AIAA Student Design Competition
Homeland Defense Interceptor: Gavial Virginia Polytechnic Institute
VersaCorp AeroSpace
...
May 12, 2006
1 Executive Summary
In response to the Request for Proposal [1] from the AIAA Foundation Undergraduate Team Aircraft
Design Competition, VersaCorp Aerospace from Virginia Polytechnic Institute and State University proudly presents Gavial, a next generation Unmanned Combat Aerial Vehicle, Homeland Defense Interceptor (UCAV-
HDI). The Gavial meets the design mission requirements of providing superior operational capability while maximizing cost-effectiveness in all facets of the design.
The Gavial is designed primarily for supersonic performance. This led to the use of a highly swept cranked arrow wing with a blended fuselage. Primary control is provided by canards and a single vertical tail. The Gavial also utilizes a single engine in the 35,000 lb thrust class. Missiles are stored externally on under-wing hard-points, and rail launched. The flyaway cost of the Gavial is ˜$15,000,000, with half˜ of that being for materials and systems.
The Gavial is capable of performing three distinct missions, each contributing significantly to the aircraft’s overall mission of ensuring homeland security. The streamlined fuselage is equipped with the M61-A1 Vulcan
20mm Gatling rotary gun which is driven by the aircraft’s hydraulic system and has a maximum rate of
fire of up to 7,200 shots per minute (SPM), providing the Gavial with excellent lethality during dog-fighting and pinpoint attacks. The wing undercarriage is capable of carrying a maximum of four AIM-120 advanced medium-range air-to-air missile (AMRAAM), four AIM-9 Sidewinder missiles, or a combination of two AIM-
120 missiles and up to two AIM-9s. Such arsenal versatility optimizes the Gavial’s performance in the RFP specified Defensive Counter-Air Patrol (DCAP), Intercept/Escort and Point Defense Intercept missions.
The option of mounting three 660 gallon exterior fuel tanks provides the capability of fulfilling the DCAP missions four hour loiter requirement. As a supersonic performer, the Gavial is designed to minimize drag and maximize maneuverability at the design altitude of 35,000ft, out-performing many operational features of modern fighters. The inclusion of the most advanced radar and UCAV communications systems, including a portable next-generation ground station, maximizes mission effectiveness while simultaneously decreasing the overall cost of the aircraft and completely eliminating pilot risk. This highly survivable and versatile aircraft highly exceeds many of the RFP requirements with an Initial Operational Capability (IOC) date of
2020 and a flyaway cost of $15 million.
i ii Summary of Requirements
Criterion Requirement DCAP I/E
Intercept Mission Radius 200 nm N/A Met
CAP Endurance at 300 nm Radius 4 Hrs Met N/A
Performance at Maneuver Weight
Maximum Mach Number at 35,000 ft 2.2 Not Met 1.7 Met 2.2
Min Sustained Load Factor -at Maximum Thrust 5 g´s Met 9 g’s Met 9 g’s
Max Instantaneous Turn Rate at 35,000 ft 18.0 ◦/s Met 21.2 ◦/s Met 21.2 ◦/s
1-g Specific Excess Power -Military Thrust
0.9M/Sea Level 200 ft/s Met 1401 ft/s Met 1500 ft/s
0.9M/15,000 ft 50 ft/s Met 983 ft/s Met 1000 ft/s
1-g Specific Excess Power -Maximum Thrust
0.9M/Sea Level 700 ft/s Met 896 ft/s Met 1053 ft/s
0.9M/15,000 ft 400 ft/s Met 468 ft/s Met 603 ft/s
5-g Specific Excess Power -Maximum Thrust
0.9M/Sea Level 300ft/s Met 354 ft/s Met 496 ft/s
0.9M/15,000 ft 50 ft/s Met 103 ft/s Met 215 ft/s
iii Contents
1 RFP Analysis 1
1.1 Missions ...... 1
1.1.1 DCAP Mission ...... 1
1.1.2 PDI Mission ...... 2
1.1.3 I/E Mission ...... 2
1.2 Stability and Control Requirements ...... 3
1.3 Performance Requirements ...... 3
1.4 Equipment ...... 4
1.4.1 Engine ...... 4
1.4.2 Government Furnished Equipment ...... 4
2 Survey of Existing Aircraft 5
3 Initial Sizing 7
3.1 Wing Loading and Thrust to Weight Ratio ...... 7
3.2 Fuel Weight Estimation ...... 10
4 Conceptual Design 12
5 Configuration Description 18
6 Operational Concept 20
6.1 Manned Vs Unmanned ...... 20
6.2 Threat Identification, Threat Tracking, Target Verification, Kill Verification, Collision Avoidance 21
6.3 Weapons Integration ...... 22
6.4 Loss of Vehicle Command and/or Control ...... 22
6.5 Handling of Enemy Countermeasures ...... 22
6.6 Aircraft Launch and Recovery Scheme ...... 23
6.7 Items/Levels of Redundancy Required ...... 23
6.8 Benefits/Limitations ...... 23
iv 7 Ground Control Station 24
7.1 Design Objectives ...... 24
7.2 The Command Chair ...... 24
7.3 Helmet-mounted Heads Up Display (HUD) Screen ...... 25
7.4 Vehicle Management System (VMS) ...... 27
7.5 Final Layout and Platform Assembly ...... 29
8 Aerodynamics 31
8.1 Wing Planform ...... 31
8.2 Airfoil Selection ...... 31
8.3 High-Lift Devices ...... 35
8.4 Drag ...... 35
9 Stability and Balancing Scheme 42
9.1 Longitudinal Control ...... 42
9.2 Control Surfaces ...... 42
9.3 Canard Sizing ...... 43
9.4 Control Authority ...... 44
10 Vehicle Performance 48
11 Propulsion Design 53
11.1 Single vs Multi Engine ...... 54
11.2 Engine Scaling ...... 54
11.3 Inlet Design ...... 55
11.4 Shock Ramps ...... 56
11.5 Nozzle ...... 58
12 Structural Design 59
12.1 Materials Selection ...... 59
12.2 Fuselage Structure ...... 60
12.3 Wing Structure ...... 61
12.4 Canard and Tail Structure ...... 63
12.5 Landing Gear ...... 63
12.6 Structural Limits ...... 65
v 13 Weapons/Fire Control System 67
13.1 Weapons ...... 67
13.2 Wide Angle Cameras ...... 68
13.3 FLIR ...... 69
13.4 Radar ...... 69
14 Internal Systems 70
14.1 General Layout ...... 70
14.2 Cooling of the Avionics ...... 73
15 Fuel and Electrical System 74
15.1 Fuel System ...... 74
15.2 Electrical System ...... 75
16 Servicing Plan 76
16.1 Forward Avionics Access ...... 76
16.2 Aft Access ...... 76
17 Weight Parameters 79
17.1 Takeoff Gross Weight ...... 79
17.2 Center of Gravity ...... 79
17.3 Weight and Balance Analysis ...... 80
18 Cost Estimation 84
vi List of Figures
1.1 Defensive Counter Air Patrol Mission Schematic ...... 1
1.2 Point Defense Intercept Mission Schematic ...... 2
1.3 Intercept Escort Mission Schematic ...... 3
3.1 Sizing Chart for Conceptual Design ...... 9
4.1 Configuration Flow Chart ...... 12
4.2 Variable Geometry Concept ...... 13
4.3 Arrow Wing Concept ...... 14
4.4 Conventional Wing Concept ...... 15
4.5 Multiple Lifting Surface Concept ...... 16
4.6 Drag vs Sweep Angle Comparison ...... 17
5.1 Gavial Interceptor Configuration ...... 18
6.1 Operation Control Concept ...... 21
7.1 Martin Baker MK-16A Ejection Seat [2] ...... 25
7.2 Aluminum Frame with Contoured Seat Panels, Controls and Computer ...... 26
7.3 Computer ...... 26
7.4 Helmet-Mounted HUD Visor ...... 27
7.5 VMS System and Housing. (Red arrows highlight directions of motion) ...... 28
7.6 Auxiliary Screen ...... 28
7.7 The Final Ground Station Layout (3 views) ...... 30
8.1 Planform Characteristics (dimensions in inches) ...... 32
8.2 Planform Characteristics [3] ...... 33
8.3 Airfoil Section Drag Polars [3] ...... 34
8.4 NACA 641-406 Airfoil[3] ...... 34
8.5 High Lift Configuration ...... 35
8.6 Increased Lift Due to Flap System ...... 36
vii 8.7 Intercept Configuration ...... 38
8.8 Defensive Counter Air Patrol Configuration ...... 39
8.9 Mach 1.2 Drag Breakdown ...... 40
