Final Report of Conceptual Design and Development of Multi-Role Fighter Aircraft (PAF-X) Muzammil Ashraf, Waleed Bin Arshad, Ahmed Ali, Muhammad Asad, Muhammad Safdar, M.Atif Raheem NUST, College of Aeronautical Engineering, Risalpur, Pakistan, 20480 The aircraft design process is the engineering design process by which aircraft are designed. These depend on many factors such as customer and manufacturer demand, safety protocols, physical and economic constraints etc. Design of an aircraft is a complex process. It is accomplished in many stages. The design process starts with the aircraft's intended purpose. A military aircraft designed primarily for air-to-air combat against other aircraft. In this report we have discussed in detail various design methods that we used to design a feasible jet fighter for Pakistan Air Force. It starts with requirements analysis proceeded by KJ, Kano, functional Analysis and quality function deployment. Finally we start our sizing and configurations based on our initial guess of weight. Towards the end of process we have applied the morphological matrix method in our design and created feasible concepts out of it and explained them in their first stages. CAD models of all concepts were made. All the concepts were tested for aerodynamics and stability analysis. Then we applied some innovative ideas from TRIZ parameters .Finally we selected our final concept based on wealth of knowledge gained by all the methods. At the end our PAF-X is compared with baseline aircraft i.e. F-22 Raptor. Keywords: Design and Development,KJ Analysis, Kano Model, House Of Quality, Initial Sizing, Constraint Analysis, Configuration Layout Morphological Charts, Morphological matrix, TRIZ method Nomenclature DAs RFP Wo GAs HOQ MDO MM STOL KJ QFD FFBD TO = = = = = = = = = = = = Design attributes Request for Proposal Empty Weight General Attributes House of quality Multi-disciplinary optimization Morphological Matrix Short Take-off and Landing Jiro Kawakita Quality Function Deployment Functional Flow Block Diagram Take Off I. Introduction The conceptual design is a very important stage in the process of aircraft development. The work in this stage has a decisive impact on the aircraft’s final capability.it is an iterative process which goes on changing during the process and finally we come up with some innovative idea. For conceptual design and development of a fighter jet we were given a request for proposal from our instructor. We were a group of six members. The process started in Oct 2015 when our instructor held our hand to teach us some interesting and important methods for designing an aircraft. First of all we inspected the RFP and listed down the requirements from it .we made some changes or the better design. Then all the requirements were divided into groups .we identified must have’s, linear satisfier and added some wow factors. HOQ was tabulated and we got the important engineering attributes for our design process .all the requirements were converted into functions through FFBD.We followed the rhymer’s approach for our sizing and configuration layout. Dessault CATIA 5 was used to make CAD models for our aircraft. MM was used to generate feasible designs and TRIZ principals were applied to bring innovations in our designs. Finally a design was selected on the basis of aerodynamic and stability analysis using XFLR, XFOIL. At the end our final conceptual design in proposed for further validation of tests and changes. 1 College of Aeronautical Engineering,NUST,Risalpur,PK II. Request For Proposal 2 College of Aeronautical Engineering,NUST,Risalpur,PK II. Requirements Analysis We have read the RFP given above and taken out following customer’s requirements. Some changes are made to accommodate the requirement of new design as it was the same RFP last semester. Type A/C : Multi-Role fighter Range : 1200 nm Payload : 4250 kg STO : 762.5 m SLand : 715 m Combat Alt : 30000 ft. Cruise Altitude : 20000 ft. Cruise Mac : 1.2 3 College of Aeronautical Engineering,NUST,Risalpur,PK Cruise Range Loiter Ceiling Max. Mach Load Factor : : : : : 150 nm 20 minutes 35000ft 2 9g’s (combat) III. KJ Analysis On the basis of our extraction of requirements from given RFP we carried out the process of KJ analysis. During the KJ analysis, the whole group sat down together, did some brain storming and wrote down the requirements on small chits and combined the chits with similar parameters together and made different groups. Later each group was given a particular name. The results of our KJ analysis are as below: Operations Design mission Specs Top speed Capable of sustaining 7gs Mach) max of 2 Ferry range of 1200 nm Landing and takeoff distance 715 m. Night ops Highly maneuverable Climb to 20000 ft. Cruise of 150 nm Cruise at Mach 1.2 for 20 minutes Constraints Service ceiling up to 35000 ft. Structural load factor of 9gs Durability Service life of at least 3000 hours Robust Survivability No fail system Fail safe system Recovery of components and crew Damage resistant User Friendly Easy flight controls Easy to assemble IV. KANO Model In the KANO model, the requirements from the KJ analysis are divided into three categories and are rated. These three categories are: • Must have’s • Linear satisfiers • Wow factors 4 College of Aeronautical Engineering,NUST,Risalpur,PK We have given ratings as per RFP focus. Maximum importance is given as number 9 and minimum as number 1. Our KANO model along with their customer ratings is as below MUST HAVE’S LINEAR SATISFIERS Easy to assemble Service ceiling less than 35000 ft. All weathers Capable of 7 g’s Cruise at 20000 ft. Endurance of 20 minutes at MACH 1.2 9 Service life of 3000 hours 8 Highly maneuverable 9 Cruise speed of 150nm 8 Ejection 9 Structural Load factor of 9 g’s 8 No fail system 8 Fail safe systems 7 Night ops 7 Maximum MACH of 2 8 Landing and takeoff distance 6000 ft. 7 Range of 2200km 8 8 8 9 7 8 WOW FACTORS Easy flight controls Robust Stealth V. Functional Analysis This design method focuses on the intended functions that our product is going to perform. It starts from end result desired by the customer. Then it’s broken down to different sub functions. Basic functions: function that is the reason of design. Secondary function: methods that are supporting the basic function In our functional analysis we have divided our basic function of multi-role fighter into 6 sub-functions that are further divided into sub levels of 2 and 3. These are given as below in functional flow block diagram (FFBD) 5 College of Aeronautical Engineering,NUST,Risalpur,PK 9 8 7 6 College