the left propeller feathered and150 KCAS for the right propeller feathered. At 120 KIAS in the TO and PAconfigurations, approximately 30 to 40 pounds pedal force was required witheither propeller feathered and approximately 2 pounds left aileron force wasrequired with the right propeller feathered. Within the scope of this test, theOV-ID/HOT BRICK Ill airplane single-engine control margins are satisfactory.

Single-Engine Minimum Control Airspeed

20. The single-engine minimum control airspeed (VMC) was evaluated in the TO,CR, and PA configurations at the conditions listed in paragraph 6. A comparisonof the VMC from the test data with the data from the operator's manual is presentedin table 4 and in figures 23 through 25, appendix F. Time histories of theapproach to VMC for the three airplane configurations at the two test altitudesof 7500 and 14,000 feet are preseaited in figures 26 through 32.

21. The VMC was defined by stall for all configurations tested. The stalls wererelatitely mild, but without warning. For all configurations, the stall airspeed washigher with the right propeller feathered; therefore, the right engine is the criticalengine in the all-stores (E) configuration with HOT BRICK III. Previous testingwithout HOT BRICK Ill had indicated that the left engine would be critical;however, the increase in asymmetric load and drag caused the change. Adequatecontrol existed about all three axes approaching the stall, except in the CRconfiguration with the right propeller feathered. For this configuration, full leftaileron control was required at the stall. Stall recovery was accomplished byreleasing the control stick back pre:ssure and reducing power on the operatingengine. There was no tendency toward poststall gyrations.

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Table 4. Single-Engine Minimum-Control Airspeed.

CalibratedMinimum-Contro 1

Propeller Cross Pressure AirspeedConfiguration Proeer Weight Altitude (kt)

Feathered (lb) (ft)

Test Operator'sData Manual

17.800 8120 87.5 96Left

17,800 13,580 89.5 98.5

TO17,760 7820 92.5 Not available

Right,.

17,680 14,640 94.5 Not available

22. The lack of cues to the approaching VMC and the stall at VMC withoutwarning is a shortcoming. As discussed in paragraph 18, lick of stall warning wouldbe hazardous during the approach to a landing and obstruction takeoff. Thiscondition is f,rther aggravated by tht higher stall airspeeds for the single-engine:onfiguration Adequate stall warning should be provided.

23. During the evaluation to determine VMC for the PA configuration with theleft propeller feathered, a rudder force reversal was encountered. In thisconfiguration, approximately 110 pounds of right pedal force were required justprior to the single-engine stall airspeed. At this point, the pedal force requiredchanged to a 50-pound left pedal force. This characteristic would increase pilotworkload in an emergency situation. The rudder force reversal in the PAconfiguration with the left propeller feathered is a shortcoming.

24. During the single-engine testing, it was apparent that the OV-ID airplane doesnot have a single-engine capability at 18,000 pounds gross weight for the conditionstested. In the event of an engine failure at the high gross weight, the fuel droptanks (including the HOT BRICK Ill device) may have to be jettisoned. Jettisonof IIOT BRICK III wou!d mean the loss of IRCM protection. Single-engineperformance should be improved.

FLIGHT FLUTTER TF-STS

25. Tests were conduicted at 5000 feet pressure ..ltitudc to determine the fluttercharacteristics of the OV-ID/HOT BRICK IIl airplane in configurations B and E.The method of excitation was a lateral stick pulse (rudder and longitudinal stickpulses did not produce adequate excitation). The test results arm presented infigures 33 through 40, appendix E. In configuration B. the damping ratio was

11

redluced at airspeeds above 260 KIAS. Testing was terminated at 300 KIAS whendamping ratios reduced to 0.04 at two locations (right wing tip forward andHOT BRICK III aft). In configuration E, the damping ratio remained above 0.05at all airspeeds tested (up to 330 KIAS) except for the forward end of the HOTBRICK III tank in the vertical direction. At this location, the damping ratio wasdecreased to 0.04 at 330 KIAS. In both configurations, there were no flutterproblems encountered and the OV-ID/HOT BRICK III exhibited satisfactoryflutter characteristics for normal flight conditions up to the airspeeds tested.

