USAAEFA PROJECT NO. 81-01-5 IFUEL CONSERVATION EVALUATION OF US ARMY HELICOPTERS, PART 5, AH-1S FLIGHT TESTING .,_ LOREN L. TODD ROBERT A. WILLIAMS MAJ, IN CW4, AV PROJECT ENGINEER PROJECT OFFICER/PILOT RICHARD T. SAVAGE GARY T. DOWNS CPT, AR MAJ, AR PROJECT ENGINEER PROJECT PILOT U RICHARD L. VINCENT MICHAEL K. HERBST CPT, AR PROJECT ENGINEER PROJECT ENGINEER A -JANUARY 1983 A FINAL REPORT , F '" a LJ Approved for public release, distribution unlimited UNITED STATES ARMY AVIATION ENGINEERING FLIGHT ACTIVITY EDWARDS AIR FORCE BASE, CALIFORNIA 93523 S .2 hi
Transcript
USAAEFA PROJECT NO. 81-01-5
IFUEL CONSERVATION EVALUATION OFUS ARMY HELICOPTERS,
PART 5, AH-1S FLIGHT TESTING
.,_ LOREN L. TODD ROBERT A. WILLIAMSMAJ, IN CW4, AV
PROJECT ENGINEER PROJECT OFFICER/PILOT
RICHARD T. SAVAGE GARY T. DOWNSCPT, AR MAJ, AR
PROJECT ENGINEER PROJECT PILOT
U RICHARD L. VINCENT MICHAEL K. HERBSTCPT, AR PROJECT ENGINEER
PROJECT ENGINEER
A -JANUARY 1983
A FINAL REPORT
, F
'"
a
LJ Approved for public release, distribution unlimited
UNITED STATES ARMY AVIATION ENGINEERING FLIGHT ACTIVITYEDWARDS AIR FORCE BASE, CALIFORNIA 93523
S .2
hi
·•·
THIS DOCUMENT IS BEST QUALITY AVAILABLE. THE COPY
FURNISHED TO DTIC CONTAINED
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I 1lC I AAq T F T imSECURITY CLASSIFICATION OF THIS PAGE (BWen Date Entered)
READ INSTRUCTIONSREPORT DOCUMENTATION PAGE BEFORE COMPLETING FORM
I. REPORT NUMBER 2. GOVT ACCESSION NO. 3. RECIPIENT'S CATALOG NUMBER
USAAEFA PROJECT NO. 81-01-5 A 1_\ w ° -
4. TITLE (and Subtte) S. TYPE OF REPORT & PERIOD COVERED
FUEL CONSERVATION EVALUATION OF US ARMY FINALHELICOPTERS, PART 5, AH-IS FLIGHT TESTING 31 JUL 82 - 21 SEP 82
6. PERFORMING ORG. REPORT NUMBER
7. AUTHOR(e) 8. CONTRACT OR GRANT NUMBER(e)
LOREN L. TODD ROBERT A. WILLIAMSRICHARD T. SAVAGE GARY T. DOWNSRICHARD L. VINCENT MICHAEL K. HERBST
9. PERFORMING ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT, PROJECT, TASKAREA & WORK UNIT NUMBERS
US ARMY AVN ENGINEERING FLIGHT ACTIVITYEDWARDS AIR FORCE BASE, CA 93523 IK-1-DT043-O1-1K-1K
11. CONTROLLING OFFICE NAME AND ADDRESS 12. REPORT DATE
US ARMY AVN RESEARCH & DEVELOPMENT COMMAND JANUARY 19834300 GOODFELLOW BOULEVARD 13. NUMBER OF PAGESST. LOUIS, MO 63120 76
14. MONITORING AGENCY NAME & ADDRESS(It different from Controlling Office) 15. SECURITY CLASS. (of this report)
UNCLASSIFIED
1Sa. DECLA SSI FICATION/DOWN GRADINGSCHEDULE
16. DISTRIBUTION STATEMENT (of this Repoet)
Approved for public release; distribution unlimited.
