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AD-AO09 714
ARMY PRELIMINARY EVALUATION II.MODEL 200 CEFLY LANCER
George M. Yamakawa, et al
Army Aviation Engineering Flight ActivityEdwards Air Force Base, California
August 1974
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1. REPORT NUMBER OVTTACCESSION No. 3. RtECIPIENT CATAL41G NUMBER
USAAEFA PROJECT NO. 74-2 1. J A O4. TITLE (and Subtitle) 5. TYPE OF' REPCRT & PERIOo coy E
FINAL REPORTARMY PRELIMINARY EVALUATION 11 27 April -15 May 1974
USAAEFA PROJECT NO. 74 217. AUTHOR(&) II. CONTRACT OR GRANT NUMBEk(a)
GEORGE M. YAMAKAWA, TOM P. BENSON,MMJ LARRY K. BREWER
9, PERFORMING ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT. PROJECT, TASK___________________________________________AREA &_______ WORK_____ UN__IT___NUMBERS____
US ARMY AVIATION ENGINEERING FLIGHT~ ACTIVITYEDWARDS AIR FORCE BASE, CALIFORNIA 93523
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US ARMY AVIATION ENGINEERING FLIGHT ACTIVITY AUGUST 1974 ___EDWARDS AIR FORCE BASE, CALIFORN7A 93523 13. NUMBER OF PAGES
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I0. SUPFý-EMENTARY NOTES
19. KEY WORDS (Conti'nue on reverse. aide it necessary and identify by block nun~bor)
Army Prelimninary Evaluatio-i
Model 200 CEFLY LANCER aircraftISuper KingAir Mode! 200Research and development tes,-bed. airrraftPerformance ttsts -~(continued)
20. ABSTRACT (Continsue on, rAver.. 21dv d necessary' end Identify by block nun ibe,)
The Urnited Slates ..rmy Aviation Engineering Flight Activity conductedPreliminary Evaluation (APE) It of the Model 200 CEFLY LANCER aircraft,manufactured by Beech Aircraft Cu.rporation, from 27 April to 15 May 1974 atthe Beech facility in Wichita, Kansas. During the test program, 7.3 productive hourswere flown. Performance, stability and control characteristics, and mifcellaneousengineering tests were conducted. During these tests, no additional deficiencies or
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19. Key Words
Stability and control testsMiscellaneous engineering characteristicsHeavy gross weight conditionsHigh temperature conditionsSingle-engine takeoff configurationAKi-weather aircraft capability
20. Abstract
shortcomings were noted. A deficiency determined during APE I and still presentduring APE II was the loss of power to primary attitude and heading gyros whenpropeller speed was less than 2000 rpm. It was not possible to duplicate theconditions under wldcn the deficiency of smoke in the cockpit and cabin areasat altitudes above 15,000 feet pressure altitude (determined during APE I) ,wasnoted; therefore, it was impossible to determine whether or not the contractormodifications eliminated this deficiency. Four previously determined shortcomingsremained uncorrected. The most significant shortcoming was the inadequate
single-engine performance under heavy grass weight or high temperature coiiditionsin the single-engine takeoff configuration.
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ERRATA
USAAEFA PROJECT NO. 74-21
ARMY PRELIMINARY EVALUATION II
MODEL 200 CEFLY LANCER
UNITEUL STATES ARMY AVIATION ENGINEERING TEST ACTIVITYEDWARDS AIR FORCE BASE, CALIFORNIA 93523
!. Page 25, photo C-1: The photograph was reversed in photo reprocessing.
Although this photo is not reprinted, readers arf. advised of this reverse pr.nting.
2. Figure in paragraph 3, page 32: Figure depicted below replaces figure presentedon page 32.
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dt
dtt
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3. Appendix D, page 40. Add Figure 1, Handling Qualities Rating Scale.
ADEQUACY KR SELECTED TASK AIRCRAFT DEMANDS ON THE PILOT PIO
IN SELECTED TASKPLTOR REQUIRED OPERATIONs' CHARACTERISTICS OR REQUIRED OPERATION" RATINC
PILOTI1ECSIO"~ Ratng Sale Ref ASA NI-S13 Mnd inimal deioopnsation oflvqght fh~ ador I
FiguT 1.ON HadlngQsita ratdn Sefracae.MM odrt
TABLE OF CONTENTS
INTRODUCTION
Background ........ ........................ ... 3Test Objectives ............. ...................... 3Description .............. ........................ 4Test Scope ............ ........................ .Test Methodology . . . . . . . . . . . . . . . . . . . . . 6 6
RESULTS AND DISCUSSION
General ............. ......................... 71Performance .......... ...........................
Genera! ............. ....................... 7
Level Flight Performance ..... ....................Climb Performance ........... .................. 8SSawtooth Climb ......... .................. 8
Continuous Climb .......... ......... .. . ... 11Handling Qtalities ................................. 11
Genera.ý ......... ....................... . 11Control System Characteristics ..... .............. .. 11
j Static Longitudinal Stability ..... ............... . .11Static Lateral-Directional Stability ................ ... 12Dynamic Longitudinal Stability ...... .............. 13Dynamic Lateral-Directional Stability ..... ........... 13
Dutcn-Roll Characteristics ..... .............. .. 13Spiral Stability Characteristics .... ............ .. 13
Stall Characteristics ........ .................. .. 14Single-Engine Characteristics ................. 14Ground Handlin? Characteristics .... ............. ... 14Instrument Flight Capability ...... ............... 15Aircraft SystemTs Failures ............... 15
Yaw Damper ...... ................... .. 15Rudder Boost ............................. 15Alternating Current Generator ................ ... 15750 Volt-Ampere Inverter .................. .. 16
Antenna Vibration Evaluatioa ...... ................ .. 16Human Factors ................................... 16
Cockpit Evaluation ...... ................... .. 16Noise......... ....................... . 16Toxicity ..................................... 17
Reliability and Maintainability ....... ................ 17
CONCLUSIONS
General . . . . . . . . . . . . . . . . . . . . . . . . . 19Deficiencies and 3hortcomings. ....... ................ 19Specification Compliance ......... .................. 19
RECOMMENDATIONS ........... ..................... 20
APPENDIXES
A. References ......... ........................ ... 21B. Description ............ ........................ 23C. Instrumentatim ........... ...................... .. 25D. Test Techniques and Dal. Analysis Methods ............. .. 31E. Test Data ........... ........................ .. 41
DISTRIBUTION
2
INTRODUCTION
BACKGROUND
I. In June 1971, a contract was initiated by the United States Army AviationSystermn- Command (AVSCOM) with Beech Aircraft Corporation (BAC) for theprocurement of three modified KingAir Model AI00 (U-2IF) aircraft. Theseaircraft were to be utilized by the United States Army Security Agency as researchand development test-bed aircraft in support of classified CEFLY LANCER missionrequirements. The original centract was modified in June 1973 to permit theprocurement of three "T"-tailed Model 200 aircraft in lieu of the modified U-21 F.These aircraft are currently being type certificated in the normal cat. -,ory of FederalAir Regulation (FAR) Part 23 (ref 1, app A) oi the Federal AviationAdminis;ration (FAA). Within this category the maximum gross weight may notexceed 12,500 pounds. The contractor has performed test and analysis whichpermits military qualification to extend the maximum gross weight to15,000 pounds with reduc-d maneuvering and airspeed limitations. This additionalgross weight capability was essential to the inclusion of all desired missionequipment for test-,e, purposes. In N~ovember 1973, the United States ArmyAviation Engineering Flight Activity (USAAEFA) was tasked by an AVSCOM testdirective (ref 2) to cor duct an Army Preliminary Evaluation (APE) on a prototypeBAC Model 200 aircrft. The APE was to be conducted in two phases. The APE itests were conducted from 20 Februa.; to 6 March 1974 at the Beech facilityin Wichita, Kansas, using the basic airplane without the external mission equipmentinstalled. The APE II tests were to Lz conducted with a mission-configured airplane.
