AI AA-88-0093 Design of a Rotary Engine-Powered Four Place Aircraft J. Smith and D. Stevenson, University of Washington, Seattle, \NA

AlAA 26th Aerospace Sciences Meeting January 11-14, 1988/Reno, Nevada

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L’Enfant Promenade, S.W., Washington, D.C. 20024

DESiGN OF A ROTARY ENGINE-POWERED FOUR PLACE AiRCFiAFT

Jared Smith Doug Stevenson

University of Washington Seattle. Washington

A W L a t

This paper addresses the requirement for a new technology general aviation personal aircraft. A detailed development is presented of the Faiconair t60TR. an affordable. high performance aircraft intended to compete with used aircraft sales. The performance advantages of utilizing a highly advanced rotary Combustion engine include high power to weight and power to volume ratios, low specific fuel consumption, and multi-fuel capability. Cabin comfort is enhanced by the inherent low vibration of the rotary engine. Surface finish and dimensional control is improved due to construction methods incorporating advanced composite materials, allowing the use of new NASA tow speed airfoils. Sound and vibration absorption along with lower manufacturing costs enhance the design. The aircraft was designed to be aesthetically pleasing to the customer.

liawwma

aspect ratio effective acceleration(fWsec2) tail lift curve slope per degree wing lift curve slope per degree engine brake horsepower coefficient of drag coefficient of drag at obstacie(50 11) coefficient of drag of flaps coefficient of drag of gear wing drag coefficient change in hinge moments with change in a change in hinge moments with change in 6 wing lift coefficient coefficient of lift in around attitude

Qmax maximujm wing lift coefficient CLSO coefficient of lift at obstacie(50 11) CM

C C

c,

CR D e f 90 H i J K KF L N n n ' Pa PO Pr R

- q Re R i

coefficient of moment coefficient of propeller power wing chord (11) mean aerodynamic chord (11) wing root chord (ft) drag (Ibs) Oswald's efficiency factor equivalent fiat plate area (It2) sea level gravity (fUsec2) height of transition (it) propeller advance ratio ground effect factor flap torsion factor l i f t (Ibs) rotational velocity (revolutionstsec) load factor excess load factor (1 .57) power available (hp) engine power (hp) power required (hp) range @mi) dynamic pressure (lbs/ft2) Reynolds number radius of transition (It)

rate of climb wing planform area (112) specific fuel consumption(lbs1hr hr) air distance travelled during landing (ft) total landing ground roil (ft) horizontal distance traveled during climb (It) flap effective area (ft2) free roll distance traveled before brakinglft) total takeoff ground roll(ft) landing field length (ft) distance traveled during transition(tt) total takeoff distance(1t) thrust (blsaO velocity (ftlsec) horizontal tail volume

Vst Vso

stall speed in takeoff configuration(fWsec) Stall speed in landing configuration(ft1sec) velocity at touchdown(ft1sec) velocity at 50 ft obstacle clearance (fWseC) weight (ibs) gross weight (Ibs) wing weight (ibs) distance from c.g. lo the a.c. distance from c.g. to root quarter chord distance from c.g. to the neutral point angle of attack (deg) angle of attack at zero lift(deg) elevator angle (deg) downwash angle (deg) propeller efficiency wing flap efficiency wing taper ratio Coefficient of fraction on ground transition anale Ideo) " I "I

density elevator effectiveness factor

This study illustrates the design process of the Falconair 16-TR rotary engine-powered tour place aircraft. The aircraft design study covers evolution of the aircraft from initial to final specifications. and included considerations of fuel efficiency, low cost, customer appeal, passenger safety and confort. The advanced rotary engine incorporated into the aircraft allows low cost alternate fuel usage, increased fuel efficiency with low vibration and sound levels. Composite airframe construction. dictated by the configuration contour, will reduce construction, manufacturing and operating costs. The increased skin rigidity and manufacturing tolerances obtainable with these materials and construction methods allows effective use of new NASAiow speed airfils. These airfoils, designated the LS(1) family, are designed expressly for use in general aviation dtype aircraft. Use of the airfoils, coupled with the Fowler flap design, results in significant aircraft performance due to their inherent high lift and low drag characteristics as compared to older NACA series airfoils Performance analysis of the Faslconair 160TR design determined SDecifications that were an imwovement over spec I cat ons o.ti ned as to achieve compktilweness n Ihe general aviation market

