UNCLASSIFIED AD NUMBER LIMITATION CHANGES toward the artillery mission of displacing a 105mm...

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Distribution authorized to DoD only;Administrative/Operational Use; MAR 1969. Otherrequests shall be referred to Army AviationSystems Command, St Louis, Missouri. Pre-datesformal DoD distribution statements. Treat asDoD only.

USAASTA ltr 12 Nov 1973

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RDTE PROJECT NÜ. USAAVSCOM PROJEC'IVvNOv/68-46 USAAVNTA PROJECT NO. 68-46

ARMY PRELIMINARY EVALUATION OF THE

PROTOTYPE BHC MODEL 211

HUEYTUG

Final Report

JOHN I. NAGATA PROJECT ENGINEER

M I*

to

THEODORE K. WRIGHTu LTC, ARTY B US ARMY PROJECT OFFICER

EH

IVAR W. RUNDGREN MAJ, TC US ARMY PROJECT PILOT

E-i

MARCH 1969

DISTRIBUTION

E

t.

^ This document is subject to special export controls and each transmittal to foreign government or foreign nationals may be made only with prior approval of the Commanding General, Hq, USAAVSCOM, ATTN: AMSAV-R-F, Main Office, PO Box 209, St. Louis, Missouri 63166.

i US ARMY AVIATION SYSTEMS TEST ACTIVITY EDWARDS AIR FORCE BASE, CALIFORNIA 93523

D D C

FOR OFFICIAL USE ONLY ^N.^

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THIS DOCUMENT IS BEST QUALITY AVAILABLE. THE COPY

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A/I-A r ) U C- i:mri'A-t>t

\ ARMY PRELIMINARY EVALUATION

OF THE PROTOTYPE BUG MOÜEL 211

(HUEYTUG), /

/ FINAL REPORT, -.-// ': /.

JOHN I. NAGATA PROJECT ENGINEER

■y/Jy pEODORE J<:. / WRIGHT ,

U&?ABMY I PRGJEGT-OF-IMC&R

IVAR W. /RUNDGREN

™rrTC 3^ US ARMY PROJECT PILOT

T y/Vo^pJ^

JJJ Marafe- ä»69 j (/^j-j huL

DISTRIBUTION

This document is subject to special export controls and each transmittal to foreign government or foreign nationals may be made only with prior approval of the Commanding General, Hq, USAAVSCOM, ATTN: AMSAV-R-F, Main Office, PO Box 209, St. Louis, Missouri 63166.

US ARMY AVIATION SYSTEMS TEST ACTIVITY EDWARDS AIR FORCE BASE, CALIFORNIA 93523

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abstract The Army Preliminary Evaluation (APE) of the Bell Model 211 prototype helicopter (Hueytug) was conducted at the Bell Helicopter Test Facility, Arlington, Texas, Edwards AFB, California, and Bishop, California, from 19 October through 7 November 1968. Flying qualities, performance, and mission suitability were evaluated to determine aircraft capabilities to carry six thousand pound sling loads at a takeoff gross weight of 14,000 pounds. Primary emphasis was directed toward the artillery mission of displacing a 105mm Howitzer M101A1 with 10 rounds of ammunition and 3 cannoneers. The helicopter had eight deficiencies which require mandatory corrections. Two of these are major design deficiencies that may require extensive engineering redesign. They are the directional oscillations in the 30 to 60 KIAS airspeed range, especially prevalent during heavy sling load missions; .lack of sufficient directional control margin during high gross weight (14,000 pounds) and high density altitude (above 4000 feet) conditions. The renain- ing six deficiencies are ineffective force trim feature at high airspeeds, excessive forward position of longitudinal control at high airspeeds, poor static engine droop compensation, tail rotor drive train torque limitations, lack of an engine power torque limiter and lack of a standby generator for IFR flight. There are seven shortcomings the corrections of which are desirable and should be accomplished as soon as possible. The prototype model 211 could marginally perform the 14,000 pound gross weight misfion at sea level. At 4000 feet density altitude the marginal tail rotor control and transmission and drive train torque limitations prevented the helicopter from satisfactorily accomplishing the mission. Correction of the deficiencies discovered during this APE coupled with the 200 horsepower increase in drive train torque limits of the design proposal should result in a superior performing helicopter. Correction of the deficiencies should be accomplished prior to a production contract.


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TABLE OF CONTENTS

iNTKonurnoN

Background I'usi übjei-l i vfs DescrLptLou Scopt' of Tivi. Mc'Liiod ot Tost Chrüaolugy

RESULTS AND DISCUSSION

Hover Performanct! Level Flight Perforni;inci;

General Range Performance Endurance

Autorotation Airspeed Calibration Stability and Control

Dynamic Lateral-Directional Stability Static Lateral-Directional Stability Static Longitudinal Stability Longitudinal Control Motion Dynamic Longitudinal Stability

Control Response Longitudinal Control Response Lateral Control Response Directional Control Response

Sideward and Rearward Flight Control Margin Simulate^ Power Failures Sling Load Operations Stability and Control Augmentation System Miscellaneous

Collective Creep Structures Vibration Power Management Drive Train Limitations Torque Limiter

4 4 5 5 6 1 8 8 8 8 9 9 10 10 10 10 11 12 13 13 14 14 15 16 16 16 16 16 16 1?

i OFFICIAL USE ONLY

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Noise Level Gearbox Temperatures Power Source Limitations Cargo Mirror

17 17 17 17

CONCLUSIONS

General Specific

RECOMMENDATIONS

19 19

21

APPENDIXwES

I. References 22 II. Flight Restrictions and Operating Limitations 23

III. Performance Test Conditions 27 IV. Test Data and Index 29 V. Test Instrumentation 93

VI. Pilot's Rating Scale 94 VII. Control Motion 95

VIII. Distribution 96

VI


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INTROD M u ^hssß ta%j

i. In 196b the Bell Helicopter Company (BHC) commented tlie devel- opment of an artillery-prime mover version oi the Ull-1 helicopter Concurrently, BHC also began developing the dynamic components for a 20ÜÜ shaft horsepower (shp) drive system. In early 1968, a con- verted model UH-1C with increased horsepower, larger rotor blades and additional modifications was first flown and introdur 'd as the BHC Model 211 (Hueytug). The prototype Hueytug was designed to transport .sling loads weighing up to 600Ü pounds at a design takeoff gross weight of 14,000 pounds. The Hueytug is also designed for battlefield recovery of downed aircraft, command aiu. control, medical evacuation and resupply missions. The US Army Aviation Systems Test Activity was directed by the US Army Aviation Systems Command (ref 1, app I) to perform an Army Preliminary Evaluation (APE) on the prototype BHC Model 211 (Hueytug). N

TEST OBJECTIVES

2. The objectives of this test were to evaluate the Nielicopter performance, stability and control characteristics witrvin the established flight envelope, and to determine mission suitability. This evaluation was conducted with internal and external loadings, with particular emphasis on known stability and control deficiencies found in the UH-1B/C (ref 3, app I).