TEST OBJECTIVES
2. The objectivc-t of APE I! were as follows:
a. Provide quantitative and qualitative engineering flight test da',a as neededto assist in substantiation of airworthiness at the 15,600-pound gross weight.
b. Verify contract compliance in appropriate areas.
c. Assist in determining the flight envelope to :m used for future test-bedflight operations.
d. Provide preliminary aircraft performance data at the military maximumgross weight for operational use.
e. Provide the data required to substantiate a statement of airworthinessqualification of the 3ircraft after mission provision modification.
DESCRIPIMN
3. The test aircraft were two BAC Moxiel 200's, serial numbers (SN) 71-21059and 71-21060, each powered by two United Aircraft of Canada, Limited (UACL)
PT-6A-I1 turboprop er4gmes. This aircraft is a prototype military version of theBAC Super KingAir Model 200 presmuized all-weather executive transport. Thepilot and copilot ame seated side by side with dual flight controls. The tricyclelanding par has dual wheels on each main gear and is retractable. The flight controlsystem is fully revessible. A pneumatic rudder boost is installed to help compensatefor asymmetrical thrust and a yaw damper system is prov;ded to improve directionalstability. Major differences between the civilian and military versions inciude theremoval of the flight director system, the autopilot, and the weather radar system-the addition of high-flotation landing gear and a fuel dump system; the installationof an 8.5-kilovolt-ampere (KVA) alternating current (AC) generator in a blisteron each engine 't- ;o provide additional electrical power for classified minionequipme':, ,CEFLY LANCER antenna army, and a 750 volt-ampere (VA) inverterto provide emergency power for the pilot attitude and heading gyro indicators(SN 71-21060 only). A detailed description of the Model 200 aircraft is containedin Beech Prime Item Development Specification BS22296A and AircraftProcurement Specification, Light Fixed Wing Reconnaissoance Aircraft (refs 3and 4, app A). Appenoix B conta•ia a furiher desc-iption of the test aircraft.
TEST SCOFE
4. The APE il tests were conducted at the BAC facility in Wichita, Kansas, from27 April to 15 May 1974. During the test program. 5 flights were conducted fora total of 9.6 hours, of which 7.3 hours were productive. Aircraft SN 71-21059was the primary test aircraft utLized for this evaluation. One flight was ronductedin aircraft SN 71-21060 to investigate the correction of specific deficiencies whichwere identified during APE i. The Model 200 aircraft was evaluated to determineperformance and handling qualities after the antenna array and mission equipmentprovisions were irstal!ed, and to verify contract compliance. Average test conditionsare shown in table I and test configurations are shown in table 2. Flight restrictionsand operating limitations applicable to this evaluation are contained in theoperator's mantual (ref 5, app A), as modified by the safety-of-flight releases (refs 6and 7).
Table 1. Test Conditiona..
Density Outside Air Gross Center-of-Gravity Trim -
CalibratedTeot Oescriptlon Altitude Tmiperature Weight Location Configuration(ft) (C) (ib) (in.) Airspeed
12,980 5.0 14,SO0 187.6 (fvd) 2110 to 210 CruiseLevel f light
performance 7200 13.5 13,920 187.0 (fid) 110 to 150 Single-enginecruise
8260 12.0 14,620 187.9 (fwd) 139
7320 14.5 14.780 188.0 (fwd) 170 Cruise
7430 15.5 14,940 188.2 (fwd) 204
8260 12.0 14.460 187.7 (fwd) 14AStatic Pover-appt oach
longitudiuial 7320 14.5 14,280! 187.5 (fid) '143stability
"800 8.0 14,360 197.1 (aft) 143
7770 8.5 14,700 197.2 (aft) 168---- '-I Cruise
7310 8.5 14,880 197.3 (aft) 202
7730 8.0 14.560 197.8 (aft)* 170
79s0 8.0 14,200 197.1 (aft) 142i , - iPower-approach
7600 9.5 14,040 197.1 (aft) 142
Statie 9040 6.0 13,800 197.2 (aft) 145 Cruise
lateral-direct tonal -
stability 8860 6.5 13,560 197.4 (aft) 144 Power-approach
Drutamiclongitudinal 10,760 4.5 14,220, 197.1 (aft) 170 Cruise
11,010 4.0 14,200 197.1 (aft) 140lOvnaml c . *JCruise
lateral-directional 11,020 4.0 i',.180 I 197.1 ' Jft) 170stability :i' . .11.02() 4.0 14,16O 197.1 (aft) 140 :Pover-approach
1 11,780 0.0 14,640 197.2 (aft) 1.15Vs CruiseS11,940 -1.0 14,800 197.2 (aft) I.5VS iv
____ P5 j ov r-approach
Stall 11,960 -1.0 14.900 197.) (aft) 1.15V Glidecharacterist ics S- - -. -" -- .S
11,940 j -1.0 14,1480 197.3 (aft) 1.15V Landing
11.66, 0.0 14,480 197.1 (aft) 1.15V Single-engine-.. . . . .. .. cruise
•11,660 0.0 ,14.480t 197.1 (aft) 6120
Single-engine 4-; Cruise
characteristics i11.640i 0.0 14,520 197.1 (.It) 140
_harcterlts -~I..7 ... 0- 14,440 - 197. 1(aft) 6120 Power-approach
'All tests periormed at propeller ,tpeedi of 2000 rpm.:Approzmatelj 10-kno' increrwnts.
ýTrlimmd at 3-degreo, a•lne of descent.%Center of gravity to simulate crensember In latrine."SVs: !tall airapeed.f'lnitial airspeed used to begin airspeed-for-minimum-control (V'%*) evaluation.
Si
Table 2. Airplane Configurations.
