Project Report

Conceptual Design of a Commercial

Tilt-rotor Aircraft

Aerospace Vehicle Design Project

Submitted by

BHOVAD PRIYANKA

YASH GUHA

UNDER THE GUIDANCE OF

Mr. Pankaj Priyadarshi

Dr. Ramanan, R. V.

IN

DEPARTMENT OF AEROSPACE ENGINEERING,

INDIAN INSTITUITE OF SPACE SCIENCE AND TECHNOLOGY,

THIRUVANANTHAPURAM

August - November 2011

Acknowledgment

We wish to express our sincere thanks to Dr. K. Kurien Issac, Head, Aerospace Department,

IIST for providing us an opportunity to work on this project as a part of our Aerospace Vehicle Design

course and giving necessary guidance though out the project work.

We wish to express our deep sense of gratitude to Mr. Pankaj Priyadarshi, Adjunct

Professor, Aerospace Department, IIST and Dr. Ramanan, R. V. Adjunct Professor, Aerospace

Department, IIST, for being our guide and helping us through the obstacles faced by us during the

completion of the project.

Abstract

This design report contains conceptual design of a tilt rotor aircraft designed for chosen requirements,

along with the comparison of various possible configurations. The various flight regimes and power

requirements associated with them is included, and based on that the initial sizing is done. According to

the aerodynamic requirements wing configuration, airfoil, and Turboshaft engines are chosen. Based on

the conceptual design results drag coefficient for cruise is also calculated.

Table of ContentsAcknowledgment ......................................................................................................................................2Abstract......................................................................................................................................................31.Mission Requirements.............................................................................................................................72.Introduction.............................................................................................................................................7

2.1Background of major possible configurations..................................................................................83.Conceptual Design..................................................................................................................................9

3.1Mission Profile.................................................................................................................................93.2Conversion process of a conventional tilt-rotor.............................................................................113.3Initial Weight estimation.................................................................................................................11

4.Aerodynamics ......................................................................................................................................154.1Aerodynamic Requirements ..........................................................................................................154.2Wing Airfoil Selection ...................................................................................................................15

5.Thrust to weight ratio and Wing Loading.............................................................................................176.Tail Sizing.............................................................................................................................................17

6.1Tail configuration...........................................................................................................................176.2 Sizing of H-tail..............................................................................................................................186.3Tail Airfoil selection.......................................................................................................................19

7.Calculation of drag coefficient..............................................................................................................207.1Component Build up method for CD0...........................................................................................20

8.Propulsion ............................................................................................................................................208.1Engine ............................................................................................................................................21

9.Gear box................................................................................................................................................2210.Seating Configuration.........................................................................................................................2311.FAR Regulations Applicable...............................................................................................................2312.Results.................................................................................................................................................2313.References...........................................................................................................................................24Appendix..................................................................................................................................................25

A1 Codes for weight estimation..........................................................................................................25

Index of TablesTable 1: Comparison of various possible configurations...........................................................................7Table 2: Parameters for performing Initial sizing....................................................................................14Table 3: Initial Sizing Results..................................................................................................................14Table 4: Tail Sizing parameters................................................................................................................18Table 5: Results for Tail configuration.....................................................................................................18Table 6: Component build up method for estimation of Drag coefficient...............................................20

List of figuresV-22 Osprey...............................................................................................................................................8Bell/Augusta BA 609.................................................................................................................................8Curtiss-Wright X-19...................................................................................................................................9Mission Profile.........................................................................................................................................10NACA 65415 Cl vs curve.......................................................................................................................15NACA 65415 Drag polar.........................................................................................................................16High lifting Devices.................................................................................................................................17NACA 0009 airfoil...................................................................................................................................19NACA 0009 lift and moment coefficient curve.......................................................................................19NACA 0009 Drag Polar...........................................................................................................................19Rolls Royce AE1107C-Liberty Turboshaft Engine..................................................................................21Prop Rotor gear box.................................................................................................................................22Typical Seating arrangement in a 50 seater aircraft.................................................................................23