8.10 Mach 0.9 Drag Breakdown ...... 40
8.11 Total Vehicle Drag Variance with Mach Number ...... 41
9.1 Logintudinal Control Planform ...... 43
9.2 Lateral Directional Control Planform ...... 43
9.3 X-plot for Canard Sizing ...... 44
10.1 Sizing Chart for Conceptual Design ...... 49
10.2 Turn Rate vs Mach Number from ACS ...... 51
10.3 Maximum Sustained Load Factor ...... 52
10.4 Excess Power Available ...... 52
11.1 GE F-414 Engine ...... 53
11.2 Thrust vs SFC for Cruise Mach Number of 0.9 ...... 55
11.3 Mach Number vs Thrust for Varying Altitudes ...... 56
11.4 Propulsion Unit - Inlets, Ramps, Engine and Power Systems ...... 57
11.5 Shock Ramp Schematic for Mach 2.2 ...... 57
11.6 Nozzle Geometry, Supersonic (left) and Subsonic (right) ...... 58
12.1 Fuselage Structural Layout ...... 61
12.2 Wing Structural Layout ...... 62
12.3 Landing Gear Schematic View ...... 63
12.4 Forward Landing Gear ...... 64
12.5 Main Landing Gear ...... 65
12.6 V-n Diagram for Maneuver and Gust Loads ...... 66
13.1 Weapons Layout ...... 67
13.2 IR360-1 Camera.[4] ...... 68
13.3 CCD Array in Canopy ...... 68
13.4 AN/ASQ-228 ATFLIR.[5] ...... 69
viii 13.5 AN/APG-79.[6] ...... 69
14.1 Bandwidth Limitations ...... 71
14.2 Systems Layout ...... 72
15.1 Fuel System (shown in green) ...... 74
15.2 Lateral Directional Control Planform ...... 75
16.1 Forward Access Panels ...... 77
16.2 Aft Access Panels ...... 78
17.1 CG Limits ...... 82
17.2 CG Excursion Chart: Defensive Counter Air Patrol ...... 82
17.3 CG Excursion Chart: Point Defense Intercept ...... 83
17.4 CG Excursion Chart: Intercept/Escort ...... 83
18.1 Airframe Fly Away Cost Breakdown ...... 84
ix List of Tables
1.1 Specific Excess Power Requirements ...... 3
2.1 Current Generation Aircraft [7] [8] [9] [10] [11] ...... 5
2.2 Pros and Cons of Comparison Aircraft[7] [8] [9] [10] [11] ...... 6
3.1 Performance Parameters ...... 10
3.2 Fuel Use by Mission Segment for Defensive Counter Air Patrol ...... 11
3.3 Fuel Use by Mission Segment for Point Defense Intercept ...... 11
3.4 Fuel Use by Mission Segment for Intercept/Escort ...... 11
4.1 Qualitative Assessment Summary...... 15
6.1 Pro/Con Comparison for Unmanned Concept ...... 20
8.1 Drag Characteristics (Planform Reference area: 566ft2) ...... 37
9.1 Logitudinal Derivatives ...... 45
9.2 Lateral-Directional Derivatives ...... 45
9.3 1-g Trim Assessment ...... 46
9.4 Maneuver (Pull-up) Assessment ...... 46
9.5 Steady Sideslip ...... 47
9.6 Time-to-Bank Summary. Military Specifications and Gavial Performance. Time in seconds. . 47
11.1 Engine Selection ...... 53
12.1 Material Selection ...... 59
12.2 Landing Gear Specifications ...... 63
17.1 Aircraft Dry Weight by Component ...... 80
17.2 TOGW by Mission ...... 81
17.3 Center of Gravity Location by Mission ...... 81
x List of Symbols
A Area CG Center of Gravity D Diameter e Spanwise Efficiency L Length LE Leading Edge M Mach Number m˙ Mass Flow Rate Ps Specific Excess Power 1 2 q Dynamic Pressure ( 2 ρV ) Re Reynold´sNumber SF Scale Factor SM (frac−CmCL) subsonic longitudinal static margin SFC Specific Fuel Consumption T Thrust t/c Thickness-to-Chord Ratio TE Trailing Edge TOP Takeoff Parameter V Velocity W Width, Weight
CD Coefficient of Drag CDfriction friction drag coefficient CDinduced induced drag coefficient CDwave wave drag coefficient CDform form friction drag coefficient CDtotal total drag coefficient CL Coefficient of Lift CL0 zero-lift lift coefficient CLmax maximum lift coefficient CLq rate of change in lift coefficient with change in dynamic pressure CLδf rate of change of lift coefficient with flaperon deflection at constant α CLδc rate of change of lift coefficient with canard deflection at constant α Clp damping-in-roll Clr rolling moment due to yawing Clα lift-curve slope, rate of change of lift coefficient with α Clβ rate of change of rolling moment with sideslip angle Clδa rate of change of lift coefficient with aileron deflection Clδr rate of change of lift coefficient with rudder deflection Cm0 zero-lift pitching moment coefficient Cmq damping-in-pitch Cmδf rate of change of moment coefficient with flaperon deflection at constant α Cmδc rate of change of moment coefficient with canard deflection at constant α Cnp yawing moment produced by rolling motion Cnr damping-in-yaw Cnβ rate of change yawing moment with sideslip angle Cnδa rate of change of normal force coefficient with aileron deflection
xi Cnδr rate of change of normal force coefficient with rudder deflection Cyβ rate of change of side force with sideslip angle Cyδr rate of change of side force coefficient with rudder deflection
α Angle of Attack β Angle of Sideslip σ Pressure Ratio ρ density
xii Chapter 1
RFP Analysis
The RFP supplied a challenging set of requirements, and wide lattitude in terms of acceptable solutions.
Initial possibilities ranged from a retrofit of existing aircraft to the development of a brand new Unmanned
Combat Air Vehicle (UCAV). Two primary missions were specified: Defensive Counter Air Patrol (DCAP),
Point Defense Intercept (PDI). There was also a secondary mission to be analyzed, Intercept/Escort (I/E)
These missions are described in detail below.
1.1 Missions
1.1.1 DCAP Mission
The DCAP mission (Figure 1.1) begins with a launch and optimum cruise out to 300 nm. Following this, a four hour loiter patrol. After four hours, the plane is required to dash 100nm at maximum velocity, Mach
2.2. A combat phase follows this, one 360◦ turn at Mach 0.9, and one at Mach 1.2, expending all missiles and retaining M61-A1 ammunition. After this, a 400 nm cruise home at best speed and velocity, with reserve fuel for 30 minutes of loiter before landing. [1]
Figure 1.1: Defensive Counter Air Patrol Mission Schematic
This is the most challenging mission from a design standpoint. The four hour loiter drives the airplane toward straight high span wings with a low wing loading. The supersonic mission segment, as well as the
1 combat requirements drive the aircraft toward short, swept, highly loaded wings. These two conflicting requirements are the major balance that needs to be struck in the design. Some options for minimizing this dichotomy are increasing the fuel efficiency of the aircraft, both by improving the engine and minimizing zero lift drag.
1.1.2 PDI Mission
The PDI mission Figure 1.2 is simpler than the DCAP, with less extreme requirements. After takeoff, a 200 nautical mile, Mach 2.2 dash at 35,000 ft. Next, the combat phase, with the same turns as the DCAP, firing all missiles, then return to base at best cruise. Again, the aircraft must maintain 30 minutes of reserve fuel.
Figure 1.2: Point Defense Intercept Mission Schematic
Since the PDI requires a long supersonic flight, with no significant loiter requirement, this pushes the overall system optimization toward supersonic performance. Many of the driving factors, like minimizing zero lift drag, stay the same. This mission will drive the wings to be smaller, and of a higher loading than if only the DCAP mission were considered.
1.1.3 I/E Mission
A third mission, Intercept and Escort Figure 1.3, is specified, but not as a primary design mission. The
Interceptor must dash out at top speed to intercept a target, which it then escorts at minimum feasible speed for 300nm, then cruise home. In this mission, the aircraft is to retain all weapons.