of Aeronautical Engineering,NUST,Risalpur,PK VI. Quality Function Deployment This ranking is determined by communicating with the customer And ranked accordingly. There are many segments in the house of quality relating various parameters with each other depending upon their relationship. Customer Requirements or WHATS Customer Requirements or WHATS represent a ranking of how important each requirement is to the customer. This ranking is determined by communicating with the customer. Engineering Attributes or HOWS The engineering attributes (called HOWS) are selected based on the customer requirements. The EAs are the measurable set of parameters that help meet the customer requirements satisfactorily. Customer Requirements and Engineering Attributes Relationship With these WHATS and HOWS occupying the left and top floor of the house, first floor on the right side is used to see their relationship to ascertain the design drivers. To make this analysis meaningful, a numbered scale of 0 1 3 and 9 is used. Correlation Matrix The roof of the HOQ shows correlation of the Engineering Attributes (EAs) with each other. The aim of establishing such a correlation is to identify all the conflicts and agreements among various EAs that would become part of the product. Competition Benchmarking The competition benchmarking is carried out to ascertain how new product competes against the existing ones in meeting each of the customer requirements Relative and Absolute Importance Finally, the bottom section of the QFD chart is used to enter a computed total of all the scores given to EAs based on how well they would fulfill the customer requirements. This is ascertained on the absolute as well as the relative scale 7 College of Aeronautical Engineering,NUST,Risalpur,PK 8 College of Aeronautical Engineering,NUST,Risalpur,PK From the HOQ deployment we have come to know that following are our main design drivers 9 College of Aeronautical Engineering,NUST,Risalpur,PK 1 .Yield Strenght 2 . L/D) max 3 . W/S 4 . CL) max 5 . Thrust VII. Baseline Aircraft F-22 raptor was taken as baseline aircraft Specifications of F-22 are as Length – 62 ft/18.90 m Wingspan – 44.5 ft / 13.56m Wing area – 840sq ft / 78.04 sq m Empty weight – 43,340 lb / 19700kg Speed – Mach 2 Range – 1600nm Engine thrust – 35000 lb / 15876 kg Airfoil – NACA 64A205 Wing loading – 77.2lb/ft2 Thrust/Weight – 1.08 Service ceiling – 65000ft / 20000 m VIII. Mission Profile The mission profile is divided into nine different segments. The Mach number, altitude and the speed of the aircraft at different segments during the whole mission was given in the RFP. The mission profile as per our RFP is as follows: 10 College of Aeronautical Engineering,NUST,Risalpur,PK 0 – 1 : Warm up and takeoff. 1 – 2 : Climb/Accelerate 2 – 3 : Cruise M 1.2 @ 20,000 ft, 150 nm 3 – 4 : Descent 4 – 5 : Combat/Weapon drop @ 15000 ft, 7g. 5 – 6 : Climb/Accelerate 6– 7 : Cruise M 1.2 @ 20,000 ft, 150 nm 7 – 8 : Descent 8 – 9 : Loiter 20 min 9 – 10 : Landing and Taxi Back IX. Initial Weight Sizing Initial takeoff gross weight is calculated from a conceptual sketch using a simplified sizing method. This method works by taking data from historical trend and can be used for all types of mission profiles. Gross Weight Estimation 11 College of Aeronautical Engineering,NUST,Risalpur,PK      Calculation of the design gross weight is the first step of the takeoff weight build up. It is the total weight of the aircraft for the designed mission. This weight should not be confused with the maximum takeoff weight. The design takeoff weight is broken down into, Payload weight (Wpayload) Crew weight (Wcrew) Fuel weight (Wfuel) Empty weight (Wempty)  Wo = Wpayload + Wcrew + Wfuel + Wempty Now the above equation can be written as,   Wcrew  Wpayload Wf We 1  Wo Wo Wo= Crew Weight Estimation Crew weight is estimated following the historical trends which suggest that for a specially suited pilot the approximate weight is 75 kg. We are given with the specifications of 2 pilots so:        (a) Wcrew = 150 kg Payload Weight The payload weight is given as a design requirement i.e. 5000 kg. So,  Wpayload = 5000 kg Empty Weight Estimation Empty weight fraction is estimated statistically from historical trends. For a jet fighter,  Where, We/Wo = A (Woc) (Kvs)  (b) A = 2.11=2.34 We will use A=2.11 because we are using metric units   C = -0.13 Kvs = 1 Kvs is the variable sweep constant which is 1.04 if the aircraft has variable sweep and 1.00 if the aircraft has fixed sweep. Our aircraft is a fixed geometry (fixed sweep) so this factor is 1.00.  We/Wo= 2.11 (Wo)-0.13 12 College of Aeronautical Engineering,NUST,Risalpur,PK  Fuel Fraction Estimation The fuel required / consumed by the aircraft depends upon many factors, some of which are:    Drag of the aircraft Thrust specific fuel consumption Size / total weight  So for estimating fuel fraction we must find different weight fractions of the segments of mission profile. For different segments we use the weight fractions as:  Weight fraction = Wi / Wi-1 o Weight Fractions of Different Segments  Segment (0 - 1): Warm-Up, Taxi and Take-Off At this stage the weight fraction for warm-up, taxi and take-off can be taken from historical trends.  W1/W0=0.97 Segment (1 - 2): Climb Again the weight fraction for climb / accelerate can be taken from historical trends. W2/W1=0.985 Segment (2 - 3): Cruise Range formula is used for the calculation of weight fraction of cruise segment. It is givens as R = Range of Cruise= 277800 m (150 nm) C = TSFC = 0.0002222/sec V = Velocity during cruise = 379.8 m/sec H = Altitude = 6080 m (20000ft) From historical estimation Swet/Sref = 4.25 The aspect ratio (AR) will be estimated to calculate the value of L/Dmax will be found by calculating Wetted Aspect Ratio, � = � / = 2.67 Wetted Aspect Ratio= � / / = 0.628 (L/D) max = 11 For cruise, L/Dcruise = 0.866 * L/Dmax L/D = 9.526 Now putting all the values in equation for weight fraction of cruise, we get, 13 College of Aeronautical Engineering,NUST,Risalpur,PK W3/W2 = 0.9834 Segment (3-4): Descent W4/W3 = 1 Segment (4-5): Combat W5/W4 = 1 Segment (6-5): Climb W6/W5 = .985  Segment (6-7): Cruise W7/W6 = 0.9834 Segment(7-8):Decent  W8/W7=1 Segment(8-9): Loiter Endurance formula is used for estimation weight fraction of loiter i.e. Where, C = SFC = 0.000194238/sec E = Endurance = 1200 sec L/Dloiter = L/Dmax (Loiter) = 11 By putting the values in equation for weight fraction for loiter we get, W9/W8=.9794 Segment(9-10):Landing W10/W9=.995 The Final Process of Estimating Wo 14 College of Aeronautical Engineering,NUST,Risalpur,PK Now we finally find the fuel fraction by using the following method: Wx Wo = W1 Wo W2 W3 W4 W5 W6 W7 W1 W 2 W 3 W 4 W 5 W 6 After putting values in above equations, we get W11/ Wo = 0.886 Wf / Wo = 0.121 Also we know, We / Wo = 2.11 (Wo)-0.13 Wo = 5150 / (1 - 0.284 - 2.11WO -.13) The final take off gross weight was found to be, Wo = 15740 kg Empty weight of aircraft, Wempty= 9455 kg Weight of fuel Wfuel= 1884kg X.  