evaluation of an oh-4oa fielicoptm with an allison … · evaluation of the performance, handling...
TRANSCRIPT
7*'* ( ~USAAEFA PROJECT NO. 74-48
EVALUATION OF AN OH-4OA FIELICOPTM
WITH AN ALLISON 25O-C2O03 EV1GBINE
FINAL REPORT
CARL F. MITTAGTOM P. BENSON MMJ, AD
PROJECT OFFICER US ARMYPROJECT PILOT
JAMES E. JENKS JRROBERT M. BUCKANIN CPT, FA
*PROJECT ENGINEER US ARMYPROJECT PILOT
Reproduced From APRIL 1975Best Available Copy 1S17
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Approved for public release; distribution unlimiAniui~
UNITED STATES ARMY AVIATION ENGINEERING FLIGHT AC11VITYEDWARDS AIR FORCE BASE, CALIFORNIA 93523
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lThe findings of this report are not to be constnied au an official Department ofthe Army position unless so diesignated by other authorized doctiments.
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Destroy this report when it is no longer needed. Do not return it to the originator.
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The one of trade names in this report does not constitute an official endorsementor approval of the use of the commercial hardware and software.
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USAAEFA7" = 7ý- 744, ____
mid OFLP~,~r~&PjIOOCOVEREDY.VALUATION OFAN OH¶S81, IELICOPTER) -FINAL REPQ3T. -*-
WITH AN ALLISON 2q 01-~ ENGIE, IT.- -w Decvv*nZ74,le ~ ~ ~ ~ N 1 6t RbMN ple/P~yýWG
S. CONTRACT OR GRANT NUMBER(&)
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US ARMY AVIATION ENGINEERING FLIGHT ACTIVITY AE OKUI UBR
EDWARDS AIR FORCE BASE, CALIFORNIA 93523
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US ARMY AVIATION ENGINEERING FLIGHT ACTIVIT ':, APRE,275EDWARDS AIR FORCE BASE, CALIFORNIA 93523 .NUMBER OF PAGES
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1S. SUPPLEMENTARY NOTES
It. KEY WORDS (Contifnue on reverse side If necessayeati edIdentify by black 8umber)
Limited performance and handling Pilot workloadqualities evaluation Allison 250-C2013 and standard T63-A-700
OH-58A helicopter with an Allison engine comparison250-C201B engine Engine/airframe compatibility characteristics
(cc-31ing and vibration levels)20. ABSTIMCT (Conilnue ow reverse side it necessary mid identilfy by block number)
The United States Army Aviation Engineering Flight Activity conducted a limitedperformance and handling qualities evaluation of a Bell Helicopter CompanyOH-58A helicopter with an Allison 250-C2OB engine installed. The evaluation wasconducted at Edwards Air Force Base and Bishop, California, from 17 Octoberthrough 6 December 1974. Twenty-two flights with 17.6 productive test hourswere required for the evaluation. Test results obtained with the Allison 250-C2OB
(continued)
DD I JAN 73 1473 zono ED~oNr I NOV 61 IS OBSOLETE UCASFESECURITY CLASSIFICATION OF THIS PAGE (V1'en Dae Eteed
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UNCLASSIFIEDSECURITY CLASSIFICATION OF THIS PAGEr"1Mh DaEte. i
20. Abstract
engine were compared with those previously obtained with the Allison 250-C20engine and the standard T63-A-700 engine. Primary performance improvement overthe standard T63-A-700 engine was an in=crase in out-of-ground-effect hover ceilingfro-n 4600 to 11,050 feet standard-day density altitude at a gross weight of3000.. pojunds. One _derficienarv ea.shJ-comcs weJre noted The deficiency
Swas the considerable pilot compensation required to maintain .arcraft control inleft lateral accelerating flight. The shortcomings consisted of (1) insufficient leftpedal in right sideward flight, (2) insufficient aft cyclic control in rearward rfght,(3) extensive pilot compensation required in left sideward flight, (4) moderate pilotcoripensation required to maintain right accelerating flight, and (5) the
I unntisfactQ-r ggvzosiizg_-caracteristics of the OH-58A helicopter equipped withan Allison 250-C20B engine. Y'h-e'jnsatisfactory handling qualities characteristicsare inherent to the basic OH-58AKheiicopter and are not associated with theinstallation of the 250-C20B engine. The engine/airframe compatibilitycharacteristics (cooling and vibration levels) of the OH-58A helicopter with the250-C20B engine are similar to the standard OH-58A helicopter with the T63-A-700engine. Within the scope of the test, the performance of the OH-58A helicopterwith an Allison 250-C20B engine installed was improved over the basic OH-58Ahelicopter. Handling qualities were essentirlly unchanged.
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SECURITY CLASSIFICATION OF THIS PAGOIr'Wý,n DeMO En*,eed)
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SPREFACE
Throughout this evaluation, technical support was provided by the enginemanufacturer, Detroit Diesel Allison Division of General Motors Corporation,Indianapolis, Indiana. Emergency fire fighting and medical support were providedby the United States Air Force Flight Test Center, Edwards Air Force Base,
California.
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TABLE OF CONTENTS
INTRODUCTION
Background . . . . . . . . . . . . . . . . . . . . . . . . 5Test Objectives .. .... .............. ............ ...... 5
TatMethodology .. .. .............. ............. .... 6
RESULTS AND DISCUSSION
General.................. . . ....... . .. .. .. .. .. ... 8Performance .. ........ .............. ................. 8
General......... ..... . . ........ .. .. .. .. .. ....Hover Performance .. .. ........ .............. ...... 10Vertical Climb Performance. .. ........ ............... 10Lateral Acceleration Performance. .. .... ............... 10
Handling Qualities .. .. ............ .............. ...... 11VGeneral .. ...... .............. .............. .... 11
Low-Speed Flight Characteristics .. .. ............ ...... 11Lateral Acceleration Handling Qualities. .. .......... .... 12
Subsystem Tests. .. .. .............. .............. .... 13Engine Characteristics .. ........ .............. ..... 13Engine Vibration. .. .............. ................. 13Engine Compartment Temperature Survey. .. ........ .... 14
Engine Governing Characteristics .. .. ............ ...... 14
CONCLUSIONS
General .. ....................... 16Deficiency and Shortcomings .. ........ ............ ...... 16Specification Compliance. .. ............ ............ .... 17
RECOMMENDATIONS
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APPENDIXEs
A. ReferencesB. General Engine Information ................. .21
C. Instrumentation................................... 25D. Test Techniques and Data Analysis Methods .............. 27SE. Test Data
. T .................................... 30
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INTRODUCTION
BACKGROUND
1. Engineering flight tests werc previously conducted by the United States ArmyAviation Systems Test Activity (USAA.STA) (since redesignated the United StatesArmy Aviation Engineering Flight Activity (USAAEFA)) to determine theperformance and handling qualities characteristics of a Bell Helicopter Company(BHC) OH-58A helicopter equipped with Allison T63-A- '00 (refs 1 and 2, app A)and 2W0.20 (ref *3) gas turt , ie engines manufactured by the De-troit Diesel AllisonDivision of General Motors Corporation. It was subsequently determined that anevaluation of the performance, handling qualities, engine/airframe interface, andengine cooling characteristics of an OH-58A helicopter with an Allison 250-C2O1)engine installed was necessary. In June 1974, the United Staz-s Army AviationSystems Command (AVSCOM) directed USAAEFA to conduct that evaluation(ref 4).