Land ing FlapConfiguration Gear Setting Power Setting
Position (%)
Takeoff Down 40 Takeoff
Cruise Zero Power for level flight
Landing I Down 100 Flight-idle
Power approach Down 40 Power for level flight
Glide Up Zero Power off, propellers feathered
Single-engine iPower for level flight on leftcruise Up Zero engine, power off and propeller
feathered on right engine
Maximum continuous power (MCP)Single-engine 40 left engine, power off andtakeoff propeller feathered on right
engine
TEST METHODOLOGY
5. Established flight test techniques and data reduction procedures were usedduring this program (refs 8 through 13, app A). The test methods are describedbriefly in the Results and Discussion section of this report. Flight test data wererecorded by hand from test instrumentation in the pi!ot and copilot panels, andfrom '.he photopanel. A detailed list ot the test instrumentation is contained inappendix C. Test techniques (other than the standard techniques described in theappropriate references), weight and balance methodology, and data reductiontechniques are contained in appendix D. A Handling Qualities Rating Scale (HQRS)(app D) was used to augment pilot comments relative to handling qualities.Airspeed calibrations were obtained from the contractor. Deficiencies andshortcoming ..e in accordance with the definitions presented in ArmyRegulation 70-10.
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RESULTS AND DISCUSSION
GENERAL
6. Performance and handling qualitles of the Model 200 aircraft wele evaluatedunder a limited variety of operating conditions with emphasis on operation in thenormal mission configuration (aft center of gravity) (cg) near the military maximumgross weight of 15,000 pounds. The test aircraft was compared to FAR Part 23(ref 1, app A), BAC Airworthmnss Qualification Specification 22301 (ref 14),military specification MIL-F-8185B(ASG) (ref 15), and military specificationMIL A-8806A (ref 16), to assist in determining future military applications. Nonew deficiencies or shortcomings were identified. Deficiencies and shortcomingsdetermined during APE I are contained in USAAEFA Final Report No. 74-21(ref 17). The deficiency of loss of power to mission equipment and primary attitudeand heading gyros when propeller speed was less than 2000 rpm remaineduncorrected. It was not possible to duplicate the conditions under which thedeficiency of smn'.e in the cockpit and cabin areas at altitudes above 15,000 feetpressure altitude (Hp) was noted during APE 1. Four previously determinedshortcr,;;.ings were uncorrected: lnad.quate 6ingle-engine performance capabilityunder aieavy gross weight or high tempt [riure conditions in the single-engine takeoffconfiguration; inability to maintain a trim airspeed within ±6 knots; fluctuatingtorque ind-icator needles; and leaking exhaust gases a'Jout the heated engine aiiinlet lips.
PERFORMANCE
General
7. The performance characteristics of the Model 200 aircraft were evaluated undervarious oplratin conditions with emphasis on operation in the normal missionconfiguration near the military maximum gross weight of 15,000 pounds at theforward cg .limit of fuselage station (FS) 188.3. Single-engine performancecapability under heavy gross weight or high temperature conditions was inadequatein the power-approach configuration and severely reduced the overall effectivenessof the aircraft. The inadequate single-engine performance of the aircraft under theseoperating conditions is a shortcoming (ref 18, app A).
Level 'ilight Performance
8. Level flight performance was evaluated at the conditions shown in table Ito determine the maximum level flight airspeed, cruise airspeed, range, andendurance capabilities. The zero thrust glide test method was used by the contractort(, obtain the base-line drag point for the aircraft. The drag polar of this aircraftw'th the external mission provision modification was then compared to the dragpolar of the clean aircraft used in APE I (fig. 1. app E). This comparison showed
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- V . . . .. ,,
a forty-drag count (ACD) (0.004) increase in parasitic drag over the clean aircral't.The constant pressure-altitude technique for determining the sitigle-engine (prop,'llerfeathered) and dual-engine drag polar was then performed to confirm thedegradation in performance due to the increase in drag of the antennas. The aircraftwas stabilized and trimmed at incremental airspeeds from the maximum airspeedfor level flight (VH) to VS.
9. Forty counts of drag were added to the drag polar equation of the cleanai, craft and were checked against the actual level flight test data from the modifiedaircraft. With the forty counts of drag confirmed, the drag polar equations of the Ivarious configurations tested in APE I were then modified to reflect the dragincrease. The aircraft performance was then calculated using the modified dragpolars and UACL engine performance data, which included installation and 1accessories losses. The results of these tests are presented in figures 1 through 5,appendix E. Aircraft specific range, recommended endurance, cruise airspeed, andVH in level flight for the cruise configuration are summarized in figures 6through 9. The level flight drag polar equations for the Model 200 aircraft withthe antenna array are presented in table 3.
10. At a maximum gross weight of 15,000 pounds at 15,000 feet, standard day,the maximum dual-engine airspeed in level flight using MCP was 251 knots true Iairspeed (KTAS). The recommended maximum range airspeed was 223 KTAS andthe recommended endurance airspeed was 175 KTAS. The maximum single-engineairspeed in level flight (right engine shut down and propeller feathered), using MCPon the left engine at 5000 feet, standard day, was 170 KTAS. The recommendedsingle-engine airspeeds for maximum range and endurance were 167 KTAS and151 KTAS, respectively. This aircraft is designed to be utilized as a research and Idevel'-ment test-bed aircraft, and no specific mission performance profile has been
designated.
Climb Performance
Sawtooth Climb:
! I. Dual and single-engine climb performance for the aircraft without the antennaarray were evaluated at the conditions shown in table 1, using the sawtooth-climbmethod of test. All dual-engine climb tests were conducted with both enginesoperating at MCP. All single-engine climb tests were conducted with the left engineoperating at flight-idle and the propeller feathered, while the right engine wasoperating at MCP. Zero sideslip was maintained for all tests. Forty counts of dragwere added to the climb drag polar equation of the clean aircraft for reasons statedin paragraphs 8 and 9. The climb drag polar equations for the Model 200 aircraftwith the antenna array are presented in table 4. Test results are presented infigures 10 through 17, appendix E.
8
r/
Table 3. Level Flight Drag Polar Coefficients. 1I
Number of ACD
Configuration Engines CD-- A B C
Operating 0 ACL 2 '
Zero 0.0365 0.04684 Zero Zero Zero
Cruise 1 0.0365 0.04684 0.857 0.070 -0.004
2 0.0365 0.04684 Zero 0.140 -0.0055+AC D C +A
'General drag equation: C D C +2 + AT + BT+C: D v° CL 2 L CCL
Where:
CD - Coefficient of drag
CD - Minimur, coefficient of drag of the propeller o
o feathered drag polar
"ACD-C-L 2 Slope of drag polar
ACL 2
CL Coefficient of lift
Tc' - Coefficient of thrust
A, B, C - Constants
Table 4. Climb Drag Polr Coefficients. 9Number of CD AC
Configuration Eagines D A B C0 2Operating ACLL L
Zero 0.0365 0.04684 Zero Zero Zero
Cruise 1 0.0365 0.04684 10.22917 0.0893 -0.00080
2 0.0365 0.0 D 0.0893 -0.0018
Zero 0.06918 0.0539 Zero Zero Zero
Takeoff 1 0.06918 0.0539 0.7500 0.0558 -0.01398
2 0.06918 0.0539 Zero 0.0558 -0.01398
W AC D 2+AT2 +CGeneral drag equation: C . CD + .,C C 2 + ATc2 + WT' + C
L
12. At a representative gross weight of 15,000 pounds, the aircraft has a calculateddual-engne rate of climb of 1735 feet per minute (ft/min) at the recomninendedbest-rate-of-climb airspeed of 133 knots calibrated airspeed (KCAS) in the cruiseconfiguration at sea level on a standard day. At the same conditions in the takeoffconfiguration, at the recommended best-rate-of-climb &irspeed of 120 KCAS, thecalculated rate of climb was 1370 ft/min. At a representative gross weight of15,000 pounds in the cruise configuration, the aircraft has a calculated single-enginerate of climb of 295 ft/min at the recommended best single-engine rate-of-climbairspeed of 127 KCAS at sea level on a standard day (I 5°C). At the same conditionsin the takeoff configuration, at the recommended best single-engine rate-of-climbairspeed of 106 KCAS, the calculated rate of climb was -93 ft/min. Single-engineperformance capability under heavy gross weight or high temp,ýrature conditionsis marginal in the cruise configuration and inadequate in the takeff configuration.The single-engine performance severely reduces the overall effectiveness of theaircraft under these operating conditions. The inadequate single-engine performanceof the Model 200 aircraft in the normal mission takeoff configuration is a
shortcoming. An Equipment Performance Report (EPR) concerning thisshortcoming was submitted (ref 18, app A).