Constraint Analysis Thrust to Weight T/W directly affects the performance of the aircraft. An aircraft with higher T/W will, accelerate and climb more quickly, reach a higher max speed and sustain higher turn rates. On the other hand it consumes much more fuel, thus increasing the aircraft’s gross weight. T/W is not constant. It varies with altitude, velocity and weight.  Wing Loading Wing loading is defined as the weight of the aircraft divided by the reference area of the wing. The term W/S refers to the takeoff condition. This affects the stall speed, climb rate, take off, landing distances, and turn performance. W/S has a large effect on the takeoff gross weight. The smaller values of wing loading results in a bigger wing, means larger drag and weight. Using the design method used by Brandts for constraint analysis given in the notes provided by the instructor we did the constraint analysis for a jet fighter aircraft.For a jet fighter we have following constraints: i. ii. iii. iv. Subsonic combat turn Supersonic combat turn Take off Landing 15 College of Aeronautical Engineering,NUST,Risalpur,PK From the data taken from these constraints, we plot a graph to get the design regime for our jet fighter which is shown below T/W= 0.94 Wo/S= 640.5 kg/m2 XI. XII. XIII. Refined Sizing Final Refined Weight The refined weight obtained after iterating the rubber sizing equation came out as follows. The weight increased slightly because of addition of combat segment which was neglected during rough weight analysis. Final Refined Weight Fractions 16 College of Aeronautical Engineering,NUST,Risalpur,PK W10/W0 0.897 Rf 1.060 Fuel Weight Fraction Wf/W0 0.1090 Empty Weight Fraction We/W0 Total Weight Fraction Reserved Fuel 0.540684 Gross takeoff weight 17580kg Fuel weight 4670kg. Empty weight 9505.5 kg. Table 1: Comparison of initial sizing and refined sizing weights PARAMETERS INITIAL WEIGHT SIZING REFINED WEIGHT SIZING Gross takeoff weight 15740kg 17580kg Fuel weight 1884kg 4670kg. Empty weight 9455 kg 9505.5kg. The refined weight has decreased from the initial weight of 17868.3 kg. The reasons for this increase are: I. The weight fractions for different mission segments were mostly calculated on the basis of historical data which gives very crude results. II. The influence of actual aerodynamic parameters such as the T/W ratio and wing loading on mission segments was not taken into account in initial sizing. III. Some mission fragments like Combat and weapon drop were not catered for in the initial weight sizing. XIV. Geometry Sizing Once the takeoff gross weight has been estimated, the fuselage, wing, and tails can be sized. In this chapter the geometry sizing for optimized aircraft is shown by making use of tail volume coefficient. Fuselage 17 College of Aeronautical Engineering,NUST,Risalpur,PK Fuselage sizing has been done using Table 6.3 of textbook for a jet fighter. The fuselage length will be calculated as L=aWoc Fuselage data calculation Length = aWoc a 0.389 Wo 17580.5 kg c 0.39 Length of Fuselage 17.6 m Fineness Ratio 12.5 from Raymer(Supersonic) Fineness ratio = Length/Diameter 12=17.6/diameter Max. Diameter 1.4 m Wing Wing is the main lifting body of the aircraft, so due respect is required for the selection of wing geometry. It depends a lot on design specifications and the mission requirements Aspect Ratio (AR) For Jet Aircraft Equivalent Aspect Ratio (AR) = � a=4.110 c=-0.622 Mmax= 2 So, the Aspect Ratio is 2.87 Wing Sweep LE Sweep = 90 – sin-1(1 / Mmax) = 90 – sin-1(1 / 2) = 60o But from the historical trends and the sweep angle of F-22 aircraft we selected, LE sweep = 320 Taper Ratio λ = 0.3(F-22 Data) Twist Wing twist = 0o Incident Angle Incident angle is Zero degrees for military aircrafts. Dihedral It is the angle of the wing with respect to the horizontal when seen from front. When wing is given dihedral angle then it helps the aircraft in rolling performance. But it is a historical trend that all aircrafts which are designed to have transport, mission or all those, which require excessive stability they have dihedral angle. Similarly all those aircraft’s, which require more maneuverability than stability they have, negative dihedral, which is actually anhedral angle. Using table 4.2 we have Wing Dihedral = -0.5o 18 College of Aeronautical Engineering,NUST,Risalpur,PK Wing Vertical Location A mid-wing configuration has been selected for our aircraft. This arrangement provides superior aerobatic maneuverability. A high-wing arrangement cannot be used as it restricts the pilot’s visibility to the rear while the lowwing arrangement would not provide ground clearance to carry armaments. Wing Tips A more effective and important effect is the influence of the tip shape upon the lateral spacing of the tip vortices. This is largely determined by the ease with which the high pressure air on bottom of the wing can “escape” around the top of the wing. There are different types of the wing tips used including; rounded, sharp, cut off, Hoerner, drooped, upswept, cut of forward sweep, endplate and winglet. The Wing Tip selected is Cut-off, Forward Swept .It has been selected because the tip is cut off at an angle equal to the supersonic Mach-cone angle and wing tip contribute a little bit increase in lift. And also the tip reduces the torsional loads applied to the wing. The actual wing size can now be determined as the takeoff weight divided by the takeoff wing loading. The parameters of the wing are given as, S=Wo/(W/S) Wo 17580 kg W/S 640.5 kg/m2 S 27.5 m2 ASPECT RATIO = b2/s 2.87 b = (A.R*s)^0.5 = 8.9 m λ 0.3 Croot 4.76 m Ctip 1.43 m Tail Volume Coefficient The value of tail volume coefficient is obtained from table 6.4 CHT = 0.4 CVT = 0.07 Area of the tail is related by following relation � SVT= � ST= Vertical Tail Geometry � Lvertical tail 7.91 m Cvt 0.07 (V Tail) SVT = 2.15 m2 ARVT = bVT2 /Svt 1.1 bvt 1.53 m Croot (Vertical Tail) = 2.15 m 19 College of Aeronautical Engineering,NUST,Risalpur,PK Ctip (Vertical Tail) = 0.65 m Horizontal Tail Geometry The calculations of horizontal tail geometry are as follows, CHT 0.4 LHT 7.91 m Shorizontal tail 4.70 m ARH.T = bH.T2 /SH.T 3.0 bHT 4.05 