1. Mission RequirementsThe mission requirements for the aircraft are following:

1. An aircraft with vertical take-off and landing capabilities.2. It should be able to take off from 300m to 4000m AMSL.3. Payload capacity: 50 passengers and an extra luggage of 1000 Kg. For medical purpose.4. Range : 1000 km5. Cruise mach no. of 0.5

Application : The major consumers of this aircraft would be the pilgrims of North India who want to visit various pilgrim places at higher altitudes in India. The transport service can be provided using this design between New Delhi to various places up to Jammu & Kashmir.

2. IntroductionThe requirements can me met by certain number of possible configurations. A comparison between them is done in next section on the basis of there viability and efficiency. A tilt-rotor aircraft configuration has the potential to combine a vertical take-off and landing capability with efficient, high-speed cruise flight.

Configuration VTOL Speed requirements Pros Cons

Conventional aircraft Not possible Achievable

High cruise speeds, less complexity in structural

design Not capable of VTOL

Helicopter Possible Not achievable Hover for long time, fly in any axis with almost equal

ease Inefficient, noisy and costly

Conventional Tilt Rotor Possible Achievable

Hover for long durations, higher cruise speed than helicopters, more fuel

efficient

Complex control system, heavy tilting mechanisms

leads to high structural weight

Quad Tilt rotor Possible Achievable

Do not require mechanical linkages to vary the rotor

blade pitch resulting in Less complex mechanical system, lesser diameter for each rotor

Reduced efficiency due to backward engines in wake

Tilt wing Possible Achievable Better climb out of hover,

better performance in helicopter mode

loses the lift generated by the wing at lower speed, dynamic stall, heavier

tilting mechanism

Table 1: Comparison of various possible configurations

2.1 Background of major possible configurations

Conventional Tiltrotors

The principle example of a modern tiltrotor is the V-22. The V-22 is a military transport and multi-role tiltrotor first flown in 1989, and has undergone many revisions and improvements over the last two decades. A similar modern design is the Bell/Augusta BA609, a corporate transport tiltrotor. The RFP specifies that the aircraft must be able to land on water; however there exists no such tiltrotor aircraft with that capability to date. Despite this, the V-22 has very similar characteristics to the RFP requirements in payload weight, range and speed. Modern day tiltrotors require a heavily computer augmented control system; this would need to be incorporated in the concept.

Figure 1: V-22 Osprey

Figure 2: Bell/Augusta BA 609

Quad Rotors

A solution to the high thrust requirements of tilt rotor aircraft is to increase the number of engines. Quad rotor aircraft have four engines, two forward and two aft of the center of gravity. This design has the advantage of being relatively stable and forgiving of large fore and aft center of gravity movements. The most notable quad rotor was the Curtiss-Wright X-19 developed in the 1960’s. This experimental aircraft used two turboshaft engines and was manned by a crew of two. The aircraft had a range of 282 nm and top speed of over 390 knots, significantly faster than conventional tiltrotors. The quad rotor does have the disadvantage of smaller, and therefore less efficient, rotors. In addition, the extra engines can increase the weight of a quad rotor design.

Since the quad-rotor technology is still in research and based on various advantages of tilt-rotor over quad tilt-rotor a comparison is made between both probable configurations and tilt-rotor configuration is chosen.

3. Conceptual Design

3.1 Mission Profile1 – 2 is vertical take off

2 – 3 is forward flight under ground effect (about 8-10 m height above ground) to gain sufficient sped before taking lift from wing.

3 – 4 is climb in which conversion from helicopter mode to airplane mode takes place

4 – 5 is steady climb in airplane mode

5 – 6 is cruise for 850 km with 155m/s

7 is loiter for 45 minutes

Under this condition when aircraft decides to land and due to some reason doesn't get permission to land, there is extra fuel for loiter.