TEST OBJECTIVES
2. Tae objectives of this evaluation were to determine !he aircraft performanceand handling qualities, and engine vibration and temperature characteristics of anOH-58A helicopter with an Allison 250-C2OB engine installed. Specific objectives
were as follows:
a. Evaluate compatibility of the engine/airframe, to include engine vibrationcharacteristics.
b. Determine the engine nacelle and lubrication system coolingcharacteristic&.
C. Determbie the performnance characteristics and handling qualities at lowand high-altitude test sites.
DESCRIPTON
3. The Allison 250-C2OB engine is a growth version of the 250-C20 engineincorporating a redesigned combustor and turbine. For this evaluation, the testengine was equipped with a Bendix fuel control system and was restricted to thehelicopter's transmission takeoff power limit of 317 shaft horsepower (shp) fromits uninstalled static sea-level takeoff power rating of 420 shp. A detaileddescription of the Allison 250.C2OB engine is contained in the installation design
manual (ref 5, app A). A general description of the engine is contained in
appendix.....................
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4. The test aircraft, a JOH-58A light observation helicopter, serialnumber 68-16706, was manufactured by BHC, Fort Worth, Texas. The single mainrotor is a two-bladed, semirigid, teetering type and the tail rotor is also of thetwo-bladed, semirigid, teetering type with delta-hinge coupling. The cockpit providesside by side seating for a crew of two (pilot and copilot/observer) and the cargocompartment has seats for two passengers. The cyclic and collective controls arehydraulically boosted and irreversible, while the directional controls on the standardconfiguration are unboosted and reversible. The modifications intcorporated in thetest helicopter which resulted in the "J" designation were installation of a BHCelectronic 3-axis stability and control augmentation system (SCAS) whichincorporates a hydraulically boosted and irreversible diirectional control and a Sperryhelicopter command information system (HCIS). A detailed description of bothsystems is contained in USAASTA Final Report No. 72-20 (ref 6, app A). Thelanding gear consists of fixed skids. The helicopter is normally powered by anAllison T63-A-700 (Allison 250-C18) free gas turbine engire with an uninstalledtakeoff power rating of 317 shp at static sea-level, standard-day conditions. Themain transmission has a rating of 270 shp for continuous operation, with a takeoffpower limit of 317 shp (5-minute rating). A detailed description of the standardOH-58A helicopter is contained in the operator's manual (ref 7).
TEST SCOPE
5. The performance, handling qualities, and engine/airframe in~efface evaluationsof the Allison 250-C20B engine installed in a JOH-58A helicopter were conductedby USAAEFA personnel. Testing was performed at Edwards Air Force Base(elevation 2302 feet), Bishop (elevation 4120 feet) and Coyote Flats (elevation9980 feet), California, from 17 October through 6 December 1974. Twenty-twotest flights for a total of 17.6 proctuctive hours were flown. Flight limitationscontained in the operator's manual and the safety-of-flight release (ref 8, app A)were observed during the testing. Test conditions are shown in table 1. The testaircraft was evaluated against the requirements of military specificationsMIL-H-8501A (ref 9) and MIL-T-25920 (USAF) (ref 10).
TEST METI'HODOLOGY
6. Established flight test techniques were used for the handling qualities andperformance testing (refs 11 and 12, app A). Test methods are described brieflyin the Results and Discussion section of this report. All tests were conducted under• '• nonturbulent atmospheric conditions M- preclude uncontrolled disturbances from
influencing the test data. A detailed de' cription of test instiumentation is containedin appendix C and a description of data reduction procedures in appendix D. Pilotcomments were used to aid in the analysis of data and to determine the overallqualitative assessment of the flying qualities of the JOH-58A helicopter with anAllison 250-C20B engine installed. The Handling Qualities Rating Scale (HQRS)used to augment pilot qualitative comments is included as figure 1 in appendix D.
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RESULTS AND DISCUSSION
GENERAL
7. A limited performance and handling qualities evaluation of the OH-58Ahelicopter with an Allison 250G20B engine installed was conducted by USAAEFA.Performance, handling qualities, engine/airframe characteristics, and miscellaneousengineering tests were evaluated under a limited variety of operating conditions,with emphasis oa operations near the military maximum gross weight of3200 pounds. Primary performance improvement over the standard T63-A-700engine was an increase in OGE hover ceiling from 4600 to I 1,050 feet standard-daydensity altitude at a gross weight of 3000 pounds. Handling qualities were evaluatedduring low-speed and lateral accelerating flight. The OH-58A helicopter with theT63-A-700 engine is normally power limited, which prevents the accompl;shmentof certain maneuvers at heavy gross weights with hot-day conditions. Theinstallation of the Allison 2504C20B engine increased the helicopter's capability,but full utilization of its maxunum gross weight capability is still not possible,due to the inadequate handling qualities. One deficiency and five shortcomingswere found. The deficiency was the considerable pilot compensation required tomaintain aircraft control with SCAS OFF in left lateral accelerating flight. A SCASis not installed in the standard OH-58A helicopter. Handling qualities shortcomingsconcisted of (1) insufficient left directional control in right sidexard flight.(2) insufficient aft cyclic control in rearward flight, (3) extensive pilotcompensation required in left sideward flight, (4) moderate pilot compensationrequired to maintain right lateral accelerating flight, and (5) the unsatisfactorygoverning characteristics of the OH-58A helicopter equipped with an Allison250-C20B engine. These unsatisfactory handling qualities characteristics are inherentto the basic OH-58A helicopter and are not associated with the installation ofthe 250-C20B engine. The engine/airframe compatibility characteristics (cc lling andvibration levels) of the OH-58A helicopter with the 250-C20B engine are similarto the standard OH-58A helicopter with the T63-A-700 engine.
PEFIZORMANCE
General
8. Performance testing of the OH-58A helicopter with an Allison 250-C20Bengine was ccnducted at a hover, in verticd climbs, and in lateral acceleratingflight. Hover test results were compared to results obtained with the previouslytested T63-A-700 ar.d 250-C20 engines. The OGE hover ceiling was increased fromA600 feet with the T63-A-700 engine to 11,050 feet density altitude at a grossweight of 3000 pounds with the 250-C20B engine. Within the scope of the test.the hover performance characteristics of the OH-58A helicopter with an Allison250-C20B engine are greatly enhanced over the basic OH-58A helicopter.