10
Continuous Climb:
13. Dial and single-engine continuous climb performance were also calculated abstated in paragraph 11. The dual-engine service ceiling in the cruise configurationwas 25,900 feet density altitude (HD) at a g'-ass weight of 15,000 pounds (fig. 10,app E). The sirgle-engine service ceiling in the cruise configuration was 8190 feetHD at a gross weight of 15,000 pounds (fig. 14).
HANDLING QUALITIES
General
14. The handling qualities of the Model 200 aircraft were evaluated under a varietyof operating conditions, with emphasis on operation in the normal missionconfiguration near the military maximum gross weight of 15,000 pounds at theaft cg limit of FS 197.4. The test results were comnpared to MIL-F-8785B(ASG).No new deficiencies or shortcomings were determined. The loss 6f electrical powerto mission equipment and primary attitude and heading gyros when propeller speedwas less than 2000 rpm was a deficiency which remained uncorrected from APE I.Inability to maintain a trim airspeed within ±6 knots was a previously determinedshortcoming which remained uncorrected.
5Control System Characteristics
15. Control system characteristics were evaluated on the ground (longitudinally
only) and in flight at the conditions shown in table 1. Control forces were measuredon the pilot control wheel and rudder pedals. Breakout forces (including friction)were determined by recording the forces required to obtain initial movement ofthe controls. There were no significani changes from the results determined andpresented in the APE I report. Moderate departures from trim conditions (6 knots)did occur with the controls free, due to the friction band encountered at trimconditions throughout the airspeed envelope. Inability to maintain a trim airspeedwithin ±6 knots was objectionable and was previously reported as a shortcomingin APE i. An EPR concerning this shortcoming was submitted (ref 19, app A).
Static Lonkitudinal Stability
16. The static longitudinal stability characteristics were evaluated at the conditionsshown in table 1. The aircraft was trim, d in steady-heading, ball-centered levelflight at the desired trim aiipeed, then stabilized at incremental airspeeds greaterthan and less than trim airspeeds. The test results are presented in figures 18through 28, appendix E.
17. The elevator control force gradient, as indicated by the variation in elevatorcontrol force with airspeed, was positive for airspeeds above and below trim,indicating stab', static longitudinal stability. The elevator control position gradient,as indicated by variation in elevator control position with airspeeds in the forward
11
cg configuration, was positive, although shallow, for airspeeds above and belowtrim. However, the gradieiit in the normal mission configuration (aft cg) wasessentially neutral. The neutral elevator control position gradie~nt was notobjectionable, due to the *strong influence of the positive elevator force gradient.AIThe static longitudinal stability characteristics met the requirements ofFAR Part 23. However, these characteristics did not meet the requirements ofparagraph 3.2.1.1 of MIL-F-8785B(ASG), in that the elevator contiol positionAgradient h., the normal mission configuration (aft cg) is essentially neutral. Withinthe scope of these tests, the static longitudinal stability of this aircraft issatisfactory.
Static Lateral-Directional Stability
18. Static lateral-directional stability characteristics were evaluated at theconditions shown in table 1. The aircraft was initially trimmed for zero sideslipflight at the desired airspeed. The aircraft was then stabilized at incremental sideslipangles left and right, holding airspeed constant until attaining full rudder pedaldeflection or until reaching sideslip envelope limits. Test results are presented inIfigures 29 and 30, appendix E.A
rudde pttcied toal storcey w as poitdive te foreaitino sideslip angles bewew0deieethan19.de Staicdietoal stablity was poitdicate byteoaitino sideslip angles wiwe 0dereethanright from trim. A lightening of rudder pedal force occurred at sideslip anglesoutside this range at airspeeds below 170 KCAS. However, the rudder pedal force
k never reduced to zero, nor did it result in any unusual flight characteristics orobjectionable increases in pilot effort to maintain aircraft control. The variationof sideslip angle with rudd'!r pedal deflection was essentially linear for sideslipangles encountered up to full rudder pedal deflection (170 KCAS and below). Thestatic- directional stability failed to meet the requirements ofparagraph 23.1 77(a)(3) of FAR Part '23, in that the rudder pedal force gradientreverses prior to obtaining the full rudder control limit below 170 KCAS. However,the static directional stability did meet the requirements of MIL-F-8785B(ASG)and was satisfactory.
20. Dihedral effect, as indicated by the variation of aileron control displacementwith sideslip angle, was positive and essentially linear. Increasing aft displacementof the elevator control was required with increasing sideslip angles in both left
and right directions. The corresponding increase in elevator control forces withincreased sideslip angles was not objectionable. The side-force characteristics, asIindicated by the variation of bank angle with sideslip angle, were positive andessentially linear. The strong side-force characteristics provided excellent cues ofout-of-trim flight conditions to the pilot and enhanced coordinated flight whilemaneuvering (HQRS 2). Within the scope of these tests, the static lateral-directionalstability characteristics are satisfactory.
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IDynamic l~onitudinal Stability
21. The dynamic longitudinal stability characteristics were evaluated at theconditions shown in table 1. The long-term dynamic charz -teristics were evaluatedby slowing the aircraft with aft elevator control to an ai.,peed 30 knots belowtuim airspeed and then returning the control to the trim position (stick-fixed) orreleasing the control and allowing it to seek the trim position (stick-free). Short-termdynamic characteristics, simulating gust response, were evaluated by rapidlydisplacing the elevator control 1 inch from trim !or a duration of 0.5 second andthen returning the control to the trim position. The long-terni aircraft responsewas oscillatory and was lightly damped. The natural frequency was 0. 1167 radiansper second (rad/sec) and damped natural frequency was 0.1164 rad/sec. Thedamping rat'io was 0.076 and the period was approximately 55 seconds. This weakdamping combined with the large friction band contributes to the poor trimmabilityof the aircraft (para 15). There is r~o FAA requirement for long-term dynamics.and the long-term characteristics met the requirements of MIL-F-8785B(ASG).Within the scope of this test, the long-term dynamic characteristics are satisfactory.