m Taper Ratio 0.3 Croot (Horizontal Tail) 1.78 m Ctip (Horizontal Tail) 0.53 m XV. Crew Station and Payload Crew Station The cockpit is designed to provide maximum comfort and vision to the pilot. Provision of easy access to all vital controls, systems and accessories without causing the pilot’s attention to be diverted is also very important. A single piece bubble type canopy is used to give easy maintenance. The canopy is hinged to open sideward. For pilot safety a united Technologies Corporation ACES II ejection seat is installed. It can eject at minimum height of 140 ft. Seat Back Angle The cockpit layout uses a 30 degree seat back angle. This can obstruct the pilot’s view but retains the advantage of pilot bearing high g’s during operation. Over Nose Vision Over nose vision is important for safety especially during landing. Following historical trends the over nose for fighter aircraft is 11-15 degrees. Vision Angle Fighters should have completely unobstructed vision above and all the way to the tail of the aircraft. Any canopy structure should be no more than 2 in. wide to avoid blocking vision. Bubble canopy is used which provides a 360 degree view. XVI. Weapon Carriage There are four types of weapon carriage which are given below. • External • Semi-submerged • Internal • Conformal Five hard points have been provided four under the wing and one under the fuselage. The 5000kg payload allows the aircraft to carry a wide variety of modern weapons for air superiority. 20 College of Aeronautical Engineering,NUST,Risalpur,PK XVII. Propulsion and Fuel System Integration Selection of the Engine The designed aircraft is basically a modification of F-22 since it was taken as the reference aircraft so one engine F-136 AFTERBURNING was used in the F-22 so the thrust requirement of our design is comparable to it. The engine selected for the design aircraft is “a low bypass turbo fan engine‟. The reason for selecting turbofan engine is that it has higher thrust than the turbojet engine due to the bypass air; it has comparatively low SFC values. It is efficient at both subsonic and supersonic speed up to M # 2. XVIII. Engine Sizing The sizing of the engines is of the two types: Fixed engine sizing Rubber engine sizing Thrust Required CALCULATION The max thrust required is calculated as: Thrust T/W * Wo T 16525.5 Therefore the required engine thrust at takeoff comes out to be 16525.5 N. Scale Factor The scale factor of an engine is the ratio of the required thrust and the actual thrust of the nominal engine. Scale Factor =1.05 Because we used fixed engine sizing and used an off shelf engine. XVII. Inlet Geometry Turbofan engines are incapable of efficient operation unless the air entering is slowed to a speed of Mach 0.4 - 0.5, thus efficient slowing down of the flow with the minimum losses is the key goal of inlet design. There common types of inlet used on the aircraft are: i. NACA Flush Inlet ii. Pitot or Normal Shock Inlet iii. Conical or Spike Inlet iv. 2-D Ramp Inlet The last two types are most suitable for supersonic aircraft. A new type of supersonic inlet has been under study and is very promising in terms of the engine performance for a large number of Mach no. range up to Mach 2.25. This Divert less Supersonic Inlet has been chosen. XIX. Inlet Geometry Size Calculations Throat Area The throat area of spike inlet should be 70-80% of engine front face area A/A*)THROAT 1.03823 A/A*)ENGINE 1.59014 A)THROAT/ A)ENGINE 0.652917 m2 Diffuser Length Length Diameter 1.90 1.82 m m 21 College of Aeronautical Engineering,NUST,Risalpur,PK Inlet Location Two side have been used for the design aircraft. Like nose inlet, it offers clear airflow as it is free of wing and fuselage distortions .Unlike nose inlet, it has a shorter length and thus reduced weight. It is good at high angles of attack as fuselage fore-body helps to turn the flow into it. Nose landing gear problems encountered in single chin inlet can be minimized as the side would be located at the side of nose landing gear. Capture Area Calculation To determine capture area mass flow is multiplied by value selected from following figure. Capture area /mass flow = 0.0057 m2/kg/s from fig 10.16 (Raymer) Mass flow = 26*(engine front face diameter) 2 Mass flow = 38.56305024 kg/s Capture Area = 0.219809386 m2 Boundary Layer Diverters Five kinds of diverters are used which have been named below  Step Diverter  Boundary layer by-pass duct  Boundary layer suction  Channel type boundary layer diverter   Diverter-less Supersonic inlet Nozzle Integration Keeping in view the above requirement, Converging Iris Nozzle has been selected. It performs the same function as a variable convergent nozzle but has an edge over the variable convergent nozzle in terms of drag reduction, as it does not introduce a base area when it is in closed position. XX. Fuel System Fuel Type Using table 10.5 the design fuel is JP-5 which has a largest Mil-spec density of 6.8 lb. /gal and thus occupies the least space. Fuel Tanks Bladder Tanks are used for design aircraft. They are made up of rubber and can occupy different shapes. They are self-sealing and if a bullet passes through a self-healing tank, the rubber will fill in the hole preventing a large fuel loss and fire hazard. XXI. Landing Gear and Subsystems Landing Gear Attachment From considerations of surrounding structure, the nose and main assembly are located such that the landing and ground loads can be transmitted most effectively, while at the same time still comply with the stability and controllability considerations. Landing Gear Arrangements The selected landing gear arrangement is “TRICYCLE GEAR”. 22 College of Aeronautical Engineering,NUST,Risalpur,PK Tyre Sizing The tires are sized to carry the weight of the aircraft. Typically main tires carry about 90% of the total weight of the aircraft, rest is carried by nose wheel. To estimate the size of our fighter we use following statistical tire sizing table from Raymer using table 11.1. Specifications of selected of main landing gear tire: Speed 230 mph Max Load 17580 kg Max Dia .74 m Rolling R .32 m Wheel Dia .40 m No Of Piles 18 Gear Retraction In high speed modern aircraft especially jet fighters retractable gears are used. Non-retractable gears are impractical at high speeds as they offer a large amount of drag. Fuselage Retraction System is selected for the design aircraft. XXII. Aerodynamics Aircraft aerodynamics is the study of the behavior of airflow passing over the aircraft and the forces generated due to it. In other words all aerodynamics lift and drag forces result from the combination of shear force and pressure forces. The drag on wing includes forces variously called airfoil profile drag, skin friction drag, separation drag, parasite drag, camber drag, drag due to lift, wave drag, interference drag and so forth. However, we will restrict our (a) Lift 1) CL vs. Mach No. For subsonic flight, πA CL = +√4+ ta 2 Λ axt AR 2 β 2 { + } β2 η2 xp r F As this is a fighter aircraft having a supersonic cruise CL is calculated for subsonic, transonic and supersonic regimes. So, after putting these values for subsonic we get a graph for CL vs. M # for subsonic lift curve slope. Similarly by using supersonic charts we calculated CL with best approximation available and the transonic regime CL vs. M # slope was derived by interpolating the values as suggested in textbook. 