DESCRIPTION

3. The prototype Model 211 helicopter is a modification of the UH-1B/C series helicopter and is designed for the external trans- portation of heavy loads. Modifications incorporated in the basic airframe are as follows:

a. T55-L-7B turboshaft engine with a takeoff power rating of 2650 shp at sea level standard day conditions.

b. Fifty foot diameter two bladed main rotor with a 27 inch chord.

c. Rotor mast extended 12 inches.

d. Eighteen hundred shp dynamic drive sxstem.

1


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e. Tail boom .structurally reinforced and extimded 45.25 inches

f. Tail rotor diameter of 9 feet, 8 inches.

g. Increased structural rigidity of the fuselage.

h. Three axis stability and control augmentation system (SCAS).

4. The design proposal of the Model 211 includes the following modifications not present in the prototype:

a. T55-L-7C Lycoming turbo shaft engine with a takeoff rating of 2850 shp at sea level standard day conditions.

b. Two thousand slip dynamic drive system.

c. Tractor tail rotor.

d. Main rotor and tail rotor blades incorporating two double sweep back blades (outboard of 80% main rotor span).

SCOPE OF TEST

5. The helicopter was evaluated as a heavy lift vehicle (14,000 pounds design gross weight) with primary emphasis on the artillery mission of displacing a 105 mm M101A1 howitzer, 10 rounds of ammunition, and a crew of three plus pilot and copilot within a 50 nautical mile (NM) radius.

6. Flight restrictions and operating limitations issued by USAAVSCOM, St. Louis, Missouri are presented in appendix II. The test conditions are presented in appendix III.

7. This test encompassed three weeks which includes ferry time and aircraft preparation. Twenty-two test flights were conducted for a total of 25.3 test hours. In addition, 15.0 hours were flown ferrying the aircraft from Arlington, Texas, to a high altitude test site at Bishop, California.

METHOD OF TEST

8. Performance and stability and control test techniques as outlined in reference 2, appendix I, were adhered to in obtaining the pertinent helicopter characteristics. Deviations to the above are clarified in paragraphs 9 and 10.

2


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9. Slow upuuil data wert obtained by stabilizing tlie helicopter in sideward, rearward, or forward flight, with the aid of a pace vehicle with calibrated anemometer. Control position data and anemometer readings were recorded.

ID. Static longitudinal stability (collective fixed) was evaluated in climbing flight by performing constant power setting climbs through, a density altitude of 5000 feet at selected airspeeds above and below the best climb speed (62 KCAS).

CHRONOLOGY

11. The chronology of this APE is as follows:

Test directive received Test plan submitted Test team arrived at contractor's

facility Flight test commenced Flight test completed Test helicopter returned to contractor Preliminary report submitted Final report

11 September 1968 5 October 1968

13 October 1968 19 October 1968 7 November 1968 7 November 1968

12 December 1968 March 1969

FOR ;!AL USE ONLY

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DISCOS MU

GENERAL

12. The procotype Lust lielicopter was evaluaLed within the proposed flight envelope for limited performance and stability and control character is t ii;;. Problem areas specified in the UH-1B/C test report were carefully compared with the flight characteristics of the llueytug. There were no contractor or military speci- fication guarantee requirements. Power available data were derived from Lyc.oming engine charts for the proposed T55-L-7C engine and for a 2000 slip dynamic drive train. The pilot's rating scale (app VI) was used for stability and control evaluation. Test instrumentation used during the conduct of the test are presented in appendix V. Power available and fuel flow data for the Lycom- ing T55-L-7C engine are presented in figures 1 and 2, appendix IV. This data were furnished by Bell Helicopter Company and is based upon the design proposal installation with the test inlet losses of figure 22, appendix IV, applied, except that inlet particle separator screens were not installed. Power required data were determined by summing the power extracted from the accessory gear- b(5k^tC ) and dividing this sum by the speed decreaser shaft effi- ciency (0.988). This correction was required because of the location of the pickup for the engine torquemeter. All stability and control testing was performed with the SCAS operating unless otherwise specified. Control motion data are presented in percent of control travel on stability and control plots. Amount of control movement with percent travel data are presented in appendix VII. Control forces are unchanged from a UH-1P/C helicopter.

HOVER PERPOMANCE

13. Hover performance tests were conducted at density altitudes ranging from 2110 feet to 10,540 feet. Tests were conducted at skid heights of 7 feet in ground effect (ICE) and 100 feet out of ground effect (OGE). The tethered hover method of test was used with an attached calibrated load cell to determine the load at various power and rotor rpm settings. Quantitative data are presented in figure 3, appendix IV, and the hovering summary for OGE capability is presented in figure 4. The summary plot was derived from the T55-L-7C engine power available charts and a transmission limit of 2000 shp. With the above criteria, the maximum altitude that the helicopter can hover OGE on a standard day at 14,000 pounds gross weight is slightly greater than 10,000 feet. On a 35 degree centigrade hot day, the maximum OGE hover altitude is


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DO feet. During a letln-Tud hover test at a density altitude of 10,540 fe't, rotor rpm I'Bl', full left directional pedal was required to maintain direct ion at 59 percent engine torque (1700 slip). Hover performance is satisfactory providing tail rotor con- i r i' power is increased to .Jlow usage of the full 2000 slip of the i..t ign proposal.

LEVEL FLIGHT PERFORMANCE

i.( ■ral

] ■ Level flight performanc • tests were conducted to determine tli*. power required as a func.ion of airspeed. Various gross weights, altitudes and sling load configurations were used to achieve a wide range of thrust coefficients (C ) . Quantitative data are presented in figures 5 through 15, appendix IV, and summarized in figures 16 and 17. Combinations of cargo doors and cargo mirror on or off were flown to determine the equivalent Irt plate area (F ) penalty. Figures 10 and 11, show that with

the cargo mirror on and cargo doors open a 5.5 square feet increase of F or r'.5 percent increase in power required occurred at 120 knots true airspeed (KTAS) as opposed to the doors closed, r-irr r off configuration. With the cargo doors open (see fig 12 ...iQ 3), the F was increased 2,0 square feet resulting in a 4,2 pert nt increase in power required at 120 KTAS. The cargo mirror by itself created 3.5 square feet of F . A 105 mm howitzer M101A1 with ten rounds of ammunition was used as a sling load in one levf 1 flight performance test. Figure 15 shows 27 percent increase in power required at the limit airspeed of 80 KTAS. Another test used a conex container as a sling load (fig 14). This conex container required an increase of 21.8 percent power at 60 KTAS.

TABLE 1. Next Page

FOR ornniAi USE ONLY


Table 1. Kquivalent Flat Plate Summary.

Configurat h';, Incremental Equivalent

Flat Plate, .',F ! e i

Cargo doors closed

Cargo mirror off

Cargo doors closed and cargo mirror off

105 howitzer and 10 pounds of ammunit ion

conex container

- 2.0 ft2

- 3.5 ft2

-5.5ft2

2 54 ft 1

2 94 ft |

The AF is based on a comparison of the helicopter with cargo doors open and cargo mirror on configuration. The AF of the sling loads is based on extrapolated data based on the above configuration.

