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AD-AO09 713

ARMY PRELIMINARY EVALUATION I.MODEL 200 CEFLY LANCER

George M. Yamakawa, et al

Army Aviation Engineering Flight ActivityEdwards Air Force Base, California

June 1974

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1. REPORT NUMBER 2. GOVT ACCESSION NO. 3. RECIPIENT'S CATALUG NUMBER

USAAEFA PROJECT NO. 74-21 A'-)" A Ct 07 '/ c4. TITLE (and Subtitle) 5. TYPE OF REPORT &APERIOD CCVERED

FINAL REPORTARMY PRELIMINARY EVALUATION 1 20 February - 6 March 1974MODEL 200 CEFLY LANCER 6. PERFORMING ORG. REPORT NUMBER

USAAEFA PROJECT NO. 74-217. AIJTHOR(a) 5. CONTRACT OR GRANT NUMBER(,)

GEORGE M. YAMAKAWA, TOM P. BENSONMAJ LARRY K. BREWER, CPT MICHAEL A. HAWLEY

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US ARMY AVIATION ENGINEERING FLIGHT ACTIVITYEDWARDS AIR FORCE BASE, CALIFORNIA 93523 Not available

II. CONTROLLING OFFICE NAME AND ADDRESS 12. R!PORT DATE

US ARMY AVIATION ENGINEERING F!,IGHT ACTIVITY _ JUNE 1974

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18 SUPPLEMENTARY NOTES

19. KEY WORDS (Cow 1itme on reverse side i! rice-nary and identify by block number)

Model 260 CEFLY LANCER aircraft

Super KingAir Model 200 aircraftResearch and dcevelopment test-bed aircraftPerformance tests (continued)

20. ABSTRACT (Continue on reverse n:de If necess iry and Identify by block number)

The United States Army Aviation Engineering Flight Activity conducted an ArmyPreliminary Evaluation of the Model 200 CEFLY LANCER aircraft, manufacturedby Beech Aircraft Corporation, from 20 February to 6 March 1974 at the Beechfacility in Wichita, Kansas. During the test program, 24.8 productive hours wereflown. Performance, stability and control charact ristics, and miscellaneousengineering tests were conducted. During these tests, three enhancing characteristics,

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19. Key Words

Stability and control testsMiscellaneous engineering testsHeavy gross weight conditionsHigh temperature conditionsSingle-engine takeoff configurationAll-weather aircraft capability

20. Abstract

four deficiencies, and eighteen shortcomings were noted. The enhancingcharacteristics incl'ided the stall warning horn automatic readjustment with flapsetting to provide a computed warning at approximately 10 knots *ndicatedairspeed above stall, the rudder boost feature during asymmetric power conditions,and the excellent braking characteristics. The deficiencies determined were pitchattitude instability during loading and ground operations of the aircraft in normalmission configuration (aft center of gravity), the requirement to maintain acontinuous 2000-rpm propeller speed during ground and taxi operations, loss ofpower to mission equipment and primary attitude and heading gyros when propellerspeed was less than 2000 rpm, and smoke in cockpit and cabin areas at altitudesabove 15,000 feet pressure altitude. The mest significant shortcoming was theinadequate single-engine performance under heavy gross weight or high temperatureconditions in the single-engine takeoff configuration. Consideration should be givento the installation of an autopilot and weather radar system in pr,,duc .on aircraftto reduce pilot workload and enhance the all-weather capability of the aircraft.

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ERRATA

USAAEFA PROJECT 1O 4-21

ARMY PRELIMINARY EVALUATION I

MODEL 200 CEFLY LANCER

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

1. Pages 46 and 53. Photos captioned B3-3 and B-10 were transposed during theprinting process. Although these pages are not reprinted, readers are advised ofthis error.

2. Appendix D, page 74. Add Figure 1, Handling Qualities Rating Scale.

ADI)IIUA(Y FOR S[I.EC'fl1I) TASK AlR 15AH II1 %IA\ I S O N I II I I'l 01I II

)R RI QUIAREl) RA' 110 HIO'. I551 I15 RI ' OR KI )d Ii ON I~lsIlON' RIN

NII~lIII ONINUS 1111")11 1 SI KAlIIS I II,

I/I

I All! .......

TABLE OF CONTENTS

INTRODUCTION

Background ............... ........................ .Test O bjectives . . . . . . . . . . . . . . . . . . .Description ................. ........................ 4Test Scope ........................ ....................... 4Test Methodology ..... ...................... 7

RESULTS AND DISCUSSION

General .................... ......................... 8Performance ................... ....................... 8

General ............... ....................... . .. 8Takeolf'r and Lmnding Performance ..... .............. 9Climb Performance ......... ................... ... 9

Sawtooth (limb ............. .................. 9Continulous Climb.... ....................... ... 11

Level Flight Performance ...... ................ .. 11Stall Performance .......... ................... ... 12

Hlandling Qualities ............ ..................... ... 14General ............... ....................... ... 14Control System Characteristics ...... .............. ... 14Static Longitudinal Stability ...... ............... .. 16Static Lateral-Directional Stability ..... . . . .......... 6Dynamic Longitudinal Stability ..... .............. ... 17Dynamic Lateral-Directional Stability .... ........... ... 18

DIitch-Roll Characteristics .... .............. 18Spiral Stability Characteristics ..... ............ .. 19

Maneuvering Stability ....... ..................... 20Roll Performnvince Characteristics ..... ............. .. 20Stall ('haracteristics ........... .................. .. 21Single-Engine Characteristics ...... ............... 21Ground Handling Characteristics ..... ............. .. 22Takeoff and Landing Characteristics ..... ............ .. 25Trim Change Characteristics ...... ............... .. 26Night Operations ........... ................... .. 28InstrumeMt Flight Capability ...... ............... .. 29Aircraft Systems Failures ........ ................ .. 29

Yaw Damper .......... ................... .. 29Rudder Boost ................... 30Alternating Current Generator ........ ............ 30

Huim an Factors . . . . . . . . . . . . . . . . . . . . . . 30Cockpit Evaluation . . . . . . . . . . . . . . . . . . . 30N oise . . . . . . . . . . . .. . . . . . . . . . . . 33Toxicity .............. ....................... .. 34

Reliability and Maintainability ........ ................ ... 34

CONCLUSIONS

General............................ ......... 37Deficiencies and Shortcomings ........ ................ ... 37

RECOMMENDATIONS ............ ..................... ... 40

APPENDIXES

A. Reterences ................ ........................ .. 41B. Description .............. ........................ ... 44C. Instrumentation .............. ...................... ... 57D. Test Techniques and Data Analysis Methods .... .......... .. 65E. Test Data, .................... ........................ 74F. Definitions. Abbreviations, and Symbols ..... ............ .. 133

DISTRIBUTION

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INTRODUCTION

BACKGROUJNID

I In June 1971, a contract was initiated by tile United States Army AviationSyste us ('ommand (AVSCOM) with Beech Aircraft Corporation (BAC) for thepro,-uremnent of three modified KingAir A 100 (U-2 IF) airc'aft. These aircraft wereto be utilized by the United States Army Security Agency (USASA) as researchand development test-bed aircraft in support of classified CEFLY LANCER missionrequirements. The original contract was modified in June 1973 to permit theprocurement of three "T"-tailed Model 200 aircraft in lieu of the modified U-21 F.The, ' aircrwft are currently being type certificated in the normal category of FederalAir ,0,'ulation (FAR) Part 23 (zef I, app A) of the Federal AviationAdministri,'ion (FAA). Within this category the maximum gross weight may notexceed 12,5•.0 pounds. The contractor has performed test and analysis whichpernits military qualification to extend the maximum gross weight to15,000 j•Oinds with reduced maneuvering and speed limitations. This additionalgross weight capability was essential to the inclusion of all desired missionequipmcnt for test-bed purposes. In November 1973, the United States ArmyAviation Engineering Flight Activity (USAAEFA) was tasked by an AVSCOM testdirective (ref 2) to conduct an Army Preliminary l-valuation (APE) on a prototypeBAC Model 200 aircraft. The APE was to be cond-,,:ted in two phases. Phase Iwa. to be conduhcted with the basic airplane without the external mission equ pmentinstalled. Phase Ii tests were to be conducted with a mission configured ai-plane.

TEST OBJ ECTIV ES

2. The objectives of APF I were as follows:

a. Provide quantitative and qualitative engineering flight test data as neededto assist in substantiation of airworthiness at the 15,000-pound gross weight.

1. Verify contract compliance in appropriate areas.

c. Assist in determining the flight envelope to be used for future test-bedflight operations.

d. Provide preliminary aircrai't performance data at the niilitary maximumgross weight for operational use.

I3

I)ES(CllI)TION

3. The test aircraft was a BAC Model 200, serial number 71-21058, poweredby two United Aircraft of Canada, Limited (UACL), PT6A-41 turboprop engines.This aircraft is a prototype military version of the BAC Super KingAir Model 200piessurized all-weather executive transport. The pilot and copilot are seated sideby side with duai flight controls. The tricycle landing gear is retractable with dualwheels on each main gear. The control system is fully reversible, A pneumaticrudder boost is installed to help compensate for asymmetrical thrust and a ya-vdamper system is provided to improve directional stability. Major differencesbetween the civilian and military versions include thc removal of the flight directorsystem, the autopilot. and the weather radar system, the addition of high-flotationlanding gear and a fuel dunip system; and the installation of an 8.5-kilovolt-ampere(KVA) alternating current (AC) generator in a blister on each engine nacelle toprovide additional electrical power for classified mission equipment. A detaileddescription of the Model 200 aircraft is contained in BAC Prime Item DevelopmentSpecilication BS22296A and Procurement Specification PS3102 (refi I and 4,app A). Appendix B contains a further description and photographs of the testaircraft.

TEST SCOPE

4. The APE Phase I tests were conducted at the BAC facility in Wichita, Kansas,from 20 Febiutary to 0 March I1974. During the test program, 16 flights wereconducted for a iotal of 32.3 hours, of which 24.8 hours were productive. TheModel 200 aircraft was evaluated to determine performance and handling qualities,and to verify contract compliance. Average test conditions are shown in tables Iand - and test configurations are siown in table 3, The contractor takeoff andlanding performance tests were witnessed by test team personnel. Representativetakeoffs ind landings wcrc perfotrmed to spot-check contractor test results. Flightrestrictions and operating limitations applicable to this evahlation are containedin the operator's manual (ret" 5, app A) as modifieu by the safety-of-flight relases(refis 0 and 7).

I4

rable 1, Performance Teit Conditions.l,2

S~TrimDensity (Outside Air Gross Center-of-Gravity Callb d

Ttest Altitude Temperature Weight Location ConfrIdguratiDescription (ft) (

0C) (lb) (in,) Airspeed

itakeof f 2,000 17 15,o0(1 197.4 (aft) Not applicable Takeoff

Landing f 2,000( 17 13,500) 197.0 (aft) I Not applicable Landing

13,260 -29 14,180 187.3 (fwd) 110, 120,130, Cruise

-2 14,360 - 187. (fI) 30 140, 150,'. 160, 170, 18(0 Cruise

, 1-1 31 . w 100, 110, 120,Sawtootl 3,350 -8 14,340 187.3 (fwd) 130, 140, 1ofecIltimbs "130, 140, 150 ''ie~

climbs

3,210 --9 (fwd) 120, 130, 140, Singte-engine3'ý)0 1,80 1877 (wd)150

2,600 -5 15j,00 197.4 (aft) 110, 115, 120 ing1e-angine' ~tak~olff

Suirfacc to2 -40 14, 700 197.2 (aft) 140 I r Ise('ont iiluous ,4 (

"i imh12 ,001 1 tLI0-a in

12,0 -17 14, 300 191.1 (alt) 10, -12,9 (I :cruiLse

-8 11, 170 183.2 (fwd) 120, 160, 20o,' 240

8,8180 -14 11,060 183.1 (fwd) 10, 130, 140,!- _ _-- -.- - - - -I - - -_. - -1 -----

.,t'idethroat 8,460 -20 14,00m 187.5 (fwd) 110

lide

9,440 -10 14,820 '37.4 (fwd) 120

11,360 7 14,760 1U7.9 (fwd) 140

15, 310 -12 11,90o 183.4 (twd) '100 to 210

13,261) -29 14,601) 187.b (fwd) I 105 to 215

f 1 ght -- +--------Single-engine

10,220) -3 I 14,7501 193.3 (mid) '120 to 160 cruise andpowei approach

All tests performed at propeller speed of 2000 rpm.2Stall tests listed in table 6.