Figure 3: Curtiss-Wright X-19

8 – 9 is further descent including deconversion to helicopter mode

9 – 10 is vertical land

Drawing 4: Mission Profile

1

23

4

567

89

1025000ft (M

SL)

Source (MS

L)

1000km

Destina tion(

MS

L)

convers ionD

e-con version

5-8 ft

Mission P

rofile

Loiter

3.2 Conversion process of a conventional tilt-rotorThe process of rotating the nacelles to transition between helicopter and airplane modes is called conversion. This process is simple, straight forward, and easy to accomplish. The amount and rate of nacelle tilt can be completely controlled by the pilot or can be performed automatically by the flight control system. The minimum time to accomplish full conversion from hover to airplane flight mode is 12 seconds. A tilt-rotor can fly at any degree of nacelle tilt.

During vertical take-off, conventional helicopter controls are utilized. As the tilt-rotor gains forward speed to between 40 and 80 knots, the wing begins to produce lift and the ailerons, elevators, and rudders become effective. At this point, rotary wing controls are gradually phased out by the flight control system. At approximately 100 to 120 knots the wing is fully effective and cyclic pitch control of the proprotors is locked out. The conversion from airplane flight to a hover simply reverses the process described above. Since the fuselage and wing are free to remain in a level attitude during the conversion, there is no tendency for the wing to stall as speed decreases. Rotor-borne lift fully compensates for the decrease in wing lift.

3.3 Initial Weight estimationThe total take-off weight is divided into fuel weight, payload weight, crew weight and empty weight or structure weight:

W 0=W crew+W fuel+W payload +W empty

Crew weight: 6∗90+6∗5=570 Kg

Payload weight:50∗90+50∗20+1000 Kg (extra for medical purposes)=6500 Kg

For calculation of fuel weight following power relations are used and the calculations are being made using a MATLAB code given in Appendix A1.

1. Power required in hover

2. Power required in cruise

3. Loiter

4. Landing Hover

5. Total fuel weight

6. Maximum take-off weight

7. Various parameters for performing Initial sizing

LbDcruise 14

FMmr 0.75

LbDm 14

kf 0.3

ηprop 0.85

Rmr 6 m

ρ1 1.225 kg/m3

ρ2 0.549 kg/m3

SFC 7.0442* 10-7 N/W/s

τair 0.97

τmr 0.93

tmr 1.3

t1 5 min

t4 2 min

tloiter 40 min

Vcr 155 m/s

Table 2: Parameters for performing Initial sizing

8. Results

Wfuel1 88.85 KgWfuel2 1285.60 KgWfuel3 478.42 KgWfuel4 30.41 Kg

W0 22135 Kg

Table 3: Initial Sizing Results

4. Aerodynamics

4.1 Aerodynamic Requirements

The 1000 km range requirement and the transition between vertical and horizontal flight modes were the driving aerodynamic considerations for the design. To achieve the range requirement it was imperative that drag be reduced as much as possible and a high lift to drag ratio (L/D) achieved. The higher this ratio, the less fuel the aircraft requires, which in turn eases vertical takeoff and reduces engine weight. To achieve a high L/D, the wing should be small but with a large wingspan. The airfoil should also have a high coefficient of lift (Cl) at zero degrees angle of attack, to eliminate the need for wing incidence. It was considered acceptable for the airfoil to have a zero angle of attack Cl of 0.3 or greater.

The transition between flight modes requires the aircraft to fly slowly without stalling. This heavily influenced the airfoil selection; when choosing airfoils for the main wing it was considered necessary for the airfoil to have a stall angle of attack around 12°. Furthermore, the airfoil chosen for both the inboard and outboard wings must have a high thickness to chord ratio; in tiltrotor applications a thin wing will be very prone to wing flutter. To mitigate flutter, the thickness to chord ratio must be at least 0.15.