2 Figure 1.3: Intercept Escort Mission Schematic
1.2 Stability and Control Requirements
The stability and control requirements mandated by the RFP are brief. The un-augmented longitudinal static margin is to be confined to the range of no greater than 10% and no less than -10%, Aircraft designs that are statically unstable in the longitudinal axis require a digital flight control system. The stability scheme must fit these requirements as well as be appropriate for meeting the mission performance requirements.
1.3 Performance Requirements
Beyond the requirements on range and endurance specified in the individual missions, there are several individual performance requirements that must be met. It must be able to sustain a maximum Mach number of 2.2. It must be able to withstand +/−5 g’s load factor, and make a maximum instantaneous turn rate of 18◦/second. The HDI must also meet the following specific excess power requirements outlined in
Table 1.3.
Table 1.1: Specific Excess Power Requirements Altitude (ft) g’s Thrust Class Specific Excess Power (ft/sec) Sea Level 1 Military 200 Sea Level 1 Maximum 700 Sea Level 5 Maximum 300 15,000 1 Military 50 15,000 1 Maximum 400 15,000 5 Maximum 50
3 1.4 Equipment
1.4.1 Engine
The AIAA supplied an engine deck for the GE F-414-400 class low bypass turbo jet. The F-414 is used in the US F/A-18E/F series of aircraft. It is required that the geometry and engine cycle be completely described, and it must provide 50 kw of power to all the subsystems, as well as divert 2% of the inlet flow to cool the avionics.[1] The use of this specific engine is not mandated. Any engine selected is required to be non-developmental, i.e. currently in production or forecast to be in production within ten years. This limited the choices to engines whose cores are currently in use, though not necessarily members of the F-414 series. The engines can also be scaled up and down depending on the needs of the design.
1.4.2 Government Furnished Equipment
The cost of the aircraft was not to exceed $15 million, and to that end the RFP specified that most of the internal equipment should be government furnished. All internal avionics will be off the shelf from currently available systems. The radar, targeting pods will all be chosen from existing products to reduce costs. It is prefereable to the design team, however, that of existing components, only top of the line equipment is to be considered.
4 Chapter 2
Survey of Existing Aircraft
The first step in the design process was a survey of existing technology, to see if any current designs could be re-tooled to meet the requirements. Many high quality current generation fighter aircraft were researched, including the majority of aircraft in active service to the USAF and US Navy. In addition, the two newest
American aquisitions were studied, as were two Russian fighters and the new Eurofighter. Table 2.1 summa- rizes the results of comparing existing aircraft to the requirements of the RFP. As can be seen, all existing aircraft, with the possible exception of the MiG-29 are too expensive to suit the RFP. The flyaway costs for aircraft are not routinely published, but the unit costs are indicative of a high flyaway cost in all but the
MiG-29 and F-16. In addition, most lacked the necessary combat radius, and all seemed to lack the CAP cpability required by the first mission. The F-14, F-16, and F-18 (all models) missed the mark when it came to maximum Mach Number. The other aircraft can, at least nominally, reach the required Mach 2.2.
Table 2.1: Current Generation Aircraft [7] [8] [9] [10] [11] Aircraft Unit Cost Where it fails the RFP F-14 $38,000,000 range, top speed, loiter, cost F-15 $35,000,000 range, loiter, cost F-16 $26,900,000 range, top speed, loiter, cost F/A-18 $40,000,000 range, top speed, loiter, cost F-22 $92,400,000 range, loiter, cost F-35 $33,000,000 range, loiter, cost Su-27 $30,000,000 loiter, cost Mig-29 $20,000,000 range, loiter, cost Euro-fighter Typhoon $90,000,000 range, loiter, cost
The aircraft summarized in table 2.1 all fail the RFP in one way or another. All of these aircraft did posess some distinguishing features that bore further scrutiny in the design of an interceptor. Table 2 summarizes the pros and cons of each aircraft. All aircraft were clearly oversized, at least in reference to maximum ordnance loads. The requirement is to carry, at most, a 20mm cannon and 4 AMRAAM missiles; the smallest of these aircraft can carry at least twice that load in missiles alone. Initially, the intention was to scale one of these aircraft down and upgrade performance. As analysis progressed, it was determined that
5 a simple scale change would be insufficient to meet the disparate requirements of the DCAP’s long loiter and the long supersonic dash of the PDI mission. Ultimately, it was determined that a new, ground-up design of the interceptor would be the optimal method of meeting the requirements.
Table 2.2: Pros and Cons of Comparison Aircraft[7] [8] [9] [10] [11] Aircraft Positive Attributes Negative Attributes F-14 variable sweep, extremely heavy ordinance capacity oversized for the mission F-15 high ordinance capacity, high speed, oversized for the mission fuel capacity F-16 low cost, ordinance capacity, short range low weight F/A-18 low drag body design, short wingspan ordinance capacity oversized for the mission F-22 ordinance capacity, high speed, high cost high wing area, thrust vectoring oversized for the mission F-35 multiple design variants, low drag body design, high cost high wing area Su-27 low cost, high speed, foreign aircraft high instantaneous turn rate oversized for the mission Mig-29 low cost, high speed, foreign aircraft high instantaneous turn rate oversized for the mission Euro-fighter Typhoon high speed, delta wing design, high cost, foreign aircraft high wing area oversized for the mission
Several design features of these aircraft were desirable for the interceptor, for instance, the high sweep angles of the Typhoon and MiG were desirable from the standpoint of reducing supersonic drag. The strake effect on the F-15, F-16, and F-18 was useful for high-speed stability. The inlet plan on the F-14 and F-15 is known to maintain performance at high angles of attack, which is necessary for an agile fighter like the HDI.
The wing-body blending in the F-22, as well as the incorporation of the inlets into the sides of the fuselage simplified the configuration and structural layout, as well as having potential for drag reduction.
6 Chapter 3
Initial Sizing
3.1 Wing Loading and Thrust to Weight Ratio
After reviewing existing aircraft, the next step was to establish a preliminary size based on mission con- straints. The first sizing cut was taken using a basic sizing chart formulated basic equations relating the wing loading and thrust to weight ratios. Parameters for e, AR, CD0 were estimated using average values for modern fighter aircraft. In order to begin the initial sizing requirements, the maximum allowable wing loading was determined. This wing loading occurs during landing, and the following equation was used to determine this value:
W S − S = l a σC = 92.1375 (3.1) S 80 L−T akeoff
In equation 3.1 Sl represents the landing field length and Sa represents the runway length needed to clear a 50 foot object. Take-off parameters were added to the sizing chart next. Equation 3.2 is the relationship that equates thrust to weight and wing loading in takeoff.
W T = T OP σ + C (3.2) S L−T akeoff W
The third criterion for the missions was to perform a sustained 5g maneuver, governed by the following:
T C W n2 = q D0 + (3.3) W W S qeπAR S A transonic cruise requirement was also needed as specified by the RFP. This requires the aircraft to sustain a Mach number of approximately 0.9. In order to plot this criterion the maximum lift to drag ratio
first needed to be determined using equation 3.4.
√ L πAReC = D0 (3.4) D Max 2CD0
7 The thrust to weight ratio at cruise is defined to be 40% of the installed thrust to weight. According to
Raymer [12] this is on the low end of the spectrum, however a high excess thrust capability is required to reach the maximum Mach number of 2.2. Equation 3.5 shows this:
L W = (3.5) D Max 0.4T0
In order to plot the requirement for transonic cruise the equation is then manipulated and is shown in equation 3.6.
√ T q 3 (5C ) = D0 (3.6) W W S Additionally, the excess power requirements were considered for sizing. Equations 3.7 and 3.8 show how the excess power is incorporated into the thrust to weight ratios and the wing loading. In this case Ps is the excess power. The RFP required various excess power conditions, as previously specified in Table 1.3.
V (T − D) P = (3.7) s W
T D P = V − (3.8) s W W
Drag and dynamic pressure were calculated using equations 3.9 and 3.10.
D = 5qCd (3.9)
CD q = W (3.10) S The drag coefficient is needed in equation3.10 and must also be solved for. Equation 3.11 shows how this is done. In equation 3.11 the lift coefficient is used and needs to be determined before solving for the drag coefficient. Using equation 3.12 to solve for the drag coefficient and using that in equation 3.10, equation
3.8 can now be plotted for each of the required excess power numbers.