17. DISTRIBUTION STATEMENT (of the abstract entered In Block 20, If different from Report)
IS. SUPPLEMENTARY NOTES
19. KEY WORDS (Continue on reverse side If noceeery and Identify by block number)
Advance Tip Mach NumberCompressibilityFuel Efficiency
2C. AuSIRACr (' mcet e am rve." hft i n cweary sad Identitf by block number)
The United States Army Aviation Engineering Flight Activity conducted levelflight performance tests of the AH-IS (Prod) helicopter to provide data todetermine the most fuel efficient operating conditions. Hot and cold weathertest sites were used to extend the range of the advancing tip Mach numberdata to supplement existing AH-IS performance data. Preliminary analysis ofnon-dimensional data identifies the effects of compressibilitA on performanceand shows a power penalty of as much as 6% at a high Nj/IV. The power
DD, •" 114 ,,1Q OF I NVSISOOLT
D J A 1473 LornOwOr NO S.SOLETE UNCLASSIFIED
SECUPITY CLASSIFICATION OF THIS PAGE (Whten Date Entered)
UNLSSFE -
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-required characteristics determined by these tests can be combined with
engine performance to determine the most fuel efficient operating
conditions.
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UNCLASSIFIED
S. SECURITY CLASSIFICATION OF THIS PAGE(Wh e
n Date ont*,ed)
1
DEPARTMENT OF THE ARMYHQ. US ARMY AVIATION RESEARCH AND DEVELOPMENT COMMAND
DA- 4300 GOODFELLOW BOULEVARD, ST. LOUIS, MO 6312
DRD)AV-D
SUBJECT: Directorate for Development and Qualification Position on the Final
Report of USAAEFA Project No. 81-01-5, Fuel Conservation Evaluation7of US Army Helicopters, Part 5, AH-IS Flight Testing
SEE DISTRIBUTION
.I1. The purpose of this letter is to establish the Directorate for Developmeuc
and Qualification position on the subject report. The report documents part 5of a 5 part effort which involves performance flight testing of the AH-IS toobtain performance data and determine the most efficient operatingcharacteristics. Part 1 involved conducting a flight operation improvementanalysis. Part 2 was initiated to develop and evaluate flight manual datadesigned for optimizing fuel conservation. Parts 3, 4 and 5 entail flighttesting of the UH-1H, OH-58C, and AH-1S which is specifically oriented towardsobtaining performance data applicable to fuel conservation. The part 5
evaluation conducted by the US Army Aviation Engineering Flight Activity(USAAEFA) consisted of obtaining detailed comprehensive performance data for
the AH-IS in both hot and cold temperatures. Future major revisions to theAM-IS Operator's Manual will consider incorporation of changes in performance
data or optimum performance operating procedures resulting from thisevaluation.
2. This Directorate agrees with the report conclusions and recommendations.
A. References ............................................... 9B. Aircraft Description .................................... 10C. Test Techniques and Data Analysis Methods ................ 13D.* Instrumentation ......................................... 20E. Test Data ............................................... 24
DISTRIBUTION
INTRODUCTION
BACKGROUND
1. The US Army is placing emphasis on achieving fuel conservationrelative to the operation of Army aircraft. The Deputy Chief ofStaff for Logistics (DCSLOG), Aviation Logistics Office/SpecialAssistant (DALO-AV) supports a program that will investigate waysto minimize fuel consumption. The Directorate for Developmentand Qualification of the US Army Aviation Research and DevelopmentCommand (AVRADCOM) and the US Army Aviation Engineering FlightActivity (USAAEFA) have jointly developed a fuel conservationprogram for helicopters which was briefed to Headquarters, US ArmyMateriel Development and Readiness Command (DARCOM) and DCSLOGon 28 January 1981. Both DCSLOG and DARCOM agreed to implementthe program. DSCLOG additionally agreed to provide necessaryOperation and Maintenance - Army (OMA) funding. AVRADCOM directedUSAAEFA to conduct the Cobra portion of the fuel conservationprogram on the AH-IS (Prod) (ref 1, app A).
TEST OBJECTIVE
2. The objective of this test program was to obtain flight testdata to determine the most fuel efficient operating characteris-tics of the AH-IS (Prod).
DESCRIPTION
3. The test aircraft, a production AH-IS serial number(S/N) 76-22573 (photo 1, app B) is a tandem seat, two-place heli-copter with two-bladed main and tail rotors. The helicopter ispowered by a Lycoming T53-L-703 turboshaft engine thermodynami-cally rated at 1800 shaft horsepower (SHP) at sea-level, standard-day conditions but limited by the main transmission to 1290 SHPfor 30 minutes at airspeeds below 100 KIAS and 1134 SHP for con-tinuous operation at any airspeed within the flight envelope.Distinctive features of this helicopter include the narrowfuselage, a modified flat plate canopy, stub wings with fourstores stations, a model 212 tractor tail rotor and the K747improved main rotor blades (K747 IMRB). A more complete descrip-tion of the aircraft is presented in the operator's manual(ref 2, app A).