4.2 Wing Airfoil Selection The NACA 65415 airfoil was selected for the wing section, its position of minimum pressure at is at 0.5c. The airfoil has a stall angle of attack of 15°. This satisfies the requirement for low speed transition performance. The airfoil’s thickness to chord ratio of 0.15 is sufficient to mitigate wing flutter. In addition, the airfoil possesses a zero degree angle of attack Cl of 0.3, which is high enough to allow cruise at zero degrees angle of attack.

Figure 5: NACA 65415 Cl vs α curve

Figure 6: NACA 65415 Drag polar

5. Thrust to weight ratio and Wing LoadingFor aircraft designed primarily for efficiency during cruise, a better initial estimate of the required T/W can be obtained by “Thrust matching.” This gives us

( TW

)cruise

= 1

( LD

)cruise

(L/D)cruise = 14

This gives (T/W)cruise = 0.0714

Clmax take-off = 2.5;

Vstall=20m/s

ρair =1.225 kg/m3

(W /S )takeoff =12ρair V stall

2 (Clmax)takeoff

(W /S )takeoff =612.5 kg /m2

Our airfoil max lift coeff is 1.5 so we have to use fowler flap to get the required lift at takeoff and landing.

6. Tail Sizing

6.1 Tail configurationThe H-tail configuration is chosen as consistent to historic data and for its advantages in this configuration. The need of H-tail was driven by the need to increase stability in terms of spin resistance. By having two vertical stabilizers, the area of the individual vertical stabilizer is reduced.

In case of a sudden cross wind, only one of the vertical stabilizers will be affected allowing the other to apply a correcting maneuver that will help stop or reduce the spin of the aircraft. Also the vertical stabilizer will be directly in the flow from the twin turboprop engines making the rudder more effective during low airspeed and taxiing. However, the configuration need high strength and that will increase the weight of the tail section which will increase the overall weight of the aircraft.

Figure 7: High lifting Devices

6.2 Sizing of H-tailThe coefficient that is used in scaling of vertical stabilizer is referred to as, CVT. The total area of vertical tail and horizontal tail is found from equations

SVT=CVT

bw Sw

lVT

S HT=CHT

CW S W

lHT

where bW and SW are the span and area of the main wing, respectively, and lVT is the distance between the quarter-chord locations of the mean aerodynamic chords of the main wing and vertical stabilizer.

The values for sizing are

CVT (For twin turboprop) 0.08bW 16.5 meterSW 36.3 meter2

lVT (Assumed to be 50% of the fuselage length) 12.84 meterCHT 0.9CW 2.2LHT 12.84 meter

Table 4: Tail Sizing parameters

This gives a vertical stabilizer area of 3.732 m2.

And for horizontal stabilizer it is 5.592 m2.

The various parameters chosen for horizontal and vertical tail are shown in table below.

Horizontal Tail Vertical TailSpan 4 m 1.903

Aspect ratio 2.858 1.940Area 5.598 3.732

Average chord 1.4 0.9807Root chord 1.4 1.4011Tip chord 1.4 0.5604

Table 5: Results for Tail configuration

6.3 Tail Airfoil selection

The airfoil chosen for tail configuration if NACA 0009. Since, it has to be symmetrical and has a tc

less than the wing. The characteristic of tail airfoil should be lower CD0 and higher dC L

d α.

Figure 8: NACA 0009 airfoil

Figure 9: NACA 0009 lift and moment coefficient curve

Figure 10: NACA 0009 Drag Polar

7. Calculation of drag coefficient

7.1 Component Build up method for CD0

The

C D0=ΣC FC FFC QC SWETC

Sref+C Dmisc+C DLP

where, CDmisc is Miscellaneous drags for special features of an aircraft such as flaps, unretracted landing gear, an up-swept aft fuselage, and base area.

CDLP accounts for leakages and protuberances.