C2 C = C + L (3.11) D D0 πARe
8 n W C = (3.12) L q S
The excess power curves were plotted along with the curves for landing, take-off, transonic cruise, and the required maneuver. Figure 3.1 shows these curves together. These sizing requirements along with the drag and other aerodynamic constraints are the governing factors on the design of the aircraft.
Figure 3.1: Sizing Chart for Conceptual Design
The take-off curve was found to be far less than the maximum thrust to weight ratio and therefore will not play a significant role in the sizing of the aircraft since the engine will have sufficient thrust to overcome the necessary requirements. The dominant curves in for sizing were the sustained g maneuver and transonic cruise requirements. The intersection of these two curves is the optimal area for design, specifying a thrust to weight of 0.8, and a wing loading of approximately 70 lb/ft2.
9 Table 3.1: Performance Parameters Take Off Parameter 300 Density Ratio, σ 0.8107 CL at Takeoff 1.3 AR 3.5 Cd0 - subsonic 0.015 Cd0 - trans/supersonic 0.03 T/Wcruise/T/Winstalled 0.4 Weight Bombs [lbs] 1308 Weight Fixed (including cannon) [lbs] 2510 Weight of combat fuel [lbs] 1750 Subsonic cruise Mach number 0.9 Supersonic cruise Mach number 1.2 Maneuver Mach Number 1.2 Design Altitude (ft) 35,000
3.2 Fuel Weight Estimation
Using Raymers text [12], certain initial parameters for fighter aircraft could be determined and are shown in Table 3.1. These values were used in a Nicolai [13] sizing routine to do initial fuel weight estimates. After the basic sizing ratios were determined, overall vehicle weight was predicted.
The fuel estimates were derived from the scaled (detailed in chapter 11) engine deck for each mission.
It was determined that the defensive/counter-air patrol was the mission that required the most fuel and the plane was sized to that criterion. The mission fuel weight was calculated using the desired speed and thrust the aircraft would require during each mission phase. The total weight and thrust scale was calculated iteratively. After several iterations, the plane was sized to a total weight of 43,000˜ lb, and a thrust class of
35,000 lbs. The thrust required was margined up from the initial estimate to leave room for growth. The optimum speed for the cruise sections was calculated to be Mach 0.9 at 40,000 ft. The amount of fuel used during the various missions was then calculated, and is broken out by mission segments in the tables.
For the other missions, the same process and speeds were used to calculate the fuel weight. The point defense intercept mission’s fuel weight can be seen in Table 3.3 and the intercept/escort mission can be seen in Table 3.4. The fuel system design point is clearly the DCAP mission. The rest of the systems, based on historical estimates from Roskam [14] made up approximately 13,000lbs. This yielded a nigh unprecedented fuel fraction for the DCAP mission of almost 75%. The group decided, however, that based on the extreme nature of the required mission, and that no other existing aircraft rivaled the scope of this mission, historical precedent was not necessarily accurate.
10 Table 3.2: Fuel Use by Mission Segment for Defensive Counter Air Patrol Mission Segment Fuel Used (lbs) Take-Off and Climb to Altitude 793.2 Cruise 300 nm 2508.0 4 Hour Loiter @ 35,000ft 15236 Dash 100 nm @ 35,000ft 5351.6 Combat 1850.8 Cruise 400 nm to Base 3344 30 Minute Reserve 1166.3 Total 30250
Table 3.3: Fuel Use by Mission Segment for Point Defense Intercept Mission Segment Fuel Used (lbs) Take-Off and Climb to Altitude 793.2 Cruise 300 nm 2508.0 Dash 200 nm @ 35,000ft 10703.3 Combat 1850.8 Cruise 200 nm to Base 1672 30 Minute Reserve 1166.3 Total 16185.6
Table 3.4: Fuel Use by Mission Segment for Intercept/Escort Mission Segment Fuel Used (lbs) Take-Off and Climb to Altitude 793.2 Dash 200 nm @ 35,000ft 10703.3 Escort 300 nm a@ M0.6 7320.348 Cruise 500 nm to Base 5350.0 30 Minute Reserve 1166.3 Total 25333
11 Chapter 4
Conceptual Design
Conceptual design began with each member of the team creating individual concepts. Each team member submitted their designs to qualitative assessment by the group. The individual designs were then analyzed for similarity and pared down. For instance two designs used delta wings, and two others used a cranked arrow configuration. These four designs were combined into a single arrow wing design. Several conventional wing designs were also proposed, and those were also unified. One design proposed variable geometry, and this design was carried forward. The one design proposed that had a forward swept wing was discarded.
Forward swept wings are historically not suited to supersonic flight. The team could not conceive of any benefits to the forward swept design, so it was discarded. The overall conceptual design flowchart is shown in Figure 4.1.
Figure 4.1: Configuration Flow Chart
12 The three unified designs were joined by a fourth concept. This was a multi-surface design intended to boost loiter efficiency by increasing Oswald’s Efficiency factor while maintaining a low aspect ratio to keep maneuverability and a high fineness ratio for dash capability. The four concepts, presented below, represent the design options chosen for investigation. All designs conformed to the initial sizing determined in chapter
3.
Figure 4.2: Variable Geometry Concept
The first concept (Figure 4.2) was inspired by the F-14 Tomcat. Variable geometry wings would have allowed for optimal flight configurations for both the extended loiter in the CAP mission, low speed efficiency for the escort as well as a tight delta for supersonic dash. Initially, the variable geometry looked like the best choice. Upon further examination of the concept, several flaws appeared. First and foremost, variable geometry wings add complexity and expense, especially in the connection pins and manufacturing of the wing itself. Additionally, the variable sweep mechanism is bulky and consumes large amounts of internal volume that would either decrease available fuel volume or increase volumetric wave drag.
The second concept was a standard notched arrow wing (Figure 4.3). It originated with a delta wing after the Eurofighter, which was cut down to bring the area into line with the desired 500 ft2. This was done instead of reducing span and length because the vehicle was sized with an aspect ratio of 3 in order to reasonably meet the CAP endurance requirement. The delta was desirable to increase supersonic efficiency, and the notched arrow modifications increase that efficiency further[13].
13 Figure 4.3: Arrow Wing Concept
The next design was a conventional wing vehicle (Figure 4.4) based loosely on the F-15 and the Su-27.
The main selling point to this design was it similarity to planes which had met the top speed requirement.
The other primary argument in favor of conventional wing vehicles is their relative simplicity as a well tested and proven design.
The final design (Figure 4.5) was a non-standard wing design that emphasized non-planar lifting surfaces to increase Oswald’s Efficiency, rather than aspect ratio as a means to increase loiter performance. The benefits of this design were believed to be an increase in fineness ratio and maneuverability without loss of loiter capability. On top of that, the aerodynamic surfaces were to be fully integrated, such that no tail or canard would be necessary, reducing weight. The downside to this design was a lack of available design tools able to handle the aerodynamics, as well as level of variation available in the concept.
Each of these four designs were analyzed at a low level and subjected to a qualitative comparison to determine which was the best choice to continue on to preliminary design. The pros and cons of each design are summarized in Table 4.
After this assessment, the Multiple Surface concept(Figure 4.5) was removed from consideration. It was determined that the lack of existing design and analysis tools for such a vehicle were prohibitive given the
14 Figure 4.4: Conventional Wing Concept
Table 4.1: Qualitative Assessment Summary. Concept Pros Cons Variable Geometry Optimum configuration over all flight ranges Expensive Bulky internal components Added mechanical weight Arrow/Delta High supersonic efficiency Poor loiter properties Conventional Simple, proven design Wings extend beyond shock cone Straight, high AR wings for loiter (increased supersonic induced drag) Multi-Wing Fewer Aerodynamic surfaces Complex, untested aerodynamics Reduced Surface Area Poorly defined design envelope time and resource constraints. In addition, there were area distribution concerns, due to the presence of four large lifting surfaces in the same place as the engine pods would yield a large maximum cross section, as well as putting this maximum farther aft on the aircraft than desirable. The variable geometry concept was removed several weeks later. The extra complexity required for swing-wing designs is quite undesirable.