t

The Falconair 160TR was designed as a personal, four place, advanced technology aircrafl intended to be sold in the depressed general aviation market. Preliminary performance specifications were outlined to achieve competitiveness within this market, while providing a means of fast, convenient. and inexpensive transportation. The initial goals were as follows:

Table 1 lists specifications for possible rotary engines for use in the aircraft. The John Deere model 1007R engine was selected as the powerplant upon which the airplane design was based. Table 2 contains a description of the engine.

Table 2. Specifications for John Deere Model 1007R

W

speed: 150 knots (Growth version Availability est. 1994) Range:

Field Requirements: 2000 ft. maximum Displacement L(in3) 0.7 (40) Meet Federal Avaition Regulation Part 23 for normal Power kW(BHP) 120 (160) and utility aircraft Rated Speed RPM 8,000 Inexpensive production and operation Weight kg(lbs) 90 (198)

Soecific Weioht ko ikW 0.75

600+ nautical miles with 30 minute reserve @ 5000 ft msl 1 rotor

Desian S e l e c m

Powerolant.

The type of powerplant used in this design is a rotary engine. The choice of a specific rotary engine is based on a limited selection of available engines since rotary engine manufacturers have not yet produced a design specifically for the general aviation market. New technology engine needs that can be met by an advanced rotary engine include the following.

Compared to Piston Power: Burn alternate fuel High HPiweight ratio Smooth operation Small frontal area

Compared to Turbine Power: Lower initial cost Improved altitude capability Low fuel consumption rates

vblume m3(<3) 5.150 (5.29) Takeoff BSFC 0.38 Cruise BSFC 0.36 Liquid cooled Multi-fuel (JP4. JP5, no. 1 and no. 2 diesel, alcohol)

Performance L imitations

Given the prescribed rotary engine choice, limitations were determined for' the purpose of narrowing down the infinite number of possible aircraft configurations to only those which would satisfy the initial performance requirements. The design selection chart shown in Fig. 1 consists of five sets of lines representing stall. cruise, takeoff. landing, and range performance limitations. These lines enclose an area on the chart representing the combinations of gross weight and wing area which were possible for the given set of performance requirements.

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Federal Aviation Regulation (F.A.R.) Part 23 requires General Needs: Reliability that stall speeds for single engine aircraft not exceed 61

e Long service life knots. Using three estimated maximum lift coefficients, the corresponding maximum wing loadings were generated.

The use of this type of high technology rotary engine results in a significant improvement in performance over existing and candidate new technology engines. NASA, Cessna, Beech and the majority of current aircraft engine manufacturers

w 1

s 2 - PV,,2 CLmax ( 1 1 - =

recognize these advantages and are pursuing usage of the stratified charge rotary engine for general aviation.

Table 1. Rotary Englnes

Specification John Deere Wedtech Telidyne model 1007R Wankel Continental

GR-36

Number of rotors

Takeoff S S F c Cruise BSFC Cooling Fuel Induction

0.7 (40) 120 (160) 8,000 90.0 (198) 0.75 0.38 0.36 Liquid Multi-Fuel2 Turbo-charged Supercharged

1 Without reduction drive. 2 CITE, JP4, JP5, alcohol, no. 1 and no. 2 diesel. 3 Version 2 uses JP4, JP5, and diesel

z z 0.816 (49.8) 0.588(35.8) 89.5 (120) 63.3 (85.0) 7,000 7,500 28.6 (63)1 50.0 (1 10) 0.3191 0.787 N.A. 0.50 N.A. 0.48 Liquid Liquid Gasoline3 Auto Gas. Normally Normally Aspirated Aspirated