Cf for laminar case is given by

C f =1.328/√ Re

Structure Characteristic length (m)

FFC SWET QC Cf cruise (10-3)

Wing 2.2(1.06+100∗(

tc)

4

)(1.34∗M 0.18)64 1 2.8

Horizontal Tail

1.4(1+0.06

( tc)

( xc )

+100( tc)

4

)∗(1.34M0.18)(cos Λm)0.2811.196 1.08 3.061

Vertical Tail 0.9807(1+0.06

( tc)

( xc)+100( t

c)

4

)∗(1.34M0.18)(cos Λm )0.28

7.464 1.08 3.245

Fuselage 25.76 1+60f 3+

f400

220 1 4

Nacelle 1.98 1+0.35f

3.52 1.3 2.9

Table 6: Component build up method for estimation of Drag coefficient

This gives us a CD0 value of 0.0213.

8. Propulsion A tiltrotor aircraft requires a unique propulsion system that must adequately perform in both Vertical Takeoff and Landing (VTOL) and conventional airplane mode. The RFP mission requirements and the tiltrotor design limitations cause many limiting factors in a propulsion system. The propulsion system selected must be able to provide adequate thrust in VTOL operations with a safe and acceptable vertical climb rate. The propulsion system must also be able to meet the RFP requirements and be able to obtain a 558 km/hr cruise speed at cruise altitude (25,000 ft).

8.1 Engine The engine selection for a tiltrotor is limited. Historically, turboprop and turboshaft engines have been used in tiltrotor aircraft. For this case, an existing engine is desired to allow for ease of production and reduction in cost. A turboshaft engine was chosen over a turboprop because of the turboshaft’s ability to provide larger amounts of shaft horsepower. The selection of a turboshaft engine is consistent with current production tiltrotors such as the V-22 and the Bell/Augusta BA609.

As a result of these selection criteria, the Rolls Royce AE 1107C Liberty engine was chosen for this design. The AE 1107C is specifically designed to power tiltrotor aircraft.

Its specifications are given below.

Rolls Royce AE1107C-Liberty Free Shaft Gas Turbine Engine

• Type: Dual spool, free shaft turbine· Inlet: Axial flow • Compressor: 14 stage axial, with variable geometry stators • Burner: Annular through flow combustor, effusion cooled • Turbine: Dual spool, two stage axial gas producer turbine, dual stage free power turbine • Exhaust: Axial diffuser • Power Rating: 6,150 shaft horsepower at 15,000 rpm • Rated Torque Output at Full Power: 2,153 lb/ft @ 15,000 rpm • Weight: 971 lbs. • Power/weight: 6.3:1 hp/lb • Pressure Ratio: 16.7:1 • Specific Fuel Consumption: 0.42 lb/shp/hr

Figure 11: Rolls Royce AE1107C-Liberty Turboshaft Engine

9. Gear boxA typical drive system for tilt rotor is shown in fig below. The system was designed by NASA as a part of their design for Large Civil Tilt-rotor (LCTR) project.

This system is just a representative of how the drive system in a typical tilt-rotor works and there are no original design parameters associated with the present design in this block diagram.

Figure 12: Prop Rotor gear box

10. Seating ConfigurationA typical seating arrangement viable for 50 passengers.

Seat pitch=95 cm

Seat width= 50 cm

Aisle height=200 cm

Aisle width =56 cm

11. FAR Regulations ApplicableFAR Part 9: Airworthiness standards - Transport category rotorcraft.

Rotorcraft with MTOW greater than 20000 lb and 10 or more passengerseats must be certified as Category A rotorcraft.

FAR Part 25: Airworthiness standards - Transport category airplanes

Propeller-driven airplanes with a MTOW greater than 19,000 pounds(8,618 kg) and 19 or more passenger seats must be certified under thisstandard.

12. ResultsFrom the conceptual design of the aircraft we have found out the gross take off weight of the aircraft as

22135 Kg. The aerodynamic parameters are being chosen to facilitate better performance of the aircraft. With the chosen airfoil for wing and tail the calculated CD0 value comes out to be 0.0213.