23 College of Aeronautical Engineering,NUST,Risalpur,PK (b) Drag 2) Parasite Drag Drag forces, which are not strongly related, to lift are termed usually as parasite drag. The skin friction drag of a flat plate of the same wetted area as the aircraft can be determined for the various Reynolds number and skin roughness. 24 College of Aeronautical Engineering,NUST,Risalpur,PK 1) Scrubbing Drag Scrubbing drag is the increase in skin friction drag due to prop wash or jet exhaust impinging upon the aircraft skin. Prop wash or jet exhaust increases the effective velocity and also assures turbulent flow over the aircraft. Both of these increases drag. 2) Form Drag Viscous separation drag is termed as form drag. This depends upon the location of the flow separation point on the body. This flow separation is due to the viscous effect. The form drag is less if the flow separates far away on the body than due to separation at a shorter distance. The location of separation depends on the curvature of the body and also on the energy in the flow. A turbulent flow has more energy than a laminar flow. So turbulent flow will have less form drag. 3) Profile Drag The subsonic drag of a streamlined non lifting body depends only upon the skin friction and viscous separation drag. This subsonic drag is called the profile drag. Sometimes profile drag is referenced to the maximum cross sectional area. 4) Interference Drag This is the drag due to the various components of the aircraft and their interference. These components affect the airflow over the body and increase drag. For example the fuselage increases the drag of wing and encourages flow separation at the wing root. 5) Wave Drag This is the drag produced from formation of shocks at supersonic and high subsonic speeds. At high subsonic speeds the shocks form first on the upper surface of the wings because the air is accelerated as it passes over the wing. 6) Induced Drag The drag that is directly related to lift is induced drag or sometimes termed as “drag due to lift”. The airflow circulation over the 3D wing of the aircraft causes induced drag. This circulation produces vortices at the rear edge of the wing. This vortex is produced by extracting energy from the airflow. This produces the drag force and is directly proportional to the square of lift 7) Equivalent Skin Friction Drag: 25 College of Aeronautical Engineering,NUST,Risalpur,PK An estimate of parasite drag for subsonic cruise of aircraft can be obtained from equivalent skin friction drag. As CD0 is a function of Re no so with an increase in Mach no, Re no increases and hence CD0 decreases. But in case of altitude, CD0 increases with increasing altitude because density is decreased. Parasite drag coefficient is different at different altitudes and it varies with Mach no also. The drag in the subsonic region is mainly due to friction. As the Mach # increases above MDD, CD0 starts increasing due to formation of shocks and the increment in supersonic regime is purely due to wave drag. Induced drag is purely a pressure drag. It is caused by the wing tip vortices which generate an induced, perturbing flow field over the wing which in turn perturbs the pressure distribution over the wing surface. Oswald efficiency e is a function of the leading edge sweep also and that’s why as the e changes so the factor K. CL is also a function of velocity as the velocity decreases the value of CL increases to sustain the same amount of lift. So the variation of K is evident from the above graph with the CL and then with the increase of Mach #. 8) Drag Polar Drag polar is the standard presentation format for aerodynamic data used in performance calculations. It is simply the plot of coefficient of lift vs. coefficient of drag. Virtually all aerodynamic information of the aircraft is wrapped up in the drag polar. Drag polar graph shows the complete behavior of CL and CD at various mach numbers. With increase in CD there is a corresponding increase in CL for the constant Mach no. And for a constant value of CL there is a corresponding decrease in CD with increasing Mach numbers. Drag Polar: CD = CD0+KCL2 The slope of the tangent line from origin to the drag polar gives the point of maximum lift to drag ratio (L/Dmax). i. With increase in CL there is a corresponding increase in CD for the same Mach No. ii. For the same values of Cl there is a corresponding decrease in Cd with increase in Mach number iii. Subsonic drag-coefficient rise with an increase in lift coefficient is more prominent as the leading edge sweep is quite high. So for the same amount of drag produced the Lift generated by the wing is less. 26 College of Aeronautical Engineering,NUST,Risalpur,PK XXIII. Stability, Control and Handling Qualities . The basic concept of stability is simply that a stable aircraft, when disturbed, tends to return by itself to its original state (pitch, yaw, roll, velocity etc.). But Stability and Control are opposite to each other, so the best combination is a good tradeoff between stability and controllability. For jet fighters, the controllability requirement is comparatively higher than the stability requirement because the aircraft is supposed to carry out extreme maneuvers. Static Stability Static stability is present if the forces created by the disturbed state push in the correct direction to return the aircraft to its original state. The requirement for good stability, control and handling quantities are addressed through the use of tail volume coefficient method and through location of aircraft center of gravity at some percent of wing mean aerodynamic chord. Dynamic Stability Dynamic stability is present if the dynamic motions of the aircraft will eventually return the aircraft to its original state. The manner in which the aircraft returns to