Systems are needed both to move the wings, and determine when and at what angle they should be moved.
More systems mean more possibilities for failure, which could mean anything from undesirable control conditions to catastrophic failure, and therefore the loss of a very expensive aircraft, and possibly a pilot.
Structurally, the variable sweep design requires much more support than a fixed-wing aircraft. Large loads will be experienced by wing movement at high flight speeds, making thicker/stronger wing ribs, struts, and
15 Figure 4.5: Multiple Lifting Surface Concept fuselage-wing connections at the root necessary. This, in turn, increases the weight of the aircraft. A third disadvantage to the variable sweep design is quite obvious: cost. One major cost factor in the F-14 Tomcat is the manufacturing of the wing box for the moving section of the wing. Since our aircraft was to remain below $15 million, it was determined that variable sweep geometry was ultimately too costly for this mission.
Since the Gavial is required to remain under the aforementioned cost, and still have a turn rate of 18 ◦/sec and reach Mach 2.2, the extra weight and internal volume (supersonic drag) associated with a variable sweep wing was highly undesirable.
This brought the conventional and arrow configurations forward to quantitative assessment. In almost all cases, the planes were equivalent. The cranked arrow planform, however, gave approximately 10% less drag than the conventional wing from preliminary estimates. The drag comparison is shown in Figure 4.6.
Since drag reduction is the most paramount requirement in supersonic aircraft, the cranked arrow (Figure
4.3) became the obvious choice. While the initial concept was designed without separate pitch control surfaces, canards were added to simplify the control design and increase the allowable CG travel, which in turn simplified overall layout and weight adjustments. The preliminary design process was dominated by the desire to increase simplicity in all systems, both to reduce manufacturing costs and save weight on unneccesary actuation. When the decision to use a cranked arrow wing instead of a pure delta was made, the two seep angles of 48˜ and 66˜ degrees showed promise as low drag angles.
16 Figure 4.6: Drag vs Sweep Angle Comparison
17 Chapter 5
Configuration Description
The Gavial was designed as an unmanned high-speed interceptor. The design was geared towards this goal, making some sacrifices in other areas to maximize high speed performance. The design process led eventually to the selection of a cranked arrow / canard configuration aircraft. The cranked arrow was chosen over a straight delta wing in order to reduce the overall wing area while maintaining a high span. Initially, the planform was based on the smoothly curved wings of the Concorde airliner, to remove the corners where strong shock waves could develop. Due to manufacturing concerns, this smoothly curved shape was abandoned in favor of the simple planform seen in figure 5.1. Adopting the straight leading and trailing edges also simplified the high-lift device design, and allowed the inclusion of leading edge devices, which became essential in meeting the turn rate requirement.
Figure 5.1: Gavial Interceptor Configuration
The Gavial begins with an Ogive nose. Ogival shapes have long been accepted as the optimum design for supersonic aircraft. The fuselage continues cylindrically at a slightly increasing diameter until just aft of the canards and canopy. At this point, the bottom and sides begin to square off as a lead in to the inlet mouth.
18 The upper surface widens and blends into the leading edge of the wing, with a flat undersurface that leads into the inlet. This flat surface houses the shock ramps to decelerate inlet airflow.
The air intake is a twin inlet system mounted on either side of the main fuselage. The inlets are square, with the lower surface flat across the whole fuselage, and the upper surface buried in the wing. The underside of the airplane tapers back into a cylindrical cross section again around the engine and into the nozzle. The inlets curve aft and up, joining together 30˜ inches in front of the compressor face of the engine.
The upper surface of the Gavial blends into the upper surface of the wing. There is a single central ridge that runs the length of the aircraft to carry power and data lines. The vertical tail grows out of this ridge.
There is a conformal fuel tank that begins just forward of the leading edge of the vertical tail and wraps the aft of the aircraft until just forward of the nozzle.
The vertical tail is a simple trapezoidal shape, with a rudder cut out of almost the whole height. There is a small atmospheric sensor package mounted on top of the vertical tail. While high-α flight could interfere with measurments, the Gavial should not spend any significant flight time in high-α flight regimes, both due to its mission profiles and α limiters in the control system. The canards are similarly simple trapezoids. The canards are all flying surfaces, mounted high on the fuselage to keep the wake out of the engine inlets.
19 Chapter 6
Operational Concept
6.1 Manned Vs Unmanned
Once the basic configuration was chosen, the next step in the design process was determination of the operational concept. Both manned and unmanned aircraft were deemed acceptable options, so this was the
first evaluation step. The team’s first instinct was to choose an unmanned aircraft. There were several expressed reasons for this, primarily because UCAV development seems to be gaining popularity in the military. The recent J-UCAS program, as well as arming the Predator, seem to indicate a high degree of confidence in unmanned air combat. Furthermore, the possibility of removing the pilot from harm’s way during dangerous combat missions was seen as nearly sufficient justification without any other reasons.
Table 6.1: Pro/Con Comparison for Unmanned Concept Pros Cons Takes pilot out of potentially dangerous situation Communications delay time Eliminating cockpit streamlines Inclement weather could attenuate design and reduces drag or distort communications Eliminating cockpit creates more UCAVs are not currently a proven internal volume for fuel and UAV systems technology in this mode of combat Weights and internal volumes of manned systems are eliminated G-forces are limited by structures rather than human factor
In order to ensure that an unmanned concept was actually acceptable from a design stand point, table
6.1 was compiled to weigh the option in as objective a manner as possible. The group decided, in this case, that the pros of the unmanned approach outweigh the communications drawbacks and inherent risk in using cutting edge concepts.
20 Figure 6.1: Operation Control Concept
6.2 Threat Identification, Threat Tracking, Target Verification, Kill Verifica- tion, Collision Avoidance
The avionics package included in Gavial is capable of managing all threat identification, tracking, target and kill verification and collision avoidance. The use of an intuitive cockpit-like Ground Control Station allows for a seamless interface between the ground crew and the aircraft. The combination of radar, cameras, and targeting systems on board the aircraft allow for a high level of pilot situational awareness. All of the data and imagery acquired by the avionics package is transmitted to the pilot and be interpreted naturally as if the pilot were truly flying the aircraft.
The ground control station, (described in detail in the next chapter), was designed to be highly modular.
The Gavial will be controllable from any location capable of providing power and some form of high band- width transmission. Initially, the control station would be located in the rear airbase where the interceptors launch from, and controlled via a satellite up link when out of direct line of sight. With some further refine- ment, the command unit could be placed into an AWACS or other forward air command aircraft, where the interceptors could be controlled using direct line of sight communication, eliminating control lag between
21 pilot input and vehicle response.
The final stage of design evolution for the command chair would be integration into the gunner or trainer seat of current or next generation two seat aircraft. This would allow the second seat pilot to control his wingman directly, both eliminating the control lag from a ground station control, and providing expendable
first strike and covering capability. While this would be highly disorienting for the remote pilot, since he would be experiencing the g-forces of his own plane and controlling the response of the remote, this is a problem that could be overcome with training.
6.3 Weapons Integration
All weapons related operations will require human interaction and authorization. This will greatly reduce the risk of friendly fire and accidental weapons discharge. Defensive actions such as the deployment of countermeasures and evasive actions will be handled autonomously. This will allow for greater survivability in the case of a temporary communications loss during combat.
6.4 Loss of Vehicle Command and/or Control
In the case of remote control loss during normal flight, aircraft goes fully automatic, obtains straight and level
flight, and remains on current course avoiding collisions. Whenever control is lost, the flight computer will also conduct a full self diagnosis to determine the state of the aircraft. If it is deemed that the communication loss is unrecoverable, the aircraft will execute an autonomous return to base. In the case of remote control loss during combat maneuvers, the aircraft will go into a self-defense mode where it is capable of evasive maneuvers and deployment of countermeasures but cannot discharge any weapons.
6.5 Handling of Enemy Countermeasures
Enemy countermeasures will be handled like any other fighter aircraft. The pilot can assess that countermea- sures have been deployed and react accordingly. In the event that electronic countermeasures, or physical systems like chaff and flares are used, the pilot can use the CCD/IR/FLIR suite to manually track the target.