22. The longitudinal short-term characteristics of the Model 200 aircraft wereessentially deadbeat for all test conditions, including flight in turbulent conditions.The short-term characteristics met the requirements of FAR Part 23 and ofMIL-F-8785B(ASG). For the conditions investigated, the short-term longitudinaldynamic characteristics are satisfactory.
Dynmic Lateral-Directional Stability
Dutch-Roll Characteristics:
23. The dynamic lateral-directional characteristics were evaluated at the conditionsshown in table 1. These tests were conducted by exciting the aircraft from acoordinated level flight trim condition with a rudder pulse and doublet, aileronpulse and doublet, and by release from a steady-heading sideslip. The Dutch-rolloscillations were lightly damped and easily excited with the yaw damper OFF.With the yaw damper ON, all oscillations were damped within four overshoots.The aircraft's lateral-directional response and controllability characteristics weregood in atmospheric disturbances with the yaw damper ON. However, considerablepilot compensation was required to overcome the sensitive gust response duringturbulent flight conditions at all cg locations with the yaw damper OFF (HQRS 5).The Dutch-roll characteristics met the requirements of FAR Part 23 andMIL-F-8785B(ASG). Within the scope of these tests, the Dutch-roll characteristicsof this aircraft with the yaw damper ON were satisfactory.
Spiral Stability Characteristics:
24. The spiral stability characteristics of the Model 200 aircraft were evaluatedat the conditions shown in table I. These tests were conducted by establishinga 20-degree bank (both left and right) from trim conditions (wings-level, zeroyaw-rate flight with the controls free) and timing the motion to a 40-degree bank
13
angle or recording the bank angle achieved after 20 seconds elapsed time. Spiralstability, as indicated by change in bank angle with el&.ýed time, was essertiailyneutral for both left and right turns and confirimed the test results report-d inAPE I. This aircraft possesses the capability of holding lateral trim in hands-off4flight for periods of time in excess of 20 seconds. The spiral stability characteristics Jmet the requirements of MIL-F-8785B(ASG) and are satisfactory.
Stall Characteritic.
25. Dual and single-engine stall characteristics of the Model 200 aircraft wereevaluated at the conditions shown in table 1. These tests were conducted byestablishing trim configuration at the desired airspeed and then making a slightpitch attitude increase and decelerating at a rate of approximately I knot persecoad until achieving a stall. Stall warning margins and recovery characteristicswere evaluated qualitatively. Test results were essentially identical to the results
determined in APE I. Within the scope of these tests. the stall characteristics aresatisfactory.
Sinkie.Engine Characteristics
26. The single-engine characteristics of the Model 200 aircraft were evaluated atthe conditions shown in table 1. These tests were conducted by establishing trimconfiguration at the desired airspeed and simsxlating sudden engine failure by movingthe left engine condition lever to the fuel cutoff position, itd by establishingsingle-engine trim conditions at the desired airspeed and slowly decelerating theaircraft to the VMC at which ei"!ter lateral or directional control could not bemaintained. Test results were identical to results determined in APE I. The Isingle-engine stall speeds were also determined to be the VMC for the testconfigurations of the Model 200 aircraft. The single-engine control characteristicsmet the requirements of FAR Pari 23 and MIL-F-3785B(ASG). Within the scopeof these tests, the sinole-engine characteristics ame satisfactory.
G ound Handling Characteristics
27. The ground handling characteristics of the Model 200 aircraft were evaluatedthroughout the conduct of these tests. The pitch attitude instability evident duringloading and ground operations of the aircraft in the normal mission configurationduring APE I was not present during this evaluation. The nose gear oleo strutpressure was decreased and the main gear strut pressures were increased to improvethe pitch attitude stability on the growid. The requirement to maintain a continuous2000-rpm propeller speed during ground operations in order to provide power foraiission equipment was reevaluated. This requirement was eliminated byincorporating a STANDBY operational mode in which the mission equipment couldbe placed following warmup. Other ground handling characteristics were identicalto those determined in APE I Within the scope of these tests, the normal groundhandling characteristics are satisfactory.
14
Instrument Flight apa( ity
28. A limited reevaluation of the instrument flight capability of the Model 200aircraft (SN 71-21060) was conducted to confirm the propei operation of the750-VA inverter in the event of the loss of AC generator power to the attitudeand heading gyros. During a simulated approach in 120-knot indicated airspeed(KIAS), 700-ft/min descent, the power levers were retarded to the flight-idleposition and the propeller speed dropped to 1900 rpm, resulting in the loss ofpower from the 8.5-KVA AC generators. The 750-VA inverter immediately begansupplying power to the primary attitude indicator gyro; however, heading gyroalignment was lost for 90 seconds. This was recorded a, h maintenance discrepimcyand prevented a further check for correction of the deficiency determined in AP' I(loss of the primary attitude amd heading gyros when the propeller speed is leIthan 200U rpm). Correction of the loss of heading gyro alignment during powersource transition should be demonstrated by the contractor prior to Armyacceptance flights. An EPR concerning the original deficiency was submitted(ref 20, app A).
Aircraft Svstems Failures
Yaw Damper:
29. Yaw damper system failure was simulated by switching the yaw damper systemOOFF. With the yaw damper system OFF, I-inch rudder control pulse inputs resultedin 8 tc 10 overshoots and long-term residual lateral-directional oscillations. Thelightly damped. easily excited residual lateral-directional oscillations which resultedwith the yaw damper OFF were objectionable but did not constitute a hazardto safe flight. The yaw damper system is required by the FAA for flights above17,000 feet, due to the weak lateral-directional damping characteristics discoveredin the commercial aircraft.
Rudder Boost:
30. Rudder boost failures were evaluated during the conduct of single-engine tests.Test results were identical to those determined during the conduct of similar testsin APE 1.
Alternating Current Generator:
31. The failure of one AC generator was simulated in flight by switching off theleft AC generator. Failure of the AC generator was indicated by a flashing MASTERCAUTION light on the glare shield and a steady light in the CAUTION panel.The opposite generator was capable of continuing to supply the necessary AC powerrequirements. The probability of a dual AC generator failure is remote. However,this condition was artificially induced by retarding the power levers to the flight-idleposition at airspeeds below 125 KIAS. This resulted in the immediate loss ofelectrical power to the primary attitude and heading gyros (para 28), with thesame results which were reported in APE 1,
15
750 Volt.Anire Inverter:
32. The failure of the 750-VA inverter was simulated by switching the inverterOFF. Failure was indicated by a flashing MASTER CAUTION light on the glare 'shield and a steady light in the CAUTION panel. Loss of this inverter had noeffect oua the aircraft instruments. However, in the event of the loss of powerfrom 'foth AC generators and the 750-VA inverter, the pilot would lose the useof the primary attitude and heading gyros, j
ANTENNA VIBRATION EVALUATION
33. The antenna &rray installed on the Model 200 aircraft was monitored visuailythroughout the test program for any vibration tendencies. No antenna vibrationswere encountered during large sideslip excursions (plus and minus ,-n estimated20 degrees) c: during dives to achieve the never-exceed airspeed (VNE) of245 KIAS. Within the scope of these tests, there are no objectionable antenna Ivibrations.