its original state depends upon the restoring forces, mass distribution, and damping forces. Damping forces slow the restoring rates. It should be noted that dynamic instability is not always unacceptable provided that it occurs slowly. Most aircraft have at least one unstable mode, the spiral divergence. This divergence mode is so slow that the pilot has plenty of time to make the minor roll correction required to prevent it. In fact, pilots are generally unaware of the existence of the spiral-divergence mode because the minor corrections required are no greater than the roll corrections required for gusts. Dynamic-stability analysis is complex and requires computer programs for any degree of accuracy. Therefore the stability analysis of design aircraft presented in this section evaluate static stability only. LONGITUDINAL STABILITY The overall Cmα of aircraft is positive indicating that aircraft is unstable in longitudinal axis. This is somewhat desirable for a fighter jet, as it makes the aircraft highly maneuverable. To cater for stability an automatic control system is employed. 27 College of Aeronautical Engineering,NUST,Risalpur,PK XXIV. CMARC and GAMBIT Results We tested the wing and tail geometry in CMARC and Gambit .but due to some errors we were unable to get require results .however the effort is shown here 28 College of Aeronautical Engineering,NUST,Risalpur,PK 29 College of Aeronautical Engineering,NUST,Risalpur,PK These screenshots shows the pressure and co-efficient of pressure distribution on the wing surface XXV. ANSYS RESULTS Ansys 12 was used for structural testing . due to selection of small AR i.e. 2.87 areas between the nodes were not large enough for meshing .so meshing through pressure distribution was used . Some errors were encountered and results are below 30 College of Aeronautical Engineering,NUST,Risalpur,PK 31 College of Aeronautical Engineering,NUST,Risalpur,PK XXVI. Morphological Charts It is a study of form or structure which means we draw various charts of our required design leading us to multiple options. So, basically morphology matrix is a concept generation process. We are interested in functions we need neglecting how those functions are performed. We combine these individual functions into overall concepts that meet all functional requirements. 32 College of Aeronautical Engineering,NUST,Risalpur,PK •These method can generate number of ideas if not controlled •Each function is considered independent •Sketches are drawn for the generated concepts Morphology Matrix for Fighter Aircraft Morphology matrix for fighter aircraft was drawn and seven concepts were generated considering configuration propulsion and structure domains. MORPHOLOGY CHART SYSTEMS C O N F I G U R A T I O N P R O P U L S I O N S T R U C T U R E S OPTIONS canopy Single double Intake location Wing root Over the wing Chin Armpit Inlet types Conical or spike 2-D ramp Pitot inlet Flush inlet Boundary layer diverter Bypass duct Step diverter Channel type Boundary layer suction Fuel tanks Internal External fixed Drop tanks Engine location Under wing Tail mounted Over wing fuselage cylindrical Area ruled oval Landing gear tricycle bicycle Quadricycle Number of engines single Twin engine Gears retraction Wing-fuselage junction Fuselage Weapon carriage internal External Weapon drop Guided unguided Semi submerged Rear fuselage conformal Nozzle Fixed convergent Variable convergent Wing position Anhydral dihedral Convergent iris Flight controls Hydraulic pneumatic electric mechanical Wing type trapezoidal delta rectangular blended Wing vertical position Mid wing High wing Low wing Horizontal surfaces Wing and elevators Canard and wing Delta wings Tail arrangement Conventional V.H tail v-tail Twin-Tail Missile launch mechanism Rail launch Ejection launch Shock Absorber Oleo pneumatic shocked strut Levered bungee engine type Turbojet Turbo fan Thrust Internal combustion IC and Afterburner Fuel system Integral bladder Discrete wing Materials Aluminum alloy Advanced composites Carbon composites Hybrid Fuselage Materials Aluminum alloy Advanced composites Carbon composites Hybrid Landing Gear Materials Titanium alloy Graphite Epoxy Low Carbon Steel Triple -Tail Triangulated 33 College of Aeronautical Engineering,NUST,Risalpur,PK XXVII. TRIZ METHOD APPLICATION TO PAF-X Problem Solving using TRIZ tools:TRIZ consists of 5 problem solving tools. These are listed below; • Inventive Principles to solve technical contradictions (the contradiction matrix) • Separation Principles to solve Physical contradictions (using available resources) • Standards for transformation of technical systems (for improving useful function and eliminating harm) • Scientific and Technical Effects (for synthesis of functions) • ARIZ - Algorithm to solve a (complex) inventive problem (with no explicit contradiction) PARAMETERS USING TRIZ METHOD 1- Weight of moving object 2- Weight of stationary object 3- Length of stationary object 4- Length of moving object 5- Speed 6- Force 7- Shape 8- Strength 9- Power 10- Reliability 11- Ease of repair 12- Ease of manufacture 13- Device complexity 14- Adaptability 15- Manufacturing precision 16- productivity Application of TRIZ method:There are 40 principles of TRIZ method. We used 08 out of those 40 principles and applied them on our design. The principles that we used are discussed below; Principle 1(SEGMENTATION): As this principle suggests to divide an object into independent parts, to make the object easy to assemble and disassemble. We used this principle of TRIZ method in our project, and we replaced our delta wing with the simple trapezoidal wing and elevators. Principle 2(TAKING OUT): This principle states the separation of an interfering part or property from an object, or single out the only necessary part. We used this principle of TRIZ method in one of our aircraft designs, and we eliminated the elevators and used the delta wing instead, which incorporates the use of elevators also. Principle 4(Asymmetry): This principle stresses upon the asymmetry of the object part under consideration. In our aircraft design we used wing root intake which gave symmetry to our aircraft then we changed side wing root intake to chin intake according to TRIZ principle 4(Asymmetry) which says ”Change from circular O-rings to oval cross-section to specialized shapes to improve sealing”. Principle 6(Universality): This principle suggests to make a part or object perform multiple functions; eliminate the necessity of other parts. Using this method, we used the effect of simple wings and elevators, and designed a delta wing, which also incorporated the function of elevators. We also used the Canard in replacement of the elevators, which increases the lift of an aircraft. 