TEST SCOPE
4. Level flight performance tests of the Ali-IS (Prod) helicopterwere conducted in the vicinity of Edwards Air Force Base,California, El Centro, California, and St. Paul, Minnesota.
4 ' " " - ' ' - ' " . . . . .. . . ., , " '
Project flying included 41 test flights which yielded 37.2 hoursof productive test time from the 77.8 total hours flown. Of thistotal, there were 18.1 hours of ferry flight. The test aircraftand the test instrumentation were maintained by USAAEFA.
5. Flight restrictions and operating limitations contained in
the operator's manual (ref 2, app A) and the airworthiness releaseissued by AVRADCOM (ref 3, app A) were observed. For this evalua-tion, maximum power-on main rotor speed was increased above thehandbook limit of 324 RPM to 329 RPM. All tests were conductedat a mid longitudinal center of gravity (cg) location, a slightlyright lateral cg, and with the engine bleed air OFF. Twenty-eightdata sets (level flight speed-power polars) were flown in theclean configuration. Eight additional data sets were flown withXM-159/C rocket pods installed (photo 1, app B). Four sets wereflown with the pods empty and the other four sets with the podsfully loaded.
6. The evaluation consisted of level flight performance testsusing referred rotor speed (NR//I) and thrust coefficient,CT, as the major variables to supplement the range of currentlyavailable data. General test conditions are shown in table 1.
TEST METHODOLOGY
7. Established engineering flight test techniques and datareduction procedures were used and are described in appendix C.Test methods are also briefly discussed in the Results andDiscussion section of this report. Test parameters were recordedfrom calibrated instrumentation by an onboard magnetic tapesystem installed and maintained by USAAEFA (photo 1, app D). Air-craft weight and balance measurements and fuel cell calibrationswere conducted by USAAEFA prior to the start of performancetesting. Airspeed and engine torquemeter calibrations utilizedare presented in figures 39 through 42, appendix E. Performancedata were not corrected for drag changes caused by addition of thenose boom system.
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ITable 1. Level Flight Performance Test Conditions
Average Average Average Average
NR//8 Average Gross WT Density Alt Air Temperature App E-RPM CT x b ft -.C Fig. No.
IClean configuration, zero sideslip, mid longitudinal cg, constant thrust coefficient and referredrotor speed method unless otherwise noted.2Flown with empty XM-159/C pods.3Flown with full XM-159/C pods.
b3
RESULTS AND DISCUSSION
GENERAL
8. This evaluation of the A-IS (Prod) helicopter obtained levelflight performance data to determine power required and fuelflow as a function of airspeed from approximately 20 knots trueairspeed (KTAS) to the maximum level flight airspeed. Theconstant referred gross weight and rotor speed (W/6,NR//8) method was used to obtain data at zero sideslip and amid longitudinal cg location. Additional data were obtained byaddition of two XM-159/C pods mounted inboard in the empty andfully loaded configurations. The power required characteristicsdetermined by these tests can be combined with engine performancedata to determine the most fuel efficient operating conditions.
POWER REQUIRED
9. The power required for level flight data were analyzed using
nondimensional power, thrust, and advance ratio (Cp, CT, andp), as described in appendix C. The matrix flown, shown intable 1, consisted of 35 sets of speed-power data that covereda range of CT x 10 from 50 to 68, and NR//r from 294 to 342 RPM.The baseline data in this evaluation were flown at a targetNR//- of 294 RPM at Edwards AFB. Increased N/Vr (324 RPMthrough 342 RPM) were flown at cold temperatures in St. Paul,Minnesota, where the maximum permitted power-on main rotor speedwas raised for test purposes from the handbook limit of 324 RPMto 329 RPM. The lower limit of 294 RPM remained unchanged.
10. A nondimensional summary of the results is shown infigures 1 and 2, appendix E, and dimensional data for the individ-ual tests (clean configuration) are presented in figures 3 through30. Table 1 provides a cross-reference of test conditions withfigure number.