At approximately 10-KCAS (knots calibrated airspeed) incruments.

5

jI).It, "' [I IdI n Quil I tIes Test Condi t ions I ,o ,

lhens ity (hvtsidv Air Cross Cvnt er-of-Gravty C TrimtLo Dt I.prlt ion Alt it it'le 'limpe rat urv We ight Locat I01mte

(ti) (00) (Ib) (In.) A irsped(kt)

2,580 -13' 14,670 187.6 (rWd) 140

?.(6w - 1 1 14,780 18, .8 %(twd) 170 C r u Ise

2,620 -14 14,950 188.0 (fwd) 200

5t;it i," 1,900 -4 14,560 187.5 (fwd) 140 Power approach11' l,, i I i d Il I

gt ýhi I[L.v 5, iOt1 8 14,780 197.2 (;t)'L 140

5, 301) 8 14,88o 197.1 (1ft) 170 Crilse

5, 12o 7 14,990 19;.4 (a ft) 200

5,2f)0 7 14,6to 197.2 (at t ) 140 Power approach

9,640 -6 14,330 1N97.2 (aft) 140

,.It, I-1 I, ,d.re -h 14,78o 1947.2 (,aIt) 17f1 Cru iS'.

"10,,0 -h 14,9(1(g 197. I3 (lL) 2(0

,80 s-6 13,970 197.1 (aft) 140

Powe'r approach1(,2 14U:'. 14,19i 19;.k7..' (Ift 1740

0lI it 2 14, 9 (I 1 9,'. 1 t 1it) 140q':01.'l,, "

;ab- i t vit'1-'. r 14,4'0 W (a1 aIt) 200

1 40 4 14 ,220 1•) .) . if 1t 10

!,it, 1,1l-a 1 , , i- ,: 0l ,1 ) 4 1,, (1 1 ,1 I (aft) 171 rr 1se

11,01:' -5 14,,(1m) 197 . (aft) 200

(a14oo f t 1 40

I. , - 14,100 1'41." 1 (alt) 17(0 Cl1uis'

1 *14 '•1 -2 14 , (t,' 19) ". (rift) 200

111,611 I 1 3,82(1 197.2 (aI t ) 120 Power approach

IoI1,i. 02 1.)r,9h' 191.4 (aft) 200 Cruise

.1 II p-t . I-f or:.:v,(i 1)t 1i opol c r ieed o 2 )000 r pm,.I. a I Ls I i.-it d in talblte 6.

oilt ri,l sv.,%;ttm ind trim chanlgo ,haractv rl'steiic'; wLIr er•, lu,lttd oirilig tie conduct of other11), i in,. (11a I it i . [it

Table 3. Airplane Configurations.

Landing FlapConfiguration i Gear Setting Power Setting

SPosition (2)

Takeoff (TO) Dowl 40 Takeoff

Cruise (CR) i Up Zero Power for level flight

Landing (L) Down 100 Flight idle

Power approach (PA) Down 40 Power for lev~l flight

Glide M) Up Zero Power off,propellers feathered

Single-engine Up Zero Power for level flightcruise (SE CR) on left engine

Single-engine Down 40 Maximum continuous powertakeoff (SE TO) on left engine

TEST METOODOLOGY

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 pilot and copilot panels, andfrom the photopanel. Additional data were recorded on a 21-channel oscillographand by a motion picture camera located on the ground and in an F-34 Bonanzachase aircraft. A detailed list of the test instrumentation is co:ntained in appendix C.Test techniques (other than the standard techniques described in the appropriatereferences), weight and balance methodology, and data reduction techniques arecontained in appendix D. A Handling Qualities Rating Scale (HQRS) (app D) wasused to augment pilot comments relative to handling qualities. Airspeed calibrationswere obtained from the contractor. Deficiencies and shortcomings are in accordancewith the definitions presented in Army Regulation 70-10.

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RESULTS AND DISCUSSION

GENERIAL

6. Performance and handling qualities of the Model 200 aircraft were evaluatedunder a variety of operating conditions with emphasis cn operation in the normalmission configuratioii near the military maximum gross weight of 15,000 pounds.SThe test aircraft was compared toFAR Part 23 (ref I, app A), BAC AirworthinessQualification Specification 22301 (ref 14), and to military specificationMIL-F-8785B(ASG) (ref. 15) to assist in determining future military applications.Four handling qualities deficiencies were identil'ed. These included pitch attitudeinstability during loading and ground operations of the aircraft in normal missionconfiguration (aft center of gravity) (cg), the requirement to maintain a continuous2000-rpm :pr'1peller speed to perform ground and taxi operations, loss of powerto mission equipment and primary attitude and heading gyros .when propeller speedwas less than 2000 rpm, and smoke in the cockpit and cabin areas at altitudesabove 15,000 feet pressure altitude (tip) Eighteen shortcomings were noted. Thesingle-engine performauce capability tinder heavy gross weight or high temperatureconditions was inadequalte in the single-engine takcoff configuration. Othershortcomings included three stability and control shortcoraings, five cockpitevaluation short,:oniing3. and nine reliability and maintainability shortcomings.Three enhancing charactcristics were (I) the stall warning horn, whichautoniatically readjusted with flap settings to produce a cot-.puted warning atapproximately 10 knots above stall, (2) the rvdder boost, which greatly reducedpilot workload during asymmetric power conditions, and (3) excellent brakingcharacteristics.

PERFORMANCE

General

7. The perfoinanic characteristics of the Model 200 aircraft were evaluated tindervarious operating conditions with emphasis on operation in the normal amissionconfiguration near the military maximum gross weight of 15,000 pounds at theforward cg limit of fuselage station (FS) 188.3. Single-engine performancecapability tinder heavy gross weight or high temperature conditions was inadequatein the single-engine takeoff configuration and severely reduced tho overalleffectiveness of the aircraft. The inadequate single-engine performance of theaircraft under these operating conditions is a shortcoming.

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". ... .. ..

L

Takeoff and Landing Performance

8. Prior to the conduct of APE 1, the takeoff and landing performance testswere performed by the contractor at Fort Bliss, Texas (field elevation 3947 feet)and monitored by USAAEFA personnel. These result,, were spot-checked at theBAC facility at Wichita, Kansas (field elevation 1387 feet) at the conditions shownin table 1. Contractor-recommended takeoff and landing airspeeds were used duringthe conduct of these tests.

9. A maximurm performance takeoff was conducted to verify minimum groundrun distance, rotation speed, and lift-off speed. Acceleration was rapid upon brakerelease. Rotation was initiated at 85 knots indicated airspeed (KIAS) and lift-offoccurred at 95 KIAS. The ground run distance was 2600 feet, which was 250 feetless than that determined by the contractor for the test day condition. Withinthe scope of the tests, the takeoff performance data provided by the contractorwere satisfactory.

10. A maximum performance landing from a power approach was conducted toverify approach airspeed, touchdown airspeed, and minimum ground roll distance.Approach airspeed was 120 KIAS. Touchdown airspeed was approximately92 KIAS with 100 percent flaps. Full reverse wqs utilized immediately after thenose wheel contacted the ground. The ground roll distance was 1200 feet, whichwas 5 feet less than that determined by the contractor for the test day condition.Within the scope of these tests, the landing performance data provided by thecontractor were satisfactory.

Climb Performance

Sawtooth Climb: -

21. Dual and single-engine climb performance were evaluated at the conditions

shown in table 1, using the sawtooth-climb method of test. All dual-engine climbtests wer, conducted with both engines operating at maximum continuous power(MCP). A2 single-engine climb tests were conducted with the left engine operatingat flight-idle and the propeller feathered, while the right engine was operating atMCP. Zero sideslip was maintained for all tests. Test results are presented infigures 1 through 13, appendix E. The climb drag polar equations for theModel 200 aircraft without the antenna array are presented in table 4.

9J

Table 4. Climb Drag Polar Coefficients. 1

Number of ACD

Configuration Engines BOperating L

0 0.0325 0.04684 0 0 0

CR 1 0.0325 0.04684 0.22917 0.0893 -0.00080

2 0.0325 0.04684 0 0.0893 -0.0018

0 0.06518 0.0539 0 0 0

TO j 0.06518 0.0539 0.7500 0.0558 -0.01398

0.06518 0.0539 0 0.0558 -0.01398

AcD 2 2'General drag equation: C = C + "- C + AT C + BTC' + C

D D 0 C-L C Co AC L-.L

Whe re:

CD = Coefficient of drag

CD Minimum coefficienL of drag of the propeller

o feathered drag polar

ACD2 = Slope of drag polar

ACL

LC L =Coefficient of lift

Tc' = Coefficient of thrust

A, B, C = Constants

12. At a maximuni gross weight of 15,000 pounds, the aircraft has a positivedual-engine rate of climb of 1800 feet per minute (ft/min) at the recommendedbct-ratc-of-climb airspeed of 138 KCAS in the cruise configuration at sea levelon a standard day. At the same conditions in the takeoff configuration, the rateof climb was 1430 ft/min. At a maximum gross weight of 15,000 pounds in thecruise configuration, the aircraft has a positive single-engine rate of climb of

10

370 ft/min at the recommended best single-engine rate-of-climb airspeed of127 KCAS at sea level on a standard day (150C). At he same conditions in thetakeoff configuration, the rate of climb was -35 ft/miu. Single-engine performancecapability tinder heavy gross weight or high temperature conditions is marginalin the cruise configuration and inadequate in the takeoff configuration. Thesingle-engine performance severely reduces the overall effectiveness of the aircraftunder these operating conditions. The inadequate single-engine perfolmance of theModel 200 aircraft in the normal mission takeoff configuration (aft cg) is ashortcoming. An Equipment Perlformance Report (EPR) concerning thisshortcoming was submitted (ref 16, app A).

Continuous Climb:

13. Dual and s;ingle-engine continuous climb performance were evaluated at theconditions shown in table I. A dual-engine continuous climb was conducted usingMCP until full forward power lever settings were ottained at the engines' criticalaltitude. An abbreviated single-engine continuous climb was performed with tileleft engine shut down and the propeller feathered, The best-r-ite-of-climb airspeedschedule determined from the sawtooth-climb data was used.

14. The dual-engine climb airspeed schedule was easy to establish and maintainat all altitudes until passing through 25,000 feet Hp. Above this altitude, aircraftresponse to control inputs was sluggish and airspeed control demanded frequentattention. The dual-engine service ceiling was determined to be 28,400 feet densityaltitude (HI)) at a goss weight of 14,500 pounds. This *2xceeded the 27,750-footlID service ceiling predicted by the contractor by 2 percen'. A single-engine climbto service ceiling was initiated by intercepting the single-engine airspeed climbschedule at 11,500 feet. The single-engine service ceiling was determined to be12.910 feet 1id at a gross weight of 14,300 pounds. This exceeded tile 11,850-footHD service ceiling predi.:ted by the contractor by 9 percent.

Level Flight Performance

15. Level flight performance was evaluated at the conditions shown in table Ito determine the maximum level flight airspeed, recommended cruise airspeed,range, and endurance capabilities. The zero thrust glide test method was used toobtain the base-line drag polar for the aircraft. The aircraft was stabilized andtrimmed at incremental airspeeds in a descent with both engines inoperative andthe propellers feathered. The constant pressure altitude technique for single-engine(propeller feathered) --ad dual-engine drag polar was used. The aircraft was stabilizedand trimmed at incremental airspeeds from the maximum airspeed for level flight(VH) to the stall airspeed (VS). Performance at conditions not specifically testedwas calculated from the drag polar and power-available data, which includedinstallation and accessories losses. The results of these tests are presented infigures 14 through 19, appendix E. Aircraft specific range, recommendedendurance, cruise airspeed, and VH in level flight for the cruise cunfiguration aresummarized in figures 20 through 23. The level flight drag polar equations forthe Model 200 aircraft without the antenna array are presented in table 5.