Range Performance

15. Range performance (fig 18, app IV) was calculated from the level flight performance data for sea level standard day conditions with cargo doors open and cargo mirror on. Radius of action for the artillery mission of displacing the 105 mm M101A1 howitzer at 80 KTAS with 10 rounds of ammunition and three cannoneers, then returning empty to home base at 140 KTAS with 10 percent reserve fuel is 54 NM and was computed as follows:

Takeoff Condition Pounds

Empty weight 5791 Crew (2) 400 105 mm howitzer plus 10 rounds

of ammunition 5840 Cannoneers (3) 600 Fuel 1369

Takeoff gross weight 14000 pound


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.In ' a ranee

u rnduranoe vaKr s for cai ô doors open, carg.i mirror on lor .h, . gross weights and two configurations are presented in table Tin tucl flow criteria were based on the T55-L-7C engine, figure .

Table 2. Endurance,

Sea Level Standard Day j 10% Reserve Fuel

82 lb Warm-up and Climb Fuel |

j Gross Weight

i (lb") Configuration

Useable Fuel

(lb)

Endurance Airspeed (KTAS)

Endurance Time (hours) |

1 * 8,000 doors open mirror on no 'ling load

1334 59 2.2 1

*■-• 10,500 S^/ac d3 ä-boye 1334 60 2.0 |

| 14,000 doors open mirror on sling load 105 mm howitzer ilO rounds ammunition piggyback

1150 60 1.4

" Gross weight based on full fuel, pilot and copilot, and mission essential equipment.

"" ..'-oss weight is the maximum allowable for internal loading.

7

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17. AuLorotat j-un ifsls were cunducted ac two gross weights (8000 and L0,550 pounds) at an average density altitude of 3000 feet. The quantitative data are presented in figure J9, appendix IV. The airspeed for minimum rate of descent was 62 KCAS which gave a rate of descent of I()()2 fpm. The airsjieed for maximum glide distance (78 KCAS) produced a rate of descent of 1825 fpm. For every 1000 feet of descent, V.UK) feet of horizontal distance is tra- versed. There were no unusual aircraft characteristics observed during these tests. At an airspeed of 62 KCAS, tests.were conducted at various rotor rpms. Figure 20 shows that the low rotor rpm (282.5) produced a rate of descent of 1A82 fpm, while the high rotor rpm (311.0) had a corresponding rate of descent of 1960 fpm. Gross weight differences did not alter the minimum rate of descent during these tests. Future tests should be conducted at heavier gross weights using external sling loads to determine how the rate f descent varies with gross weight. o

AIRSPEED CALIBRATION

18. The pace method (UH-1C) was used for airspeed calibration. This was performed by comparing the sensitive calibrated boom airspeed systems installed on both the test and pace helicopters. The airspeed calibration data are presented in figure 21, appendix IV. The standard aircraft airspeed and altimeter were not calibrated .

STABILITY AND CONTROL

Dynamic Lateral-Directional Stability

19. Qualitative results of the dynamic lateral-directional stability characteristics were obtained by releasing from steady heading sideslips, directional control "doublets," and flight evaluation during gusty atmospheric conditions. The helicopter exhibited a lateral-directional oscillation which was primarily present in the 30 to 60 KIAS band. The motion was essentially a yaw oscillation which was easily excited during gusty conditions. During a sling load test sequence the helicopter transmitted the yaw oscillation to the piggyback load (10 rounds of ammunition slung below a 105 mm howitzer). The ensuing lateral oscillation (neutrally damped) was severe enough to cause side forces resulting in full ball deflection of the turn and slip indicator. Air- speed and power changes were required to stop the oscillation. The present directional axis SCAS capability is Inadequate to cope with the subject lateral-directional oscillation. The lateral- directional characteristics of the helicopter are adequate to

8

FOR OFFICIAL OSE OMLY

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FOR OFFICIAL USE ONL perform the intended mission. However, as presently configured, this characteristic may cause some pilots tu jettison their sling loads prematur ly. The lateral-directional oscillation in the '50 to 60 KIAS band is a deficiency the correction of which is mandatory (PRS U7) Dynamic short period tests revealed an essentially deadbeat oscillation in both the lateral and directional axes (PRS-A3).

Static Lateral-Directional Stability

20. The static lateral-directional stability tests were conducted under the configurations and conditions listed in appendix III, and the test results are presented in figures 23 through 31, appendix IV. The test helicopter exhibited positive static lateral- directional stability, that is, right pedal for left sideslip and vice versa. The neutral to slightly positive lateral cyclic gradient is indicative of limited effective positive dihedral; however, this characteristic presented no problem to the pilot. The gradient of directional control position with sideslip angle is strongly positive and indicates good apparent directional stability characteristics. Steady heading sideslips to the left at 100 KCAS were restricted to 20 degrees due to contacting the right directional limit. The linear variation of bank angle with sideslip angle is advantageous and reveals a linear side force characteristic. The longitudinal control gradient reveals a significant nose down- moment during left sideslip and a slight nose up-moment during right sideslip. At higher airspeeds the pitching-moment characteristic becomes more pronounced. In left sideslip at 100 KCAS and above, the nose down pitching-moment combined with the neutral lateral cyclic gradient resulted in a cyclic control position which was awkward for the pilot to control. During normal operational useage this condition should not be encountered; therefore, this characteristic presents no problem to the pilot. Static lateral-directional stability is suitable for operational use (PRS A3).

Static Longitudinal Stability

21. Static longitudinal, collective-fixed stability was evaluated in climbing flight during constant power (1300 shp) climbs through a density altitude of 5000 feet at various airspeeds around the best climb speed of 62 KCAS. Static longitudinal stability data are presented in figure 32, appendix IV. The static longitudinal gradient is slightly positive. This shallow longitudinal gradient coupled with nose up-pitch, which occurred when high power settings were applied, made stabilizing on a particular climb airspeed extremely difficult. Pitch attitude was the best pilot cue to desired airspeed. However, once stabilized in a climb it was not difficult to maintain the desired airspeed (PRS A3).

9

FOR OFFÎAl USE ONLY

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FOR OFFICIAL USE ONLY Longit.ud Lnal Control MoLiun

22. The variation of Longitudinal control motion with trim airspeed in love] flight with a mid eg and various gross weights is presented In figures Ti through 37, appendix IV. The longitudinal control position gradient is neutral LO slightly positive at airspeeds from 30 to 60 KCAS. At airspeeds above 60 KCAS the gradient becomes more positive. The neutral to slightly positive gradient at: the slower airspeeds effectively eliminates longitudinal control position as a cue to airspeed desired and forces the pilot to rely on pitch attitude as the only reliable reference with which to select a desired airspeed. These airspeeds are on the backside of the power required curve where no speed stability exists and compli- cates the pilot's task of stabilizing on a particular airspeed below 60 KCAS. Even though it is difficult to stabilize on an exact airspeed within this airspeed band, the aircraft can be flown through tills band with little pilot effort, and does not adversely affect mission accomplishment. At airspeeds above 60 KCAS where a positive stick gradient exists and speed stability is present, stick position is useable as a cue to airspeed desired (PRS A3).