11. For all values of 1j, the fairings of Cp versus CT infigures 1 and 2, appendix E, do not vary with referred rotor
'0speeds between NR/v'O-= 294 through 314 RPM. However, forincreasing values of p, a divergent trend from this baseline
appears for the highest NR/I§ flown (342 RPM) starting at= 0.22. As p increases above 0.22, the fairings for
NR/ f = 342 eventually form a separate family of curves withhigher values of Cp than those of the baseline. This sameseparate family of curves for NR/*8= 342 also begin to
t appear at p = 0.12 and is more pronounced for decreasingvalues of vi. Similar trends emerge for other referred rotorspeeds with increasing p: NR/Ir= 334 separates from the
4
baseline at p = 0.24 and 324 at 1 = 0.26. As 4 increases,
the divergent trends generally appear at lower CTs for thehigher values of NR//V and produce larger Cp increases athigher values of CT. From w = .14 to p = .20 the curves areidentical for all values of N //M Power required is affectedby compressibility above NR/A = 314 RPM in amounts which varywith CT and V. The difference attributed to compressibilityeffects between baseline and high NR/,6 amounted to 1.0% atCT x 04 = 50 and varied to as much as 6.0% at CT x 4 = 58.
ROTOR SPEED
12. The nondimensional summary shows a Cp penalty for
NR/6/- above 314 RPM as a function of CT and p. Thedimensional data must be used to determine whether performancefor varying gross weights, altitudes, temperature, and airspeedscan be maximized by reducing rotor speeds to operate at
lower NR/, . It should be noted that there is the possibilitythat decreasing NR/-r- beyond some condition will incur a
Cp penalty due to the increase of CT and P.
CONFIGURATION VARIATION
13. Eight sets of data were flown to compare performance of the
clean configuration with two XM-159/C pods installed inboardempty and fully loaded with dummy rockets (ten pound warheads).The data indicate an increase in drag of 3.3 square feet for the
empty pods and 5.5 square feet for the pods fully loaded. Figures31 through 38 show the data acquired in these configurations forvarious values of CT. This data compares favorably with data
taken in USAAEFA Project No. 66-06 (ref 4, app A).
ENGINE CHARACTERISTICS
14. Fuel flow characteristics of the Lycoming T53-L-703 engine
were derived from Lycoming Engine Model Specification computerprogram, US Army Model T53-L-703, Model Spec 104.43, datedI May 1974 (ref 5, app A) using installation losses described in
para 11, appendix D. Representative engine characteristics areshown in figures 40 through 42, appendix E. The consistency ofthese data indicate that engine power, based on the enginetorquemeter system, did not change throughout the program. The
engine, S/N LE12158Z was calibrated at Corpus Christi Army Depoton 30 June 1981.
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FUEL EFFICIENCY
15. Specific range (nautical air miles per pounds of fuel) was'. calculated for an installed specification engine for each of the
level flight performance tests and is shown in figures 3 through38. The fuel flow data for the cold weather tests were unreliableand therefore not presented. Specification values agree closelywith the measured data.
16. Maximum endurance occurs at minimum fuel flow rate, and bestrange at maximum nautical miles per pound fuel (specific range).Fuel efficiency for level cruise flight can be maximized at anytemperature by flying at the right combination of airspeed,altitude and rotor speed. These conditions can be determined byexamining the dimensional performance calculatd by combining theengine characteristics with the power required from the summaryin figures 1 and 2, appendix E. Aircraft performance and enginespecification data from this report should be combined withexisting data and presented in the operator's handbook in aformat suitable for use by the aviator.
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CONCLUSIONS
17. The power required characteristicb determined by these datatests can be combined with engine performance to determine themost fuel efficient operating conditions. Specific conclusionswere:
a. Power required is effected by compressibility aboveNR//'O= 314 RPM in amounts which vary with CT and ui (para 11).
b. Empty XM-159/C pods increase the flat plate area of theAH-S (Prod) by 3.3 square feet; fully loaded pods increase theflat plate area by 5.5 square feet (para 13).
- .
K'.7
K
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RECOMMENDATION
18. Aircraft performance and engine specification data from thisreport should be combined with existing data and presented inthe operator's handbook in a format suitable for use by theaviator (para 16).
8
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APPENDIX A. REFERENCES
1. Letter, AVRADCOM, DRDAV-DI, 30 June 1981, subject: Fuel
Conservation Evaluation of US Army Helicopters, Part 5, AH-IS
6. Engineering Design Handbook, Army Material Command AMC
Pamphlet 706-204, Helicopter Performance Testing, 1 August 1974.