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Table 5. Level Flight Drag Polar Coefficients. 1

, Number of ACDCur'figuration Engines 2 A

Operating o AL2

0 0.0325 0.04684 0 0 0

CR 0.0325 0.04684 0.857 0.070 -0.004

2 0.0325 0.04684 0 0.140 -0.0055

1 General drag equation: C. = C + ACD + AT2' + BT + C

D 0 C2 L C C

16. At a representative gross weight of 15,000 pounds at 15,000 feet, standardday, the maximum dual-engine airspeed in level flight using MC(l was 256 knotstrue airspeed (KTAS). The recommended maximum range airspeed was 225 KTASand the recommended endurance airspeed was 175 KTAS. Thie maximumsingle-engine airspeed in level flight (left engine shut down and propeller feathered),using MCP on the right engine at 5000 feet, standard day, was 180 KTAS. Therecommended single-engine airspeeds for maximum range and endurance were173 KTAS and 151 KTAS, respectively. This aircraft is designed to be utilizedas a research and development test-bed aircraft, and no specific mission performanceprofile has been designated.

Stall Perlormance

17. Stall performance was evaluated at the conditions shown in table 6. Thesetests were conducted by estabishing trim configuration at the desired airspeed,thien making at slight pitch attitude increase and decelerating at a rate of

approxiniately I knot per second until achieving a stall. Test results are presentedin table 6.

18. At a representative gross weight of 15,000 pounds, the Model 200 aircraftstall airspeed was 78 KIAS in the landing configuration, 79 KIAS in thepower-approach configuration, and 104 KIAS in the cruise configuration. Initialstall warning was provided by the stall warn?'g horn at approximately 10 KIASabove the stall airspeed. Additional warning was provided by a moderate buffetat approximately 2 KIAS in advance of the stall. A detailed discussion of stallcharacteristics is presented in paragraphs 36 through 38.

12

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....................

Wt4.4 4-4 4 144 14" 144 4,4. .1

q4Ca 0 0 0 0 0 0 0

0~' J C C) I)) C) 0 w 0 C) 4) cu Q) (V~4I' N fl) cn N en en N m c"" N N N N N

(u co Q)

m' oo 'A a, IC)It 0 'a va ~ ) V0 -T

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4" )

0j, Nl 4y\ 'a V4 It) a a a

C) ~~~ ~C C) C) CQC) C ) 0 ) C4"~ ~~~~~~~ ~~~~ >"". * ~ ' ) C a 0 . a ' ) N

ID 4 'a o) ma ' " a 4 I) I) I) I) ''d~-, -P44 '

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L4.0 .P..w r44 4 N1N .,4 M a4

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HANI)LING QUALITIEIS

General

19. 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 rounds at theaft cg limit of FS !97.4. The test aircraft was compared to MIL-F-3785B(ASG)to assist in determining future military applications. Three deficiencies wereidentified: pitch attitude instability during loading and ground operations of theaircraft in normal mission configuration (aft cg), the requirement to maintain acontinuous 2000-rpm propeller speed and perform ground and taxi operations, andloss of electrical power to mission equipment and primary attitude and headinggyros when propeller speed was less; than 2000 rpm. Three shortcomings identifiedwere poor long-term trimmability; elevator control force gradient reversal in thepower-approach configuration; and lightly damped, easily excited, Dutch-rolloscillations. Three enhancing characteristics were noted: the stall warning horn,which automatically readjusted with flap settings to produce a computed warningat approximately 10 knots above stall; the rudder boost, which greatly reducedpilot workload during asymmetric power conditions, particularly duringsingle-engine tests; and the excellent braking characteristics.

Control System Characteristics

20. Control system characteristics were evaluated on the ground (longitudinallyonly) and in flight at the conditions shown in table 2. 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. Control systý,m positions in trimmed forward flight are presented infigures 24 through 27, appendix E. The results of the control system evaluationare summarized in table 7. There was no detectable lag in aircraft response foreither small or large control input amplitudes along any control axis. Elevator andaileron lorce and displacement sensitivities were compatible and intentional inputsto one control' axis did not cause inadvertent inputs to another axis. Controlharmony was good and there was no tendency for the pilot to induce undesirablemotion. However, moderate departures from trim conditions (6 knots) did occurwith the controls free due to the friction band encountered at trim conditionsthroughout the airspeed envelope, and required moderate pilot compensation tomaintain adequate airspeed control (HQRS 4). The poor long-term trimmabilitywas objectionable and constitutes a shortcoming. An EPR concerning thisshortcoming was submitted (ref 17, app A).

14

Table 7. Control System Characteristics.'

Test Description Test Results

Control travel 6.6 in. from full forward to full aft

Breakout force4 lb (MIL-F-8785(B)ASG min: 0.5 lb; max: 4 lb)(including friction)

65 lb/in. Steep positive force gradient;Gradient however, mild departures from trim (±6 KIAS)

Longitudinal did occur with the controls free

(elevator) None noticeable. Does not result inFree play objectionable flight characteristics

Positive at any normal trim setting; friction

Centering prevents absolute centering

Control dynamics Well damped

1400 from full right to full left

(1 2 .85-in. arc from full right to full left)

Brea:-out force(including friction) 2 lb (MIL-F-8785(B)ASG min: 0.5 ib; max: 3 lb)

Lateral Gradient 9 lb/in.(aileron)

0.125 in. left and right from trim; does notFreeplayresult in objectionable flight characteristics

Centering Positive; friction prevents absolute centering


Control travel 7.14 in. from full left to full right

Breakout forceBrincludn forict ) 13 lb (MIL-F-8785(B)ASG min: 1 lb; max: 14 lb)(including friction)

Directional Gradient 90 lb/in.

(rudder) None detected. Does not result in objectionableFree play flight characteristics

Centering Positive


'Evaluation conducted at 170 KCAS.15

Statir lonzitludinal Stability

2 I. The static longitudinal stability characteristics were evaluated at the conditionsshown in table 2. The aircraft was trimmed in steady-heading, hall-centered leveli'ight at the desired trim airspeed, then stabilized at incremental airspeeds greaterthan and less than trim airspeeds. The test vsults are presented in figures 28through 35, appendix E.

22. The elevator control force gradient, as indicated by the variation in elevatorcontrol force with airspeeds, was positive for airspeeds above and below trim inthe cruise configuration, indicating stable static longitudinal stability, In the powerapproach conliguration, at 140.KCAS trim conditions, the elevator control forcegradient becamei negative at airspeeds below 120 K('AS in the normal missioncon figuration (a Ft cg). This lightening of control force gradient became negativeat iiirspeeds below 120 K('AS. This lightening of control force was objectionableand required moderate pilot compensation fot adequate airsped control (IIQRS 4).The elevator control position gradient, as indicated by variation in elevator controlposition with airspeeds in the forward cg configuration, was positive, althoughshallow, for airspeeds above and below trim. However, the gradient in the normalmission configuration (aft cg) was essentially neutral. The neutral elevator controlposition gradient was not objectionable, due to the strong influence of the positiveelevator force gradient on the pilot. The static longitudinal characteristics failedto inect the requirements of paragraph 23.175d of FAR Part 23, in that theelevator control force gradient reverses in the power approach configuration below120 KCAS, and the elevator control position gradient in the normal missionconfiguration (aft eg) is essentially neutral. These characteristics also did not meetthe requirements of paragraph 3,2.1.1 of MIL-F-8785B(ASG). The elevator controlforce gradient reversal in the power approach configuration is a shortcoming. AnEPR concerning this shortcoming was submitted (ref 18, app A). The aft cg limitof the normal mission configuration should be moved forward.

.Static lateral-directional Stability

23. Static lateral-directional stability characteristics were evaluated at theconditions shown in table 2. 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 rc,,ching siaeslip envelope limits. Test results are presented infigures 36 through 40, appendix E.

24. Static directional stability, as indicated by the variation of sideslip angle withrudder pedal force, was positive for sideslip angles between 10 degrees, left andright, from trim. A lightening of rudder pedal force occurred at sideslip anglesoutside this range at airspeeds below 170 KCAS. However, the rudder pedal forcenever reduced to zero, nor resulted in any unusual flight characte.,ristics orobjectionable increases in pilot elfort to maintain aircraft control. The variationof sideslip angle with rudder pedal deflection was essentially linear for sideslipangles encountered up to full rudder pedal deflection (170 KCAS and below) or

16

until achieving a I 80-pound rudder pedal force (200 KCAS). The static directionalstability failed to meet the requirements o1 paragraph 23.177(a)(3) ofFAR Part 23, in that the rudder pedal force grradient reverses prior to obtai:ingthe full rudder control limit below 170 KCAS. Hlowever, the static directionalstability did i1eet the requirements of MIL-F-8785B(ASG) and was acceptable.

25. l)ihedral effect, :. 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 leftand right directions. The corresponding increase in elevator controi forces withincreased sideslip anJes was not objectionable. The side-force characteristics, asindicated by the varition of bank angle with sideslip angle, were positive andessentially linear, The strong side-Forcc characterislics pr-Jided excellent cues ofout-o*'"trim flight con~ditions to the pilot and enhanced coc.rdinatL c flight whilemaneuv_ýring tIIQRS 2). Within the scope of these tests, the static lateral-directionalstability characteristics are sttisfactiry.

l)vnanmie Longitudinal Stabilitv

2(,. The dynamic longitudinal stabilitý characteristics were evaluated at theconditions shown in tabl ... The long-term dynamic characteristics were evaluatedby slowing the aircrm't with aft elevato; control to an airspeed 30 knots beiowtrim airspeed and then returning the cointrol to .he trim position (stick-fixed) orreleasing the control ahal allowing it to seek the trim position (stick-free>,. Short-term0 . namic characteristics, simulating gust response, were evaluated by rapidlydisplacing the elevator control I inch irom trim for a duration of 0.5 second andMen returning the control to the trim position. Time histories of representativedynamic response characteristics are presented in figures 41 and 42, appendix E.Test resilts a-e summnaiized in table 8.

Table A. Dynamic Longitudinal Stability Characteristics.

Trim Damping Undamped DampedTtst Calibrated Rapnion Natural Natural

Airspeed i Frequency Frequency(kt) (W a - rad/sec) (wd - rad/sec)

140 0.70 I 2.51 1.79

Shcrt-period 168 0.75 2.74 1.81

198 0.93 2.45 0.90

140 0.025 0.114 0.114

Long-period 168 0,029 0.113 0.1 3

1%i 0.036 0.106 0.106

11

2/. The •ong-tern, aircraft response was oscillatory and was very lightly damped.The period was approximately 60 seconds. This weak damping, combined withthe large friction band, contributes to the poor trimmability of the aircraft. Thereis no FAA requirement for long-term dynamir's. but the Io";:-•rai chwracteristicsdx '1 ; 1.., ,,,c rO.Iutrements of paragraph 3.2.1.2 of MIL-F-8785B(ASG), inthat the damping ratio of the phugoid oscillation was less than 0.04. Within thescope of thi test, the long-term lonýitudinal dynamic characteristics are acceptable.

28. The I,,ngitudinal short-term characteristics of' the Model 200 aircraft wereessentially deadbeat for all te,-t conditions, including Ilight in turbulent conditions.