2

By equating thrust lo drag in cruise flight, the maximum wing loadings for cruise performance were generated for three aspect ratios by:

EHP qp 325 w2 = fq + 1.3cowqs + - ( 2 ) -

V xq(AR)Se

Note that consideration of empennage interference and cooling drag were approximated as 30% of the wing contribution as this was a preliminary requirement. The maximum wing loadings for established landing performance were generated for three aspect ratios. The restrictions included: a rate of sink less than or equal to 1000 fVmin over the 50 foot obstacle and use of no more than 60% of a 2000 fl. landing field. The landing field length is defined as:

SLFL = (SA + SFR + Ss) i 0.60 ( 3 )

Using the balance of potential and kinetic energy of the aircraft above a 50 ft. obstacle and the ground, the following equation was derived:

( 4 )

where:

L CLmaxil.69

D f - =

( C m d ,6912 TA

S n(AR)e qs . - - + 1.3cDW + A c ~ g + ACDI +

- At touch down, a two-second free roll was assumed lo allow ground attitude lo be attained before brake application. Free roll distance is defined as:

( 5 ) SFR = ~ V T D

Distance covered on the ground during brake application assumed a constant effective deceleration, which was defined at 0.707 !ITD, The equation for braking distance is given by:

- q (L + 1 . 3 C ~ w + ACog + ACof + w s

Figure 2. Landing

Maximum wing loadings for takeoff performance, were generated for three aspect ratios, with the reStriCtlOn Of clearina a 50 ft. obstacle at the end of a 2000 ft. field. The [akeoffbistance is defined as:

STO = S G + S T + S C ( 7 )

Distance covered on the ground before rotation was calculated at a constant effective acceleration, which was defined at 0.707 VTO. The equation for ground roil is calculated by:

vT02

SG = - (8) 2aen

Figure 1. Preliminary Design Performance Limitations

3

Aircraft transition and climb distance are calculated with the following equations:

ST = R T O ~ ( 9 )

'TO2 RT = - 0.521 go

where:

t.3Cow + ACDQ + ACDI + x(AR)eK

W VTO

O T =

t.3Cow + ACDQ + ACDI + cLG2 - \I7,. dARleK . .

\ - . I "

W O T =

and ST @T

HT = - 2

+\ V-(l.3)Vai I \ R \

Figure 3. Takeoff

The preliminary range calculation was determined by the Breguet range equation with propeller efficiency, specific fuel consumption and velocity held constant for three aspect ratios. The derivative form of this equation is given by:

325qp L dW dR = - --

SFE D W

where:

D = WZ + n(AR)eSq2 (1.3ScDW + ftmdy + fgear)

n(AR)eSq

The above equation was integrated numerically and successively iterated to obtain the minimum allowable wing loading which met the range objective.

The baseline configuration was chosen from within the allowable region of the aircraft performance limitations.

wing planform area of 100 f t2 to be used as the basis for further design choices.

This configuration yielded a gross weight of 2300 Ibs and a w

This configuration matrix depicted in Table 3 was used in order to define an airplane which would satisfy performance and designers' expectations. The conventional wingttail arrangement was chosen due to distinct advantages indicated in the matrix and the designers' preference.

Enaine Placement.

The liquid cooling of the SCORE 70 1007R rotary engine allows integration of the powerplant anywhere within the airframe. Placing the engine in the tapering aft section of the aircraft allows greater forward visibility and a distinct appearance. Final aircrafl layout drawings indicate the advantages of the pusher design. Olher reasons for choosing an aft engine, pusher arrangement include:

a cabin noise isolation e propeller isolated from cockpit access

engine fire isolation from cockpit ease of fuel line routing exterior to cabin

Accurate determination of center of gravity travel is an essential part of the aircrafl design. Center of gravity travel of the Falconair 160TR is restricted to less than 20% Mean Aerodynamic Chord (MAC) for handling considerations.