HUMAN FACTORS
Cockpit Evaluation
34. Results of this evaluation were idventical with the results determined in APE IwitV the exi-cotions noted below. A capability to adjust each lap belt and keepthe shoUlder strap attachment point to the lap belts centered in the pilot's lapwas provided on the test aircraft. The propeller feather range on the control consolewas properly marked on the test aircraft. A three-position push-to-talkcommunications -.witch was provided for both the pilot and copilot. The copilotwas able to monitor very-high-frequency (VHF) communications radios selectedon his signal distribation panel while his transmit-select switch was in the intercomposition on the airc,-aft. The communications cord to the pilot control wheel waspositioned so '1:t i.t did not interfere with the ice vane handles dtiring electrical4,Vensioti.
Noise
35. A limited noise level survey was conducted at various airspeeds using apropeller speed of 2000 rpm in the cruise configuration. Measilrements were takenusing the General Radio Type 1565-B sound level meter and utilizing the measuringprocedures contained in the instrument instruction manual (ref 21, app A). Thein-flight measurements are presented in table 5. It was determined that noise levelsare acceptable at the pilot/copilot stations.
16
Table 5. Cockpit Noise Level Measurement.
Trim !Indicated MIL-A-8806A Test Rasults (db)'adctdLimit_____%.irsp~ed
(kt) Scale A Scale B Scale C
140 106 91 95 98
170 106 93 97 100
200 113 98 103 105
'Level flight, cruise configuration, 2000-rpm propeller speed.
ToxMcity
36. Smoke in the cockpit and cabin areas at altitudes above 15,000 feet Hp wasdetermined to be a deficiency in APE I (ref 22, app A). During the brief evaluationconducted during this evaluation, it was not possible to duplicate the conditionsunder which smoke was detected during APE I. It was therefore impossible todetermine whether or not contractor modifications eliminated the deficiency.Co.-ection of smoke in the cockpit and cabin areas at altitudes above 15,000 feetHp should be demonstrated by the contractor prior to Army acceptance flights.
RELIABILITY AND MAINTAINABILITY
37. Factors affect~ng the reliability and maintainability of the Model 200 aircraftwere evaluated throughout the conduct of the flight test program. Evaluatedcharacteristics included ground support equipment, accessibility, interchangeability,servicing, fasteners, cables, connectors, &nd safety. Available contractor technicaldocuments, historical data, and current maintenance procedures were reviewed. Aqualitative evaluation was performed because the limited number of program flighthours minimized the opportunity to observe component repair and replacement.Formal removal or replacement tests were not performed. The aircraft was fullyinstrumented, which resulted in maintenance complications that s,-.'uld not existon an operational aircraft.
38. The items listed below are shortcomings originally dete mined during APE Iand which were unco-rected during this evaluation. These shcrtcomings will affectthe reliability and maintainability of the Model 200 aircraft. EquipmentPerformance Reports concerning these shortcomings were submitted (refs 23and 24, app A,.
11
a. T!,- torque needles continually fluctuated ±25 ft-lb durilng flight (fullscale was 2250 ft-lb).
b. Engine exhaust gases used to continuously heat the engine air i,•let lips
leaked at random locations about the periphery of thes lips, resulting in "2scoloredand blistered paint areas on the inlet cowling.
I
II
III
CONCLUSIONS
GENERAL
39. The following conclusions were reached upon completion of APE !i:
a. The Model 200 aicraft with the mission provision modification hasrediuced performance and ewentufly the 1me handling qualities characteristic asthe clean ahrcmft at the same ioading and cg conditions.
b. One deficiency and four shortcomings noted during APE I were stillpresent during this evaluation.
DEFCENCIES AND SHORTCOMINGS
40. The following deficiency remains uncorrected: Lost of the primary attitude Iand heading gyros when the propeller speed is less than 2000 rpm (pars 28).
41. The following shortcomings were still identified:
a. The single-engine performance of the aircraft in the normal missiontakeoff confliguration was inadequate (pan 12).
I Ib. The inability to maintain a trim airspeed within t6 knots (para 15).
c. The torque needles continually fluctuated ±25 ft-lb during flight(para 38a).
d. Engine exhaust pses used to continuously heat the engine air inlet lipsleaked at random locations about the periphery of these lips, resulting in discoloredand blistered paint areas on the inlet cowling (parn 38b).
SPECIFICATION COMPLIANCE
42. The static longitudinal stability characteristics failed to meet the requirementsof paragraph 3.2.1.1 of GIL-F-878SB(ASG), in that the elevator control positiongradient in the normal mission configuration (aft 1g) is essentially neutral (para 17).
43. The static directionp! stability failed to meet the requirements ofparapraph 23.177(aX3) of FAR Part 23, in that the rudder pedal torce pradientreverses prior to obtaining the full rudder control limit below 170 KCAS (pars 19).
19
RECOMMENDATIONS
44. The deficiency identified during this evaluation must be corrected (prars 40).
45. The shortcomings should be correctd (par. 41).
46. Correction of the lam of beading gyro alignmet during power source transitionmhould be den -estrated by the contractor prior to Army aceiptance flights(pms 28).
47. Correction of smoke in the cockpit and cabin areas &t altitudes above15,000 feet HP should be demonstrated by the contractor prior to Armyacceptance flight& (pmr 36).1
20
IAPPENDIX A. REFERENCES
1. Federal Aviation Administration, Federal Air Regulation FAR Part 23,Airworthiness Standards; Normal, Utility and Aerobatic Category Airplanes,13 March 1971.
2. Letter, AVSCOM, AMSAV-EFT, 5 November 1973, subject: Test Directivefor CEFLY LANCER Army Preliminary Evaluation.
3. Specification, Beech Aircraft Corporation, BS 22296A, BeechcraftModel A)00-1 Pressurized Turboprop Executive Transport 15,000 poundsMaximum Takeoff Gross Weight, 23 February 1973, revised 30 April 1973.
4. Procurement Specification, PS 3102, Aircraft Procurement Specification LightFixed Wing Reconnaissance Aircraft, 30 April 1973.
5. Operator's Manual, Beech Aircraft Corporation, Super King Air Model 200,
2 November 1973.
6. Message, AVSCOM, AMSAV-EFT, 141645Z, 14 February 1974, subject:Safety-of-Flight Release for CEFLY LANCER, A 100-1 Aircr ft for Pilot Trainingat 12,500 pounds Gross Weight.
7. Message, AVSCOM, AMSAV-EFT, 231735Z, 23 February 1974, subject;Safety-of-Flight Release for CEFLY LANCER, A 100-1 Aircraft for Gross Weightsup to 15,000 Pounds.