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APPENDIX B. AIRCRAFT DESCRIPTION
GENERAL
* -1. The test helicopter, S/N 76-22573, was an AH-1S (Prod)(photo 1) with no modifications other than test instrumentation.
" The external configuration was slightly modified by the installa-tion of a pitot-static boom incorporating angle-of-attack, angle-of-sideslip vanes and a total air temperature probe. The boomwas iriunted below the telescope sighting unit at fuselage station(FS) 45. It extended forward of the nose 89 inches. The airtemperature probe was mounted aft of the boom mounts at FS 57.
POWER PLANT
2. The T53-L-703 turboshaft engine is installed in the AR-IShelicopter. This engine employs a two-stage, axial-flow freepower turbine; a two-stage, axial flow turbine driving a five-stage axial and one-stage centrifugal compresso ; variable inletguide vanes; and an external annular combustor. A 3.2105:1reduction gear located in the air inlet housing reduces powerturbine speed to a nominal output shaft speed of 6604.3 RPM at100 percent N2. A T7 interstage turbine temperature sensorharness measures interstage turbine temperatures and displaysthis information in the cockpit as TGT on the cockpit instruments.
Principal Dimensions
3. The principal dimensions and general data concerning theAH-IS (Prod) helicopter (photo 1) are as follows:
Solidity 0.0625Number of blades 2Planform Trapezoidal chord 30.0"
tapering to 10.0" at tip.Blade twist -0.556 deg/ftMain rotor speed 324 RPM (100%)
* 'Blade tie-down fixture is not included in the diameter.
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Tail Rotor
Diameter 8 ft, 6 in.Disc area 56.75 ft2
Solidity 0.1436Number of blades 2Blade chord, constant 11.5 in.Blade twist 0.0 deg/ftAirfoil NACA 0018 at the blade
root changing linearlyto a special camberedsection at 8.27 percent
of the tipTail rotor speed 1655.1 RPM (100%)
Fuselage
Length, rotor removed 44 ft, 7 in.Height:
To tip of tail fin 10 ft, 8 in.Ground to top of mast 12 ft, 3 in.Ground to top of transmissionfairing 10 ft, 2 in.
Width:Fuselge only 3 ftWing span 10 ft, 9 in.Skid gear tread 7 ft
Elevator:Span 6 ft, 11 in.Airfoil Inverted Clark Y
Vertical Fin:Area 18.5 ft2
Airfoil Special camberedHeight 5 ft, 6 in.
Wing:
Span 10 ft, 9 in.Incidence 17.0 degAirfoil (root) NACA 0030Airfoil (tip) NACA 0024Airfoil Inverted Clark Y
-II
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4j44
cc
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E-4
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APPENDIX C. TEST TECHNIQUES ANDDATA ANALYSIS METHODS
General
1. Conventional level flight performance test techniques wereused to conduct this evaluation (ref 6, app A). Speed-power datawere obtained in increasing increments of airspeed from 20 KIASuntil reaching an operating limitation (either VNE, transmissiontorque, TGT, or gas producer speed). Specific points would berepeated as judged appropriate by using an onboard plot of ind-icated torque versus airspeed for each point taken. All testswere conducted under nonturbulent atmospheric conditions topreclude uncontrolled disturbances influencing the results.Data were recorded on magnetic tape once a stable condition wasachieved, and each point used was held for 60 seconds.
Weight and Balance
2. Prior to testing, the aircraft gross weight and center-of-gravity location were determined with calibrated scales(electrical load cells placed under the aircraft jack points).The aircraft was weighed in the configurations flown withinstrumentation installed. The empty gross weight including fulloil and trapped fuel was determined to be 6647 pounds with
a longitudinal center-of-gravity (cg) at FS 201.6 inches andlateral cg at 0.1 inches right.
3. A manometer-type external sight gauge was calibrated and used
to determine fuel volume. Fuel specific gravity was measured witha hydrometer. The fuel loading for each test flight was determinedboth prior to engine start and following engine shutdown. Fuelused in flight was recorded by a totalizer and verified with thepre- and post flight sight gauge reading. Fuel cg based on fuelvolume contained in the fuel cell (259 gallon capacity) had beenpreviously determined, and this measurement was used as a basisto calculate aircraft cg for each test point. Aircraft grossweight and cg were also controlled by ballast installed at variouslocations in the aircraft. All tests were flown at a mid longi-tudinal cg location.