Dynamic Longitudinal Stability

23. Dynamic longitudinal stability tests were conducted under the conditions listed in appendix III. The longitudinal SCAS effectively eliminates the long period oscillation. With longitudinal SCAS "OFF" the long period oscillation is not easily excited; therefore, it is not a problem to aircraft control. Once forced into a long period oscillation by trimming in level flight and then slowing the airspeed 15 KIAS and returning the controls to trim, the helicopter exhibited a divergent phugoid oscillation. Dynamic short period tests revealed an essentially deadbeat oscillation in the longitudinal axis (PRS A3).

CONTROL RESPONSE

Longitudinal Control Response

24. Longitudinal control response tests xîth SCAS "on" and SCAS "off" were conducted during OGE hover and stabilized forward flight using step inputs from approximately 1/2 to 1 inch. Tests were conducted under the conditions specified in appendix III. Longi- tudinal control response data are presented in figures 38 through 42, appendix IV. After initial longitudinal displacement the resulting angular acceleration was in the proper direction within 0.2 seconds. During SCAS "on" testing, pitch damping was satisfactory in all conditions tested. During SCAS "off" testing pitch damping was minimal, resulting in pitch rates which built rapidly but

10


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FOR OFFICIAL USE ONLY were not objectionable. The longitudinal control response and control power are satisfactory for operational use (PRS A3).

Lateral Control Response

25. Lateral control response characteristics were evaluated under the test conditions outlined in appendix III. The data were obtained in OGE hover and stabilized forward flight using step inputs of approximately 1/2 to 1 inch. The data are presented in figures 43 through 47, appendix IV. During SCAS "on" testing lateral step inputs produced satisfactory roll rates in both directions; roll rates to the right were slightly greater than roll rates to the left. With SCAS "off," roll rates built rapidly at an ever increasing rate as shown on figure A. Lateral control response and control power (SCAS "off" and "on") are satisfactory for operational use (PRS A3).

FIGURE A. TIME WISTORY OF RIGHT LATERAL INPUT MODEL 211 S/N N6256 N • HUEYTUG

GROSS WEl&HT-10,300 LB C.G.STATION- 129.7 IN. ROTOR SPEED- Z9B RPM SCAS OFF DENSITY ALTITUDE-e660 FT FLIGHT CONDITION - HOVER.

i!1

■0

I 10

I 10

S 0 k

^ '«I >»

K 6

I / r

0 0 2 0 4 0 6 0 6 I 0 1 2 1 4 1 & 1.8 TIME- SECONDS


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FOR OFFICIAL USE ONLY Directional Control Kesponse

26. Directional control response characteristics were evaluated under the conditions outlined in appendix III. The data were obtained in OGE hover and stabilized, forward flight using step inputs of approximately 1/2 to 1 inch. The data are presented in figures 48 througn 53, appendix IV. At a hover, step Inputs in the directional axis produced acceptable yaw rates to the right with angular acceleration in the proper direction within 0.2 seconds after control displacement. Step inputs to the left, with SCAS "on," were characterized by acceptable initial yaw rates which quickly approached zero rate as shown in figure; ß.

FIGURE B TIME HISTI ..V OF LEFT DIRECTIONAL INPUT MODEL 211 S/N N625GN

aCffi5SS5Sitco

0.2 0.4 0.6 0.8 1.0 \Z TIME-SECONDS

1.4- 1.6 1.8 2.0

12


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FHP OFP""'!. fSE ONLY Th'.1 Inability to generatf a constant yaw r:\ic iu the left wliilc at a hover is a shortcoming, the correct uf whicli is desirable (PRS

A5).

SlUF.WrVRD ANÜ RF.ARWARJ)_!■ LI}.'•}IT

27. Sideward and rearward flight was evaluated under the condl- Lions outlined in appendix 111. The variation of directional and literal control positions versus airspeeds in sideward flight is presented in figures 54 through 58, appendix IV. The gradient of Lateral cyclic control with airspeed was slightly positive throughout the airspeed band tested. The directional control gradient was positive. From zero to 15 K1AS it was difficult to stabilize at a constant airspeed and constant heading because the motion of the helicopter was characterized by random yaw oscillations which required large and rapid movements of the directional control. During sideward and rearward flight at airspeeds above 1J KIAS the helicopter was easily controlled. These tests were conducted during calm nonturbulent atmospheric conditions. During a sling load test at a gross weight of 13,700 pounds and at an approximate density altitude of 4000 feet, left sideward flight at airspeeds greater than 10 KIAS could not be achieved due to tail rotor torque limitations. The control margin at this condition was less than 10 percent. The limited control margin at these conditions is a deficiency, the correction of which is mandatory (PRS 17).

28. Rearward flight test results are presented in figures 59 through 62, appendix IV. While at a maximum internal loading condition the maximum rearward velocity achieved was 20 KTAS. The longitudinal control position gradient was positive from hover to 15 KTAS rearward, and then changed to a neutral gradient from 15 to 20 KTAS rearward. At 20 KTAS the margin of longitudinal control remaining was 40 percent. The rearward flight characteristics are satisfactory for operational use (PRS A3).

CONTROL MARGIN

29. During stabilized level flight at V , while at a mid eg and sea level condition, there remained 6 percent of forward longitudinal control. This insufficient longitudinal control margin is a shortcoming, the correction of which is desirable (PRS A6). Addi- tionally, the pilot was required to stretch uncomfortably forward in order to achieve the required forward longitudinal control for V flight. The force-trim feature at airspeeds greater than 125 KIAS was ineffective. At airspeeds greater than 125 KIAS the pilot was required to physically overcome longitudinal trim spring pressure to obtain the desired incremental airspeed change. The


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sively firing. An ineffective force trim aystum at liigh alrsijecds and die excessive forward contruJ travel at lii^li airspeeds are deficiencies, the corrections of which are mandatory (l'KS 1)7).

SI^LATHf POWER FA II.11RKS

30. Simulated power failures with a sling load of A5Ü0 pounds were conducted from stabilized, climbing and level flight at a gross weight of 13,000 pounds with a density altitude of 5000 feet. Table 2 summarizes the test results. Simulated power failures resulted in a minimum of pitch and roll attitude changes. For the airspeeds investigated, yaw-attitude change was observed 0.2 to 0.5 seconds after initiation of the simulated power failure. The initial and immediate yaw attitude change of approximately 5 degrees is an acceptable cue in alerting the pilot to an engine failure situation. Simulated power failures at higher torque values resulted in a more rapid decay of rotor speed. The rotor speed time decay interval was measured from 296 rpm to 280 rpm. The simulated power failure characteristics of the test helicopter are satisfactory for operational use (PRS A3). Additional testing at a light gross weight configuration and high power climb condition is recommended to further define flight envelope restrictions.