22 6.6 Aircraft Launch and Recovery Scheme
Launch and recovery can be conducted in either manual or fully autonomous mode. The mode of takeoff and landing can be chosen by the pilot depending on the situation and environment. The takeoff and landing curves in the sizing plot were not size drivers. This indicates that the Gavial can operate from very short runways. An automated system can be implemented using the range finding gear to determine safe runways to land and launch from, including unprepared or ersatz fields like highways.
6.7 Items/Levels of Redundancy Required
As with any unmanned aircraft, a high level of redundancy will be utilized in Gavial. One of more of the following modes of communication will be available at all times: SATCOM, RF, and/or LOS (Line-of-Sight).
Additional reliability for communications can be provided by utilizing AWACS for reconnaissance and relay of communications.
6.8 Benefits/Limitations
There are several benefits to an unmanned approach to the outlined missions. The first and most obvious is that the pilot is out of immediate danger. Placing the pilot in a ground control station rather than a traditional cockpit eliminates the loss of personnel associated with the loss of aircraft. Another benefit of unmanned combat aircraft is the ability to use such aircraft as first-strike vehicles. A UCAV could take the lead when entering a combat situation and be flanked by a squadron of traditional fighter aircraft. By taking the brunt of the initial attack, the UCAV provides additional protection for the other aircraft. When considering an air-patrol mission with a four hour loiter time, pilot fatigue becomes an important concern.
Utilizing an unmanned aircraft allows the pilot to place the aircraft under autonomous control during such long loiter periods allowing him to stay sharp in case a critical situation arises.
Possible limitations of an unmanned approach include limitations of bandwidth and communication loss as well as the consideration that unmanned combat vehicles are still a relatively unproven concept.
23 Chapter 7
Ground Control Station
7.1 Design Objectives
The pilot-aircraft interface, or the ground station, is one of the most critical components of a successful unmanned aircraft. At the start of the Gavial project, two major objectives were proposed for the ground station and these remained essentially unchanged throughout the design process. The first of these objectives was that the ground station should be functionally identical to a conventional cockpit. This likeness would make the transition from flying conventional fighters to flying the Gavial smooth for any skilled pilot. The second objective was that the ground station should be re-locatable. In theory, the Gavial’s ground station could be effectively deployed in military bases, transport aircraft, or even in the gunner (trainer) seat of an existing, manned aircraft. This last location, thought it would require additional training by the pilot, would allow for the Gavial to make the first strike in air-to-air combat while both the pilot and wingman safely together in the trailing aircraft.
7.2 The Command Chair
With these objectives in mind, it was decided that the best way to functionally replicate a conventional cockpit would be by modeling the command chair after an existing ejection seat. The Martin Baker MK-
16A ejection seat, shown in figure 7.1, was selected to serve as the model because it is considered to be the world’s most advanced ejection seat presently in production [15].
The aluminum frame for the command chair, shown in figure 7.2, is designed to look and feel like the
MK-16A ejection seat while also providing mounting locations for the various components of the ground station. The size and location of the components of the ground station are based on careful measurements taken at Virginia Tech’s on-campus flight simulator. Aluminum is used for the construction of the frame because it is both strong and lightweight, which contributes to the mobility of the ground station. Contoured seat panels, modeled after the panels on the MK-16A, are designed to fit into the aluminum frame. The hands-on throttle and stick controls and the rudder peddles have computer-controlled, digital force feedback
24 Figure 7.1: Martin Baker MK-16A Ejection Seat [2] which completes the look and feel of a conventional cockpit. The computer, shown in figure 7.3, provides the processing power for the entire ground station and is located under the seat of the command chair.
7.3 Helmet-mounted Heads Up Display (HUD) Screen
The heads up display (HUD) screen for the Gavial Interceptor, shown in figure 7.4, is a helmet-mounted, electronically-dimmed, electrochromic plastic visor designed to function much like the electronically dimmed sunroofs and mirrors in automobiles [16]. The HUD visor is a self-contained unit that can be mounted onto any existing helmet by means of flexible rubber mounting surface that contours to the shape of the helmet.
Additionally, the HUD visor communicates wirelessly to the ground station’s computer giving the pilot the maximum range of motion. When activated, a voltage is passed through the electrochromic plastic visor causing it to dim to a black-colored, opaque state. This allows the display, projected from the projectors located just behind the visor, to bounce off the opaque plastic and then relay to the pilot’s eyes. When de-activated, the visor returns to a transparent state allowing the pilot to utilize the auxiliary controls of the seat.
The display projected onto the HUD visor is a ”real-time flight view” created through the computer compilation of the images from the Gavial Interceptor’s onboard CCD and infrared cameras. Other vital information such as weapons targeting, air speed, altitude, attitude and a simplified radar display are overlaid
25 Figure 7.2: Aluminum Frame with Contoured Seat Panels, Controls and Computer
Figure 7.3: Computer
26 Figure 7.4: Helmet-Mounted HUD Visor over the image. Additionally, to keep the pilot informed about the status of the aircraft during manual flight, the onboard computer is pre-programmed with a damage assessment algorithm that displays a warning on the HUD to alert the pilot of any sustained damage. Solid-state silicon gyroscopic sensors, like those used in the Segway Human Transporter, are mounted within the helmet sense the motion of the pilot’s head [15].
The head-orientation information is relayed through the ground station’s computer, to the Gavial’s on-board cameras which adjust their line of sight to correspond to the direction in which the pilot is looking. This constant adjustment of the line of sight of the cameras, coupled with the full-face, wrap-around design of the visor, completes a sense of realism for the pilot that is unmatched in existing unmanned aircraft.
7.4 Vehicle Management System (VMS)
The main operational center of the command chair consists of the Vehicle Management System (VMS) and the Auxiliary Screen. Coupled with the high-performance processing power provided by the onboard computer, these two components allow the pilot to quickly and efficiently obtain the overall status of the aircraft at any time during a flight (in real-time). The VMS, shown in figure 7.5, is an LCD, touch-screen display that utilizes a tabbed menu format to allow the pilot to select which function to display. Adjustable in four directions, the location of the VMS is fully customizable for each pilot as pre-set locations can be programmed into the on-board computer. The movement of the VMS is controlled by a system of electronic motors contained within the VMS housing which is mounted to the command chair. The concept behind the adjustment of the VMS is that it functions like the adjustment of the driver’s seat in an automobile with each direction controlled separately. When activated, the VMS repositions itself from the storage location
(inside the housing) to either a preset location, or to the last location that was used. When de-activated,
27 the VMS returns into the housing in order to minimize interference with the manual control of the aircraft and to avoid damage to the screen during the transport the ground station. The Auxiliary Screen, shown in figure 7.6, is another LCD touch-screen display, but this one functions as the activation control for the autopilot as well as the activation and positioning controls for the VMS.
Figure 7.5: VMS System and Housing. (Red arrows highlight directions of motion)
Figure 7.6: Auxiliary Screen
The primary function of the VMS is autopilot control, specifically ”point-and-click” navigation. Utilizing global positioning satellite (GPS) imagery, a map of the region through which the Gavial is flying is displayed on-screen. Once the autopilot is activated through the Auxiliary Screen, the aircraft flies straight and level until the pilot changes the settings using the VMS. The degree to which the autopilot is autonomous is fully customizable from simple navigational assistance to full ”point-and-click” navigation where the pilot selects a destination and then allows the onboard computer to optimize altitude, airspeed and path to reach
28 the destination in the most efficient manner possible. In addition to a GPS map of the region, the VMS can also display a detailed radar display to assist in enemy identification and weapons targeting. Aside from control of the aircraft and radar, the VMS also functions as the pilot’s means of obtaining the overall status of the aircraft. The VMS displays key quantities such as air speed, remaining fuel, weapons status, remaining ammunition, flap positions, landing gear position and it can also assist in damage evaluation. As mentioned before, the onboard computer is programmed with a damage assessment algorithm that senses damage to the aircraft and then alerts the pilot through a warning on the HUD. To supplement this warning, the pilot is able to access a more detailed damage report through the VMS. The logic behind the use of the
VMS as opposed to conventional displays lies in the compactness and adjustability of the unit. Because the entire VMS unit is able to retract inside the housing located under the right side panel, the LCD screen is both protected from damage and out of the way during travel. This furthers the goal of maximizing the mobility of the ground station. The adjustability of the VMS unit allows the pilot to customize the location of the screen to meet his or her personal preferences in order to further facilitate the transition from flying conventional fighters.