34 College of Aeronautical Engineering,NUST,Risalpur,PK Principle 7(Nested Doll): This principle tells us to place one object inside another; place each object, in turn, inside the other. Using this principle, we designed our intakes. The location of these intakes was at wing roots. We also used this concept foe designing the T-tail. Principle 8(Anti-Weight): Following Anti-weight principle which states that” To compensate for the weight of an object, merge it with other objects that provide lift” we shifted our elevators in the form of T-tail causing reduction in weight and producing more lift. While designing intake we changed our intake from wing root to chin type reducing aircraft weight. We used composite materials and Aluminum alloys which give more strength with less weight. Principle 24(Intermediary): Intermediary principle allows us to design removable parts so that extra weight drag and fuel consumption can be avoided. This principle tells “Merge one object temporarily with another (which can be easily removed)”. We used drop tanks in our design which can be removed according to prevailing situation to the pilot. Principle 26(Copying): TRIZ method suggests that instead of developing a total new product, which can be difficult to design, we can take a pre-designed product as a reference to start with. WE designed a chin intake, and we used F-16 aircraft as our reference for the design of chin intake. We used RAFAEL aircraft as our reference for the design of canard, and we designed canard for our aircraft. We used MIRAGE aircraft for the reference of delta wing, and we designed the delta wing for our aircraft using this reference. XXVIII. CONCEPT GENERATION These are the different concepts that were driven out of the main morphological matrix. TRIZ principles were also applied on these concepts. The catia model and the morphological matrix of all our concepts are shown below; Concept 1 CONFIGURATION canopy single Intake location wing root Inlet types pitot type Boundary layer diverter channel type Fuel tanks internal Engine location rear fuslage fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position Anhydral 35 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Flight controls electric Wing type delta Wing vertical position mid wing Horizontal surfaces canard and wing Tail arrangement conventional Vertical tail Missile launch mechanism ejection launch Shock Absorber oleo-pneumatic engine type turbo-jet Thrust internal combustion Fuel system bladder type wing Materials Advanced composites Fuselage Materials Advanced composites Landing Gear Materials titanium alloy 36 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 2 CONFIGURATION Canopy single Intake location wing root Inlet types pitot type Boundary layer diverter channel type Fuel tanks internal Engine location rear fuslage Fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided 37 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Nozzle fixed convergent Wing position dyhedral Flight controls pneumatic Wing type trapozoidal Wing vertical position mid wing Horizontal surfaces wing and elevators Tail arrangement conventional V.H tail Missile launch mechanism ejection launch Shock Absorber oleo-pneumatic engine type turbo-fan Thrust internal combustion Fuel system bladder wing Materials Aluminum alloy Fuselage Materials Aluminum aaloy Landing Gear Materials low carbon steel 38 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 3 CONFIGURATION Canopy single Intake location wing root Inlet types pitot type Boundary layer diverter channel type Fuel tanks internal Engine location rear fuslage Fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position dyhedral Flight controls electric Wing type trapozoidal Wing vertical position mid wing Horizontal surfaces wings and elevators Tail arrangement T-tail 39 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Missile launch mechanism rail launch Shock Absorber triangulated engine type turbo-jet Thrust internal combustion Fuel system integral type wing Materials Aluminum alloy Fuselage Materials aluminum alloy Landing Gear Materials low carbon steel 40 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 4 CONFIGURATION canopy single Intake location wing root Inlet types pitot type Boundary layer diverter channel type Fuel tanks External Engine location rear fuslage fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position dyhedral Flight controls electric Wing type Delta 41 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Wing vertical position mid wing Horizontal surfaces wings and elevators Tail arrangement V-tail Missile launch mechanism rail launch Shock Absorber triangulated engine type turbo-jet Thrust internal combustion Fuel system integral type wing Materials Aluminum alloy Fuselage Materials aluminum alloy Landing Gear Materials low carbon steel 42 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 5 CONFIGURATION canopy single Intake location Chin Inlet types pitot type Boundary layer diverter channel type Fuel tanks external Engine location rear fuslage fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position dyhedral Flight controls electric Wing type trapozoidal Wing vertical position mid wing 43 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Horizontal surfaces wings and elevators Tail arrangement T-tail Missile launch mechanism rail launch Shock Absorber triangulated engine type turbo-jet Thrust internal combustion Fuel system integral type wing Materials Aluminum alloy Fuselage Materials aluminum alloy Landing Gear Materials low carbon steel 44 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 6 CONFIGURATION canopy single Intake location chin Inlet types pitot type Boundary layer diverter channel type Fuel tanks internal Engine location rear fuslage fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position anhedral Flight controls electric Wing type delta Wing vertical position mid wing Horizontal surfaces wings and canards Tail arrangement V-tail 45 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Missile launch mechanism ejection launch Shock Absorber triangulated engine type turbo-fan Thrust internal combustion Fuel system discrete type wing Materials Advanced Composits Fuselage Materials Advanced Composits Landing Gear Materials Graphite epoxy 46 College of Aeronautical Engineering,NUST,Risalpur,PK Concept 7 CONFIGURATION canopy single Intake location wing root Inlet