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If the center of gravity were located aft of the neutral point, the airplane would be inherently unstable. This neutral point is found where the slope of the pitching moment coefficient versus lift coefficient is zero. The stick free neutral point is obtained directly by:

Xnp Xac dCM _ = - . - c c dCL aw

The percent Mean Aerodynamic Chord travel of the c.g. is determined by:

Lcg . xac %MAC = 100

C

Table 3. Aircraft Configuration Matrix

I Configuration n Advantages I Disadvantages il I

Provides e~~el lenl longitudinal stability

Oustanding lowaid Vsib~lity O w n load on tail with high wing lilt PlOVB" CUSfOrnD, appeal

i ~~~~~

High Stalling s p e d Pwr fuel storage With center

Inadequate loward visbiliiy

Higher drag ihcrdse d m to mulnpls surlaces P w r faward visibility

MOdern popularity Win@Canard Tail am as iimng wr~am 01 gravity location

A11 surfaces lilting

due to limng S U ~ ~ ~ C B S W

4

Figure 4 depicts the center of gravity envelope achieved through minimization of the c.g. travel. The total c.g. travel is 15.6% of the MAC. The main gear is located such that the nose wheel returns to the ground if the airplane is displaced on the ground when empty. Centers of gravity of basic empty and gross weights of 1253 and 2300 pounds are 110.0 and 97.7 inches aft of datum, respectively. v

2300

2100

1 9 0 0 ~ ~

1 7 0 0 ~ ~

1500~-

1300.-

wo . 0-

0 5 10 15 20 25 30 35 40 % MAC

Figure 4. Center of Gravity Envelope

A family of modern low speed airfoils modified expressly for general aviation applications was chosen for a comparative study. These airfoils, designated the LS(1) and LS(1) modified series, were chosen due to their extremely high lift and low drag characteristics. The design lift coefficient of 0.4 for these airfoils is Suitable for the Falconair 160TR.

Tests of the LS(1) airfoils conducted by Langley Research Centers indicate maximum lift coefficients that are substantially greater than the older NACA airfoils of comparable thickness ratios. These results are shown in Figure 5. Tests of the airfoils were conducted over a Mach number range of approximately 0.10 lo 0.28 and a Reynolds number range from approximately 2 x i o G to 9 x I O G . aThe geometric angle of attack was varied from -10" 10 22". Further results obtained are as follows:

v

The LS(1)-0417 modified airfoil provides the best performance of the six airfoils tested.

Increasing the airfoil thickness ratio results in an average increase in drag wetlicient of about three drag counts (0.0003). for each percent increase in thickness ratio at the design lift coefficient with a fixed transition near the leading edge.

At a Reynolds number of 6 x IOG and fixed leading edge transition, the 2-D lift to drag ratios at design cruise are approximately 46, 40 and 35 for the 13-, 17.. and 21- oercent unmodified airfoils resoectivelv. For the

a

modified series, results indicaie ratios' of 57. 59, and 57, respectively.

Maximum lift coefficient is insenstive to roughness

The optimum aspect ratio was chosen for the best combination of speed and range using the equations below, where gross weight. wing area, thickness and taper ratio are held constant. From equation 15 wing weight and thus fuel load are computed. Speed is then determined using equation 2 and the resulting range determined from equation 12. v

Ww = L F ~ F S + A)[., + (1+7?.)(AR)

80.000 1 +A 16 Vc

From this procedure a plot of range vs. speed for various aspect ratios was produced (See Figure 6).

Re = 6.0 x 10 0 LS(1)-04XX family

+ NACA230 series

Q max 19 1.6

.~

+ * NACA44 i + series

2 $ + - NACA24 x series x + 2 X NACA65

series

1 , 2 k - v A o 0.04 0.08 0.12 0.16 0.2 0.24 0.28

tic Figure 5. Comparison of Maximum Lift Coefficient

155 -- Speed : (Mas) :

150 -.