7.5 Final Layout and Platform Assembly
After the individual components of the ground station were designed, considerable thought went into the organization and layout of these components in the final ground station assembly. The final layout, shown in figure 7.7, is designed to minimize interference with the manual controls and to maximize the mobility of the ground station. In order to minimize the width of the ground station, the VMS housing is positioned under the right side panel and is constrained to rotate no wider than the width of the command chair. The onboard computer is positioned under the seat in order to minimize the depth of the ground station and also to allow the frame of the chair to provide a buffer against possible damage during the transport of the ground station. By constraining the overall dimensions to match those of a conventional ejection seat, the
Gavial ground station is able to fit anywhere a conventional ejection seat can. This design allows transition from the standard ground stations to combat airfram integrated command station. Finally, the entire ground station is mounted onto a steel base plate in order to facilitate transport and deployment as this plate can be outfitted with mounting points for almost any application.
29 Figure 7.7: The Final Ground Station Layout (3 views)
30 Chapter 8
Aerodynamics
Aerodynamic analysis and decisions were made to best fulfill the RFP requirements while maximizing per- formance and minimizing manufacturing and operational cost. Particular focus was placed on the four hour loiter capability of the aircraft where the greatest fuel consumption would occur; thus fuel storage and decreased drag were paramount in wing design. Efforts were directed to establish a delicate balance between supersonic performance and internal wing fuel storage. Other considerations included subsonic wing performance and maneuverability.
8.1 Wing Planform
Performance analysis combined with research of existing fighters indicated an appropriate wing loading equal to 70 ft2. Since the build up of wave drag during supersonic flight was of main concern even during preliminary sizing analysis, Raymer´s[12] method for wave drag determination directed that the total span and wing area be maximized in an effort to keep drag coefficients low. As noted previously, research of supersonic aircraft indicated the established success of high leading edge sweep, most commonly in the delta configuration for supersonic transport applications. The need for fighter maneuverability led to the decision to use a modified delta wing shape, meeting the needs for high area, span and leading edge sweep while affording the opportunity for weapons and fuel storage, both internal and external. The final planform and geometry are presented in Figure 8.1.
8.2 Airfoil Selection
Airfoil selection was made with primary focus on both the turn requirement set forth in the RFP and wing thickness. A variety of airfoils from the Abbott [3] text were evaluated based on both considerations. The
Anderson[17] text provided a method of determining the necessary maximum lift coefficient for a turn of specified speed, turn rate, altitude, wing loading and structural capabilities. The wing loading of 70 ft2 and a maximum structural load limit of +9 g´swere used to determine a maximum lift coefficient of 2.1 to sustain
31 Figure 8.1: Planform Characteristics (dimensions in inches)
32 the 18 ◦/s turn at 35,000 ft, comparable to the maneuverability requirement for modern fighters. Airfoils with high lifting capabilities were considered to meet this requirement; high-lift devices were an additional consideration to maximize lift performance.
The Abbott [3] text was referenced for all airfoil data sets simulating the addition of high lift devices; the lift polars for the six airfoils under preliminary consideration are presented in Figure 8.2.
Figure 8.2: Planform Characteristics [3]
The Figure shows the NACA 631-412 airfoil displaying the highest lift coefficient of 2.06, still underper- forming according to the necessary value of maximum lift through the specified turn. The NACA 641-406 reaches a section lift coefficient of 1.93; all airfoils presented in the Figure will require the implementation of high-lift devices to perform the RFP specified turn requirement. Wing thickness is a consideration when regarding supersonic drag build-up; thickness-to-chord ratio (t/c) was also evaluated for wing fuel storage purposes. While the t/c of 12% would provide the best opportunity for internal fuel storage, the section drag coefficients presented in Figure 8.3 indicate a higher drag build-up at the highest available Reynolds number (Re) of 9 million [3].
This large drag adversely affects high speed travel in the form of wave drag; thus, the NACA 641-406 is chosen because it best provides both high lift and the lowest value of drag. Figure 8.4 displays the chosen airfoil as graphed by the Abbott text [3].
With the decision to equip the airframe with such a thin airfoil, the need for external fuel tanks was considered specifically for the defensive counter-air patrol mission, which requires a four hour loiter period at 35,000ft.
33 Figure 8.3: Airfoil Section Drag Polars [3]
Figure 8.4: NACA 641-406 Airfoil[3]
34 8.3 High-Lift Devices
A plain leading-edge flap and slotted trailing edge flapevator were implemented because they are the simplest and most cost-effective method of meeting the lift need during the turn requirement, while simultaneously providing use as a control surface. Choosing a simple configuration reduces manufacturing costs that would be required of a more complex lifting system. Figure 8.5 shows the flap configuration as well as the deflection angles for maximum lift.
Figure 8.5: High Lift Configuration
Figure 8.6 shows the lift coefficient gain on a 6% thick airfoil with the implementation of the flap system in Figure 8.5. The Figure indicates the maximum lift coefficient needed for the turn can be reached with the use of the high lift devices deployed at the specified deflection angles for Mach numbers corresponding to various mission elements.
8.4 Drag
Drag analysis was conducted at specific flight configurations congruent with the RFP mission elements, with and without the addition of the three 660 gallon external fuel tanks.
Various drag analysis codes were used to determine the drag values for the Gavial´s dimensional charac- teristics at each Mach number and altitude. The Ives program [13] determined a planform lift distribution of
0.93 which provided induced drag values at each Mach number. The Butler MATLAB code [13] investigated the form and form friction drag based on aircraft dimensions and Mach number. The Raymer text [12] provided a method of determining the wave drag contribution.
In order to reduce wave drag, there was significant effort invested in attempting to mimic the Sears-
Haack [13] body distribution. Due to the configuration requirements and internal fuel demands, this was not entirely possible. As in many aircraft, the CG is a popular location for systems, especially expendable
35 Figure 8.6: Increased Lift Due to Flap System stores like missiles and fuel. Unfortunately, in the Gavial, the CG was also the point of maximum cross section of the wing. The choice of body design did not leave much room for bottlenecking. The fuel layout was stretched longitudinally as much as possible to avoid unneccessary bulges.
A conformal tank was added to the aft, over the engine, to smooth out the area distribution at the end of the wing and begininning of the vertical tail. This area had a fairly jagged distribution, before the inclusion of the tank. The placement of the ATFLIR pod was also selected to ease the transition between the canopy/canard area and the beginning of the wings.
The intercept configuration is fairly clean, and maintains an acceptable maximum cross section, though it is 2-3 ft2 higher than originally hoped. The addition of the fuel tanks adds almost another 12 ft2 to the cross section, and are soley responsible for the massive drag rise, and subsequent top speed detriment in the
DCAP mission.
The choice was made to attempt a high degree of wing/body blending after the style of the F-22. This resulted in thick wing roots, which blend smoothly into the upper fuselage. The lower fuselage is defined by the duct inlet shape. Similar to the F-14 and F-15, a mostly square cross section was used for the majority of the body, blending into a cylindrical foreboy, and terminating in an Ogive nose.
36 Table 8.1: Drag Characteristics (Planform Reference area: 566ft2) Mach Number Altitude[ft] CDfriction CDinduced CDwave CDform CDtotal Interceptor Configuration 0.9 0 0.00523 0.003864 0.01622 0.00107 0.026384 1.2 35,000 0.00559 0.002766 0.02535 0.00115 0.034856 2.2 35,000 0.00422 0.000245 0.02535 0.00087 0.030685 DCAP Configuration 0.9 0 0.00523 0.003864 0.03423 0.002034 0.045358 1.2 35,000 0.00559 0.002766 0.05439 0.001782 0.064528 2.2 35,000 0.00422 0.000245 0.05439 0.001497 0.060352
Figures 8.9 and 8.10 display the contribution by component to the overall form and friction drag elements and the total drag coefficient contribution at two Mach numbers and elevations which are representative of both subsonic and supersonic flight conditions.
Figure 8.11 displays the total drag build up at sea level and 35,000ft over a range of subsonic and supersonic Mach values.
Table 8.1 displays the particular drag values at various flight conditions which correspond to the RFP specified mission elements.