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"Hover Performance
9. Hover performance tests were conducted at skid heights of 4 and 10 feetIGE and 50 feet OGE under the conditions shown in table I. The free-flight hivermdtb.od was used to determine hover performance. Skid height was measured byvua:, reference to a measured weighted cord attached to the left skid. Incrementalamounts of ballast were added to the helicopxer until the gross weight was suchthat either the engine temperature or transmission power limit was reached. Testresults are presented, together with T63-A-700 and 250-C20 hover data, in figures 1through 3, appendix E. A comparison of the standard-day and 35*C hot-day OGEhover performance with the 250-C20B and T63-A-700 engines is presented infigure A. The OGE hover ceiling was increased from 4600 feet for the T63-A-700engine and 10,000 feet for the 250-C20 engine to 11,050 feet for the 250-C20Bengine. The increased hover performance of the 01-58A helicopter with the250-C20B engine enhances its operational capability. Within the scope of the test,the hover performance of the OH-58A helicopter with the 230-C20B engine isimproved over the basic OHl-58A helicopter.
Vertical Climb Performance
10. Vertical climbs (zero horizontal airspeed) of approximately 20 secondsduration were made from an OGE hover at various constant collective controlsettings at the conditions shown in table I. A radar alt;meter was used to measurethe rates of climb and an Elliott low-airspeed system was used to aid the pilotin maintaining zero horizontal airspeed while in a vertical climb. The vertical rateof climb for a given power increment was defined as that portion of the climbafter the aircraft has achieved a steady unaccelerated rate-of-climb condition.Vertical climb testing had not been previously performed on the OH-58A helicopter,which precluded any comparison. A detailed description of the test techniquesand data analysis methods used is contained in appendix D. To ensure hoverdata validity, the test hover power required was compared with the OGE hoverperformance data for the OH-58A helicopter and the results a;- presented infigure 3, appendix E. Test results are presented in figures 4 through 9. At21200 pounds gross weight and standard-day, sea-level conditions, the maximumvertical rate of climb was calculated to be 660 feet per minute. The rates of climbthat were achieved ranged from approximately 50 to 1200 feet per minute.
Lateral Acceleration Performance
I1. The lateral acceleration performance of the OH-58A helicopter was evaluatedby conducting lateral accelerations OGE (skid height approximately 50 feet) atthe conditions shown in table 1. The lateral acceleration was accomplished byincrementally rolling the aircraft to a predetermined roll attitude (5-degreeincrements) with a rapid lateral control motion while maintaining constant skidheight with power and maintaining a constant heading. A ground pace vehicle wasuwd to determine limit sideward airspeed. Surface winds were less than 3 knots.Lateral flight performance data are presented in figure 10, appendix E.
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12. The maximum roll attitude tested in right sideward flight was 18.4 degrees,at which point the never-exceed transmission torque pressure limit of 106 psi wasinadvertently exceeded, which terminated further testing. The operationaltransmission torque pressure limit (92 psi) was reached repeatedly at roll attitudesgreater than 10 degrees. In left sideward flight a maximum roll attitud,. of16.7 degrees was achieved before termination of this test. Maximum accelerationachieved in left and right sideward flight was i 1.2 and 13.9 feet per second persecond (ft/sec 2), with ,, time to maximum acceleration of 7.5 and 5.3 seconds,respectively. Representative time histories for left and right accelerating flight areshown in figures I I and 12, appendix E. There was no cue to alert the pilotof reaching limit sideward velocity (35 KTAS) other than the pilot's judgmentof ground speed. Because of the lack of ground speed cues and the high probabilityof exceeding the transmission torque limits in right lateral accelerating flight, thismaneuver should be limited to roll attitudes of less than 9 degrees during furthertesting of the lateral flight maneuver. Maximum lateral acceleration capability ofthe aircraft could not be used because the transmission torque limit would havebeen exceeded.
HANDLING QUALnITES
General
13. A limited handling qualities evaluation of the OH-58A helicopter with anAllison 250-C20B engine was conducted under day visual flight conditions. Lateralacceleration and low-speed flight characteristics were evaluated under the conditionsshown in table I and the results were compared to those obtained during theT63-A-700 and 250-C20 engine evaluation (refs 2 and 3, app A). One deficiencyand five handling qualities shortcomings were noted. The deficiency determinedwas the pilot compensation required to maintain aircraft control in left lateralaccelerating flight with SCAS OFF. The handling qualities of the OH-58A helicopterequipped with an Allison 250-C20B engine are similar to a standard OH-58Ahelicopter, except that the standard OH-58A helicopter does not have a SCAS.
Low.Speed Flight Characteristics
14. Low-speed flight tests were conducted to determine the hovering capability( •(IGE) of the OH-58A helicopter in winds of various speeds and azimuths. Wind
azimuths of zero, 90, 180, and 270 degrees relative to the nose of the helicopterwere used. Test results are presented in figures 13 and 14, appendix E.
15. Previous right sideward flight tests at a thrust coefficient (CT) of 0.00343resulted in a left pedal margin of 5 percent at 35 KTAS (ref 3, app A). Resultsof the test with the 250-C20 engine (ref 3) showed that at 35 KTAS. thedirectional control margin rapidly diminished with increased CT's. With the250-C20B engine at a CT of 0.00416, the maximum airspeed in right sidewardflight was 32 KTAS with no directional control margin. Directional control in rightsideward flight failed to meet the requirements of paragraph 3.3.2 of MIL-H-8501A.
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in that the l0-percent directional control power was no-c present at 35 KTAS.Within the scope of this test, the sideward flight characteiistics are aggravated bythe installation of the 250-C20B engine because of the increased gross weight andaltitud& capability. Minimum required lateral low-speed flight capability could notbe achieved because of tail rotor performance and aircraft instability. Insufficientleft directional control in right sideward flight is a shortcoming. Considerationshould be given to upgrading tail rotor performance and installing a SCAS.
16. In left sideward flight with SCAS OFF, as previously reported (ref 2, app A),at airspeeds from 15 to 25 KTAS, the helicoptc ý was directionally unstable, withyaw excursions of approximately 10 degrees which required extensive pilotcompensation to maintain adequate directional control (HQRS 6). The extensivepilot compensation required during left sideward flight at 15 to 25 KTAS is ashortcoming.