8. Flight Test Manual, Naval Air Test Center, FTM No. 104, Fixed WingPerformance, 28 July 1972.
9. Flight Test Manual, Naval Air Test Center, FTM No. 103, Fixt 'Wing Stabilityand Control, 1 August 1969.
10. Handbook, USAF Aerospace Research Pilot School, FTC-TIH-70-1001,Performance, September 1970.
!1. Handbook, USAF Aerospace Research Pilot Schonl, FTC-TIH-68-1002, '1Stability and Control, September 1968.
12. Flight Test Manual, Advisory Group for Aeronautical Research andDevelopment, Volume 4 Perforiiiance, Pergamon Press, Los Angeles, California,1959.
13. Flight Test Manual, Advisory Group for Aeronautical Research andDevelopment, Volume II, Stability and Control, Pergamon Press, Los Angeles,California, 1959.
21
-4
14. Specification, Beech Aircraft Corporation, BS 22301, AirworthinessQualification Specification, Beechcraft Model A100-1, 19 September 1973.
iI
15. Military Specification, MIL-F-8785B(ASG), Flying Qualities of PilotedAirplanes, 7 Apgust 1969 with Interim Amendment I, 31 March 1971.
16. Military Specification, MIL-A-8806A, Acoustical Noise Level in Aircraft;General Specification, 11 July 1966.
17. Final Report, USAAEFA, Project No. 74-21, Army Preliminary Evaluation I,
Model 200 CEFLY LANCER, June 1974.
18. Equipment Perfurmance Report, USAASTA, SAVTE-TB, EPR 74-21-21,"Army Preliminary Evaluation I, Model AIO0-1 CEFLY LANCER," 27 April 1974.
19. EPR, USAASTA, 74-21-18, 12 March 1974.
20. EPR, USAASTA, 74-21-3, 21 March 1974. 121. Manual, General Radio Company, "Instruction Manual, Sound Level Meter IType 1565-B," September 1971.
22. EPR, USAASTA, 74-21-4, 20 March 1974.
23. EPR, USAASTA, 74-21-12, 20 March 1974.
24. EPR, USAASTA, 74-21-22, 21 May 1974.
22
APPENDIX B. DESCRIPTION
I The Model 200 aircraft has the general structure and space arrangements ofthe BAC Super KingAir Model 200 aircraft. The test aircraft is shown in photo I.General specifications are listed below.
Dimensions
Wing span 54 ft, 6 in.Horizontal stabilizer span 18 ft, 5 in.Length 43 ft, 9 in.Height to top of vertical stabilizer 15 ftPropeller diameter 8 ft, 2.5 in.Propeiler/fuselage clearance 29.6 in.Propeller/ground clearance 14.5 in.Distance between main gear 17 ft, 2 in.Distance between main and nose gear 15 ft
Cabin Dimensions
Total pressurized length 264 in.
Cabin length, partition to partition 128 in.Cabin height 57 in.Cabin width 54 in.Entrance door 21.5 in. x 26.7 in.
Wing Area and Loading
Wing area 303.0 ft 2
Wing loading 4).5 lb/ft2
Power loading 8.8 lb/hp
Weights
Maximum takeoff weight 15,000 lbMaximum ramp weight 15,090 lbMaximum landing weight 13,500 lbMaximum zero fuel weight 12,500 lb
23
Ground Turning Clearance
Radius for inside gear 4 ftRaiu ornoewheel 19 ft, 6 in.
Radius for nousie ger21 ft, I in.Radis fo outide ear39 ft, l0 in.
Radius for wing tiP
2. A More detailed deacription of the test aircraft is presented in references 5
and 17, appendix A.
photo 1. Model 200 CEIFLY LANCER-.
24
APPENDIX C. INSTRUMENTATION IThe instrumentation in the BAC Model 200 aircraft, SN 71-21059, war nistalled,calibrated, and maintained by BAC personnel. hi addition to the instrumentationlisted, the aircraft was equipped with a pitot-static boom. Photos C-I and C-2show the instrument panel and photopanel. A list of test instrumentation showing 1the manufacturer, calibration range, ari,. parameter accuracies follows.
* .
'1j
* Ii-I
Photo C-I. Pilot Instrument Panel.
2
25
Iot I-.Iotpnl26I
C44
.4-e4
0 .0
el elo o o o - t
4-C 4- 0 04-
00 t ~0 t ~ o o4
(~ en en N N ~ N - . ~ - -7
-14.
_0 .0
00
0 0 00
0. 2 W 0-
-ý 0 4) 4 '
'0 '0 rA' ' I I- 0 0 '1%iA 61% -t = v1 i. 1, .
28 U
S 0.0 o0 .0
-~~~ - - - - q
o - -
%0 \0 10 %Q\- 1
0% 0% 0% 0% 0%
%0 '. 0 00 '. '. ' 0 '0 o. '.
- 0 S4 0b 0 0"
U2
In 0
00
0 0
qu( cEor.N
II~
03
APPENDIX D. TEST TECHNIQUES ANDDATA ANALYSIS METHODS
GENERAL
1. This appendix contains some of the data reduction and analysis methods tisedto evaluate the BAC Model 200 CEFLY LANCER aircraft. Although specifictests were not performed by the test team for propeller feathered glide andsawtooth-climb performance during APE 11, the propeller feathered glide techniquewas used. Beech Aircraft Corporation onducted the above test, including levelflight performance, to identify the additional drag caused by the antennainstallations. The test team confirmed the drag count by conducting single anddual-engine level flight performance tests. The drag count increment was then addedto the base-line drag polar of the clean aircraft (ref 17, app A), and the aircraftperformance data at conditions not specifically tested were predicted.
PERFORMANCE,
2. Past programs generally developed a drag polar relationship for specific flight
conditions. However, the test points showed large deviations from the faired lineat extreme altitudes (low versus high). The deviations are attributed to power effectswhich caused an apparent change in equivalent flat plate area (f) and Oswald'sspan ei-'iency factor (e) due to differences in engine thrust at varying altitudes.To elimiriatt, *hese effects, the propeller feathered glide method was used to evaluatethe Model 200 17EFLY LANCER aircraft.
3. The propeller feathered glide technique was used to define the base-line dragpolar. The aircraft was stabilized in a descent at a constant airspeed, with bothengines inoperative and propellers feathered. Airspeed, pressure altitude, outsideair temperature, gross weight, and elapsed time were recorded. The entire airspeedrange (.1IVS to VNE) was investigated for a target altitude band. The followingtechnique was used to develop the base-lne drag coefficient equation.