Table 2. Simulated Power Failure Characteristics,

Rotor Decay I Flight Airspeed Torque Time |

Conditions (KCAS) (%) (sec) 1

Level 60 34 2.25 j

Level 80 40 2.25 |

Climb 80 44 2.0

Climb 80 48 1.80 |

j Climb 80 52 1.70 j

SLING LOAD OPERATIONS

31. During the conduct of this test the four types of sling loads carried were as follows:

14

FOR 0FF!(m ÜSE OMLY

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a. Piggyback - 105 nmi lunvitzer M1U1A1 wiLli U) rounds of ammuniLion (6000 pounds).

b. Cones container with L'AUO pounds of haJlust for a LoLal of 39ÜÜ pounds.

c. Siiiiulaiod military veliicle - automobile! (1950 pounds).

d. Lead weights of varying dimensions (from 1000 to A500 pounds).

i-. A Lateral directional oscillation was experienced during a piggyback, sling load test as explained in paragraph 19.

33. Both a directional and a longitudinal oscillation were experienced while carrying the conex container and simulated military vehicle. The maximum useable velocity attained with the vehicle was 90 KIAS while 80 KIAS was the maximum useable for the conex container in smooth air. In light to moderate turbulence with SCAS "on," pitch oscillations transmitted bv the conex sling load resulted in increased pilot effort and limited the maximum useable speed to 50 KIAS (PRS AS). Dense objects such as lead weights presented no sling load problems. Because of the increased capability of this aircraft co sling load various items, further tests should be conducted to determine the optimum cable types, lengths, and rigging conditions for these items to reduce oscillations and possibly increase airspied limits.

STABILITY AND CONTROL AUGMENTATION SYSTEM

34. The SCAS, as incorporated in the prototype test helicopter, reduced pilot workload and was especially helpful during heavy sling load operations. The entire mission profile can be conducted with the SCAS inoperative; however, pilot effort approaches a maximum because of the high roll sensitivity and low roll-damping characteristic. Pilot induced oscillations (PIG) are very prevalent with SCAS "off." Commitments involving prolonged operations require a properly functioning roll channel (PRS A5). The pitch channel results in no significant reduction of pilot workload and is, therefore, not necessary for satisfactory operational use (PRS A3). The yaw channel as presently configured has insufficient gain to satisfactorily prevent yaw oscillations in slow speed flight (zero to 60 KIAS). This is especially noticeable during heavy gross weight, sling load operations. The yaw channel exhibits excessive gain in high speed flight (60 to 140 KIAS) which causes large yaw accelerations following gust disturbances with small yaw attitude changes. As presently configured the yaw SCAS is not useable. The yaw SCAS to provide proper gains to prevent yaw oscillations at all airspeeds and loading conditions is a shortcoming for which correction is desirable (PRS A5).

15


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MISCELLANEOUS

Co He dive Creep

33. During the conduct of the test, principally during periods of high vibration levels (2/rev), the collective control tended to creep upward. At V , flight the collective control had to be locked into position'by the collective friction adjustment to prevent inadvertent power changes. Correction of the collective creeping tendency is desirable for improved operational use.

Structures

36. During the APE the left-forward-engine mount failed (cracked rod end) , and the elevator-bellcrank-attaching bracket (located below the engine) fatigued and cracked. Prior to the Army test, the tail boom structure itself developed cracks which were re- paired and the tail boom was structurally reinforced. Recommend that the airframe area, surrounding and supporting the 1-55 engine and the tail boom structure with its mountings, be investigated for structural integrity prior to future Army testing.

Vibration

37. A 2/rev vibration is prevalent throughout the airspeed envelope. This vibration is a shortcoming and is especially noticeable and bothersome at high-density altitudes and at heavy gross weight conditions. Reduction in the 2/rev vibration is desirable for improved operational use.

Power Management

38. The rpm governor control characteristics of the test helicopter were undesirable. Continuous manipulation of the rpm governor beep switch was required during engine power output changes. This characteristic required an unusual amount of pilot attention. Correction of these engine-droop characteristics is mandatory for satisfactory operational use.

Drive Train Limitations

39. Full left directional control was restricted due to tail rotor gearbox torque limits. The last 10 percent of left directional pedal travel was not useable. Correction of the tail rotor dynamic drive system to permit full and effective pedal deflection during any flight condition is mandatory.

16

FOR OFFiCiAL USE ONLY

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Tor gut' Limit: er

■'tO. Due Lo the large pilot workload during hover at maximum gross weight, Che pilot and rhe copilot were unable to continuously monitor the engine torque. At this same flight condition, the power requirements approach the drive train limits on many sling load missions. To prevent inadvertent overtorque of the dynamic drive train components, installation of a torque limiter is mandatory.

Noise Level

't L. The noise level at V (140 KTAS) was excessive due primarily to vibration in die airframe (doors, etc) which made outside radio communication difficult. Correction of this shortcoming is desirable for improved operational use.

(iearbox Temperatures

42. The 42-degree tail rotor gearbox exceeded its temperature limits (166 F) twice during the test program. Once during flight at V (by 5-10 degrees F) with an OAT of 95 degrees F and once during high density altitude tethered hovering. The tendency of the 42-degree tail rotor gearbox to overheat is a shortcoming, the correction of whicli is desirable for improved operational use.

Power Source Limitations

43. The transmission mounted generator is not available. In its place a dual-source-hydraulic system has been installed. A standby generator for instrument flight rules (IFR) flight is not available. Correction of this deficiency is mandatory for an IFR flight capability.

Cargo Mirror

44. The cargo mirror was practically useless due to high airframe vibration levels during V flight and heavy gross weight/sling load operations. The mirror was useful only during the hookup sequence. Recommend that the cargo mirror be more rigidly secured to the airframe and in conjunction with the reduction of vibration levels a more serviceable mirror should result. Recommend remote controls be installed to allow for pilot adjustment of the cargo mirror in flight in order to monitor the oscillations of the sling load.

17


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GENERAL

45. Hover and leveJ flight performance is sufficient to accomplish the intended mission; however, an increased range capability is desirable.

46. Tail rotor control power of the prototype was not sufficient to accomplish the intended mission.

47. The structural integrity of the area beneath the T55-L-7 engine (engine mounts and control-bell-crank brackets) and the tail boom and mountings should be scrutinized closely prior to a production contract.

48. Correction of the deficiencies discovered during this APE coupled with the 200 shp increase in drive-train-torque limits of the design proposal should result in a superior performing helicopter.

SPECIFIC

49. Within the scope of this test, correction of the following deficiencies is mandatory for satisfactory operational use:

a. Lateral-directional oscillations in the 30 - 60 KIAS airspeed band (para 19).

b. Lack of sufficient directional control margin during high gross weight (14,000 pounds) and high density altitude (above 4000 feet) conditions (para 27).

c. Ineffective force trim feature at airspeeds greater than 125 KIAS (para 29).

d. Excessive forward position of the longitudinal control during V flight (para 29).

e. Poor static engine droop compensation characteristics (para 38).

f. Restrictions on the last 10 percent of left directional pedal travel (para 39).

19


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FOR OFFICIAL USE ONLY g. Lack of a torque limiter to prevent inadvertent overtorque

of the dynamic drive train components (para 40).

h. Lack of a standby generator for an IFR flight capability (para 43).