Drawing 13: Typical Seating arrangement in a 50 seater aircraft

13. References• Daniel P. Raymer, 'Aircraft Design: A conceptual Approach'

• J. Seddon, 'Basics Helicopter Aerodynamics', BSP Professional Books.

• Antonio Filippone, 'Flight Performance of Fixed and Rotary Wing Aircraft', Elsevier Books.• Engineering Design Handbook, Helicopter Engineering. Part One, Preliminary Design, Army

Materiel Command, Alexandria, Virginia distributed by NTIS, U.S. Department of Commerce• Thomas C. Corke, 'Design of Aircraft'• Summary of the Large Civil Tiltrotor (LCTR2) Engine Gearbox Study, NASA/TM—2010-

216908• “Performance and trade studies of a Mono tilt-rotor Design”, Robin Preator, J.Gordon Leishman

at University of Maryland.• http://en.wikipedia.org/wiki/Bell_Boeing_V-22_Osprey • www.rolls-royce.com • Airfoil investigation database, www.worldofkrauss.com/foils

Appendix

A1 Codes for weight estimation% wt estimation rho=1.225; % densityV=50:180; % airspeedfair= 200; % total equivalent flat plate areafhel=200; % wing areaAR=7.5; % aspect ratioe=1.78*(1-0.045*AR^0.68)-0.64; % oswald efficiency factorCL=0.3; % coeff of liftVcr=650*1000/3600; %cruise speedWto=22135; % take off wtRmr=6; % rotor radiusmu=V/Rmr/10000*60; % advance ratioomega=2*pi*10000/60; % rotor omegak=1; % rotor induced power coeffCd0= 0.0017; %profile drag coeffsigma= 0.8*3/pi/Rmr; % rotor solidityAmr=2*pi*Rmr^2; %main rotor area

%wing arearho1=1.225;Clmax=2.5;Vst=20;rho2=0.549;Clcr=0.3;Swst=Wto*2/rho1/Vst^2/Clmax; Swcr=Wto*2/rho2/Vcr^2/Clcr;bw=sqrt(Swst*AR); %wing span

%initial sizing Wpl= 50*90+(50*20)+ 1000 ;Wcrew=6*90+(6*5);SFC=190*9.81*10^-3/(0.735*10^3)/3600;%gm/hp/hr to N/W/snprop=0.85;tauair=0.97;LbDm=14;DL=100; % disc loading Rmr1=sqrt(Wto/DL/2/pi); % main rotor radiusW0=2.2135e+004;

%first hover in heli modetmr=1.3; % main rotor thrust recovery factortaumr=0.93;Hrho1=0;FMmr=0.75;Cpow=1.1;t1=5*60;P1=(W0*tmr)^1.5/(sqrt(pi/2*rho1)*FMmr*taumr*2*Rmr); %power required to hoverkalt1=(1-2.23*10^-4*Hrho1)^-1;Peng=P1*kalt1/2;

Pnom=Peng*Cpow;Wfuel1=P1*SFC*t1

%cruiseW2a=W0-Wfuel1;W2=0.985*W2a; % climb considerationLbDcr=14;Lcr=1000*10^3;W3=W2*(exp(-Lcr*SFC/LbDcr/nprop/tauair));Wfuel2=W2-W3

% loitert3=40*60;W4=W3*exp(-t3*SFC*Vcr*0.8/nprop/0.866/LbDm);Wfuel3=W3-W4

% landing hover P4=(W4*tmr)^1.5/(sqrt(pi/2*rho1)*FMmr*taumr*2*Rmr);Pav3=Pnom*2/kalt1;t4=2*60;Wfuel4=P4*SFC*t4

% total fuel wtKf=0.3;Wfuel=(1+Kf)*(Wfuel1 +Wfuel2+Wfuel3+Wfuel4)Wempty=0.57*W0;

Wtakeoff=Wpl+Wcrew+Wfuel+Wempty % final MTOWWfinal=Wtakeoff-Wfuel;WF=Wfuel/WtakeoffWf=Wfinal/Wtakeoff