The short-term characteristics met the requirements of FAR Part 23 and oflMIL-F-8785B(ASG). For thle conditions investigated, the short-term longitudinaldynamic characteristics are satisf'actory,

IDynamic Lateral-Dirertional Stability

hltehh.l1oll Characteristics:

2). The dynamic lateral-directional characteristics (lateral-directional damping andDuitch-roll characteristics) were evaluated at the conditions shown in table 2. Theyaw damper installed in this aircraft was inoperative for the duration of the APE.These tests were conducted by exciting the aircraft from a coordinated level flighttrim condition with a rudder pulse and doublet, aileron pulse and doublet and1w release from a steady-heading sideslip. Time histories of representative dy-,amiclateral-directional characteristics are presented in figures 43 through 46, apnendix E.Test rest'lts are sunmmarized in table 9.

Table 9. Dutch-Roll Characteristics.

S Trim - i Daniping Roll/Yaw I Damped

Calibrated Ratio Ratio NaturalAirspeed (A) 4/\ Frequency

(kt) ''(II d - rad/sec)

118 0.065 0.82 1.26

140 0.068 0.90 I 1.51

168 0.071 0.98 1.80

198 0.075 1.22 2.21

223i 0.081 1.12 J 2.29

18

30, The Dutch-roll oscill, ions were lightly damped and easily excited. At therecommend,:d maneuver airspeed of 170 KIAS, the period was 3.5 seconds andthe damping ratio was 0.071. The aircraft's latc-rl-directional response andcontrollability characteristics were poor in atmespheric disturbances. Considerablepilot conmpensaion was required to over,;ome the sensitivc gust response duringturbulent Ilight conditions at all ,-g locations (HQRS 5). The Dutch-rollcharacteristics failed to ,,ocot the requirements of paiagraph 23.177(a)(d) ofFAR Part 23 and of' paragraphs 3.3.1.1 and 3.3.2.1 of MIL-F..87851(ASG). inthat any short-period oscillation must be heavily damped with the primary controlsfixed and free, The lightly damped, easily excited [hutch-roll oscillations areobjectionable and are considered to be a shortcoming. An EPR concerning thisshortcoming was submited (ref 20, app A). A reliable yaw damper or an autopilotsystem which would reduce lateral-directional pilot workload should be installed.

Spiral Stability Characteristics:

31. The spiral stability characteristics of the Model 200 aircraft were evaluatedat the conditions shown in table 2. 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 bankangle or the bank angle achieved after 20 seconds elapsed time. Test results areshown in table 10. Spiral stability, as indicated by change in bank angle withelapsed time, was essentially neutral for both left and right turns. This aircraftpossesses the capability of holding lateral trim in hands-off flight for periods oftime in excess of 20 seconds. The spiral stability characteristics met therequirements of' MIL-F-87MSIVASM) and are satisfactory.

Table 10. Spiral Stability Character i;tics.

T'[rim Calibrated Roll Attitude Roll AttitudeAirspeed at Release at 20 Seconds

(kt) (deg) (deg)

19.6, left 24.9, left168

21.6, right 23.0, right

18.2, left 13.7, leftI198

17.9, right 20.5, right

19

.32. Maneuvering stability characturistics were evaluated at the conditions shownin table 2. The variation of' elevator control force and control position with nornialacceleration was determined by trimming the aircraft in coor-tinated level flightat the desired trim airspeed and then stabilizing at incremental batik angles in steadyturns, both left and right. Airspeed was held constait and the aircraft was allowedto descend diring the maneuver. Data were obtained at each stabilized bank angle,Symmetrical pu1l-p maileuver% were required to obtain load Factors in excess of 2:Symmetrical pushover maneuvi , were required to obtain data below +lg. The loadfactor was varied incrementally to the maximum allowable during these maneuvers.The results of the maneuvering stability evaluation are presented in figures 47,48, and 49, appendix F.

33. Tile stick-free mianeuvejing stability, as indicated by the variation of elevatorcontrol force with normal acceleration, was positive (increased aft devator controlForce with increased load factor) and was essentially linear for all conditions tested.The elevator control Iorce gradient (stick force per g) was approximately32 pounds per g. The stick-fixed maneuvering stability, as indicated by thevariation of elevator control position with normal acceleration, was positive(increased aft elevator control motion with increased load factor) and essentiallyiinear. The control position gradient was approximately 0.4 inch per g.

34. Buffeting was encountered while attempting to achieve load factors ir, excessof 2 ý:1 140 KIAS. The maneuvering control forces were high enought to preventcontrol inputs which might give abrupt aircraft responses and therc was no tendencyfor the pilot to overcontrol. The maneuvering stability characteristics of theModel 200 aircralt met the requirements of MIL-F-8785B(ASG) and aresat isfactory.

Roll Performiance Characteristics

35, Roll performance characteristics were evaluated at tile conditions shown intable 2. These tests were initiated from a trimmed unaccelerated flight conditionby applying a rapid maximum deflection aileron control input without changingeither elevator or rudder pedal control position. Representative test results aresummarized in figures 50 and 51, appendix E. The adverse yaw was small (lessthan 2 degrees of sideslip) and was not objectionable. No roll cross-couplingcharacteristics were noted. The roll performance characteristics met therequirements of MIL-F-8785B(ASG). Within the scope of these tests, the rollperformance characteristics are satisfactory.

20

mom',.

Stall Characteristits

3(Mi, Dual and single-engine stall characteristics of the Model 200 aircraft wereevaluated at tile conditions shown in table 6. 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 persecond until achieving a stall. Stall warning margins and recovery characteristicsw-re evaluated qualitatively, Test results are presented in table 6.

37. l)ual-crigine stalls were evaluated with power OFF and with power for levelflight. The initial stall warning was provided by the sta.l warning horn. A liftcomputer incorporated in the stall warning system provided a programmed constantstall warning mnargin witlh various flap settings. Initial stall warning was providedby the stall varning horn at appru..imately 10 knots above the stall airspeed.Additional warning was provided by a moderate buffet at apnroximately 2 knotsin advance of tLe stall. Lateral control effectiveness temained good throughoutthe approach to J•e stall and no discernible nonlinear increase in elevator controlforce occurred prior to the stall. In general, all stalls were characterized by a s!ightpitch-up, followed immediately by a nose-down pitch attitude. Stalls conductedat maximum continuous power settings were characterized by an uncontrollableleft roll to approximately a 35-degree roll attitude. Prompt recovery from all stallswas readily accomplished by relaxing elevator control force pressure and establishinga 15 to 20-degree nose-down pitch attitude (HQRS 3). During recovery, rates ofdescent in excess of 2000 ft/min were frequently achieved. Lateral controleffectiveness returned immediately, using this recovery technique. Applying aftelevator control force prior to firmly establishing full recovery airspeed(approximately 10 I • 1 5 kn )ts above the stall airspeed) resulted in a secondary stalland additional altitude loss. The deep stall characteristics normally associated withT-tailed aircraft were nevel encountered during these tests. The dual-engine stallcharacteristics of this aircraft were generally mild.

38. Single-engine stall characteristics were evaluated with the critical engine (leftengine) inoperative and MCI" on the remaining engine. The stall warning, stall, andstall recovery were essentially the same as the dual-engine stall characteristics. Aslight roll to the left accompanied the stall. Recovery was accomplished utilizingthe technique described in paragraph 37, and could be accomplished more quicklyby reducing power on the remaining engine immediately after the stall. The stallcharacteristics of the Model 200 aircraft met the requirements of FAR Part 23and MIL-F-8785B(ASG). The programmed stall warning system provided anexcellent stall warning margin for each of the various test configurations and isan enhancing design feature. Withir the scope of these tests, the stall characteristicsare satisfactory.

Single-Engine Characteristics

39. The single-engine characteristics of the Model 200 aircraft were evaluated atthe conditions shown in table 6. These tests were conducted by establishing trimconfiguration at the desired airspeed and simulating sudden engine failure by

21

selecting the left engine condition lever to the f'uel cutoff position, and byestablishing siigle-engine trim conditions at the desired airspeed and slowlydecelerating the aircraft to. the VMC at which either lateral or directional controlcould not be maintained. Test results are shown in table 6. Sudden engine failure(rudder boost OFF) resulted ill rudder pedal forces of 145 pounds (propellerwindmilling) and 130 pounds (propeller feathered) with MCP on the remainingengine. With the rudder boost ON, rudder pedal forces were reduced to 85 pounds(propeller windmilling) and 35 pounds (propeller feathered). With the rudder boostON, the average pilot could easily maintain directional control following unexpectedengine failure immediately after takeoff, due to •ne reduction of high rudder pedalforces. The rudder boost greatly reduced pilot workload during asymmetric powerconditions (HQRS 2) and is an enhancing feature. Transient forces resulting from-a sudden engine failure were not excessive. The pedal fkrces encountered couldbe trimmed out completely at all power settings with the rudder boost ON, greatlyredlucing pilot workload during singlc-engine operation (HQRS 3). With the rudderboost OFF, pedal forces could not be trimmcl out for power settings within,10 percent of MCP; however, the remaining pedal force required minimal pilotcompensation (HQRS 3).

40. Tae single-engine VMC was evaluated duiring single-engine stall tests and duringseparate VMC testing. The test aircraft always stalled prior to reaching single-engineVMC. Lateral and directional control effectiveness retrinend throughout theapproach to stall and returned immediately after performing the stall recoverytechniques described in piragraph 37. The single-engine stall spleeds shown intable 6 were also deter-.ined to be the VMC lor the tst. configarati, ns of theModel 200 aircraft. TIl- single-engine control characteristics met the requirementsof FAR Part 23 and MIL-F-8785B(ASG). Within the scope of these tests, thesingle-engine charactcvLstics are satisfactory.

Ground HIandling Cl'aradteristics

41. The ground handling characteristics of the Model 200 aircraft were evaluatedthroughout the conduct of these tests. In the normal mission configuration (aftcg), two people standing inside the aircraft in the vicinity of the swing-down doorentrance caused the nose gear to lift off of the ground. The fin on the 'oweraft portion of the fuselage had been bent and cracked by the contractor duringnormal ground handling in the normal mission configuration (photo A). Theswing-down door ground clearance was less than I inch in the normal missionconfiguration (photo B). In this configuration, the nose gear strut was fullyextended (photo C), in contrast to a forward eg (photo D). The pitch attitudeinstability evident during loading and ground operations of the aircraft in the normalmission configuration is a deficiency. An EPR concerning this deficiency wassubmitted (ref 20, app A).

22

Photo A. Lower Aft Portion of Fuselage.

II

ItI

Photo B. Swing-Down Door.

23

Photo C. Nose Gear Strut at Normal Mission Configuration.

Photo 1). Noise Gear Strut at Forward Center of Gravity.

24

42. An indicated 15U foot-pounds (ft-lb) of torque and 900 rpm propeller speedwere required to initiate taxi during no-wind conditions. Nose wheel steeringcharacteristics were good. Minimal pilot effort was required to maintain directionalcontrol during ground operations (HQRS 2). Use of the Beta range (propeller pitchsetting) on the power lever control console allowed low taxi speeds and reducedbraking requirements. Braking characteristics were excellent, with no fading oroverheating. No difficulty was encountered when using reverse thrust to back upfor short distances while in the aft cg configuration (this technique would notbe used under normal circumstances, due to excessive propeller blade erosion).Field of view from the cockpit was good during all giound and taxi operations.Within the scope of' these tests, the normal ground handling characteristics areacceptable.

43. The operation of CEFLY LANCER mission equipment requires power froman external auxiliary power unit (APU) prior to starting engines. After startingthe engines, it is necessary to mainta!i- 2600-rpm propeller speed to obtain the400-hertz requirements from the 8.5-KVA AC generators, which provide powerfor the mission gear when the APU is disconnected. The mission gear cannot acceptany intermittent power outages after being warmed up and power losses wouldbe experienced at propeller speeds less than 2000 rpm. It is impossible to meetthis requirement and still perform complete mandatory engine run-up checks.Attempts were made to perform normal ground and taxi operations and stillmaintain a continuous 2000 rpm propeller speed. Adequate performance was notattaiiable, using a variety of techniques, and maximum tolerable pilot compensationwas required (HQRS 7). The requirement to maintain a continuous 2000 rpmpropeller speed during ground operations is a deficiency. An EPR concerning thisdeficiency was submitted (ref 21, app A). The requirement to operate at theseconditions should be reevaluated.