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6I

CONCLUSIONS

(ENERAL

26. The following conclusions were reached upon completion of testing:

a. The OV-I) airplane used during this evaluation exhibited substantiallydegraded takeoff performance as compared to the operator's manual. The reasonfor this discrepancy could not be determined within the scope of this test.

b. The handling qualities and performance of the OV-lD airplane have notbeen significantly changed by installation of the HOT BRICK IIl device.

c. The right engine inoperative is the critical engine in the all-stores (E)configuration.

d. One apparent deficiency associated with the operator's manual was notedand four shortcomings were identified with the airplane in the all-stores (E)configuration.

IFICIENCY AND SHORTCOMINGS

27. The following apparent deficiency associated with the operator's manual wasidentified. If the takeoff performance obtained during this evaluation isrepresentative of the OV-I D airplane, then the takeoff performance chart presentedin chapter 14 is extremely optimistic and, if relied upon, could result in takeoffaccidents (para II).

28. The following shortcomings with the OV-ID/HOT BRICK III airplane in theall-stores (E) configuration were identified:

a. Apparent inadequate takeoff performance at an 18,400-pound grossweight (para 11).

. Lack of stall warning at high gross weights (para 18).

c. Single-engine minimum control airspeed occurs at the stall airspeedwithout adequate cues to the approaching stall (part 22).

d. A rudder force reversal occurs in the PA configuration when approachinga stall with the left propeller feathered (para 23).

13

SPE~CIFICATION COMWPLIANCE

29. Within the %cope of this test, the OV-ID/HOT BRICK III airplane failed tomeet the requirement of paragraph 3.6.3 of MIL-F-8785(ASG), in that theapproach to stall was not accompanied by a stall warning, which should occurbetween 1.05 and 1.15 times the stalling speed in the CR configuration andbetween 1.05 and 1. 10 times the stalling speed in the PA configuration (para 18).

14

RECOMMENDATIONS

30. The apparent deficiency identified duning tis evaluation must be corrected(Para IDl.*

31. The shortcomings should be corrected (paras 11, i8, 22, and 23).

32. Further testing is recommended to provide accurate takeoff performance data(Para 11).

33. Adequate stall warning should be provided (panas 18 and 22).

34. Single-engine performance should be imprrved (pars 24).

APPENDIX A, REFERENCES

1. Letter. AVSCOM, AMSAV-EFTr, 30 October 1973, subject: Test Request forOV- ID/HOT BRICK III Evaluation.

2. Technical Manual, TM 55-15 10-204-10/5, Operator's Manual, OV-ID Aircraft,February 1970.

3. Letter, AVSCOM, AMSAV-EF, 4 February 1974, subject: Safety-of-FlightRelease for OV- ID/HOT BRICK Ill.

4. Military Specification, MIL-F-8785(ASG), Flying Qualities of NoktedAirplanes, Amendment 2, 17 October 1955.

5. Final Report, USAASTA, Project No. 68-43, Army Preliminary Emiuwtion,Production 0 V-ID (Mohawk), March 1970.

6. Final Report, USAASTA, Project No. 70-03, Army Preliminary Eialuation 1,Production 0 V-iD (Mohawk). Performance and Handling Qualities, March 1971.