17. In rearward flight at the conditions tested, the aircraft was limited to 32 KTASbecause the longitudinal control limit was reached (fig. 14, app E). Therequirements of paragraph 3.2.1 of MIL-H-8501A were not met, in that 10 percentof the control power was not avai'able in 30-KTAS rearward flight. Insufficientaft longitudinal control in rearward flight is a shortcoming.
Lateral Acceleration Handlinl Qualities
18. The lateral acceleration handling qualities, with SCAS ON and OFF, wereevaluated during the lateral acceleration performance testing at the conditionspresented in table 1, using the methods outlined in paragraph 11. Lateralacceleration testing had not been previously conducted on the OH-58A helicopter.Represent.,tive time histories are shown in figures II and 12, appendix E. WithSCAS ON, minimal pilot compensation was required in left lateral accelerationto maintain heading, roll attitude, and pitch attitude up to approximately 18 KTAS(HQRS 3). Extensive pilot compensation was required to maintain flight path,heading, and roll attitude between 18 and 21 KTAS (HQRS 6). After acceleratingthrough 21 KTAS, flight path, heading, and roll attitude were easily controlledas the aircraft accelerated to the 35-knot limit sideward airspeed (HQRS 3). Lateralacceleration to ht.l right required minimal pilot compensation to smoothlyaccelerate to limit airspeed at roll attitudes up to 10 degrees (HQRS 3). At anglesgreater than 10 degrees, moderate pilot compensation was required to maintain
( •a constant heading (HQRS 4). The pilot's difficulty in maintaining heading occurredat approximately 2 to 5 seconds after initiation of rollover to roll attitudes greaterthan 10 degrees. Because of the moderate compensation required to maintainheading at higher roll attitudes, the pilot's attention was diverted to controllingthe aircraft and the transmission torque limits were inadvertently exceeded(para 12) bwfore pilot corrective a tion could be taken.
19. Representative lateral accelerations to the right and left were qualitativelyevaluated with SCAS OFF. Roll attitudes of 10 degrees werc used for thisevaluation. Lateral accelerations to the left required considerable pilot compensationto maintain a steady roll attitude avu heading at airspeeds up to approximately
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18 KTAS (HQRS 5). From 18 to 21 KTAS, considerable pilot compensation wasrequired to retain control of the helicopter because of the severe roll and yawexcursions at thcse airspeeds (HQRS 8). These roll and yaw excursions precludedtesting beyond 18 to 21 KTAS. Lateral accelerations to the right requiredmoderate pilot compensation throughout the acceleration to maintain a constantheading and roll attitude (HQRS 4). The moderate pilot compensation requiredto maintain right lateral accelerating flight is a shortcoming, and the considerablepilot compensation required to maintain al-craft control during left lateralaccelerating fl;ght is a deficiency.
20. The addition of a SCAS improves the handling 4aalities of the OH-58Ahelicopter by reducing the severe yaw and roll excursions in left lateral accelerationsufficiently to allow accelerations through the 18- to 21-knot regime. Withoutinclusion of a SCAS the capability to perform lateral accelerations is considerablyreduced. Extreme caution should be used if further testing is conducted at rollattitudes greater than 10 degrees. The following CAUTION should be incorporated
* in the operator's manual:
CAUTION
Lateral accelerating flight to the left at roll attitudes greaterthan 10 degrees will result in severe roll and yaw excursionswhich make control of the aircraft difficull. Additionally, rightlatural accelerating flight at roll attitudes greater than10 degrees may easily result in transmission overtorque.
SUBSYSTEM TESTS
Engine Characteristics
21. Engine characteristics of the Allison 250-C20B engine, including poweravailable and fuel flow, were determined from Allison Computer Source Deck 847.Power required was calculated using the torquemeter calibration performed by
*i Allison. Referred engine characteristics are presented in figures 15 through 20,appendix E. Engine shp available ard specification fuel flow are shown iiifigures 21 through 24. Inlet and exhaust losses were determined from datapreviously obtained with the T63-A-700 engine (rt'f 1, app A).
Engine Vibration
22. Vibration data were gathered concurrently with performance and handlingqualities tests. Vibration sensors were installed at the following enginelocations: compressor section, gearbox, turbine and combustion section, top enginemount pad, and fuel nozzle, as called out in Allison Installation Design ManualNo. lOW5 (ref 5, app A). Tail rotor gearbox vibration levels were also monitoredthroughout the evaluation.
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23. The vibration data were compared to the installed engine vibration limitsspecified on the installation drawing. The vibration characteristics at all the specifiedlocations will not be discussed in detail but, in general, show the presence of low-and moderate-amplitude low-frequency vibration levels (10 to 2000 Hz) at all ofthe accelerometer locations. Results are shown in tabular form in figures 25through 2i, appendix E. Maximum acceleration values below 10 percent of themaximum limit value specified by Allison were not presented. All the vibrationlevels were within Allison's specified limits. The ma:;imum average accelerationmeasured for the tail rotor gearbox was 1.8g at 1720 Hz in an IGE hover.
Engine Compartment Temperature Survey
24. A limited engine compartment temperature survey was performed during theevaluation. Temperatures were recorded at locations specified in Allx-3n Installation"Design Manual No. lOW5. The temperature probes were located :.t the followinglocations: compressor section, top engine mount pad surface, ignition harness,thermocouple harness, and oil cooler. Temperature data for all conditions testedare presented in figure 28, appendix E. When the component/fluid tinperaturedata are corrected to the Army's design requirement maximum ambient temperatureof 1250F (52°C), an overtemperature condition is indicated for the oil cooler outlettemperature, as is shown in figure 28. The temperature data were corrected bythe following equation:
52+ 273T = Tmeasured Tabient + 273
- (all temperatures are in degrees Centigrade)
The effect of altitude-temperature variation was not determined. Within the scopeof this evaluation, the engine compartment cooling was satisfactory, but correctingthe engine oil cooler temperature to 1250F ambient temperature indicates the oilcooler temperature would not be adequate.
Engine Governing Characteristics
25. Static and dynamic droop characteristics of the Allison 250-C20B enginegovernor were evaluated under the conditions listed in table 1. Test results arepresented in figures 29 through 48, appendx E. Dynamic stability characteristicsof the 250-C20r- engine were qualitatively and quantitatively investigated. Nocompressor stalls were experienced even with large rapid power demands. Therewere no undamped engine oscillating tendencies (figs. 29 through 32). There wereno undamped engine oscillating tendencies (figs. 27 through 30), although thetorque pressure was lightly damped, with a damping ratio of 0.06 at a dampedfrequency of 2.86 cycles per second. Dynamic stability ciiaracteristics weresatisfactory throughout the flight envelope tested.