Takeoff and Landin•r Characteristics

44. The takeoff and landing characteristics of the Model 200 aircraft wereevaluated on the 5000-foot hard-surface runway of the BAC facility at Wichita,Kansas at the conditions shown in table 1. Runway conditions were varied, andincluded wet, dry, ice, snow, and slush conditions. Takeoff characteristics wereevaluated using normal techniques (gradual application of power until achievingtakeoff power) and maximum performance techniques (set power at takeoff powerand release brakes). The time required for engine acceleration from flight-idle totakeoff power was approximately 4 seconds. Due to the torque generated duringmaximum performance takeoff, the aircraft longitudinal axis was lined upapproximately 5 degrees right of the runway center line as a technique to permita smooth takeoff roll. The brakes held well during application of takeoff' powerand simultaneous brake release was accomplished easily. Aircraft torque broughtthe aircraft onto the center line and directional control was satisfactory throughthe rudder effective speed of 45 KIAS. Normal takeoff techniques were utilizedfor normal takeoff power applications. Rotation was accomplished withoutexcessive aft elevator control pressure at approximately 85 KIAS and lift-offoccurred at 95 KIAS. The gear was raised at 105 KIAS. Travel time was

25

.. .. ....

5.3 seconds. Tie flaps were raised at 120 KIAS. Travel time was 3.0 seconds.Minor longitudinal trim changes (I or 2 degrees, nose down from takeoff trimsettings) were required during transition to initial climb conditions. Field of viewduring takeoff and initial climb was good. Placard limits for gear and flaps wouldbe difficult to exceed during normal operating conditions. Takeoffs during wetand icy conditions were conducted with caution, but required minimal pilotcompensation to achieve satisfactory performance (HQRS 3). Within the scope cfthese tests, the takeoff characteristics were satisfactory.

45. Landing characteristics were evaluated using normal landing techniques at theconditions shown in table 1. Power approaAlIes to landing were perform, .,

primarily due to the high gross weights used during these tests. Downwind airs : .dwas 140 KIAS, with transition to 120 KIAS while turn'ng on the final appro,,in.Touichdown airspeeds were approximately 80 KIAS vith 100 percent flaps,90 KIAS with 40 percent flaps, and 100 KIAS with no flaps. Brakingcharacteristics were excellent, with no fading or overheating. Full reverse propellerpitch was utilized on several landings. Roll-out distances were reduced and nocontrollability problems were encountered. A single-engine go-around to a closedtraffic pattern was executed on short final approach with no difficulty, afterretracting t!he landing gear and flaps. Failure to retract the landing gear immediatelywill resuit in a negative rate of climb. A single-engine landing was accomplishedanul naximium revetse thrust onl the operating engine was Utilized during roll-out.

There Was no tendncyIC to overcontrol, although considerable pilot effort wasrequired to m~inaiain proper brake application in response to the asymmetric reversetlhrust (llQRS 5). The malin landing gear braking characteristics were excellent andare an enhancing eafturt Of' this aircraft. Within the scope of these tests, the landingcharacteristics are satisl'a-tory..

Trim Change• ( haratel•r;s•lic,

46, Trim change characteristics were evaluated at the conditions shown in table 2.The aircraft was trimined in steady-heading, ball-centered level flight at the desiredinitial trim condlitions, then a configuration change was made while holding oneor imore of the initial trim parameters constant. Variations in tower, flap position,and gear position utilized di'riig the conduct of' this test are specified inparagraph 23.145 of FAR Part 23 and in paragraph 3.8.6.1 of Bk', SpecificationBS 22301 (ref 14, app A). The test results are presented in table 11. All controlforce variations with tri.n changes were light, ranging from 4 to 28 pounds. Onlyone force exceeded the specification limitations. In descending flight at 140 KCASwith gear down, 100 percent flaps. and power off, the rapid addition of maximumcontinuous power resulted in a nose-up pitch attitude which required a forwardstick force of' 13 pounds to maintain 140 KCAS. Although this control forceexceeded the requiremnent (If paragraph 3.8 6.1(5) of the BAC specification limit(10 pounds) by I pounds, it was not an excessive increase and required minimalpilot compensation to maintain the desired flight condition (HQPS 3). Within thescope of these tests, the trim change characteristics of the Model 200 aircraft aresatisfactory.

26

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Night Operations

47. The night operational capability of the Model 200 aircraft was evaluated briefly

during night operations at the BAC facility. Cockpit lighting features were good.

The pilot's fresh air ventilation window reflected the instrument panel and center

console lights, but visibility was not reduced and the slight glare was not

objectionable. Other lighting features included wing tip and tail strobe lights, wing

ice lights, navigation lights, rotating beacons, landing lights, and taxi light. All

switches were easily accessible and the systems functioned properly. The taxi light

provided adequate illumination for ground operations and the landing lights

provided good illumination for takeoff and landing operations. Tile landing lights

are located adjacent to the taxi light on the nose gear (photo E) and no operating

time limit is specified for ground operation. Within the scope of these tests, the

night operational capability of the aircraft is satisfactory.

!L

isi

Photo E. Landing Lights.

28

Instrument Flight Capability

48, A limited evaluation of the instrument flight capability of the Model 200aircraft was conducted at the end of several engineering flight test periods. Asimulated instrument takeoff was conducted and minimal pilot effort was requiredto overcome takeoff power torque and maintain runway heading (HQRS 3).Hlolding patterns were flown at 140 and 170 KIAS under simulated instrumentconditions. Standard rate turns were easy to perform. The aircraft would holdtrim in smooth atmospheric conditions for periods of time long enough to allowthe pilot to copy clearance changes or perform other cockpit duties. However,under conditions of light-to-moderate atmospheric turbulence, it was necessary tomonitor trim closely (para 20). A practice precision radar approach was flownat McConnell Air Force Base, Kansas, using the contractor-recommended transitionairspeeds (140 to 160 KIAS) and the contractor-recommended approach airspeed( 120 KIAS). F~light path stability appeared to be good. and adjustments to headingand glide path were accomplished with minimal pilot effort (HQRS 3).Missed-approach procedures were accomplished satisfactorily.

49. During a normal approach and landing at the BAC facility, the power leverswere retarded to the flight-idle position to make an altitude correction -nd thepropeller speed dropped to 1900 rpm, resulting in the loss of the 8.5-KVA ACgenerators and the corresponding loss of the primary attitude and heading indicatorgyros. This phenonmenon was investigated at altitude and it was determined thatait airspeeds of 125 KIAS or less (5 knots above the recommended approachairspeed of 120 KIAS). the AC generators would be lost if the power levers wereretarded to Ilight-idle. In actual instrument conditions, the loss of the primaryattitude and heading gyros could result in loss of aircraft control (HQRS 10). Theloss of the primary attitude and heading gyros when the propeller speed is lessthan 2000 rpm is a deficiency. An EPR concerning the deficiency was submitted(ref 22, app A). The absente of a weather radar capability restricts this aircraft'sahility to perform its all-weather mission in an environment where ground radarcontrol advisory facilities are not available. Consideration should be given to'nstal!ing a weather radar system in production aircraft to enhance the all-weathercapability of this aircraft,

Aircraft Systems Failures

Yaw Damper:

30. Repeated malfunctioning of the yaw damper system occurred during theconduct of these tests. D)ivergent lateral-directional oscillations were encounteredon several occasions with the yaw damper system ON. These objectionableoscillations were stopped by switching the yaw damper system OFF. On otheroccasions, steady 5-degree bank angle coordinated turns to the let't resulted withthe system ON. On the one flight when the yaw damper appeared to be operatingproperly, lateral-directional oscillations excited by I-inch rudder control pulseinputs were damped after four overshoots. With the yaw damper system OFF,these same inputs resulted in 8 to 10 overshoots and long-term residual

29

.................--.... -- -

-..

lateral-d irectional oscillations. The lightly da inped, easily excited residuallateral-directional oscillations which can result with the yaw damper OFF havealready been described as a shortcoming in paragraph 30. The yaw damper systemis required by the FAA for flights above 17,000 feet, due to the weaklateral-directional damping characteristics discovered in the commercial aircraft. Areliable yaw damper should be installed.

Rudder Boost:

5 1. Rudder boost failures were evaluated during the conduct of single- ':ngine tests.

Sudden engine failures with the rudder boost OFF resulted in a rudder pedal forceof 145 pounds (propeller windmilling) and 130 pounds (propeller feathered). Ifan engine were lost during takeoff with the rudder boost OFF, it would be necessaryto overcome sudden high rudder pedal forces to maintain aircraft control. Thissudden increase in rudder pedal force would require moderate pilot effort tomaintain ade(uate control (HQRS 4). With the rudder boost OFF, pedal forcescould not be trimmed out for power settings within 10 percent of MCP: however,the remaining pedal force required mininml pilot compensation for desiredperformance (HQRS 3).

Alternating Current Generator:

52. The failure of one AC' generator was simulated in flight by switching off theleft AC generator. The opposite generator was capable of continuing to supplythe necessary AC power requirements. The probability of a dual.AC generatorfailure is remote. However, this condition was artificially induced by retarding thepower levers to the flight-idle positions at airspeeds below 125 KIAS. This resultedin the immediate loss of electrical power to the primary attitude and heading gyros.The results of this loss have been stated in paragraph 49. This would also resultin the loss of all power for the CEFLY LANCER mission gear. The loss of powerfor this mission gear, when the propeller speed is less than 2000 rpm, is adeficiency. An electrical power system should be installed on this aircraft to permituse of lower propeller speeds for taxiing, cruise economy, and reduced noise levels.

IItJMAN F'A(CT'ORIS

Cockpit Evahiation

53. The cockpit area (Mlight decK) and ingress/egress areas were evaluatedthroughout the test program. Entrance to the aircraft is accomplished through aswing-down airstair door. A hydraulic damper allows the door to swing down slowlywhen opened. In the normal mission configuration (aft cg), the door-groundclearance was I inch. With one persona on the step, this clearance was furtherreduced to 1/2 inch. An emergency exit door, placarded EXIT PULL, is locatedon the right cabin side wall just aft of the copilot seat (photo F). Approximatelyi0 inches of clearance between the pilot and copilot seats made entrance intothe cockpit area awkward. The seat and flight control adjustment were adequate,

30

but 3 to 4 inches more aft seat travel would greatly facilitate ingress/egress. Thereis no capability to adjust each lap be!t and keep tile shoulder strap attachmentpoint to the lap belts centered in the pilot's lap. Tile single lap belt adjustmentprovided is inadequate fbr use with shoulder straps and is a shortcoming. An EPRconcerning this shortcoming was submitted (ref 23, app A). Seat comfort isadequate and the fresh air ventilation ducts were conveniently located and easilyadjusted.

Photo F. Emergency Exit Door.

54. The test aircraft cockpit area is shown in photo G. The grouping andreadability of the flight and engine instruments were adequate. The caution,advisory, and warning annunciator panels are well placed and easy to read. Thelocation of the red WARNING and yellow CAUTION flasher lights at eye level,directly in front of the pilot and copilot on the glare shield, is a good designcharacteristic. The fuel panel arrangement is accessible and easy to read. The circuitbreaker side panel is also accessible and easy to monitor. The overhead panel systemcontrols. switches, and placards are easily accessible and their systematicarrangement allows utilization with little or no confusion. The pedestal-mountedpropulsion system controls and associated controls are adequate, with goodarrangement and accessibility. A feel identification feature is provided on the major

31

''. M

controls (power levers, propeller levers, condition levers, flaps, trim wheels, andfriction adjustment). The propeller feather detent is 3/4 inch above the presentmarkings. The improper marking of the propeller feather range on the controlconsole is a shortcoming. An EPR concerning this shortcoming was submitted(ref 24, app A).

,6

Photo G. Control Console.