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APPENDIX S. DESCRIPTION

I. The test aircraft were production OV-ID airplanes, serial numbers 69-17018and 69)-17000, modified to accept the HOT BRICK III stores and controlsdescribed below. The 1OT BRICK III system is an open loop IRCM set utilizinga mechaniclly mounted IR source (fig. I). The IR transmitter assembly is coupledwith a modulator assembly and is mounted on a modified 150-gallon fuel tank(fig. 2). The IR source consists basically of a ceramic radiating element heatedby the combustion of JP-4 fuel with ambient air. Both combustion rnd coolingair are drawn from a common inlet mounted on thc pod shell. A flow controlvalve maintains an approximately constant mass flow through the combustor,regardle-.s of flight airspeed or altitude. A fuel group pumps and regulates the fuelsupply. fhc fuel for the equipment is drawn from a 15-gallon fuel tank mountedinside the modified 150-gallon fuel tank. The small internal tank is filled fromthe larger tank as long as there are more than 100 gallons of fuel in the largetank. When the fuel level of the lirge tank drops below 100 gallons, there is stillsufficicnt fLe! for the system in the small tank. With the modified 150-gallon fueltank the maximum fuel available to the engines is 135 gallons irom that tank.

2. The physical characteristics of the OV-lD/HOT BRICK IIl system are asfollows:

Device Physical Dimensions

Basic diameter (no scoops) 17.2 in.Overall length (no scoops) 46 in.

Device Basic Weithts

Device 230 lbDevice fuel tank and fuel 128 lb

Modified Extemal Stores

Overall length 178.25 in.Weight of modified stores 140 lbWeight of ballast 164 lbWeight empty 534 lbWeight of fuel (JP-4) 971 lbWeight loaded 1505 lbWeight increase loaded +382 lb

(HOT BRICK store vs standard150-gallon drop tank)

17

Pilli 911,1- 110 inPflJNI

GAssMMeW -

W011"Moo4U1

Ut, WUWUS

srt.

Iiur . OV-ID HOT BRICK System.

I COWPIT

II

"OT NICK IIVOV-1

Mpgre 2. Aircraft Installation.

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APPENDIX C. INSTRUMENTATION

(;FNF.RAIL

1. InstrLinentation for the OV-'j)/HO1 BRICK II! airplane was installed.cailibratcd, arid maint-'ined by po!rsonnel of the 'ILst and Evaluation Commind(TLICOM) at Foit Rucket, Alabama, and by USAAEFA at Edwards Air ForceBast, C~aliforia.

TFSTIN; AT FORr RUCKER

2. During icsting accompiishcd at Fort Ru'-ker, the instrumentation listed belowwas installed. The instrumentation package used an' 05lograph recorder.Supplemental data were obtained from standard cockpit Oicators and voicerecording. In addition, photo roverape from the chase aircraft was provided.

3. The existing ship's systemn instruments were used duriaig this test to recordengine and flight data. In addition, a panel-mounted maneuvering accelerometerwas installed. These instruments were calibrated prior to the test by TECOM. Acassette voice recorder was used to record pilot qualitative omments.

Instrumentation Paokacre

4. Paramnetcrs recorded were coordinated with TECOM to minimizeinstrumen tation changes after the USAAEFA tests. The followiag parameters wererequired for the USAAEFA tests:

HOT BRICK III Device:

High-frequency modulationLow-frequency modulationCombustion indicationCombu~tio 1 temperatureRan indicationFault indicationIgnition indication

Aircraft Flutter Tebts:

HOT BRICK store normal acceleration forwardHOT BRICK store normal acceleration aftHOT BRICK store lateral acceleration forwardLeft wing tip normal acceleration forwardRight wing tip normal acceleration forwardRight wing tip normal acceleration aftCenter-of-gravity normal accelerationCorrelation counter

TESTING AT EDWARDS AIR FORCE RASE

5. During testing at Edwards Air Force Base, the following instrumentation wasinstalled. A magnetic. tape recorder wa installed in the aircraft. A boom wasmounted on the SLAR antenna exten(..,g aproximately 5 feet forward from thenose of the SLAR (photo I). A g c-ot-sideslip and angle-of-attack vanes and ahigh-speed pitot-static tube were mounted on the boom. The parameters recordedand/or displayed together with the location are listed below.

Photo 1. SLAR Mounted Airspeed Boom.