types pitot type Boundary layer diverter channel type Fuel tanks internal Engine location rear fuslage fuselage Area Ruled Landing gear tricycle Number of engines single Gears retraction wing-fuselage junction Weapon carriage external Weapon drop guided Nozzle fixed convergent Wing position anhyderal Flight controls electric Wing type delta Wing vertical position mid wing Horizontal surfaces wings and elevators Tail arrangement V-tail + lateral fins 47 College of Aeronautical Engineering,NUST,Risalpur,PK PROPULSION STRUCTURES Missile launch mechanism ejection launch Shock Absorber triangulated engine type turbo-fan Thrust internal combustion Fuel system discrete type wing Materials Advanced Composits Fuselage Materials Advanced Composits Landing Gear Materials Graphite epoxy 48 College of Aeronautical Engineering,NUST,Risalpur,PK XXIX. Cost Analysis Life-cycle cost analysis The other cost reduction technique is Life Cycle Cost Analysis to influence design. To avoid products being developed with a low acquisition cost but high operation and support costs, or vice versa, life cycle cost analysis considers the full cost of ownership of the products in the engineering and decision making process, to establish the correct balance of acquisition and in-service cost, whilst maintaining the required effectiveness of the product. Production costs In considering the life cycle cost elements of a project, which relate to the costs of design, development, manufacture, operation and support of the aircraft, this study now focuses on the manufacture element of the life cycle cost with particular reference to the ‘in-house’ production manufacture cost. LABOUR COST 49 College of Aeronautical Engineering,NUST,Risalpur,PK The planned method of manufacture, known as the process and mini statement of work, will form the basis of the labour cost. It will initially be generated in man-hours. This activity is undertaken via the Work Measurement Group who are part of the Production Engineering Department. Their task is to measure in man hours, the length of time required to complete the activity identified in each of the planned operations. Each of the operation times will conform to British Standards, thereby making sure that the values are accurate and consistent and this will be done for all of the parts within the assembly. DAPCA MODEL DAPCA (Development and Production Costs for Aircraft), a computer program designed to compute development and production costs for airframes, engines, and avionic systems. DAPCA generates cost-quantity unit data and cumulative-average improvement curves. The principal inputs relate mainly to aircraft physical characteristics, such as gross takeoff weight, speed, engine type, number of engines per aircraft, and thrust or shaft horsepower. Various options are also provided to the user when making a run. Data are retained and reused unless specifically deleted or superseded by new data, so that only changes in the inputs need to be entered for succeeding runs. Each computer run takes approximately three to four seconds. A sine-power function is used to generate the sustaining tooling production rate curve. Appendixes are included that give the costing equations used, flowcharts, and a complete listing of the FORTRAN IV source program. Using DAPCA IV model of cost estimation we have calculated different costs involved in product development. Engineering hours was calculated by It came out to be $2068760.7 Tooling hours was calculated by It came out to be $3664664.5 Manufacturing hours was calculated by It came out to be $2441392 Development support cost was calculated by It came out to be $3445789.25 Flight test cost was calculated by It came out to be $42947829.5 Manufacturing material cost was calculated by 50 College of Aeronautical Engineering,NUST,Risalpur,PK It came out to be $53422.5 Engineering production was calculated by It came out to be $1320472.2 XXX. PARAMETRIC MODEL VIEWS Bottom View Front View 51 College of Aeronautical Engineering,NUST,Risalpur,PK Side View XXXI. Conclusion We got RFP for MUTT and revised it for our fighter aircraft. Starting up from requirement analysis we did KJ and Kano analysis the QFD. After that we did initial and refined sizing and sizing of different aircraft components like fuselage, wings, horizontal and vertical tail sizing besides landing gear sizing. After the sizing we selected seven design concepts via morphology matrix and drew them parametrically in catia after applying Triz method which have been shown earlier in the report. Triz method are innovative principles which are common in our daily life. Triz method lists down the solutions to the basic daily life problems. We found the parametric modelling technique for aircraft design very designer friendly because the designs can be altered any time according to requirements. The different sizing parameters which we calculated were approximate to our reference aircraft F-22. We tried our best to design most efficient aircraft just as our reference F-22. Then during the aerodynamic analysis phase we used XFLR software for testing airfoils and stability analysis. In the beginning we came across many problems because the analysis was giving many problems but it provided us with a guideline to proceed further. At last, for structure and CFD analysis we had to use C marc and fluent. We tried but unfortunately we could not get accurate results, in accurate results of the analysis have been shown in the report. In a nutshell, parametric modelling is a handful tool for design and XFLR, C marc and fluent are the directive softwares for aerodynamics, stability and structural analysis. We as a group, learnt a lot during this design phase because the techniques we came across, would be much helpful in our future projects also. 52 College of Aeronautical Engineering,NUST,Risalpur,PK Acknowledgments We are thankful to Almighty for giving us strength to do little work. We would like to acknowledge our instructor Dr. Liaqat for able guidance in applying these methods in our design. We are thankful to 80 (A) E.C for their support. References 1. 2. 3. 4. 5. 6. Aircraft Design: A Conceptual Approach, 4th Ed., Daniel P Raymer Airplane Design I, J Roskam Aircraft performance and Design by John D. Anderson Jr. http://m-selig.ae.illinois.edu/ads/coord_database.html Concept design; TRIZ (Teoriya Resheniya Izobreatatelskikh Zadatch ) “Automated conceptual design of mechanisms using improved morpholigical matrix”, Y Chen, P Feng, ASME Journal of mechanical engineering,2006. 53 College of Aeronautical Engineering,NUST,Risalpur,PK