55. The communications panel console was inadequate in several areas. There wasno provision for a three-position pusl-to-talk communications switch for eitherthe pilot or copilot in the test aircraft, The absence of a dhree-position push-to-talkswitch is a shortcoming, and fails to meet the requirements ofparagraph 3.12.4.3.4.1 of the Aircraft Procurement Specification (ref 4. app A).The communications p mel and transmission selection procedures were nonstandardand confusing. On several occasions the copilot inadvertently transmitted to thepilot on intercom and blocked out tower transmissions to the pilot. The copilotwas unable to monitor very-high-frequency (VHF) communications radios selectedon his signal distribution panel while his transmit-select was in the intercomposition, which is a shortcoming. During electrical extension of the ice vanes, thecommunications cord to the pilot control wheel prevented the ice vane handlesfrom extending fully and is a shortcoming (photo H). Three EPR's concerning

32

these communications panel shortcomings were submitted (ref's 25, 26, and 27,app A). An ultra-high-frequency (UHF) radio in lieu of' one of' the three installedVHF radios would increase flight communications tleibility and provide analternate radio for use with military Lommunications networks.

-r-I

Photo II. Instrument Panel, Pilot Side.

Noise

56. Noise in the cockpit area was evaluated qualitatively throughout the testprogram. Engine a-id propeller noise levels were considered to be excessive whileoperating at a continuous propeller speed of 2000 rpm. Noise levels were veryuncomfortable after 2 hours of' operation. even when wearing a properly fittedSPH-4 flying helmet. Noise levels at 1800 rpm propeller speed, the operating rpmof the commercial version Mode! 200 aircraft. were much lower, with no noticeablehearing discomfort. Further testing should be conducted to quantitatively measureand assess noise levels encountered during normal mission operations.

33

Tloxic'it y

57. On every flight at altitudes above 15,000 feet lip, a very light smoke filledthe cabinl anid cockpit areas. Exposure times in excess of' 30 minutes resulted inmild eye and throat irritation. Smoke in tlvh cockpit and cabin areas is a deficiency.An FPR coticcining this deficiency was submitted (relf 28. app A). Further testingshould be condutCed to determine the cause and the method of elimination ofthis ansatislfictory characteristic.

I'ELIABILITY AND MAINTAINABILITY

58. Factors affecting 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, I'istencrs, cables, connectors, and safrty. Availabile contractor technicaldocuments, historical data, anid current maintenance proce'dures were reviewed. Thisreview was a limited, noninterference evaluation. Primarily, a qualitative evaluationwas performed because the minimal number of program flight houls limited theopportunity to obser ve component repair and replacement. Formal removal orreplacement tests were not performed. The aircraft was fully instrumented, a

condition that resuhed in maintenance complications which would not exist onan operational aircraft.

59. The items listed below arc shortcomings which will affect the reliability andmaintainability of the Model 200 aircraft. Equipment Performance Reportsconcerning these shortcomings were su')mitted (ret's 29 through 37. app A).

a. The engine fuel transfer caution lights came on erroneously during flight.

N b. The engine exhaust stacks craLked several times during testing (photo 1).

c. The torque needks fluctuated ±-25 ft-lb during flight.

d. Tie engine fire warning lights came on erroneously several times duringflight (sunlight reflected inside the engine nacelle triggered one of thephotoconductive cells).

e. The rudder boost pneumatic cylinder hoses to the flow control boxconnection points are identical, making it possible to connect these hosesbackwards.

f. The installed yaw damper system malf1unctioted repeatedly during theconduct of' these tests.

g. Ground towing of' the test aircraft in the normal mission configurationresulted in the bending of the nose gear steering idler link.

34

h. The propeller synchrophaser could not match propeller speeds at2000 rpm (the primary governor stop limit), resulting in the requirement formanual adjustment of the propeller levers to obtain a synchronized rpm setting.

i. Engine exhaust gases used to continually 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 (photo J).

60. Normal maintenance procedure requires a 10-foot-high stand or ladder toperform the preflight and postflight inspections of the T-tail area. In addition,a tail stand was required to prevent the airplane from rocking back on its tailwhile performing ground maintenance operations during the normal missionconfiLuration (aft cg) tests.

35

PhIoto 1. Engine ExAhaust Stack.

Phloto j. ngine Inilet Cowlinig.

36

CONCLUSIONS

GENERAL

61. The following conclusions were reached upon completion of the BACModel 200 aircraft APE I:

a. In the normal mission configuration at the military maximum gross weightof 15,000 pounds and an aft cg limit at FS 197.4, the Model 200 aircraft hasinadequate single-engine performance and degraded longitudinal handling qualities.

b. The programmed stall warning system provided an excellent stall warningmargin for each of the various test configurations and is an enhancing design feature(para 38).

c. The rudder boost greatly reduced pilot workload during asymmetricpower conditions and is an enhancing feature (para 39).

d. The main landing gear braking characteristics were excellent and are anenhancing feature of this aircraft (para 45).

e. Four deficicncies and eighteen shortcomings were noted during thesetests.

DEFICIENCIES AND SHORTCOMINGS

"62. The following deficiencies were identified:

a. The pitch attitude instability evident during loading and groundoperations of the aircraft in the normal mission configuration (aft cg) (para 41).

b. The requirement to maintain a continuous 2000-rpm propeller speedduring ground operations (para 43).

c. The loss of the primary attitude and heading gyros whe, the propellerspeed is less than 2000 rpm (para 49).

d. Smoke in the cockpit and cabin areas (para 57).

63. The following shortcomings were identified:

a. The single-engine performance of the aircraft in the normal missiontakeoff configuration was inadequate (para 12).

b. The long-term trimmability was poor (para 20).

37

c. The elevator control force versus airspeed gradient was reversed in thepower approach configuration (para 22).

d. The Dutch-roll oscillations were lightly damped and easily excited(para 30).

e. The single lap belt adjustment provided is inadequate for use withshoulder straps (para 53).

f. The propeller feather range markings on the control console were

improper (para 54).

g. There was no three-position push-to-talk switch (para 55).

h. The copilot was unable to monitor VHF communications channelsselected on his signal distribution panel while his transmit-select was in the intercomposition (para 55).

i. During electrical extension of the ice vanes, the communications cordto the pilot control wheel prevented the ice vane handles from extending fully(para 55).

j. The engine fuel transfer caution lights came on erroneously during flight(para 59a).

k. The engine exhaust stacks cracked several times during testing (para 59b).

I. The torque needles continually fluctuated ±25 ft-lb during flight(para 59 c).

m. The engine fire warning lights came on intermittently several times duringflight (sunlight reflected inside the engine nacelle triggered one of thephotoconductive cells) (para 59d).

n. The rudder boost pneumatic cylinder hoses to the flow control boxconnection points are identical, making it possible to connect thesc hoses backwards(Para 59e).

o. The installed yaw damper system malfunctioned repeatedly during theconduct of these tests (para 59f).

p. Ground towing of the test aircraft in the normal mission configurationresulted in the bending of the nose gear steering idler link (para 59g).

38

q. The propeller synchrophaser could not match propeller speeds at2000 rpm (the primary governor stop limit), resulting in the requirement for

manual adjustment of the propeller levers to obtain a synchronized rpm setting(para 59h).

r. Engine exhaust gases used to continually heat the engine air inlet lipsleaked at random locations about the periphery of these lips, resulting in discoloredand blistered paint areas on the inlet covering (para 59i).

39

RECOMMENDATIONS

64. 'the deficiencies identified during this evaluation must be corrected (para 62).

65. The shortcomings should be corrected (para 62).

66. The aft cg limit of the normal mission configuration should be moved forward(para 22).

67. Consideration should be given to installing an autopilot system to reduce pilotworkload (para 22).

68. A reliable yaw damper or an autopilot system which would reducelateral-directional pilot workload should be installed (para 30).

69. Consideration should be given to installing a weather radar system inproduction aircraft to enhance the all-weather capability of the aihcraft (para 49).

70. A variable-frequency AC generator should be installed on this aircraft to permituse of lower propeller speeds for taxiing, cruise economy, and reduced noise levels(para S2).

71. A UlHF radio should be installed in lieu of one of the three installed VHFradios, as this would increase flight communications flexibility and provide analternate radio for use with military communications networks (para 55).

72. Further testing should be conducted to quantitatively measure and assess noiselevels encountered during normal mission operations (para 56).

40

APPENDIX A. REFERENCES

I. 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. Prime Item Development Specification, Beech Aircraft Corporation,BS 22296A, Beechcraft Model A100-1 Pressurized Turboprop Executive Transport15,000 pounds Maximum Takeoff Gross Weight, 23 February 1973, revised30 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, AI00-1 Aircraft 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, AI00-1 Aircraft for Gross WeightsLip to 15,000 Pounds.

8. Flight Test Manual, Naval Air Test Center, FTM No. 104, Fixed Wing

Performance, 28 July 1972.

9. Flight Test Manual, Naval Air Test Center, FTM No. 103, Fixed Wing Stabilityand Control, 1 August 1969.

10. Handbook, USAF Aerospace Research Pilot School, FTC-TIH-70-1001,Performance, September 1970.

I1. Handbook, USAF Aerospace Research Pilot School, FTC-TIH-68-1002.Stability and Control, September 1968.

12. Flight Test Manual, Advisory Group for Aeronautical Research andDevelopment, Volume 1, Performance, Pergaman Press, Los Angeles, California,1959.

41

13. Flight Test Manual, Advisory Group for Aeronautical Research andDevelopment, Volume II, Stabilit, and Control, Pergaman Press, Los Angeles,California, 1959.

14. Specification, Beech Aircraft Corporation, BS 22301, AirworthinessQualification Specification, Beechcraft Model A 100-1, 19 September 1973.

15. Military Specifiation, MIL-F-8785B(ASG), Flying Qualities of PilotedAirplanes, 7 August 1969, with Interim Amendment 1, 31 March 1971.

16. Equipment Performance Report, USAASTA, SAVTE-TB, EPR 74-21-21.

"Army Preliminary Evaluation 1, Model A 100-1 CEFLY LANCER," 27 April 1974.

17. EPR, USAASTA, 74-21-18, 12 March 1974.




21. FPR, USAASTA, 74-21-2, 20 March 1974.


23. LPR, USAASTA, 74-21-5, 20 March 1974.

24. EPR, USAASTA, 74-21-6, 20 March 1974.25. EPR, USAASTA 74-21-7, 20 March 1974.



28. EPR, USAASTA. 74-21-4, 20 March 1974.






42

34. FPR, USAASTA, 74-21-15, 20 March 1974.


36. EPR, USAASTA, 74-21-17. 20 March 1974.

37. EPR, USAASTA, 74-21-22, 21 May 1974.

4i

43.

e •

APPENDIX B. DESCRIPTION

GENERAL

I The Model 200 aircraft has the general structure and space arrangements ofthe BAC Super KingAir Model 200 aircraft. Three views of the test aircraft areshown in photos B-I, B-2, and B-3. General specifications are listed below.

I)imensions

Wing span 54 ft, 6 in.Horizontal stabilizer span 18 ft, 5 in.Length 43 ft, 9 in.Heigh' to top of vertical stabilizer 15 ftPropeller diameter 8 ft, 2.5 in.Propeller/fuselage clearance 29.0 in.Propeller/groun d clearance 14.5 in.Distance between main gear 17 ft, 2 in.Distance between main and nose gear 15 ft

Cabin Dimensions

Total pressurized iength 264 in.Cabin length, partition to partition 128 in.Cabin height 57 in.Cabin width 54 in.Fntrance door 21.5 in. x 26.7 in.

Wing Area and Loading

Wing area 303.0 ft2

Wing loading 49.5 lb/ft 2

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

Ground Turning Clearance

Radius for inside gear 4 ftRadius for nose wheel 19 ft, 6 in.Radius for outside gear 21 ft, I in.Radius for wing tip 39 ft, 10 in.

44

Photo B-i. Right Side View, Model 200 Aircraft.

Photo B.2. Front View, Model 200 Aircraft.

45

Photo B-3. Lef Sid (e Front Quiarter View, Model 200 Aircraft.