50. Correction of the following shortcomings is desirable for en- hanced helicopter operational suitability and mission effectiveness:

a. Inability of the directional control to generate a constant yaw rate to the left during hover (para 26).

b. Insufficient forward longitudinal control margin remaining at V cruise (para 29).

c. Inability of the SCAS yaw channel to provide proper gains in order to prevent yaw oscillations at all airspeeds and loading conditions (para 29).

d. Collective creeping tendency (para 35).

e. A 2/rev vibration throughout the airspeed envelope (para 37).

Excessive noise level at V„TT, (para 41). NE

g. Tendency of the 42-degree tail rotor gearbox to overheat (para 42)

20


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REC ATIONS

51. The deficiencies, corrections of which is mandatory, should be corn 'ted prior to a production contract.

52. Th' shortcomings, correction of which is desirable, should be correct prior to operational employment.

53. Fu ler testing of this model helicopter should include auto- rotatio . tests conducted at heavier gross weights using an external sli load to determine rate of descent variation with gross weight ira 17).

54. Fu ler testing of simulated power failures should be conducted a light gross weight and high power climbs to further define ght envelope restrictions (para 30).

55. Fu' .:ier testing should include evaluation of various cable lengths i.id types, and rigging procedures for optimization of the sling lo id capability (para 33).

56. The airframe area surrounding and supporting the T-55 engine and the tail boom structure and its mountings should be investigated for structural integrity prior to further Army Testing (para 36).

57. That the pilot should have the capability of adjusting the cargo mirror in flight in order to monitor the oscillations of the sling load (para 44) .

21


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APPENDIX I. REFERENCES

i. Letter, AMSAV-R-(EF), llq. US Army Aviation Materiel Command (USAAVCOM), Subject, "USAAVCüM TEST DIRECTIVE 68-46," 17 September 1968. USAAVCOM Project No. 68-46.

2. Test plan, USAAVNTA Project No. 68-46 "Army Preliminary Evaluation (Hueytug)," October 1968.

3. Report, USAAVNTA Project No. 64-28, "Engineering Flight Test of the UH-1B Helicopter Equipped with the Model 540 Rotor System Phase D," December 1966,

22


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APPEIMDIX IS. FLIGHT RESTRICTIONS AND OPERATING LIMITATIONS

AMSAV-R-EF 12 October 1968

SUBJECT: Safety of Flight Release for APE of BHC Model 211

Commanding Officer I'S Army Aviation Test Activity ATTN: SAVTE-P Edwards Air Force Base, California

1. This letter constitutes a safety of flight release for an Army Preliminary Evaluation (APE) of the Bell Helicopter Company (BHC) Model 211 per USAAVNTA Test plan, "Flight Evaluation of the Bell Proposed Model 211," dated October 1968, in accordance with AVCOM Test Directive 68-46. Helicopter N6256N will be used for these tests .

2. Gross weight limitations are as follows:

a. Internal loadings are permissible up to a gross weight of 10,500 pounds, however, intentional power-off landings should not be performed above a gross weight of 10,100 pounds.

b. External loadings are permissible up to a gross weight of 14,000 pounds with a maximum sling load weight of 6,000 pounds.

3. Airspeed, altitude and sideslip limitations are specified in figures C thru F respectively. These limitations apply with the Stability Augmentation System operative or inoperative, with the cargo doors open or closed, and with or without the cargo mirror installed. The low speed operation of the helicopter in or near hovering flight (sideward and rearward flight) is limited as follows:

a. Sideward flight.

(1) Up to 10,500 pounds internal

(2) Up to 14,000 pounds external

b. Rearward flight.

23

30 knots true airspeed

15 knots true airspeed

FOR OFRCIAL USE ONLY

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AMSAV-R-EF 12 October 1968 SUBJECT: Safety of Flight Release for APE of BHC Model 211

(1) Up to 10,500 pounds external - 30 knots true airspeed.

(2) Up to 14,000 pounds external - 15 knots true airspeed.

Hovering turns In excess of A0 degrees per second should not be performed and rapid hovering turns and large rapid rudder pedal inputs should be avoided in order to preclude damage to the tail rotor drive system. Steady state or transient left pedal inputs within the left 10 percent of the total pedal travel should be avoided for the same reason.

4. The Maneuver limits are shown in figure F in terms of normal flight load factors. In addition, with an external sling load, a 30 degree bank angle shall not be exceeded. Maximum power climbs shall not be performed above an indicated airspeed of 100 KIAS re- gardless of altitude or gross weight.

5. The allowable center of gravity limits are as follows:

a. Internal configuration

Most Forward

Most Aft

b. External configuration

Most Forward

Most Aft

6. Rotor speed limits are as follows:

a. Power on minimum 280 rpm

Power on maximum 311 rpm

b. Power off minimum

Fus. Sta. 128

Fus. Sta. 135

Fus. Sta. 132

Fus. Sta. 134

Power off maximum

The drive system limitations are as follows

280 rpm (240 rpm transient)

327 rpm

Main Transmission 18,000 in lb (59% torquemeter reading)

24


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AMSAV-R-EF 12 October 1968 SUBJECT; Safety of Flight Release for APE of BHC Model 211

One 11me Limit - Operation at or above a tail rotor horsepower of 300 lip is cause for the removal of the 9Q level gears in the sump of the main transmission and inspection for scuffing or other damage.

b. 90 Tail Rotor Gearbox: 250 hp

Time - Power (Accumulated Fatigue Damage) - The time - power limits for the 90 tail rotor gearbox level gears are established as follows:

Accumulated Time Tail Rotor SHP

2 min 375

20 min 325

3.3 hrs 275

10 hrs 255

Endurance Limit 250

Opt ration in excess of this time envelope is cause for retirement of gears. In addition, any operation within 0% to 10% of full left rudder pedal will require powers in excess of the 250 HP Endurance Limit as previously indicated in paragraph 3.

c. Oil Temperature Limits are as follows:

110OC Main Transmission

42 Gearbox

90° Gearbox

Oil Pressure Limits are as follows:

Main Transmission

110OC

110OC

30 psi (min) 70 psi (max) (40-60 psi normal)

25


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AHSAV-R-EF 12 October 1969 SUBJECT: Safety ol Kllght Release fur MM-: uf BUG Model 211

8. I'he Lycoming T55-L-713 Engine Limitation.s are as follows:

RPM 98% (100% N = 18,720 rpm) 1

RPM See Rutor Speed (100% N = 15,330 rpm)

59Z (18,000 in/lb transmission limit) j

0c 816 Starting and Acceleration | 735° 30 minutes |

Oil Pressure psig 300-635 normal i 50-90 normal (90 max)

| Oil Temperature 0c 1380C max I

In addition, the allowable measured exhaust gas temperature during starting or accelerations shall be 816 C maximum not exceeding 5 seconds and 746 C for the remainder of the transient time.

9. This safety of flight release is contingent upon the maintenance of the aircraft being performed by the Bell Helicopter Com- pany. Since helicopter N6256N is neither military qualified or FAA certified at this time, all maintenance procedures and safety inspections beyond those listed in this flight release are the responsibility of BHC. Limitations imposed in this release are in no way an indication of the ultimate capability of the Model 211 but merely interim limitations pending further test and analysis.