1351 - . . : . , . , I , , , , ; , , , : I . ~ . . : l ~ ~ i : r , I I : . . . .

820 840 860 880 900 920 940 960 5 Range (naut. miles)

Figure 6. Aspect Ratio Sensitivity

Based on a design decision of improving on the mission requirement for a cruise velocity of 150 knots. an aspect ratio of 8.5 was chosen. This decision resulted in a preliminary cruise velocity of 153 knots and a range of 896 nautical miles.

A similar procedure was used for thickness ratio optimization. From section data of the chosen family of airfoils. 2-D maximum oift coefficients are converted lo 3- D by:

0

where: s, 42 5 (flap affefleo area, Dl 75 ('lap elfcency

Using equations 1. 2, 12, 15. and 16 the velocity and corresponding range for the three thickness ratios considered were determined (See Figure 7). From these observations. the LS(1)-0417 modified airfoil was selected as the preliminary choice for the final wing design.

5

Throughout the calculatioils for wing design, it is assumed

chord (MAC) AGiVEN BY:

2 ( P + ? . + l )

3 (X + 1) L’ MAC = c = -CR ( 1 7 )

where:

ZS”2 CR =

0-

ARlQ(X + 1)

148 Location of the aerodynamic center with the respect to the quarter chord of the root Section is found using equation 18, below. Mean aerodynamic chord and the location of the aerodynamic center are depicted in Figure i o , along with final Parameters determined through wing design,

840 850 860 870 91° Range (naut. miles)

F igure 7. Th ickness Ratio Sensitivity

( 1 8 )

MAC = 2.54 n NACA 0012 section

-----.+- ~ , , , ,

F igure 8. Variat ion of C, max wi th Reynolds Number

rri W 2.5

0 50 100 150 200 250 2.- Station (inches)

Figure 10. Aircraf t Des ign Parameters

Q

The selection of the flap system has a profound impact on aircraft performance. By allowing a smaller wing for i

0.5.- a given gross weight and field length requirements, cruise 9 performance is greatly enhanced. Modest power and high

cruise speeds necessary to minimize operating costs make flap performance critical. These considerations resulted in the decisin to incorporate an 80% span, 30% chord, full Fowler flap system. Although this choice adds complexity to systems design, the increased performance benefits outweigh negative aspects (See Table 4).

. e [C, Distribufion_(CC=a

x/b/2 ~i~~~~ 9, spanwise Sect ion Lift Distr ibut ion

Table 4. W i n g Flap Charac tar is l l s

Re = 4 x 106

Airfoil Configuration Clmax Cl(oa) Cdmax Cd(Oa)

Basic LS(1)-0417 MOD Section 1.75 0.42 0.028 0.008

25% Plain Flap

25% Slatted Flap

30% Fowler Flap

10 deg 2.05 0.65 0.155 0.055 40 deg 2.50 1.25 0.20 0.115

10 deg 2.75 1.55 0.145 0.036 40 deg 3.40 2.44 0.172 0.135

10 deg 3.03 1.35 0.12 0.023 40 deg 3.85 3.02 0.173 0.095

v

6

Longitudinal stability and control are dependent on the locations of the aircraft center of gravity, aerodynamic center of the wing and aerodynamic center of the tail. The analysis is based on the Summation of moments about the center of gravity. Requirements for stability and control are dictated by the worst case for each: stick-free and stick- fixed, respectively. The design met the required minimum 5% static longitudinal Stability margin,

-

v

The horizontal tail surface and elevator arrangement must be designed in such a way as to satisfy the equilibrium flight condition (CU = 0) for a desired lift coefficient. The limiting case for control occurs during the landing phase, as this is when maximum lift and thus maximum moment occurs. Airplane longitudinal stability change due to downwash at the tail produced by the wing is significant. Downwash calculations at locations forward and aft of the wing were carried out by applying the Biot Savart law with a double horseshoe vortex modeling the basic lift distribution of the wing. Ground effect was included in the control analysis, where the limiting case for control is the landing configuration

Using the results obtained from the above procedure, the pitching moment contributions of the wing, body, and tail were calculated. Stabilizing propeller moment contributions were ais0 included. The sum of these individual contributions results in a stable aircraft as indicated in Figure 11.