20

Pilot Panel

Airspeed (boom)Altitude (boom)Angle of sideslipAngle of' attack('Center-of-gravity normal acceleration[levator trim positionAileron trim positionRudder trim positionLeft engine output shaft torqueRight engine output shaft torque

,Magnetic Tape

Airspeed (boo )Atitude (boom )Free air temperatureControl positions:

Longitudinal stickLateral stickPedal

Control forces:Longitudinal stickLat-ral stickPedal

Control surface positions:LIcvatorLeft outboard aileron('enter rudder

Aircraft attitude;PitchRoll

Aircraft angular velocity:PitchRollYaw

Angle of attackAngle of sideslipAcceleration:

Center-of. gravity normalCenter-of-gravity lateral

Engine gas producer speed (left and right)IEngine power turbine speed (left and right)Engine exhaust gas temperature (left and right)TunePilot event

21

6. Calibration of tihe boom-mounted pitot-static system was accomplished by useof the National Aeronautics and Space Administration's radar space positioningequipment. The airspeed system position error is presented in figures 1 and 2.

22

F spso aLI.TII

OV-10 Uso A /Jfi-17o0O$Tout Cc*&PIGUSIPr 1014 &. billIe KasIcx

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Sy~ISL SAK W1~IT G.FAV Pfts"UMi AL WTIUD 6 ~mufacMaome

G, 1* w 2 .4, 10.4s 1 0 5 0 0 T A K POWFF

0,

cm C

06 /1"11SRIV& I/,tj%'' ~j~tb (a

APPENDIX D. TEST TECHNIQUES ANDDATA ANALYSIS METHODS

TEST TF(IN IQU.S

Takeoff Performance

I. Takeoff p.:rformance tests were conducted from a concrete runway. Distanceswere estimated by aligning the airplane opposite a runway-remaining marker andobscrving the closest marker at liftoff or at 50 feet on the airplane radar altimeter.The e markers were spaced at 1000-foot intervals along the runway and distanceswere cstinated to the nearest 500 feet. During the takeoff roll, altitude andambient temperature were recorded from the aircraft's stand-rd service indicators.The pilot technique and procedure used for takeoff and climb-out were thosepresented in chapters 3 and 14 of the operator's manual.

C ntrol Margino

2. Control margins were evaluated with asymmetric loads of full fuel in bothdrop tanks: the left drop tank half full, right drop tank full; left drop tankonc-quarter full, right drop tank full; and left drop tank empty (except for trappedfuel), right drop lank full. In the CR configuration, airspeeds of approximately115 to 185 K(AS were evaluated. In the PA configuration, airspeeds ofapproximately 95 to 150 KCAS were evaluated.

3. The aircraft was trimmed at the desired conditions in level unaccelerated flightor maximum-power descending flight, starting at the maximum airspeed for eachtest and decreasing in approximately 10-knot increments. The airspd at whichfull aileron trim was required was noted.

Static Lateral-Directional Stability

4. Static lateral-directional stability tests were conducted by trimming the aircraftat the desired airspeed in zero sideslip. Power, airspeed, trim settings, and aircraftground track were held fixed. Sideslips were increased incrementally, both left andright, up to the flight envelope limits.

Stalls

5. Stall characteristics tests were conducted by trimming the aircraft atapproximately 1.2Vs (determined from the operator's manual) in level flight forthe desired condition. The stall was approached at an airspeed reduction of lessthan I knot per second. After the stall occurred, the back pressure on the stickwas reduced and for dual engines the aircraft was allowed to fly out of the stall.For single-engine stalls, the power on the operating engine was also reduced,recovery made, and power on both engines increased.

Single-Fngine Comtrol Margin.

6. The single-engine control margin tests were conducted with either engine atidle and the propeller feathered. The aircraft was stabilized at incremental airspeedsbetween approximately 150 and 120 KCAS. At each stabilized airsri', all control

25

forces were trimmed to zero, or maximum trim used, while maintainingsteady-heading flight. The airspeed at which full trim was required was noted. Thesetests were accomplished for wings-level and for 5 degrees of bank toward theoperating engine.