FOR THE COMMANDER:

4 Incl as

CHARLES C. CRAWFORD, JR. Chief, Flight Standards Office

26


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FIGURE C. GROSS WEIGHT -AIRSPEED ENVELOPE MOPEL 211 5/N N6Z56 N

160i

WO

120

I 100

80

60

40

20

"N

/HTBfiMAL.

LOAD

NOTE: REDUCE LIMIT AIRSPEED 4 KNOTS P&R. IOOO FTA&OVE 3000 FT DENSITY ALTITUpe .

SXTSJUAMC LOAD

GROSS WEIGHT ~ IOOO LB 14

FIGURE D. GROSS WEIGHT • ALTITUDE LIMITS MODEL 211 6/N N6a56N

LOW SPEED FLIGHT LEVEL FLIGHT (SIDEWARD, REARWARD 6 HOVER) INTERMAL LOAD — EXTERNAL LOAD

10

^

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GROSS WEIGHT

IZ M-

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26A

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FIGURE E. SIDESLIP • AIRSPEED LIMITS MODEL 211 S/N N6256N

ZO 4^) 60 SO 100 120

AIRSPEED ~ KNOTS (CAS)

FIGURE F. ACCELERATION LIMITS MODEL 211 5/NN6256N

1.5

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APPENDIX III. PERFORMANCE TEST CONDITIONS

.Test

i I r o s s Weight Lb

Cg Lucalion

in.

Dens i Ly AltiLude

iL Leading

Ai rspeed calibrat ion 9500 132.0 3000 Clean

Level flight power requiret 7900 131.8 1400 Clean

ir 9570 132.0 950 Clean

9 390 132.0 30 50

10,^05 131.8 b230

10,450 131.9 9900

9500 132.1 2930

9450 132.0 10,100

7910 131.8 1500

9380 132.1 3300

12,740 131.8 48/u Conex container

13,750 131.9 1710 105iTm Howitzer M101A1 with 10 rounds of ammunition

Hover Minimum G.W. to Limit parameter 132.0 2110 Tethered hover

ii 4120 II

it 10,540 II

Autorotation 8000 132.0 5000 Clean

10,550 132.0 3000 Clean

27


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Gross eg Density 1 Wei glit Location Altitude

Test lb in. ft ! Loading

Static iongitudinai j 7900 i 131.80 1400 Clean stability ! 8085 ! 133.00 5000 Clean \

10,450 1 131.90 9900 1 Clean |

1 Dynamic iongitudinai j 8000 132.00 3000 Clean j

| stability 8200 132.00 1000 Clean { 10,550 132.00 3000 Clean j

Static lateral-directional 8075 133.69 5350 Clean

j stability 1 7835 131.75 4950 Clean ! 9585 132.12 5015 Clean j

Dynamic lateral-directional 8085 133.00 5000 Clean j

stability

Sideward flight 10,775 131.94 9855 Sling load 14,030 131.97 4545 Sling load 13,965 131.95 3745 Sling load 13,100 131.68 1200 Sling load

Rearward flight 10,715 131.91 9855 Sling load 13,100 131.68 1200 Sling load 13,965 131.95 3745 Sling load 14,030 131.97 4545 Sling load

Control response

I Longitudinal 1 7870 | 132.95 5155 Clean 1 10,350 j 129.74 2660 : Clean 10,485 129.82 2660 Clean

12,465 130.03 2590 Sling load

| Lateral 7785 132.93 5155 [ Clean |

10,305 I 132.93 5155 Clean i 10,485 129.82 2660 Clean |

12,465 130.03 2590 Sling load

| Directional 7725 132.91 5155 Clean \ 7775 132.92 6400 Doors open

mirror on

12,995 132.78 5480 Sling load 10,250 129.68 2660 Clean '| 10,385 \ 129.76 2660 Clean |

12,375 129.99 2590 Sling load

28


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APPENDIX IV. TEST DATA

t- IGURE TITLE

1 2

3 4 5 6 7 S 9

LÜ 11 1.2 13 U 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42

Engine shaft horsepower aval lab]e Spuci[icaLion fuel flow Nondimensional hovering performance OGE hovering ceiling Level flight performance

Nondimensional level flight performance ii ii n ii

Range performance Autorotational descent

II II

Airspeed calibration Inlet performance Static lateral-directional stability

II

II

II

Static longitudinal stability Control position trim curves

Longitudinal response

29

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43 44 45 46 47 48 49 50 51 52 53 54 55 56 57 58 59 60 61 62

Lateral response

Directional response

Control positions in sideward flight

ii

Control positions in rearward flight II n II II II

II

II

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APPENDIX V. TEST INSTRUMENTATION

The following test instrumentation was used throughout the conduct ot the test:

Performance (hand recorded) Engine torquemeter Sensitive rotor tachometer Outside air temperature Altimeter Calibrated airspeed (nose boom) Fuel used N tachometer (production) Exhaust gas temperature Compressor inlet temperature (A probes) Compressor inlet total pressure (4 probes) Main and tail rotor shaft torque Time

Stability and control (oscillograph recorded) Control positions Longitudinal cyclic Directional pedal position Lateral cyclic Collective control

Rate gyros Pitch Roll Yaw

Attitude gyros Pitch Roll

Accelerometers Center of gravity vertical Pilot vertical

Sideslip angle SCAS actuator positions Main rotor-flapping angle Rotor rpm Engine torque

93


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APPENDIX VI. PILOTS RATING SCALE

.ON'ROL L»ALE

C4PABLf Of BEING

COHIROUEO OR

H*N*0t0 IH C0MTE<T

je HlbSlOH. 11 fH

musBU PlLOI

ATTENTION

4CCEPIABL E

•■'IT HAVE

IE ICIENCIES WHICH

H4RRANT iWfSOVLMFNI.

BUI «DEQUAIf FOR

HIS5I0M,

PUOI IOHPENSAI ION.

If REpUIRED TO

ACHIEVE ACCEPTABLE

PERfORMANCE. IS

FEASIBLE.

UNACCEPTABLE

DEFICIENCIES WHICH

REgUIRE MANDATORY

IMPROVEMENT.

INADEQUATE PERFORMANCE

FOR MISSION EVEN WITH

MAXIMUM FEASIBLE

PILOT COMPENSATION.

,»I ■ ,. A c 10 R r

Mii ", All REQUlRfMfMTS

AND IXPECIAIIOHS. GOOD

INOUGH WlIHOUI

IMPROVEMENT

CLEARLY ADEQUATE FOR

MISSION.

UNSATISFACTORY

RELUCTANTLY ACCEPTABLE.

OEFICIENCIES WHICH

WARRANT IMPROVEMENT.

PERFORMANCE ADEQUATE

FOR MISSION WITH

ffASIBLE PILOI

COMPENSATION,

ULiUlSI. "KiHir DCjlRABl I

GOOD, PLEASANT. WELL BEHAVED

FAIR. SOME MILDLY UNPLEASANT CHARACTERISTICS.

GOOD ENOUGH FOR MISSION WITHOUT IMPROVEMENT.

SOME MINOR BUT ANNOYING DEFICIENCIES. IMPROVEMENT IS RtQUiSTFD.