26. Ile tests specified for the applicable sections of MIL-T-25920B (USAF) wereperformed to demonstrate suitability of the ground and flight operational and
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performance characteristics of the propulsion systemn of the aircraft installation(figs. 33 through 47, app E). Constant manipulation of the power turbine
- .. speed-select "beep" switch was required to maintain a desired rotor speed duringpower changes. This characteristic is the result of poor static droop characteristicsand is shown in figure 48. From a normal stabilized flight condition of 45 psiengine torque and a rotor speed of 354 rpm, application of collective control causedrotor speed variation fh .m 348 rpm durhig power increases to 362 rpm duringpower decreases. Returning to the trim power setting resulted in a permanent rotorspeed droop of 6 rpm, which is within I rpm of the minimum power-on rotorspeed. The static rpm droop characteristics of the 250.C2OB engine/airframecombination (the change in rotor speed versus engine power output) wereobjectionable for all flight conditions tested (fig. 48). The engine governingcharacteristics are a shortcoming.
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CONCLUSIONS
GENERAL
27. The following conclusions were reached upon completion of the Allison250-C20B engine evaluation:
a. Within the scope of this test, the performance of the OH-58A helicopterwith an Allison 250-C20B engine was improved over the basic OH-58A helicopter,while the handling qualities were essentially unchanged.
b. The OGE hover ceiling at 3000 pounds gross weight was increased to
11,050 feet.
c. The engine oil cooler temperature when corrected to 125T ambienttemperature indicates the oil cooler performance would not be adequate.
d. One deficiency and five shortcomings were noted.
DEFICIENCY AND SHORTCOMINGS
28. The following deficiency was identified: Considerable pilot compensationrequired to maintain aircraft control in left lateral accelerating flight (para 18).
29. The following shortcomings were identified:
a. Insufficient left pedal in right sideward flight at 32 KTAS (pars 15).
I b. Extensive pilot compensation required to maintain left sideward flight(para 16).
c. Insufficient aft cyclc control in rearward flight at 30 KTAS (para 17).
d. Moderate pilot compensation required to maintain right accelerating flight(para 19).
e. Objectionable engine governing characteristics of the OH-58A helicopterwith an Allison 250-C20B engine (para 26).
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SPECIFICATION COMPLIANCE
30. Within the scope of this test, the OH-581ý helicopter with an Allison 250-C20Bengine installed failed to meet the requirements of pargraphs 3.2.1 and 3.3.2 ofMIL-H-8501A, mn that sufficient aft cyclic control was not available during 30-KTASrearward flight and sufficient left directional control was not available during35-KTAS right sideward flight (paras 17 and 15).
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RECOMMENDATIONS
31.: The deficiency identified during this -evaluation shiould be corrected beforethe OH-58A helicopter is released to perform the lateral acceleration maneuverat roll attitudes greater than 10 degrees (pana 18).
32. The shortcomings sh' uld be corrected.
33. The following CAU'IION should be incorporated in the applicable sections- - of the operator's manual:
CAUTION
Lateral accelerat; I flight to the left at roil attitudes greaterthan 10 degrees vill result in severe roll and yaw excursionswhich make cot 'ol of the aircraft difficult. Additionaflly, rightlateral accelera, .Ag flight at roll attitudes greater than10 degrees may easily result in transmission overtorque.
34. An improved tail rotor should be installed to increase the operational capabilityof the OH-58A helicopter when equipped with a 250.C2OB engine.
35. A SCAS should be provided with both lateral and directional axes to improvethe low-speed and lateral flight, capability of the aircraft (paras 15 and 20).
.PPENDIX A. REFERENCES
i. Final Re, ort, USAASTA, Project No. 68-30, Airworthiness and Flight
Characteristics Test, Production OH-58A Helicopter, Unarmed and Armed with theXM27E1 Weapons System, Performr.Lce, September 1970.
2. Final Report, USAASTA, Project No. 68-30, Airworthiness and FlightCharacterist cs Test, Production OH-58A Helicopter. Unarmed and Armed with theMX27EI Weapons System, Stability and Control, October 1970.
3. Final TA!r..-rt, USAASTA, Project No. 71-24, Evalvation of the OH-58AHelicopter ': an Allison 250-C20 Engine, December 1972.
4. Letter, AVSCOM, AMSAV-EFT, 14 June 1974, subject: Evaluation of theOH-58A Helicopter with the Allison 250-C20B Engine Installed, Project No. 74-48.
5. Installation Design Manual, Detroit Diesel Allison Division of Geaieral MotorsCorporation, No. lOW5, Commercial Turboshaft Engine, Model 250-C20B,12 March 1974.
6. Final Report, USAASTA, Project No. 72-20, Handling Qualities Evaluationof 'he OH-58A Helicopter Incorporating the I odel 570B Stability and ControlAugmentation System, February 1973.
7. Technical Manual, TM 55-1520-228-10, Operator's Manual, Army ModelOH-58A Helicopter, 13 October 1970.
8. Letter, AVSCOM, AMSAV-EFT, 4 October 1974, subject: Safety of FlightRelease for AVSCOM/USAAEFA Project No. 74-48.
9. Military Specification, MIL-H-8501 A, Helicopter Flying and Ground HandlingQualities; General Requirements For, September 1961, with Amendment 1, 3 April1962.
10. Military Specification, MIL-T-25920B (USAF), Test, Ground and ,light,Aircraft Ges Turbine Propulsion System Installation, 10 February 1966.
11. •light Test Manual, Naval Air Test Cent'r, FTM No. 101, Helicopter Stabilityand Control, 10 June 1968.
12. Flight Test 'anual, Naval Air Test Center, FTM No. 102, HelicopterPerformance Testing, 28 June 1968.
19
- itl I ...!
f -11
13. Final Report, USAAF-A, Project No. 68-55, Flight Evaluation, ComplianceTest Techniques for Army Hot Day Hover Criteria, April 1974.
14. Final Report, USAAEFA, Project No. 68-25, Flight Research Investigationof Autorotational Performance and Height-Velocity Testing, in preparation.
20
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* I
1 20IQ
APPENDIX B. GENERAL ENGINE INFORMATION
GENERAL
1. The Allison Model 250-C20B engine is an internal ;nombustion turboshaftengine of the free turbirn type. The gas producer is conposad of a combinationsix-stage axial single-stage centrifugal flow compressor directly coupled to atwo-stage turbine. The power turbine is a two-stage free turbine that is gas coupledto the gas prodicer turbine. The integral reduction geaw-box provides an internalspline output drive at the front of the gearbox. The engine has a single combustionchamber. The output shaft center line is located below the center line of the enginerotor and the exhaust is directed upward. The performance rating for the uninstalledstandard-day static sea-level conditions is shown in table 1.
COMPRESSOK
2. The compressor assembly consists of a compressor front support, caseassembly, rotor wheels with blades, centrifugal impeller, front diffuser assembly,rear diffuser assembly, diffuser vane assembly, and diffuser scroll. Air enters theengine through tile compressor inlet and is compressed by six axial compressorstages and one centrifugal stage. The compressed air is discharged throug'. thescroll-type diffuser into two external ducts which convey the air to the combustionsection, as shown in figure 1.