Level Flight Performance and Specific Range
4. The helicopter level flight performance data were generalizedby the following nondimensional coefficients:
SHP = Engine output shaft horsepower550 = Conversion factor (ft-lb/sec/shp)p = Air density (slug/ft )
Po Standard day sea level density (0.0023769 slugs/ft 3)
6 = Ambient pressure ratio (test point to sea level standard)A = Main rotor disc area (ft2) - 1520.53SI = Main rotor angular velocity (radian/sec) 2_ x RPM
60
R = Main rotor radius (ft) = 22.0W = Gross weight (lb)6 = Temperature ratio = (T + 273.15)/288.15T = Ambient air temperature (°C)1.6878 = Conversion factor (ft/sec/knot)VT - True airspeed (knot)a - Speed of sound (ft/sec) - 1116.45 /8
5. Each speed power was flown at a predetermined constant Or bymaintaining a constant referred gross weight (W/6) and referredrotor speed (NR//O). A constant W/6 was maintained by increasingaltitude to decrease ambient pressure ratio (6) as aircraftgross gross weight decreased due to fuel burnoff. Rotor speedwas also varied to maintain a constant NR//" as the ambient airtemperature varied.
6. Standard iterative carpet and cross-plotting techniques wereapplied to each set of data to provide smooth fairings in non-dimensional format and develop consistent families of curvescontinuous with each dimension (Cp, CT, and V). Sets of data for
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4-
each configuration and NR//W were independently processedin this way, followed by comparison with each other to identifytrends with referred rotor speed. Final adjustments to the fair-ings were made using combined data to arrive at a family ofnondimensional curves (fig. I and 2, app E) that summarize theentire matrix of test results and include effects of each param-eter varied.
V 7. Test-day (measured) level flight power was corrected to
presentation flight conditions for each speed-power data point byassuming the test-day dimensionless parameters Cp , Cr , and"" t t
.t are identical to Cp , CT , and us respectively.• 5
From equation 1, the following relationship can be derived:
(Ps)SHPs = SHPt (5)
~Pt
Where:
Subscript t = test day (measured for each data point)Subscript s s presentation day for each set of speed powerdata point
8. Test specific range was calculated using level flightperformance data and the measured fuel flow.
VTSR - (6)
Wi
Where:
SR - Specific range (nautical air miles per pound of fuel)VT = True airspeed (knot)Wf - Fuel flow (lb/hr)
Shaft Horsepower Required
9. The engine output shaft torque was determined from the engine;.j manufacturer's torque system using a calibration obtained at
Corpus Christi Army Depot on 30 June 1981 (para 14). The outputshp was determined from the engine output shaft torque androtational speed by equation (7).
10. Changes in the equivalent flat plate area due to change inaircraft configuration were calculated from the followingequation.
2ACp AA Fe = (8)
113
Where:
AFe = Change in equivalent flat plate area (ft2)ACp = Differential power coefficient (based on engine power)A = Main rotor disc area (ft2) = 1520.53
= Advance ratio
No power corrections were made to the data to account for noseboom drag.
Specification Fuel FlOe and Shaft Horsepower
11. Specification fuel flow and shaft horsepower were obtainedfrom Lycoming Engine Model Specification computer program, US ArmyModel T53-L-703, Model Spec 104.43, dated 1 May 1974 (ref 4,
app A). The installation losses used are described in reference 4,appendix A (USAAEFA Project No. 66-06). Engine accessory shplosses were assumed to be a constant 4 shp with customer bleedair losses computed at 0.9%. All computations were made for ableed air OFF condition.
Indicated Airspeed and Pesta-et Altitude
12. Airspeed and pressure altitude were measured from sensorsmounted on a flight test boom installed on the nose of the air-craft. The output signals were recorded on magnetic tape, andthe following expressions were used to calculate the parameters:
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Ia. Instrument corrected airspeed (Vic):
( r qcic 2/7 1/2Vic = ao 5 + 1) -i (9)
b. Instrument corrected pressure altitude (HPic):
[lPaic 1HPic = ) 1/5.255863 (6.8755856 x 10-6) (10)
Where:
Vic = Indicated airspeed corrected for instrument error (kt)ao = Speed of sound at standard day, sea level = 661.479 ktqci- = Indicated differential pressure corrected for instru-
ment error (in. Hg)Pao = Atmospheric pressure at standard day, sea level =
29.92125 in. HgHPic = Indicated pressure altitude corrected for instrument
error (ft)Paic = Indicated pressure altitude corrected for instrument
13. The boom pitot-static system was calibrated using the trailingbomb method to determine the airspeed position error. This cali-bration is shown in figure 1, appendix D. Calibrated airspeed(Vcal) was obtained by correcting indicated airspeed (Vi) usinginstrument (AVic) and position ( AVpc) error corrections.