. .

Fuselage -Wing

0 025 0 5 075 1 - c M

0 5 -025

Figure 11. Pitching Moment Contributlons

incorporating a fully flying stabilator with a geared tab results in reasonable tail volumes for the required stability and control. The minimum horizontal tail volume i S determined wi !I Figure 12.

Foward CG

Stability Requirement Requirement

-0.5 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 I

xaic Figure 12. Tail Volume Sizing

The lateral yawing moments developed because of sideslip or yaw must be compensated by an oppostie moment created by the vertical tail and other positive contributing airplane components. The procedure is then to sum up the stability contributions of the component parts of the aircraft. Directional stability is then determined by amroximations for each comoonent in terms of the rate of ~ ~ ~~~ ~ ~ , ~ , ~ ~ ~ ~ ~ ~ ~~~

change of yawing moment coefficient per degree angle of yaw.10 The resulting tail area is 11.8 R2 and the required stabilator dihedral angle is 25".

Aircrafi layout drawings include two figures: a three view, and a cut away plrofile view.

Composite construction of the aircraft is intended. allowing construction of a three-dimensionally contoured aerodynamic configuration which would not be feasible with an aluminum skin. Fiberglass and epoxy or vinylester fuselage and wing skins, along with a graphite composite and epoxy wing spar will result in an airframe weight considerably lower than present aircraft. Total component weight of the fuselage, wing and empennage is under 410 pounds. By reducing the number of parts in the aircraft. assembly costs will be greatly reduced. Partially offsetting this saving will be the increased cost of tooling required as opposed to aluminum construction.

Table 5. Tall Configuration Matrix for Pusher Type Alrcraft

7 Advantages Disadvantages

I I1

7

Seating arrangement was determined with passenger comfort in mind. The two aft passenger Seats are positioned such that they are facing toward therear of the aircraft. This arrangement was considered for the follwing reasons.

Provides easy communication between front and real passengers More efficient use of space, (legs of the rear passengers tie in the tapering section of the al l body) Passenger luggage can be placed between front and rear back to back seats Excellent center of gravity placement Survivability of the aft facing passengers is increased in the unlikely event of a crash

Cockpit accessibility is enhanced by the innovative entry design. The lower section of the door folds out and down providing a convenient step-up platform for both front and rear access. Pilot and co-pilot enter the forward seats of the airplane via the split opening doors on either side of the aircraft. Aft passengers utilize the same step arrangement to gain access to the single rear door over the left wing.

Fixed landing gear in the baseline configuration provides low cost and maintenance, minimum weight and complexity, and increases safety as the chance of gear up landings are eliminated. Drag due to the fixed landing gear was minimized by utilizing faired wheel pants and a NACA 632 -015 airfoil section for the struts.

Engine maintenance. is simplified by accessibility to the engine compartment. Access panels are provided above and on either side of the engine for periodic and major engine overhaul maintenance.

The powertrain includes an extended propeller shaft and a planetary gear reduction unit. The gear reduction System is intended to be manufactured of light weight materials. A fixed pitch propeller was chosen for low cost and light weight.

Control Systems for the aircraft are designed to use common cable-pulley, bellcrank-torque tube and push-pull rod combinations.

Radiator placement is governed by its proximity to the engine compartment in order to minimize coolant plumbing. Underbody location of the engine air intake and exhaust nozzle provide adequate coolant air mass flow rate with minimal drag effects.