COMBUSTION SECTION
3. The combustion s..ction consists of the outer combustion case and thecombuston liner. A spark igniter and a fuel nozzle are mounted in the aft endof the outer combustion case. Air enters the single combustion liner at the aftend through holes in the liner's dome and skin. The air is mixed with fuel sprayedfrom the fuel nozzle and combustion takes place. Combustion gases move forwardout of the combustion liner to v.ie first-stage gas producer turbine nozzle.
TURBINE
4. The turbine consists of a gas producer turbine support, a power turbine
support, a turbine and exhaust collector support, a gas producer turbine rotor,and a power turbine rotor. The turbine -s mounted between the combustion sectionand the power and accessory gearbox. The two-stage gas producer turbine drivesthe compressor and accessory gear train. The two-stage power turbine furnishesthe ounput power of the engine. The expanded gas discharges in an upward directionthrough the twin "ucts of the turbine and exhaust collector support.
21
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Figure 1. Allison 250-C20B Engine Cutaway.
23
POWER AND ACCESSORY GEARBOX
5. The main power and accessories drive gear trains are enclosed in a single gearcase. The par case serves as the structure support of the engine. All enginecomponents, including the engine-mounted accessories, are attached to the case.A two-stage helical and spur gear set is used to reduce rotational speed from33,290 rpm at the power turbine to 6016 rpm at the output drive spline.Accessories driven by the pcwer turbine gear train are the power turbine governorand an airframe furnished power turbine tachometer-generator. The gas producergear train drives the compressor, fuel pump, gas producer fuel control, and anairframe furnished gas producer tachometer-generatcr. The starter chive and a sparedrive are in this gear train.
FUEL SYSTEM
S6. The principal components of the fuel system are a fuel pump, a gas producerfuel control, a power turbine governor, and a fuel nozzle. The fuel control andgovernor are located in the system between the fuel pump and the fuel nozzle.The engine can be equipped with either a Bendix or Chandler Evans (CECO) fuelsystem. The test aircraft was equipped with a Bendix fuel system.
j .
24
7 __ 7 ------- -- _______-7__7__7_
I
APPENDIX C. INSTRUMENTATION
PILOT/ENGINEER PANEL
Airspeed (boom)Altitude (boom)Angle of sideslipRotor speedCenter-of-gravity normal accelerationFree air temperatureTotal fuel usedLongitudinal control positionLateral control positionDirectional control positionCollective control positionTurbine outlet tcmreratureEngine compartment temperaturesEngine torqueGas producer speedEngineer event markerRecord counter
MAGNETIC TAPE RECORDER
Longitudinal control positionLateral control positionDirectional control positionCollective control positionPitch attitndeRoll attitudeYaw attitudePitch rateRoll rateYaw rateAltitude (boom)Airspeed (boom)Altitude (radar)Vertical speed (radar)Free air temperatureAngle of sideslipCenter-of-gravity normal acceleration
7 4
Rotor speedThrottle positionEngine torqueGas producer speedEngine vibrationEngine com'partment temperaturesFuel flow rateEngine fuel i.'ozzle pressureEngineer event maurker
26
APPENDIX D. TEST TECHNiQUES ANDDATA ANALYSIS MEFTHOD3
1. This appendix contains some of the data reduction and analysis methods z.d
to evaluate the OH-58A helicopter equipped with an Allison 250-C20B engine.
The topics discussed include hover, vertical climb, and lateral flight performance.
2. The helicopter performance test data were generalized through the use of
nondimensional coefficients. The purpose is to accurately obtain performance at
conditions not specifically tested. The following nondimensional coefficients were
used to generalize the hover and vertical climb test results obtained durinr' this
flight test program.
a. Coefficent of power (Cp):
SlHP x 550 (1)!,p A(f2R)
b. Coefficient of weight (CW):
W '(2)
pA (•fR)During stabilized hover, coefficient of thrust equals coefficient of w-iMgt.
c. Vertical ,Avance ratio:
VvOR (3)
Where:
SHIP Engine output shaft horsepower
550 = Conversioni factor (ft-lb/sec/shv)
p - Air density (lb-sec 21ft4 )
A - Main rotor disc area (ft 2 )
t" 1 -" Main rotor angular velocity (rad/sec)
R - Main rotor rr-lius
W - Gross weight (lb)
Vv = Vertical ve!. city (ft/sec)
27
.. ]
3. Engine output shp was determined from the engine torque pressure. Torquepressure as a function of the power output of the engine was obtained from theengine manufacturer's test cell calibration. The shp required was determined bythe following equation:
i! 2"x X x GR x x Q
here: SUiP ,33,000
. I Where:
SHP Shaft horsepower
Kt Conversion factor to change measured engine torque pressure (psi)to ft-lb (from contract( r's engine acceptance data)
GR = Gear ratio of the output shaft rotational speed to the mainrotor rotational speed
NR = Main rotor speed (rpm)
.- j Q Engine torque pressure (psi)
.J •33,000 Conversion factor (ft-lb/min per shp)
HOVER
4. Hover performance was determined IGE and OGE by the free-flight hovertechnique. Formulas 1 and 2 were used to define the hover capability. A plotof Cp versus CT was constructed for each skid height tested. Hover performancecharacteristics may be extracted from these curves in preparing tables or curvesfor flight manuals for any combination of conditions.
VERTICAL CLIMB'
5. The vertical climb technique used by the pilot was to stabilize in a 50-foot
OGE hover and then to increase engine power by a predetermined increment cfgas producer speed (NI) over the hover Ni speed to the transmission torque limit.Two cues were used by t;,-., pilot to maint ain vertical climbing flight. An Elliottlow-airspeed sensor sensitive to I knot in horizontal a!,-speed was used to providecues of lateral translation during the climb. To provide cues to any forward oraft movement, the pilot used a narrow taxiway as a visual reference.
6. Vertical climb performance was J.etermined by using equations 1, 2, and 3.Each vertical climb was flown at a predetermined W/6 and N/V7'. To maintain
28
W/6 approximately constant, the aircraft was periodica!.y reballasted as fuel was
consumed. N/V'F was held constant by increasing or decreasing rotor speed as theambient air temperature increased or decreased, respectively.