46

FLIGHT CONTROL SYSTEM11

2. The Model 200 aircraft is provided with conventional dual controls for thepilot and copilot. The flight control system is reversible. The elevator and ruddercontrol surfaces are of conventional design. The aileron control surface has a28 inch x 1-112 inch metal sandwich added to the trailing edge adjacent to thetrim tab to aid lateral control effectiveness (photo B-4). The elevators and aileronsare operated by conventional control whcels interconnected by a T-column. Therudder pedals are interconnected by a linkage below the floor. These systems areconnected to the control surfaces through closed cable bell crank systems. Rudder,elevator, and aileron trim are adjustable with controls mounted on the centerpedestal. Position indicators for each of the trim tabs are integrated with theirrespective controls. An elevator bob-weight and downspring has been incorporatedto lighten longitudinal control forces in flight. A control lock is provided whichpermits positive locking of the control column, rudder pedals and the engine powercortrols.

Photo B-4. Right Aileron Control Surface Modification.

3. A rudder boost system is provided to assist in maintaining directional controlduring asymmetrical thrust conditions, such as engine failure or a large variationof power between the engines. Incorporated in the nidder cable system are twopneumatic rudder boosting servos that actuate the cables to provide rudder pressureto help compensate for asymmetrical thrust. The system is operated by sensing

41

differential pressure between each of' the engine bleed air systems. The systemis operated by a toggle switch located on the pedestal below the rudder trim wheel.A functional check of the system may be obtained during the conduct of normalengine run-up procedures.

4. A yaw damper system is provided to assist in maintaining directional stability.The system coraponents include a yaw sensor, amplifier, and control valve.Regulated air pressure from the control valve is directed to the same pneumaticservos used for thfl rudder boost system. The system is controlled by a toggleswitch adjacent to the rudder boost switch on the pedestal. The circuit of theyaw damping system is interrupted by the landing gear safety switch while theairplane is on the ground and will not operate in this condition. The system mayne used at any altitude and is required for flight above 17,000 fct.

ELECTRICAL SYSTEM

5. The aircraft electrical system is a 28 VDC single conductor system with thenegative lead of et.ch power source grounded to the main aircraft structure. DCelectrical power is provided by one 34 amlpere-hour, air-cooled, 20-cellnickel-cadmium battery and two 300-ampere starter/generators connected inparallel. Two three-phase, 8.5-KVA generators provide AC power to the aircraftand special equipment busses. These generators are housed in blisters, one on eachengine nacelle (photos B-5, B-6, and B-7). The system is capable of supplyingpower to all subsystems that are necessary for normal operation of the airplane.A hot battery bus is provided for emergency operation of certain essentialequipment and the cabin entry threshold light. Power to the main bus from thebattery is through the battery relay, controlled by a switch placarded BAT-ON-OFF,located on the left subpanel. Power to the main bus system from the DC generatorsis controlled by a generator control panel which includes the followingfeatures: generator voltage regulator, generator paralleling, generator line contractorcontrol, generator overvoltage protection, generator feeder ground fault protection,and reverse current protection. The DC generators are controlled by switches,placarded L DC GEN and R DC GEN, located on the left subpanel. Each ACgenerator is controlled by a generator control panel located beneath the floor aftof the main spar. Each control panel includes a voltage regulator and currenttransformer which regulates the voltage output of the AC generator, and a powermonitor that isolates the AC generator whenever voltage is not within 95/1 17 voltsAC and the frequency is not between 375/425 Hz. The AC generators arecontrolled by switches, placarded L AC GEN and R AC GEN, located on theleft subpanel.

48

IKPhoto B-5, AC (Ceierator (cover removed).

Photo B.6. AC (Cenerator. Frowl View.

49

LA

Photo B-7. AC Generator, Side View.

ENVIRON' ENTAL SYSTEM

6. The environmental system consists of the bleed air pressurization, heating andcooling systems, and their associated controls. The cabin pressure vessel is designedfor a normal working pressure differential of 6.0 psi, which will provide a cabinpressure altitude of 3870 feet at an airplane altitude of 20,000 feet. It will providea nominal cabin altitude of 9840 feet at an airplane altitude of 31,000 feet. Amixture of bleed air from the engines and ambient air is available for cabinpressurization at a rate of approximately 10 to 1 pounds per minute. This airmixture also passes through a heating flow control unit in each nacelle and isducted into the cabin to provide heating. An air-to-air heat exchanger helps regulatethe temperature of the bleed air. Cabin air conditioning is provided by a refrigerantgas vapor-cycle refrigeration system consisting of a belt-driven engine-mountedcompressor installed in the right engine (photo B-8). An environmental controlsection on the copilot subpanel provides for automatic or manual control of theenvironmental system. Pressurization controls are located on the pedestal andconsist of a pressure setting rheostat and a cabin pressure/dump toggle switch,

50

Li

Photo B-8. Air Conditionitig Compressor.

PROPULSION SYSTEM

7. The PT6A-41 engine, manufactured by United Aircraft of Canada, Ltd, hasa three-stage axial, single-stage centrifugal compressor, driven by a single-stagereaction turbine. The power turbine, counterrotating with the compressor turbine,drives the output shaft. These engines produce 850 shaft horsepower each.Maximum continuous speed of the engine is 38,100 rpm, which equals101.5 percent N1. Prior to gear reduction, the turbine speed on the power sideof the engine is 30,000 rpm at 2000 rpm propeller speed. 'The two engines installedon the test aircraft were experimental prototype engines (left engine serialnumber X70014. right engine serial number X7001 1).

8. The Hartzell propeller is of the full-feathering, constant-speed, counterweightedreversing type, controlled by engine oil pressure through single-action, engine-drivenpropeller governors. The propeller is threc-bladed and flange-mounted to the engineshaft. Centrifugal counterweights, assisted by a feathering spring, move the bladestoward the low rpm (high pitch) position and into the feathered position. Governorboosted engine oil pressure moves the propeller to the high rpm (low pitch)hydraulic stop and revcrsing position. The propellers have no low rpm (high pitch)stops; this allows the blades to feather after engine shutdown.

"51

9. The propulsion system is operated by three sets of controls: the power levers,propeller levers, and condition levers (photo B-9). The power levers provide controlof engine power from idle through takeoff power by operation of the gas generator(Ni) governor in the fuel control unit. When the power levers are lifted ov,.•r theidle detent they control engine power through the beta and reverse ranges. Thepropeller levers are operated conventionally and control the constant-speedpropellers through the primary governor. Normal operating range is 1600 to2000 rpm. However, the propeller levers must remain full forward (2000 rpm)throughout all flight conditions on the CEFLY LANCER aircraft due to presentconfiguration electrical power supply requirements (2000 rpm are required tooperate the direct-drive 400-cycle 8.5-KVA generators). The condition levers controlthe flow of fuel at the fuel control outlet and select fuel cutoff, low-idle(52 percent NI), and high-idle (70 percent NI) functions.

II/IPOWI '1M

Photo 1-9. Propulsion Control Console.

FUEL SYSTEM

10. The fuel system consists of two separate systems connected by avalve-controlled cross-feed line. The separate fuel system for each engine is furtherdivided into a main and an auxiliary fuel system. The main system consists ofa nacelle tank, two wing leading-edge !anks, two box-section bladder tanks, andan integral (wet cell) tank, all interconnected to flow into the nacelle tank bygravity. This system of tanks is filled from the filler located near the wing tip.

52

The auxiliary fuel system consists of a center section tank with its own filler openingand an automatic fuel transfet system to transfer the fuel into the main fuel system.When the auxiliary tanks are filled, they will be used first. During transfer ofauxiliary fuel, which is automatically controlled, the nacelle tanks are maintainedfull. A swing check valve in the gravity feed line from the outboard wing preventsreverse fuel flow. Normal gravity transfer of the main wing fuel iuto the nacelletanks will begin when the auxiliary fuel is exhausted. The two systems are ventedthrough a recessed ram vent coupled to a protruding heated ram vent on theunderside of the wing adjacent to the nacelle.

11. A fuel dump system is provided on the CEFLY LANCER aircraft(photo B-10). A guarded toggle switch on the fuel control panel activates the dumpsystem. A check valve is opened and fuel is dumped from each of the separatefuel systems, utilizing gravity feed. Approximately 1500 pounds of fuel can bejettisoned during a 10-minute period. The dump may be terminated at "he pilot'sdiscretion, using the dump toggle switch.

• ,,

Photo B-10. Fuel Dump Outlet.

53

. .-- ...- ~ .......

MISCELLANEOUS SYSTEMS

Landing Gear

12. The tricycle high-flotation landing gear is retracted and extended by a 28-voltsplit field motor. A close-up view of the dual main gear wheels is shown inphoto B-I I. The gear motor is controlled by a switch located on the pilot subpanel,which must be pulled out of a detent to initiate retraction or extension. In theretracted position, the nose gear is fully enclosed and scaled by the nose geardoors. The main gear is partially exposed, with 8 inches of the dual main gearwheels exposed through a cut-out portion of the main gear doors. Manual extensionmay be accomplished in the event of a failure of the electrically operated system.The dual main gear tires and nose wheel tire are identical in size (6.50 x 10),but the main gear tires are 6-ply and the nose gear tire is 4-ply. Dual hydraulicbrakes are operated by depressing the toe portion of either the pilot or copilotrudder pedals. Shuttle valves permit braking by either pilot or copilot.

Photo 1-I 11. Mai, (;ear Wheels.

Annunciator System

13. The annunciator system consists of a warning annunciator panel (with redreadout) centraily located in the glare shield and a caution/advisory annunciatorpanel (CAUTION yellow, ADVISORY grecn) located on the center subpanel.

54

Adjacent to the warning annunciator panel on the glare shield is a press-to-testswitch to test the lights, and the pilot and copilot red WARNING and yellowCAUTION flashers. Individual function lights are of the word readout type. Inthe event of a fault, a signal is generated and applied to the respective channelin the appropriate annunciator panel. If the fault requires the immediate attentionv)f the pilot, the fault warning lights on the glare shield will flash. The flashing,I.lt warning lights may be extinguished by pressing the face of the light to resetthe circuit. The illuminated fault indication on the warning annunciator panel willremain on if the fault is not, or cannot be, corrected. If an additional fault occurs,the appropriate light on the annunciator panel will illuminate and the warningflashing light will again illuminate.

Fire Detection System

14. A fire detection system is installed to provide immediate warning in the eventof fire at the engine compartment. The system consists of three photoconductivecells in each engine nacelle, a control amplifier on a panel on the forward pressurebulkhead, two annunciator warning lights, placarded FIRE L ENG andFIRE R ENG, two fire extinguisher control switches, with lenses placarded L ENGFIRE - PUSH TO EXT, R ENG FIRE - PUSH TO EXT, located on the glareshield, a test switch on the copilot subpanel, and a circuit breaker placardedFIRE DET, on the right sidewall. Photoconductive cells, sensitive to infrared rays,are used as flame detectors. These cells are positioned in the engine compartmentsto receive direct and reflected rays, thus covering the entire compartment withthree cells. Heat level and rate of heat rise are not factors in the sensing method.The cell emits an electrical signal proportional to the infrared intensity and ratioof the radiation striking the cell. To prevent stray light rays from signaling a falsealarm, the control amplifier activates only when the signal reaches a preset alarmlevel, which illuminates the appropriate warning light in the warning annunciatorpanel. When the fire has been extinguished, the cell output voltage drops belowthe alarm level and the control amplifier resets. No manual resetting is requiredto reactivate the detection system.

Emergency Locator Transmitter

15. An emergency locator transmitter is installed on the right rear side of thefuselage to provide automatic emergency homing signals in the event of a crashor forced landing (photo B-i 2). The transmitter has a three-position toggle switch,ARM-ON-OFF, and the normal operation mode is the ARM position.

55

Photo H-12. Emergency Loc(ator rrannsmitler Switch.