Performance

Drag analysis included a drag breakdown of various components of the airplane. Items of the airframe included in the analysis were: fuselage, landing gear, wing, tail and flaps. Interference effects due to the wing on the body, body on wing, and empennage were accounted lor,” along with cooling drag of the radiator design.’D,’‘ This study yielded the following results:

fbody 0.80 (fuselage only) faear 0.15 (struts, nose, main) 6cDwing 0.0075 DCDlail 0.0025

0.110 DCDnaps DCDint 0.0052 D C b l i n g 0 . 0 0 2

Aircraft performance was determined by the power available from the engine. The power available for flight at various airspeeds was the product of the power delivered to the propeller and the propeller efficiency, i.e.,

(19) Pa = Psllp

The choice of a fixed pitch propeller dictated the calculation of propeller efficiency for varying airspeed and propeller rpm. The method of calculation is as follows:

A blade angle of 25’ was chosen in order to achieve the best cruise performance with an acceptable takeoff distance. The power to the propeller was assumed to be directly proportional to the engine rpm. Coefficients of power are calculated for the rpm range of interest (2300 - 2700) by:

P k cp = - =-

p n3 D5 p n2 D5

From a plot of Cp vs. J for an RAP-6 two bladed propeller,2’ valaues of J corresponding to each Cp were found along the line of constant blade angle. A plot of h vs. J determined values of propeller efficiency.2’

e Power available is then determined by:

( 2 1 ) Pa = l lP

The above procedure produced the power curve up to the maximum rpm. Beyond this point the constant and the efficiency was a function of velocity only. The power required to maintain altitude is determined by:

Figure 13 illustrates power curves for sea level and 10.000 feet.

Rate of climb is found from:

550 (Pa - Pr)

WO

wc =

250 -

/ zoo Pr

full power

cruise power a-100

50.-

150 / full power

cruise power a

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50 100 150 ZOO 250 300 350 400 i

Velocity (Wsec) Figure 13(a). Power Curve at 10,000 ft Altltude

8

o c , , I , , , , I , 1 , I : , , , : , , ~ ' I ' " ' " " ' ' " 50 100 150 200 250 300 350 400 450

Velocity (Wsec)

Figure 13(b). Power Curve at Sea Level

Figure 14 depicts maximum rate of climb with altitude. The final performance envelope is shown in Figure 15.

25rMO -- I

Slope due to fixed pitched propeller

Turbocharger limit

W

3 toooo

5000

Slope due to fixed pitched propeller

Turbocharger limit

W

toooo

5000

0 0 100 200 300 400 500 600 700 800 900 1000

WC (Wmin) Figure 14. M a x i m u m Rate of C l imb with Alt i tude

v

W U .s c E z

Figure 15. Aircraf t Performance Envelope

--.

Performance parameters were determined and used to calculate landing and takeoff performance. These values are:

Landing C ~ m a x 2.66

* DC5ea3 0.00125 DCDfiaps 0.11 CL (in ground altitude) 2.0

a Aspect ratio 8.5

a Oswald efficiency factor 0.75 Effective wing area with flaps(sq fl) 0.75

v Takeoff e Weight (Ibs 2300.0 Wing area (fP) 100.0

C h a x t .a

f (body.radiator) (1P) 1 .o a f (landing gear) (fP) 0.15

DCDnaps 0.0 aspect ratio 8.5 Oswaid efficiency factor 0.9

s CL (ground attitude)

%wing 0.0075

1 .o

The design goal to produce a general avaiation aircraft design that would be atlractive to both owner and pilot has been accomplished. The design of the Falconair 160TR has proven that it is possible to build a low cost. high performance personal aircraft that is as distinctive in performance as it is in appearance.

€lame

Range is determined using equation 12, where:

D = W2 + n AR e S q2 (body + fgear + ACDwing t ACCcooling + AcDint + AcDtail) ( 2 4 )

Performance values for landing, takeoff and range of the Falconair 160TR are listed in the specification sheet given in Table 6.