7. The climb rates were measured after the aircraft was stabilized in anonaccelerating vertical climb by means of a radar altimeter. The initial rate ofclimb (dh/dt) was corrected to tapeline rate of climb (R/CT) by the equation:
dh Ta SR/c x h thT/ dt Tat
Where:
Tat = Test ambient air temperature (*K)
Tas = Standard ambient air temperature (*K)
8. The standard rate of climb was determined by correcting the tapeline rateF of climb for gross weight differences and by the power-einergy equation for
nonsteady flight conditions. The adjusted momentum theory discussed inreference 13, appendix A, was used to facilitate curve fairing. A power adjustmentfactor (Ks) of 3.5 and equivalent flat plate area of 50 square feet were foundto adequately fit all data sets.
U 9. After the raw data were reduct.d to calibrated engineering units, it waspresented in referred terms of SHPI6/V, Vvt0"O, W/8, and NR 0Vr. For conveniencethe test data were also presented in dimensionless parameters, Cp, CW, and PNv.
~ r ~Vertical climb performance cliaracteristicz may be extracted from these curves Lnpreparing tables or curves for flight manuals or further engineering evaluation forany combination of conditions.
LATERAL FLIGHT PERFORMANCE
10. Lateral flight performance characteristics data were reduced from parametersrecorded on board the aircraft. The test helicopter was equipped with aircraft(body) axis accelerometers and by measuring Euler angles, it was possible to
V I' transform the aircraft axis accelerations into earth (ground) axis accelerations. Thismade it possible to measure the lateral aircraft acceleration without the use ofspace positioning equipment. A detailed explanation of this technique can be found
. in reference 14, appendix A.
11. The HQRS pre-ented as figure 1 was used to augment pilot comments relative*i to handling qualities.
Ii'
, 41
V 29-1
DEMANDS ON THE PILOTADEQUACY FOR SELECTED TASK AIRCRAFT IN SELECTED TASKPIO
OR REQUIRED OPERATION* CHARACTERISTICS OR REQUIRED OPERATION* RATING
Attainable~ WitELEN A-EUR AORCniea
pilot compensation no atrequore
PILOT ECISINS Raing Sal. (Rf GOA D TN-S1 APdilVot comensgation noflgt pas factr/or
FigureSIRBL .desln Qaiie aingd Scrorale.
I;S. FAI - Iua dtcmarssnrqie o
SOME, IID. de-d promneUUNLEAAN
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*I.
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APPENDIX E. TEST DATA
INDEX
F"r Wi,.re Number
PerformanceHover I through 3
Vertical Climb 4 through 9
Lateral Acceleration 10 through 12
Handlinz Qualities
Low-Speed Flight Characteristics 13 and 14
Sutsystem Tests
Engine Characteristics 15 through 2,1,Engine Vibration 25 through 27Engine Compartment TemperatLre 28Engine Governing Characteristics 29 thrcugh 46
31
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H >1 ~~ FIGURE 2VrL0L'1! ~SHAFT "$ORWEOWERAVIBL_
ii::::ALLISON 250-C20EIENGIN4EI ffQ~.,NRMAL RATED POWER.TTSi51J~:~:
--------- -. :STATIC. cCtONDTIONS. . :2* 2:. ZERO BLEED ANDOANTI-ICE OFF., .... 77I .3~KUREXTRACTED. EQUALS, 24) SHP...__--
I .4. _IN~LE AND EXHAUST PRESSURE LOSSESOBTAINED'1....R0MREFERENCE1 APPEN4DIX A..
* .'.53S SHAFT HORSEPOWER BASED ON ALLISON COMUTER. DECK NO. :847.1
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FIGURE 25ACCESSORY GEARBOX LONGITUDINAL AXIS VIBRATION
MAXIMUM AVERAGEFLIGHT CONDITION SOURCE ACCELERATION(g) (Hz)
Ground-idle Tail rotor 8/rev 1.25 350.4
Flight-idle Note1 ......
Hover (IGE) Tail rotor 12/rev 2.08 525.6
Normal takeoff Tail rotor 12/rev 1.95 525.6
Level light Tail rotor 12/rev 1.87 525.6at 80 kt
50-kt climb at500 ft/min
80-kt climb at500 ft/min
80-kt climb at Tail rotor 12/rev 1.98 525.6maximum power
Main rotor Fundamental 0.52 5.9Main rotor 2/rev 0.42 11.8Main rotor 4/rev 0.34 23.6Main rotor 6/rev 0.31 35.4Main rotor 8/rev 0.39 47.2
50-kt descent Main rotor 10/rev 0.38 59.0at 500 ft/min Main rotor 12/rev 0.44 70.8
Tail rotor 2/rev 0.42 87.6Unknown 0.42 121.0
Engine andtail rotor Fundamental 0.43 121.0
shaft
80-kt descent Main rotor 6/rev 0.8 35.4at 500 ft/min Tail rotor 12/rev 1.87 525.6
S-turns at 100 kt Mand 450 bank angle Main rotor 2/rev 0.05 11.8
IDashes indicate values less than 10 percent.
1 I8
+ _ .. 58.:
FIGURE 26250C208 ENGINE LATERAL AXIS VIBRATION
• ~~~MAXIMUM AVERAGE FEUTCACCELEROMETER MXI'JAVRGEFREQU ENCYP FLIGHT CONDITIONl LOCATION SOURCE ACCELERATION (Hz)
Gearbox Note1 ......Ground idle Compressor ---........
Fuel nozzle ---.. ..
Gearbox Main rotor 2/rev 0.20 11.8Hover (IGE) Compressor Main rotor Fundamental 0.08 5.9SFuel nozzle N1 Fundamental 4.32 852.0
Fuel nozzle Main rotor 12/rev 0.34 70.8
Gearbox Main rotor Fundamental 0.19 5.9Gearbox Main rotor 2/rev 0.17 11.&Gearbox Main rotor 4/rev 0.14 23.6
NFuel nozzle Main rotor Fundamental 0.48 5.9Fuel nozzle Main rotor 2/rev 0.43 11.8Fuel nozzle Main rotor 4/rev 0.34 23.6Fuel nozzle Main rotor 6/rev 0.29 35.4Compressor Main rotor Fundamental 0.15 5.9
Gearbox Main rotor Fundamental 0.34 5.9
Level flight at Compressor Main rotor 2/rev 0.29 11.880 kt Fuel nozzle Main rotor 6/rev 0.34 35.4
Fuel nozzle Main rotor Fundamental 0.23 5.9
Fuel nozzle N1 Fundamental 4.2 852.0
50-kt c lim b at Com p re s so r .. ..... . .500 ft/mm Gearbox500 Fuel nozzle .........
80-kt climb at Compressor ............500- ft linb a Gearbox ---..........-500 ft/mn Fuel nozzle ---........