EFFECT ON PERFORMANCE IS EASILY COMPENSATED FOR BY PILOI.

MODERATELY OBJECTIONABLE DEFICIENCIES. IMPROVEMENT IS NEEDED.

REASONABLE PERFORMANCE REQUIRES CONSIDERABLE PILOT COMPENSATION.

VERY OBJECTIONABLE DEFICIENCIES. MAJOR IMPROVEMENTS ARE KEEDED.

REQUIRES BEST AVAILABLE PILOT COMPENSATION TO ACHIEVE

ACCEPTABLE PERFORMANCE.

UNCONTROLLABLE

CONTROL WILL BE LOST DURING SOME PORTION OF MISSION.

MAJOR DEFICIENCIES WHICH REQUIRE MANDATORY IMPROVEMENT FOR

ACCEPTANCE. CONTROLLABLE. PERFORMANCE INADEQUATE FOR

MISSION, OR PILOT COMPENSATION REQUIRED FOR MINIMUM

ACCEPTABLE PERFORMANCE IN MISSION IS TOO HIGH.

CONTROLLABLE WITH DIFFICULTY. REQUIRES SUBSTANTIAL PILOT SKILL

AND ATTENTION TO RETAIN CONTROL AND CONTINUE MISSION.

MARGINALLY CONTROLLABLE IN MISSION. REQUIRES MAXIMUM AVAILABLE

PILOT SKILL AND ATTENTION TO RETAIN CONTROL.

»3

A6

A6

U7

U8

UNCONTROLLABLE IN MISSION.

U9

94


-"*"»«!»■*-'.'" ,


APPENDIX VII. CONTROL MOTION

Amount of contro] movement with percent travel of all flight controls is as follows:

Longitudinal Lateral Directional Collective

1% = 0, ,115 inches 1% = 0. .115 inches 1% = 0, ,0625 inches 1% = 0. ,1025 inches

95


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APPENDIX VIII. DISTRIBUTION

Agency

Commanding General US Army Aviation Systems Command ATTN: AMSAV-R-FT PO Box 2Ü9 St. Louis, Missouri 63166

Defense Documentation Center Cameron Station Alexandria, Virginia 22314

Test Plans

■Iquipmerit Failure Interim Final Reports Reports Reports

IS

20

96


■mmmmmwmmi üiüü '^»^tteto.teT.â«.'.^

mmmmimmm

UNCLASSIFIED FOUU Secunly t'ljssifu .itmi

DOCUMENT CONTROL DATA -R&D SvcurHy i («> w/n u/(u/i ../ (Itta. lunly ol nt^trat t itnJ imh-iiu,] cf fee ofiffrt-ti u/i.-n //»f ,.,i<,l\ rfport (.-. . tuh-itieäj

\û, M t t'OH I * I N A T N <i Ac

I'S Army Aviation Systems Test Activity Edwards Air Force Hast', California 'JTilTj

bt i UMIlT C \ AbSlt-lCAliOtJ

I'NCl.ASSIl IF!) POUO

) R E POW l

ARm' PRELIMINARY EVALUATION OF THE I'ROTOTYPF. 211 (HUEYTUC)

4 DE SC Rl P . i NO ' t 5 f T^-p«' iii rvfjort (tin/ inclumvf itatcs)

Final Report^ Septomher i c.'hR i hrnnph March 1969 1 »u r MORiSi (Fi/»( (mm«, widdU initial, lam name I

Theodore K. Wright;, I.TC, ARTY, US Army, Project Officer Ivar W. Kundgren, MAJ, TC, US Army, Project Pilot John I. Nagata, Project Engineer

o REPORT LjAIt

March 1969 7/1. TOTAL NO OF MAGLb 7h.HO OF R E F S

8a. CONTRACT O W G W A N I NO

b. PROjf-OT NO

9«, OHIGir-jA TOR'S HFFÔW I NUM B E R(S)

USAAVNTA Project No. 68-A6

USAAVSCOM Project No. hB-Ab ')b. OTHLR REPORT fJO(Sl (Any ulhcr numhors that muy be uti signed

this report)

N/A

10 DISTRIBoTION 5TATEWE.NT

This document is subject to special export controls and each transmittal to foreign government or foreign nationals may be made only with prior approval of the Commanding General, Hq, USAAVSCOM, ATTN: AMSAV-R-F, Main Office, P0 Bnx 209, St. T.mnc- Missouri ill6ft

V S U P P L E. M F N ^

j /

NOTES 12- SPONSORING MILITARY ACTIVITY

Commanding General US Army Aviation Systems Command ATTN: AMSAV-R-F F0 Box 209, St-. Touis, Missnnri 63166

X3 ABSTRACT

The Army Preliminary Evaluation (APE) of the Bell Model ^ll prototype helicopter (Hueytug) was conducted at the Bell Helicopter Test Facility, Arlington, Texas, Edwards AFB, California, and Bishop, California, from 19 October through 7 November 1968. Fly- ing qualities, performance, and mission suitability were evaluated to determine aircraft capabilities to carry six thousand pound sling loads at a takeoff gross weight of 14,000 pounds. Primary emphasis was directed toward the artillery mission of dis- lacing a 105mm howitzer MlOlAl with 10 rounds of ammunition and 3 cannoneers .-^ ""he elicopter had eight deficiencies which require mandatory corrections. Two of^hese are major design deficiencies that may require extensive engineering redesign. \ They are the directional oscillations in the 30 to 60 KIAS airspeed range, especially prevalent during heavy sling load missions; and lack of sufficient directional control margin during high gross weight (14,000 pounds) and high density altitude (above 4001 feet) conditions. The remaining six deficiencies are ineffective force trim feature at high airspeeds, excessive forward position of longitudinal control at high airspeeds, poor static engine droop compensation, tail rotor drive train torque limitations, lack of an engine power torque limiter and lack of a standby generator for IFR flight. There are seven shortcomings the corrections of which are desirable and should be accomplished as soon as possible. The prototype model 211 could marginally perform the 14,000 pound gross weight mission at sea level. At 4000 feet density altitude the marginal tail rotor control and transmission and drive train torque limitations prevented the helicopter from satisfactorily accomplishing the mission. Correction of the deficiencies discovered during this APE coupled with the 200 horsepower increase in drive train torque limits of the design proposal should result in a superior performing helicopter. Correction of the deficiencies should be accomplished prior uo a production contract.

DD FORM 1 N O V 6 ,1473 UNCLASSIFIED FOUO

i'lhi'iMmrni'ili'tllffhMMi^MT ■ .„.,wMJ<.m(.-ii«..,... -....Mfrjitv.Tm T-nVi' rftinini'iininiiil n■TiimnlMHi' ■îMMüM'^^^-'"'^"*^'^'"'^^''"^^"^'' - -i-iJtlBiliittiiiiiiMaaaiMajAMaiaia

UN l US SI KIM FOUO Security Classification

K E1 WO RDS

Bell Model 211 Prototype Helicopter Army Preliminarv Kvaluation Flying Qualities Performance Mission Suitability

UNCLASSIFIED FOUO Security Classification

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