31
"10
e
L - W coq y (1)
D - T + W sin y (2)
DV - TV + WV sin y (3)
-V sin y - dh/dt TV DV
I
32
I ~ ~ ~~ ~Where:-.h Q ~ Z : n nL LiUft force (0b)
W Aircraft go weiglht (Vb)
*= ement an* (del)
T=Net thrust 0Ib) -- Zero in a descnt
D - Drag force = Level flight di8 (lb). Net thnust required
V - Aircraft velocity on descent path (ft/min)
dh/dt - Tapeline rate of descent (ft/min)
Cousideiring the drag and lift fore equations and applying power-off glide
conditions, the following relationship can be developed:
q9 (3)
CD qs (6)
CLmq (7)
CL (8)
Where:
CD - Coefficient of drag
q - 1/2 p V2 (lb/ft 2 ) dynamic pressure
s = Wing area (ft2 )
CL 2 COcfficient of lift
33
The base-line coefficient of drag (CDL) wa then developed by plotting CD veraCL 2 and fitting a first-order equation to the test points.
CD
SLOPE~
0 L_-C
o I_= CD + -
4. During powered 111Hght, the drag of the aircraft increased with thrust. To reflectthe change, the basic dSrag] equation was modified.
ACD PF - BL mC DpF -CDBL (10)
Where:&CDpF BL = Increased drag due to thmst effect
CDpF = Total coefficient of drag for powered flight
CDBL = Base-line coefficient of drag
34
Coefficient of thrust (Tc), thrust (T), thrust horsepower (THP), and shafthorsepower (SHP) were calculated as follows:
S2T
PSVT
550 x THPT VT (12)
Fn xVTTHP - n x SHP + 550 (13)
27r
SHP - Q x x 33,000 (14)
Where:
= Coefficient of thrust
T = Thrust (lb)
p = Air density (slug/ft 3 )
S = Wing area (ft2 )
VT = True airspeed (ft/sec)
THP = Thrust horsepower
rip =Propeller efficiency
SHP = Shaft horsepower;
Fn = Jet thrust (Ib)
Q = Engine torque (ft-lb)
Np Propeller speed (rpm)
35
The values of ACDPF . BL and TC' were then plotted to develop a generalized
equation that represented the change in drag due to thrust. A second-order fittingwis used.
D PF-BL.
Ca
AC T + B T + C (15)I DPF L - CI
From equation 10,
C =C +ACD DDPF BL DpF- BL
Or
C -C +A T12 + B T I + C (16)D D C CI
Equation 16 represents the generalized equation for all level flight and climbperformance in either single or dual-engine operation. When an externalconfiguration change is made to the aircraft that may affect its performance, th-propeller feathered glide test is the only flight required to find the difference indrag count.
36
I-OPE=CC
LOP£L
0
The slopes, ACD/ACL, for the basic and modified aircraft are identical. The CD,
of the modified aircraft contains the difference in drag which will be reflectedin equations 9 and 16.
5. Level flight performance tests (single an•d dual-engine) were conducted usingthe constant pressure altitude method. The aircraft was stabilized and trimmedat incremental airspeeds from VS to VH while maintaining a constant pressurealtitude throughout the entire flight. The coefficient of drag (CD), lift (CL), and ]thrust (TO' were obtained from the recorded test data to determine the coefficientsfor the generalized equations.
6. The shaft horsepower available, fuel-flow rate, and net thrust of a PT6A-41specification engine, including all installation losses, were furnished by the airframemanufacturer. The UACL-furnished computer program was used to calculate theperformance for an installed specification engine. The computer program is basedon the minimum performing engine that has the maximum allowable time beforeoverhaul. For this reason, the calculated aircraft performance data, which are basedon a specification engine, were always less than the observed test data. The testengines, SN X70012 and SN X70009, used for this evaluation were uncalibratedexperimental engines and the torque conversion factor for each engine was notavailable. The specification engine torque constant of 30.57 ft-lb per psi was used.The propeller efficiency table was furnished by BAC and is presented in table 1.
37
3 Bh4 AO~C qeIt POE.4rA7
AC7'I1-a-/r A.4C7-OR-2O -
Co
-04 -Y 6 - W-- -PO-w . --- Arm AWISA- f
-~ Oro .4M WO3A .79*2. 432 184 0*7 & W.W s .7? ."! j.07 .4 "M6-6 .- .2/1820.3 W13 .M -7M60
* 4 NO 7o3 .WO 0 .1m, fts .0U2 .wsD.19.
./0 .324 42.& JZ 7r? 7 f2 7.d4/2 N AW.RO .& .v "
Si/ -mNSP -~ SA(4.47 AiWdVROWC
Oý A- APAOQsdW &AFA- (epOW~)
//zý- 145- 0 ' IV- PR~L.CIC DA*Ar7-CqI (a. 2 08 f),OC Pe I- rMIYC A/f'SdTO (KhNITS)
J AD V,4A-Cd R'A 77
Table 1. Propeller Efficiency Table.I
38
7. Ambient test temperatures (Ta) were obtained by correcting the indicatedtest temperature (Ti) for instrument error (ATic) and for compressibility (ATC).
Ta - Ti + ATic + ATc
8. Pressure altitudes were obtained by correcting indicated pressure altitudes(Hpi) for instrument error (AHpic).
H- HP, + AHPIo
9. The density ratio (o) was determined from the following relationship:
a = To/T Pa/P0 a a 0
Where:
To = Standard-day, sea-level temperature
Po = Standard-day, sea-level pressure
10. The density altitudes were determined from the test density ratio (a test)and the US Standard Atmosphere, 1962 tables.
AIRSPEED CALIBRATION
II. The boom and ship's standard pitot-static systems were calibrated by thecontractor, using a low-altitude ground speed course to determine the airspeedposition error (fig. 3, app E). Calibrated airspeeds (Vcal) were obtained by.correcting indicated airspeed (Vi) for instrument error (AVic) and position error(AVpc).
Vca V£ + AVc + AlC
12. Equivalent airspeed was used to reduce the flight test data, as it is a direct
measure of the free stream dynamic pressure (q).
Ve = Vca + AVc
Where:
AVc is the compressibility correction, q 0.00339 Ve 2
39
13. True airspeeds (VT) were calculated from the equivalent airspeed and densityratio.
VVT e
Where:
a =Density ratio (R= where p is the actual ambient density)Po
14. Mach numbers (M) were determined from the test altitude absolute
temperature (Ta) in degrees Kelvin (K) and the true airspeeds.
Where: aVTaT =Sedo on nkos(897~
S.Teairat wpeigh afsond onidina got were96 deTeriea ro)t ahwih
and/or cg configuration change. Weighing was accomplished using electronic scaleslocated under the aircraft jack points with the crew on board at their designatedstations.
DYNAMIC STABILITY
16. Dynamic stability characteristics were tested by using the techniques describedin references 9, 11, and 13, appendix A. Analyses of the test data were performedto determine the resuilting damping ratios (ý) and damped natural frequencies (d).The damped natural frequencies and the damping ratios were derived by thelogarithmic decrement method.
17. The undamped natural frequencies (wn) of the motion in radians pcr secondwere calculated from the following equation.
WA (J d_
'~' ~i40
APPENDIX L. TEST DATA
INDEX
Fi~re Figure Number
Performance:
Level Flight 1 through 9Climb 10 through 17
Handling Qualities:
Static Longitudinal Stability 18 through 28Static Lateral-Directional Stability 29 and 30 I
Airspeed Calibration 31
'iS?
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