Sinile-Enine Minimum Control Airspeed

7. The single-engine minimum cortrol airspeed tests were conducted by stabilizingthe airplane at the desired conditions using military power and then reducing poweron the desired engine to idle and feathering the propeller. The airspeed wasdecreased at a maximum rate of I knot per second while maintaining constartheading and wings-level until a single-engine stall occurred. The critical engine wasdetermined by conducting tests with either the left or right propeller feathered.

Flight Flutter

8. Flight flutter tests were conducted by stabilizing the airplane at the desiredc:onditions and attempting to excite the flutter mode, usin6 a lateral stick pulse.Data were recorded for several seconds to enable an analysis to be made.

DATA ANALYSIS METHODS

Takeoff Performance

9. The estimated takeoff performance was compared with data obtained fromfigure 14-I I of the operator's manual at the altitude, ambient temperature, grossweight, and height above the runway encountered during each test. The powerajailable was assumed to be equal to or in excess of the minimum power availableas contained in the engine model specification. This assumption was substantiatedby the method explained in the Power Available section.

Power Available

10. Power available from the engines installed in the test aircraft was evaluatedto determine if the aircraft was a suitable sample for this test. Shaft horsepowerwas obtained from the aircraft torquemeters at a variety of test conditions usingthe torquemeter calibration from the engine test run after overhaul.

I I. The gas producer speed and the shp were referred to sea-level, standard-daystatic conditions using the following expressions:

1SHiP SHP - ASHP- and

) t

t t t t tWhere:

NI Gas producer speed

26

/*

SHP SHP corrected for ram

ASLIP = Ram correction

= SHP available at test true airspeed

- SHP available at zero airspeed (based on nower availableversus true airspeed obtained from the OV-ID APE 1Ireport (USAASTA Project No. 70-03, April 1972)

St = Presure at test altitude/sea-level pressure

0t = Absolute temperature at test altitude/absolute sea-level,standard-day temperature

The referred values foi cach engine were plotted and a curve faired through therespective data. These curves were compared with curves obtained from the enginetest stand run after overhaul for the respective engine. This comparison for theNo. 2 engine is presented in figure 1.

I0GLJSCb I

jiEQIWIL CI4ARAC'TfISTICSoV- I D USA % (641 -rTooOTS- "O A /M3OO156

9i .ILN- -

: /

IN

I 4o . .. STTTT AM

to

4o so #0 V'O tZO

RIF&IBMO GAS PftOCfER ?,PSO-MI/4i -SP4

27

12. This figure shows that the engine was apparently developing 300 to 400 sipless than during the test stand run. To verify this, a lift-drag polar was calculatedfrom the data obtained during the airspeed position error calibration in the cruiseconfiguration, using the following equations:

= L (2)

/2 p VT 2 S

and

CD =- D/ p T2 S (3)CD 1/2 p VT2

Where:

CL = Lift coefficient

CD = Drag coefficient

L = Lift (assumed equal to grow weight (Ib))

D - Drag (lb)

p = Air density (lb-Wc 2 /ft4 )

VT = 1lrue airspeed (ft/sec)

S Wing area (ft 2 )

THP 550 (4)VTWhere:

T Thrust (lb)

THP = Thrust horsepower

THP - ? SHPt (5)

Where:

17 - Propeller efficiency

The propeller efficiency was obtained from the propeller efficiency chart for the53C51/7125-6 propeller including the blocking effect of a T-53 engine but notincluding compressibility corrections for high tip ,peed conditions.

28

13. A comparison of the calculated lift-drag polar based on er~gine torquemneterreadings was made. with a lift-drag polar contained in the APE 11 performancereport for the same conditions, except without the HOT BRICK III device. Thiscomparison is presented in figure 2.