Table 6. Ai rcraf t performance speci f icat ions

Est imated equipped price

Seats (incl. pilot)

Engine Manufacturer Model Power EHP Critical Altitude feet

Design we igh ts kg (Ib) Max gross Basic empty Useful load Zero fuel

Dimensions feet Length Height Width Wingspan Wing area feet2 Aspect ratio

Wing Ibsifoot2 Loading

Power ibshhp

Per formance (SL, ISA, 100%) Maximum rate of climb fpm Best rate of climb - Vy ktas Maximum speed - Vmax Mas Minimum speed - Vmin Mas Stalling speed, clean ktas Stalling speed, flaps ktas Range nm Absolute ceiling feet Design Cruise Altitude feet

Operat in field incl 60% safety foot o%staciej . Takeoff distance feet Landing distance feet

~~

$ 55.000+

4

John Deere Score 70 1007R 160 8,000

1,043 (2,300) 544 (1,200) 499 1,100) 897 11,977,

19.5 8.0 3.6 2 9 2 100.0 8.5

23.0 14.4

861 .O 803.0 170.0 69.0 64.0 49.0 920.0 20,000.0 8,000

margin and a 50

1,966.0 1.61 3.0

9

3. Jones, C., and Remount, RE&LLQI a Hiah P- m t a r v Stratified Cha rae Aircraft Enoing, A lAA paper,

The performance analysis of the Falconair 160TR 84-395. June 1984. rotary engine-powered four place aircraft confirms a design that should be extremely competitive in today's personal aircraft market. Aircraft performance meets or exceeds all specifications determined for effective competitiveness within the general aviation market. Use of the John Deere SCORE 70 advanced rotaly engine and composite construction improves performance and reduces operational costs over

pleasing aerodynamic configuration. Use of the new generation NASA LS(1) series airfoils and a Fowler flap design also resulted in significant performance advantages over currently available general aviation aircraft that incorporate the older NASA series airfoils.

4, w,H,, ~ i ~ ~ k ~ , K , ~ , , and

a 10% S l o m , NASA Contractor Report 3081, 1979.

v

-, NASA . . 5 . McGhee. R.J., Beasley, W.D., currently available aircraft. while allowing an aesthetically 210, TK

TM 78650. 1978.

f the GAIWl -2 Airfoil EY&bl&! Flao. 30% Fo V&J Flaa

6 . Wentz. h 200 W.H.. Aileron Wind Tunnel Tests 0

and a 10% Slot-LipSMiler, CR-145349, 1979.

7 . McGhee, R.J.. Beasley, W.D., WQ& Wind T ! z u ~ l > led 13 P-, NASA~

TM x-74018. ~~ ~

. . 8 . McGhee. R.J., Beasley, W.D., A e r o d v n a m i c t e r i s t i c s of an in- Familv of Alrloils for General Aviation Amlicationg.

rl 1 1 . Hoerner, S.F., EMdJynamic Draa. , Hoerner Fluid

12. Hoerner. S.F.. i ' , Hoerner Fluid

13 . Kays. W.M.. London, AL, v,

Dynamics, Brick Town, NJ, 1965.

Dynamics, Brick Town, NJ. 1985. L'

National Press, 1955.

14. Eckert and Drake, -. McGraw- Hil l .

15 . Federal Aviation Regulations. Part 23 - Airwort- ,-- d A- m. International Aviation Publishers, Inc.. WY,

Revised February 1984.

. . 16. Irving, F.G.. lo the v of LOW S@2&Aiu& Pergamon Press Ltd.1

1966.

17. NACA L-212 Wartime Report, E a p e r i m e n W

Iails.

1 ,

2

. . References 18 . NASA-CR-174812, Rotarv T- r

Mount, Robert E., Paente, Anthony M.. Hady, William F.,

ASME paper, 86-GT-IBI, June 1986. 19 . NASA-TM-83699, A n w of lhe NASA

, 1984. . .

20. NASA-TM-85794, Noise Generated bv a Prowslier in a Huggins, G., and Ellis, D.R., -, Cessna Aircraft Co., Wichita, KS, Cessna-AD-217. (NASA CR- 165564), 1981.

E!&LB& , .

21. NACA TR-640, Propeller Data. L

10