Gearbox Main rotor Fundamental 0.18 5.9Fuel nozzle Main rotor Fundamental 1.4 5.9Fuel nozzle Main rotor 2/rev 1.36 11.8Fuel nozzle Main rotor 4/rev 1.18 23.Cmaximum power Fuel nozzle Main rotor 6/rev 0.72 35.4"Fuel nozzle Tail rotor Fundamental 0.53 43.0Fuel nozzle Main rotor 12/rev 0.98 70.8
50-/t descent Gearbox Main rotor 2/rev 0.06 11.8Gearbox Main rotor 2/rev 0.07 11.8
at 500 ft/mn Fuel nozzle ---.........
Gearbox Main rotor Fundamental O.l7 5.980-kt descent Gearbox Main rotor 2/rev 0.13 11.8at 500 ft/min Compressor Main rotor 2/rev 0.09 11.8
Fuel nozzle Main rotor 2/rev 0.14 11.8
Fuel nozzle Tail rotor Fundamental 0.97 43.8
S-turns at 100 kt Gearbox Main rotor Fundamental 0.08 5.9and 450 bank angle Compressor Main rotor Fundamental 0.04 5.1
Right 900 turn at Fuel nozzle Main rotor Fundamental 0.2 5.9100 kt and Compressor Main rotor Fundamental 0.15 5.9450 bank angle Gearbox Main rotor Fundamental 0.16 5.9
IDashes indicate values less than 10 percent.
59
FIGURE 27250C208 ENGINE VERTICAL AXIS VIBRATION
MAXIMUM AGERAG EQECACCELEROMETER FREQUENCY
FLIGHT CONDITION LOCATION SOURCE ACCELERATION (Hz)(9)
Gearbox Note] ......G r o u n d - i d l e T u r b i n e - - -. . .. . .. . .
Fuel nozzle ...
Fuel nozzle Main rotor Fundamental 0.12 5.9
Flight-idle Compressor Main rotor 2/rev 0.08 11.8Fuel nozzle Unknown 3.8 635.0
Turbine ---...Gearbox ---....
Hover IGE Compressor -. -- ... ---
Fuel nozzle Main rotor Fundamental 0.21 5.9
Gearbox Main rotor Fundamental 0.16 5.9Turbine Main rotor Fundamental 0.39 5.9Turbine Main rotor 2/rev 0.31 11.8Turbine Main rotor 4/rev 0.18 23.6
Normal takeoff Fuel nozzle Main rotor 2/rev 1.0 11.8Fuel nozzle Main rotor 4/rev 0.80 23.6Compressor Main rotor Fundamental 1.0 5.9Campressor Main rotor 2/rev 0.46 11.8Compressor Unknown 0.18 18.0
Fuel nozzle Main rotor Fundamental 0.64 5.9Fuel nozzle Main rotor 2/rev 0.69 11.8Fuel nozzle Main rotor 4/rev 0.62 23.6Fuel nozzle Main rotor 6/rev 0.60 35.4
Level flight Fuel nozzle Main rotor 10/rev 0.94 59.0at 80 kt Fuel nozzle Unknown 0.94 121.0
Fuel nozzle Tail rotor 4/rev 0.84 175.0Turbine Main rotor Fundamental 0.07 5.9
Compressor Tail rotor Fundamental 0.13 43.8Gearbox ......
50-kt climb at500 ft/mln Compressor ---.......
F u e l n o z z l e - - -. . .. . .. . .
Gearbox ---...8O-kt climb at Compressor ---...500 ft/min Fuel nozzle ---...
Turbine Main rotc: Fundamental 1.76 11.8Turbine Main rotor 2/rev 1.46 23.6Turbine Main rotor 4/rev 1.04 35.4
80-kt climb at Turbine Tail rotor Fundamental 0.68 43.8maximum power Turbine Main rotor 10/rev 0.60 59.0
Turbine Main rotor 12/rev 0.42 70.8Turbine Tail rotor 2/rev 0.29 87.6Turbine Main rotor 8/rev 0.65 47.2
50-kt descent Gearbox ---at500 descen Compressor .......
Fuel nozzle ---
Turbine Main rotor Fundamental 0.52 5.9Turbine Main rotor 2/rev 0.32 11.8
80-kt descent at Turbine Tail rotor Fundamental 0.44 43.8500 ft/min Turbir.. Main rotor 10/rev 0.91 59,0
Turbine Unknown 0.86 121.0Compressor Main rotor 6/rev 0.23 35.9
Gearbox Unknown 2.1 121.0S-turns at 100 kt Turbine ---and 950 bank angle Fuel nozzle ---.........
IDashes indicate values less than 10 percent.
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DISTRIBUTION
Director of Defense Research and Engineering 2Assistant Secretary of the Army (R&D) IChief of Research and Development, DA (DAMA-WSA) 3US Army Materiel Command (AMCPM-UA, ADCRD-FQ, AMCSF-A, AMCQA) 9US Army Aviation Systems Command (AMSAV-EQ) 12US Army Training and Doctrine Command (USATRA DOC/CDC LnO,SATCD CM) 22US Army Test and Evaluation Command (AMSTE-BG, USMC LnO) 3US Army Electronics Command (AMSEL-VL-D) IUS Army Forces Command (AFOP-AV) Ii " •US Army Armament Command (SARRI-LW) 2
US Army Missile Command IUS Army Munitions Command IHq US Army Air Mobility R&D Laboratory (SAVDL-D) 2
V US Army Air Mobility R&D Laboratory (SAVDL-SR) IAmes Directorate, US Army Air Mobility R&D Laboratory (SAVDL-AM) 2
J' Eustis Directorate, US Army Air Mobility R&D Laboratory (SAVDL-EU-SY) 2Langley Directorate, US Army Air Mobility R&D Laboratory (SAVDL-LA) 2Lewis Directorate, US Army Air Mobility R&D Laboratory (SAVDL-LE-DD) IUS Army Aeromedical Research Laboratory I
F US Army Aviation Center (ATZQ-DI-AQ) IUS Army Aviation School (ATST-AAP, ATST-CTD-DPS) 3US Army Aviation Test Board (STEBG-PR-T, STEBG-PO, STEBG-MT) 4US Army Agency for Aviation Safety (FDAR-A, IGAR-MS/Library) 2US Army Maintenance Management Center (AMXMD-MEA) IUS Army Transportation School IUS Army Logistics Management Center I
i US Army Foreign Science and Technology Center (AMXST-CB4) 1US Military Academy 3
SUS Marine Corps Development and Education Command 2
US Naval Air Test Center IUS Air Force Acronautical Systems Division (ASD-ENFDP) IUS Air Force Flight Dynamics Laboratory (TST/Library) I
I. 77'
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US Air Force Flight Test Center (SSD/Technical Library, DOEE) 3US Air Force Special Communications Center (SUR) IDepartment of Transportation Library 2US Army Bell Plant Activity (SAVBE-F) 5Bell Helicopter Company 5Detroit Diesel Allison Division of General Motors Corporation 2Defense Documentation Center 12
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