FIGURE Mo. Z

OOAL ENGI LmVEL PLIGIA1 PERrORMI.C.OV-ID USA%0d-1OOO0

STORt COWFIGUR A11014 E WITIA WO75gtCKIMCRUISE CONVIGURAMIOJ

o SH4P BASED ON M., SPEEDC 51P B&ASE Ou TVORQDEMETER

AP G. [ OJECT TO-03,

U

7 0 0w

0' 0.102a

9..9

JU

14. Figure 2 shows that a considerable increase in performance (ie, less powerrequired for a given gross weight and airspeed) is indicated when basing poweron the engine torquemetcr. Also shown on figure 2 is a lift-drag polar calculatedfrom the data obtained during the airspeed position error calibration in the cruiseconfiguration, using power based on the engine test stand run after overhaul andgas producer speed. This power was obtained in the following manner. The gasproducer speed recorded in flight was referred to sea-level, standard-day conditions.This referred gas producer speed was used to enter a plot of referred gas producerspeed versus referred shp derived from the engine test stand run after overhaulto obtain referred shp. Shaft horsepower at the test conditions was determinedfrom referred shp using the following equation:

S (S__.) (6 t t) + SHP (6)

This shp was t'ln used to calculate the drag coefficient using equations 3through 5 , ,;fore.

15. The ,omparison of the calculated lift-drag polar based on gas producer speedand ,he engine test stand run with the APE II performance data showed a slighti,,.rease in power required for the test aircraft. When considering the increaseddrag of the HOT BRICK IlI device and the decrease in power available with theengines installed in the aircraft, the comparison seemed reasonable. It was thereforeassumed that the engines in the test aircraft were developing the appropriate powerand should be representative.

Flight Flutter

16. The oscillations at each flutter instrument location, resulting from the lateralstick pulse, were reduced to damping ratio using the ratio-of-maximums method.The first 0.4 second following the stick pulse apparently contained effects ofaileron movement and was not used in determining damping ratio.

30

i

I I - -

APPENDIX I. TEST DATA

INDEX

Pure Fgre Number

Minimum Trim Airspeed IControl Margins 2 through 5Static Lateral-Dircctional Stability 6 through 12Dual-Engine Stalls 13 through 16Single-Engine Control Margins 17 through 22Single-Engine Minimum Control Airspeed 23 through 25Single-Engine Stalls 26 through 32Flight Flutter 33 through 40

31

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DISTRIBUTION

Director of Defens. Research and Engineering 2

Assistant Secretary )f the Army (R&D) I

Chief of Research and Development, DA (DAMA-WSA) 3US Arm' Materiel Command (AMCPM-UA, AMCPM-,SE-TM,

AMCRD-FQ. AN:CSF-A, AMCQA) 14

US Army Aviation Systems Command (AMSAV-EQ) 12

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ATCD-CM) 22

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US Army Forces Command (AFOP-AV) IUS Army Armament Command (SARRI-LW) 2US Army Missile Command I

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Ames Directorutc, US Army Air Mobility R&D Laboratory (SAVDL-AM) 21istis Direclorate. US Army Air Mobility R&D Laboratory (SAVDL-EU-SY) 2

Langley Directorate, US Army Air Mobility R&D Laboratory (SAVDL-LA) 2Lewis Directorate, US Army Air M'obility R&D Laboratory (SAVDL-LE-DD) I

US Army Acronwdical Research Laboratory IUS Army Aviation Center (ATZQ-DI-AQ) I

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1US Maric (orp% Ivrlopment and F.ducation Command 2IW; Nival Atr Vcn'kt ( iI

US Air Force Aeronautical Systems Division (ASD-ENFDP) IUS Air Force Flight Dynamics Laboratory (TST/Libm~ry)IUS Air Force Flight Test Center (SSD!/Techiiicat Library, DOEE) 3US Air Force Special Communications Center (SUR) IDepartment of Transportation Library 2US Army Grumman Plant Activity 2Beech Aircraft Corporation 5Grumman Aerospace Corporation 5AVCO Lycoming Division 5Sanders Associates 5Sargent-Fletcher Company SUnited Aircraft of Canada Ltd SDefense Documentation Center 12

4.S .WWOW**I^-Awi,