Vcal = Vi + AVic + AVpc (11)
14. True airspeed (Vt) was calculated from the calibrated airspeedand density ratio.
Vcal
Where: (k)
a Density ratio Po
17
Corrected Pressure Altitude and Altitude Position Error
15. HPic was corrected for altimeter position error by using
AV c . The assumption was made that a pressure position error
, (App) was produced entirely at the static source. Since both
airspeed and altitude systems utilize the same static source, the
following relationships were used:
V 2 ] 3.5qc = .2( I Pao (13)
Pp qc - qcic (14)
Pa = Paic - APp (15)
pa a /5.255863
Hp=[1.0 (Pa) # (6.8755856 x 10- 6) (16)
Where :
qc = Differential pressure corrected for position and instru-ment error (in. Hg)
qcic = Indicated differential pressure corrected for instru-
ment error (in. Hg)
Vcal = Calibrated airspeed (knots)
ao = Speed of sound at standard day sea level = 661.479 knots
Pao = Atmospheric pressure at standard day, sea level29.92125 in. Hg
APp = Pressure position error (in. Hg)Pa = Atmospheric pressure at corrected altitude (in. Hg)
Paic - Indicated pressure altitude corrected for instrument
* 16. Static temperature was obtained by correcting the measured
total temperature for temperature rise due to compressibility.
The assumption was made that the temperature probe recovery
factor is equal to I.
[The following expressions were used:
TTic -OATic + 273.15 (17)
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TTicTa = (18)(qc 1)2/7
OAT = Ta - 273.15 (19)
Where:
OATic = Indicated ambient temperature corrected for instrumenterror (C*)
TTic = Indicated temperature corrected for instrument error(KO)
Ta = Static temperature (K0 )
Pa = Atmospheric pressure at corrected altitude (in. Hg)qc = Differential pressure corrected for position and
instrument error (in. Hg)OAT = Static temperature (CO)
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APPENDIX D. INSTRUMENTATION
1. The test instrumentation system was designed, calibrated,installed, and maintained by USAAEFA (photo 1). Digital andanalog data were obtained from calibrated instrumentation andwere recorded on magnetic tape and/or displayed in the cockpit.The digital instrumentation system consisted of various trans-ducers, signal conditioning units, a ten-bit pulse coded mod-ulation (PCM) encoder, and the Ampex AR 700 tape recorder. The
digital data were telemetered to a ground station for in-flightmonitoring. Time correlation was accomplished with a pilot/engineer event switch and on-board recorded and displayed Inter-Range Instrumentation Group (IRIG) B time. Various specializedtest indicators displayed data to the pilot and engineer continu-ously during the flight. A boom with the following sensors wasmounted on the nose of the aircraft: swiveling pitot-statichead, sideslip vane, angle-of-attack vane, and total-temperaturesensor (photo 1, app B). Boom airspeed system calibration isshown in figure iC. The ship's airspeed calibration is shown infigure 39.
2. Calibrated cockpit monitored parameters and special equipment
are listed below.
Pilot Station
Airspeed (boom)Airspeed (ship's system)Altitude (boom)Altitude (ship's system)Rate of climb (ship's system)Rotor speed (sensitive)Engine torqueMeasured gas temperature (TGT)Gas generated speed (N1 )
Fower turbine speed (N2 )Angle of sideslip
Outside air temperature (ship's system)
Event switch
Copilot/Engineer Station
Airspeed (boom)Altitude (boom)
Engine torque* Event switch
Fuel used (totalizer)Gas ge.erated speed (ship's system)Instrumentation controls and displays
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-, '-- .- .- --
Measured gas temperatureRecord CounterRotor speedTime of dayTotal air temperature (boom)
3. Parameters recorded on magnetic tape were as follows:
PCM Parameters
Airspeed (boom)Airspeed (ship's system)Altitade (boom)Altitude (ship's system)Angle of attackAngle of sideslipControl position