56

APPENDIX C. INSTRUMENTATION

1. The instrumentation in the BAC Model 200 aircraft, serial number 71-21058,was installed, calibrated, and maintained by BAC personnel. In addition to theinstrumentation listed, the aircraft was equipped with a pitot-static boom whichincorporated angle-of-attack and angle-of-sideslip vanes. Photos C-I, C-2, and C-3show the instrument i.1-,v' p.otopanel, and oscillograph, respectively.

2. The left airspeed bo),,; .;:as a nonswivel type manufactured to Air Force FlightTest Center Drawiz7.g No. 5!'1DD261, Model No. 50-260. The right airspeed boomwas a swivel type designed and manufactured by BAC. Photos C-4 and C-5 showthe left and right airspeed booms. A list of test instrumentation showing themanufacturer, calibration range, and estimated parameter accuracies follows.

Photo C-1. Pilot Inistrumient Panel.

57

~~~1

rhoto C-2. Plxotopalnel

58

Photo C-4. Airspeed Boont, Left Side.

IP'hoto C-5. Airspeed Bo1om, Right Side.

59

kn 41)

0 4- 0. 0 ~ -'

0 0 0. CD -D

00 010 0 0W) 0n Cd Itu C 0 -

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-64

APPENDIX D. TEST TECHNIQUES ANDDATA ANALYSIS METHODS

GENERAL

I. This appendix contains some of the data reduction and analysis methods usedto evaluate the BAC Model 200 CEFLY LANCER aircraft. The topics discussedinclude level flight, glide and climb performance, and dynamic stability. Pastprograms generally developed a drag polar relationship for specific flight conditions.lHowever, the test points showed large deviation from the faired line at extremealtitudes (low versus high). The deviations were attributed to power effects whichcaused an apparent change in equivalent flat plate area (f) and Oswald's spanefficiency factor (e) due to differences in engine thrust at varying altitudes. Toeliminate these effects, the propeller feathered glide method was used to developthe base line drag polar for the BAC Model 200 CEFLY LANCER aircraft. Levelflight performance tests were conducted using the constant pressure altitude methodand the sawtooth climb method was used for climb performance. All test datawere converted into nondimensional coefficients which were used to develop thebase line drag polar and the final generalized equations. The equations were thenused to predict aircraft performance data at conditions not specifically tested.

Performance

2. The propeller feathered glide method 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 (I. I VS to VMO) (mnaximunL operating airspeed) was investigated for a targetaltitude band. The following tec~hnique was used to develop the base line dragcoefficient equation.

65

i -..

"= W o (Co

D = T + W sin(

DV = TV + WV sin 0

TV- DV (4)-V sin0 dh/dt w

66

Where:

L = Lift force (Ib)

W Aircraft gross weight (lb)

0 = Descent angle (deg)

T = Net thrust (Ib) zero in a descent

I) = Drag force = level flight drag (Ib) = net thrust required

V = Aircraft velocity on descent path (ft/min)

dh/dt = Tapeline rate of descent (ft/min)

Considering the drag and lift force equations and applying power-off glideconditions, the following relationship can be developed.

DcD (5)D qs

W sin 0(6)D qs

CL = q1 (7)

W cos 0L qs (8)

Where:

CD = Coefficient of drag

q = 1I/ p V2 (lb/ft 2 ) dynamic pressure

s = wing area ([t2)

CL = Coefficient of lift

The base-line coefficient of drag (CDBL) was then developed by plotting CD versus

CL 2 and fitting a first-order equation to the test points.

61

_f/

C1

SLOPE -

0 .

ACD

D o ACL 2 L (9)

3. During powered flight, the drag of the aircraft increased with thrust. To reflectthe change, the basic drag equation was modified.

Lc c - c (10)D D DPF -BL PF BL

Whcre:

-B = Increased drag due to thrust effect

CI )1,1 : [otal coetticient of drag for powered flight

Base-line coefficient of drag

68

Coefficient of' thrust (Tc'), thrust (T), thrust horsepower (THP), and shafthorsepower (SHP) were calculated as follows:

T 2T (11)

PSVT

T=550 x THP (12)T

F x VTHP = nl x SHP + -5 T (13)p 550

SHP = Q x N x 30 ) (14)

Where:

Tc' = Coefficient of thrust

T = Thrust (Ib)

p = Air density (slug/ft 3 )

S = Wing area (ft 2 )

VT = True airspeed (ft/sec)

THP = Thnrst horscpower

7p = Propeller efficiency

SHP = Shaft horsepower

Fn= Jet thrust (ib)

Q = Engine torque (ft-lb)

Np = Propeller speed (rpm)

69

"rhe 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 fittingwas used,

APF-BL+

0 - T •

AC A T' ,2 + B TC' + C (15)DPF - BL C C

From equation 10,

CD CD + CDPF BL PF -BL

or

CD + A TC' + B T ' + C (16)DpF C CBL

lEquation 16 represents the generalized equation for all level flight and climb Iperformance in either single or dual-ngine operation.

70

4. Level flight performance tests (single and 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), andthrust (Tc') were obtained from the recorded test data to determine the coefficientsfor the generalized equations.

5. Climb performance tests (single and dual-engine) were conducted using thesawtooth climb method. All dual-engine climb tests were conducted with bothengines operating at MCP. All single-engine climb tests were conducted with theleft engine operating at flight-idle and propeller feathered, while the right enginewas operating at MCP. The aircraft was stabilized and trimmed at incrementalairspeeds from 1. 1 to 1.8VS for ± 1000 feet of the target altitude. The tape-linerate of climb and coefficients of drag, lift, and thrust were obtained from therecorded test data to determine the coefficients for the generalized equation.

6. The shaft horsepower available, fuel-flow rate, and net thrust of a PT6A-41specification engine, including all installation losses, are presented in figures 52

through 55, appendix E. Figures 56 and 57 present the engine inlet pressurerecovery data which were furnished by the airframe manufacturer. TheUACL-furnished computer deck was used to calculate the performance for aninstalled specification engine. The computer deck is based on the minimumperforming engine that has the maximum allowable time before overhaul. For thisreason, the calculated aircraft performance data, which are based on a specificationengine, were always less than the observed test data. The test engines, serialnumbers X70014 and X7001 1, 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 chart was furnished by BAC and is presented in table D-l.

Ambient test temperatures (Ta) were obtained by correcting the indicatedtest temperature (Ti) for instrument error (ATic) and for compressibility (ATC).

T = T + AT + ATa i c c

8. Pressure altitudes were obtained by correcting indicated pressure altitudes(Hpi) for instrument error (AHpic).

Hp HPi + AHp~c

71

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,T,'C A.'SPEZ- (KNOTS)J -,-, A-', ,',/v:CL .19,4 77O

Table D-I. Propeller FEfficieuiey.

72

9. The density ratio (a) was determined from the following relationship:

a- (To/Ta)(Pa/Po)

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.

11. True airspeeds (VT) were determined from the test altitude air density ratio(a) and calibrated airspeed as follows:

VealVT -

Airspeed Calibration

12. 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. 58, app E). Calibrated airspeeds (Vcal) were obtained bycorrecting indicated airspeed (Vi) for instrument error (AVic) and position error(AVpc).

Vcal = Vi + AVic + AVpc

"Weight and balance

13. The aircraft weight and longitudinal cg were determined prior to each weightand/or cg configuration change. A typical internal ballast loading is shown inphoto D-1. Weighing was accomplished using electronic scales located under theaircraft jack points with the crew on board at their designated stations.

13

............

AI!//

Photo D-I. Internal Ballast.

I)yNaniic Stability

14. Dynamic stability characteristics were tested by using the techniqcles describedin references 9, i1, and 13, appendix A. The data recorded from dynamic testingwere presented as time histories of the pertinent parameters tiat describe themotion of the aircraft. Analyses of these time histories were performcd to determinethe resu~lting damping ratios (') and damped natural frequencies (cod). The dampednatural frequencies and the damping ratios were derived by two iuiwthods for allthe v onditions tested. These were the logarithmic decrement and time ratio method.

15. The undamped natural frequencies (con) of the motion in radians per secondwere calculated from the following equation.

W 2i..• =•d t 1 •n14

14

APPENDIX B. TEST DATA

INDEX

MFigture Number

Performance

Climb Performance I through 13Level Flight Performance 14 through 23

Handling Qualities

Control System Characteristics 24 through 27Static Longitudinal Stability 28 through 35Static Lateral-Directional Stability 36 through 40Dynamic Longitudinal Stability 41 and 42Dynamic Lateral-Directional Stability 43 through 46Maneuvering Stability 47 through 49Roll Performance Characteristics 50 and 51

Miscellaneous Engineering Tests

Engine Characteristics 52 through 57

Airspeed Calibration 58

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-13

APPENDIX F. DEFINITIONSABBREVIATIONS, AND SYMBOLS

This list includes most of the symbols used in this report, However, certain portionsof the report use special or unusual abbreviations and symbols. The meaning ofthese is made clear in the text of the report and, when that is the case, theabbreviation or symbol will not be found in this list. Also, certain symbols havemore than one meaning, however, the context should make the meaning clear.

Symbols andAjbbreviations l)efiition Unit

ANA Air Force Navy Aeronautical

AC Alternating current

) Wing span teet

Minimum coefficient of drag of thepropeller feathered drag polar

CD Coefficient of drag

CDBL Base-!ine coefficient of drag

CDPF Powered flight coefficient c' drag

SCp Coeffi ient of power

C1 Coefficient of lift

Cont Continuous

D Drag

lDe Degree °C

e Oswald's span efficiency factor

f Equivacnt flat plate area ft2

FN Jet thrust pounds

SAcceleration. of gravity Itt/sec-

Il) l)ensitv altitude 'ec l

134

I lp Indicated pressure altitude feet

lip Pressure altitude feet

Ilpic Instrument corrected pressure altitude feet

J Advance ratio

L Lift pounds

MAC Mean aerodynamic chord -

Max Maximum

MCP Maximum continuous power

Mill Minimum, minute

Np Propeller s i rpm

N( Gas producer speed percent

N-2 Power turbine speed rpm

NAMPP Nautical air miles per pound of fuel --

NU Nose up

ND Nose down

OAT Outside air temperature T

p Roll rate radians/sec

t'a Ambient pressure in. of mercury

P) Standard-day, sea-level pressure in. of mercury

psi Pounds p--r square inch lb/in. 2

q I)ynamic pressure lb/ft2l

Q Torque ft-lb

ref referred, reference

R/C Rate of climb ft/min

135

S Wing area ft 2

SE Single engine

SliP Shaft horsepower

SL Sea level

S/N Serial number

S'[lI) Standard

Ta Ambient air tcmperature 0C

TC' Coefficient of thrust

Ti Indicated air temperature

T Thrust lb

Tic Instrument corrected on temperature °C

THP Thrust horsepower liP

To Sea-level, sta'ndard-day statictemperature OK

UHF Ulltra high lrcttuIencv

Vcal Calibrated airspeed knot

VHF Very high frequency

Vi Indicated airspeed knot

Vic Instrument corrected airspeed knot

VT TruC airspeed knot

VMC Airspeed for minimum control knot

VS Stall airspeed knot

VH1 Maximuma airspeed for level flight knot

VMO Maximum operating airspeed knot

V True airspeed ft/sec

136

Wa Elgine airflow lb/hrW Weight

pounds"C Degrees Centigrade degreesOF Degrees Fahrenheit

degrees

OK Degrees Kelvin degrees

Difference

CDpF - BE Difforence in coefficient of dragdue to thrust effects

AVpC Airspeed position error correction

I)amping ratio

0 Temperature ratio, descent angle -, degrees

Pressure ratio

(3 Density ratio

P Air mass density slug/sec3Wdt D~amped natural frequency radian/sec

)n Undampled natural frequency radian/sec

a Angle of attack degrees

,1, Roll or bank angle degrees

7p Propeller efficiency

Roll to yaw ratio

dh/dt Tape-line rate of descent ft/.,it,

7r 3.1415)

qin. Inlet duct efficiency percent

137

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