37 Figure 8.7: Intercept Configuration
38 Figure 8.8: Defensive Counter Air Patrol Configuration
39 Figure 8.9: Mach 1.2 Drag Breakdown
Figure 8.10: Mach 0.9 Drag Breakdown
40 Figure 8.11: Total Vehicle Drag Variance with Mach Number
41 Chapter 9
Stability and Balancing Scheme
The RFP requires a subsonic longitudinal static margin in the interval of -10% to 10%. The stability scheme for the Gavial was determined with information from [18] and by investigating the stability of current
figher/interceptor aircraft. Moderate to high performance aircraft typically have inherent static stability paired with dynamic instability, which requires a closed-loop feedback control system, or may be both statically and dynamically unstable. A major issue associated with balancing a high performance aircraft stably is the issues associated shift of the aerodynamic center at high speeds. As the aircraft speed increases, the aerodynamic center shifts aft of its initial position increasing the static margin. At these high speeds with large positive static margin, the aircraft may be too stable to perform maneuvers without the use of unnecessarily large control surfaces. This also reduces the drag associated with control deflections which is particularly important during the supersonic regime. Balancing unstably could still yield static stability at high speeds, which aids in maintaining control of the aircraft without being overwhelmingly stable.
9.1 Longitudinal Control
Taking these behaviors into consideration, the Gavial was balanced with static instability. The static margin was evaluated using JKayVLM, a vortex-lattice code. The longitudinal input planform is shown in Figure
9.1. The most unstable configuration is when the CG is at its aftmost limit, 52%˜ MAC. In this most extreme case, the subsonic longitudinal static margin of the Gavial is -9.8%. A digital feedback control system is used by the Gavial due both to its unstable nature and unmanned operation.
9.2 Control Surfaces
The Gavial uses a canard, single vertical tail, ailerons and flapevators for additional pitch control. Informa- tion on existing military fighter aircraft was used in shaping the canard and vertical tail, as well as in sizing the vertical tail. [19] The vertical tail is 55 ft2 with a rudder of 13 ft2.
Ailerons, providing the primary source of roll control, are located outward of the flapevators and extend
42 Figure 9.1: Logintudinal Control Planform to the wing tips. This placement was chosen to maximize the roll moment for a given aileron size. The aileron has a span of 50 inches with a constant chord of 16 inches. The lateral-directional input planform for JKayVLM is shown in Figure 9.2, this was used to size the vertical tail.
Figure 9.2: Lateral Directional Control Planform
Flapevators were chosen over separate flap and elevator devices to simplify design and manufacturing and to keep costs down. The flapevators were sized for aerodynamic high-lift characteristics. The flapevators may also be differentially deflected to provide additional roll support under conditions where the aileron alone is insufficient.
9.3 Canard Sizing
The canard was sized using the longitudinal constraint plot (X-plot) shown in Figure 9.3. The vertical axis represents the total planform area of the canard, as opposed to the area of one canard surface. The -10%
43 Static Margin line was determined using JKayVLM with the location of the canard essentially fixed from previous balancing.
Figure 9.3: X-plot for Canard Sizing
The hashed region above the -10% Static Margin represents the infeasible region with less than -10% stability. The vertical lines represent the fore and aft limits on the center of gravity based on loading configurations. The canard size was governed by the static margin limits and the aft limit on the center of gravity. From Figure 9.3, the canard size is 40 ft2. This size does not violate the stability constraints during any flight conditions or loading conditions.
9.4 Control Authority
The stability and control derivatives of the Gavial were estimated using JKayVLM at subsonic speeds and the United States Air Force’s DigitalDATCOM for supersonic speeds. A summary of the stability and control derivatives is shown in tables 9.1 and 9.2.
The control authority was assessed using a spreadsheet created by Jacob Kay [13]. The stability and control derivatives, flight conditions and various aircraft geometry parameters were used as input for the spreadsheet. The control authority for basic maneuvers and trim flight is shown in tables 9.3, 9.4, and 9.5.
In all of the conditions, the required control surface deflections are feasible. The control surfaces were deemed to have sufficient control authority for basic maneuvering and trim flight. The roll performance is governed by MIL STD 1797 [20] in terms of the time needed to achieve a certain bank angle change at
44 Table 9.1: Logitudinal Derivatives Mach 0.2 0.6 1.4 CL0 0.00 0.00 0.00 Cm0 0.00 0.00 0.00 CLα 1.457 1.637 3.515 Cmα 0.119 0.0179 -1.220 Cmq -2.216 -1.012 -2.829 CLq -0.261 -0.319 3.030 CLδf 0.873 0.976 0.322 Cmδf -0.349 -0.397 -0.199 CLδc 0.070 0.085 0.077 Cmδc 0.217 0.223 0.201 SM(-Cm/CL) -0.082 -0.048 0.027
Table 9.2: Lateral-Directional Derivatives Mach 0.2 0.6 Cyβ -0.188 -0.194 Cnβ 0.042 0.044 Clβ -0.023 -0.024 Clp -0.205 -0.192 Cnp -0.069 -0.068 Clr 0.017 0.018 Cnr -0.053 -0.055 Cnδa -0.004 -0.004 Clδa 0.023 0.23 Cyδr 0.119 0.127 Clδr 0.014 0.015 Cnδr -0.044 -0.048
45 Table 9.3: 1-g Trim Assessment Altitude (ft.) Mach α (deg.) Deflection (deg.) Flapevator Only 0 0.6 4.67 0.92 Flapevator Only 35000 0.6 15.03 2.97 Flapevator Only 35000 1.4 1.89 0.13
Table 9.4: Maneuver (Pull-up) Assessment Altitude (ft.) Mach g’s α (deg.) Deflection (deg.) Flapevator Only 0 0.6 2 11.85 -2.24 Flapevator Only 35000 0.6 2 38.26 -7.47 Flapevator Only 35000 1.4 2 3.94 -2.02 Flapevator Only 35000 1.4 7 13.72 -7.38
various speeds. A summer of the Gavial’s roll performance is shown in table 9.6. The Gavial has excellent roll authority for Level 1 flying qualities at high speeds. At lower speeds, the roll authority suffers slightly and the Gavial achieves Level 2 flying qualities which are still adequate by military standards. It is for these strict roll requirements that the flapevators must be used to supplement the ailerons. The decision to use ailerons as opposed to using only the flapevators for roll control is due to the sufficient authority provided by the ailerons in basic maneuvering. By using the ailerons for common flight conditions and using the
flapevators for roll control only in the extreme cases, the large drag associated with the deflection of the
flapevators is kept to a minimum.
46 Table 9.5: Steady Sideslip β (deg.) Altitude (ft.) Mach δ Aileron (deg.) δ Rudder (deg.) Bank Angle (deg.) 18 0 0.2 7.81 16.45 1.06 18 0 0.6 8.86 15.82 10 18 35000 0.6 18.86 15.82 2.36 30 0 0.2 13.01 27.42 1.77 30 0 0.6 14.76 26.37 16.67 30 35000 0.6 14.76 26.37 3.93
Table 9.6: Time-to-Bank Summary. Military Specifications and Gavial Performance. Time in seconds. High (1600 fps) 90◦ 180◦ 360◦ Level 1 1.4 2.3 4.1 Level 2 1.7 2.6 4.4 Gavial 1.04 1.63 2.68 Medium (1300 fps) 90◦ 180◦ 360◦ Level 1 1 1.6 2.8 Level 2 1.3 2 3.4 Gavial 1.19 1.88 3.16 Low (600 fps) 30◦ - - Level 1 n/a - - Level 2 1.3 - - Gavial 1.28 - -
47 Chapter 10
Vehicle Performance
The two most stringent requirements were the required maximum Mach number of 2.2, and the instantaneous turn requirement of 18 deg/sec. The engine size was driven almost solely by the maximum Mach number requirement, making meeting the specific excess power requirements easy in comparison to the thrust needed to reach the dash Mach number. In the case of the maximum turn rate requirement, CLmax was the limiting factor. The optimum speed to meet the turn requirement was determined to be 0.9M at 35,000 ft. The conceptual sizing plot and load limit of 7g’s set the rest of the parameters for turn rate, calculated by:
"r #"r #! dψ ρSC n2−1 = g L (10.1) dt 2W n
To get the required 18 deg/sec, this size indicated a CLmax requirement of almost 3, which was far outside the feasible range. To bring CLmax down to a feasible range, the g-limits were increased, and the wing loading was reduced. A new sizing plot was created using the process described in chapter 3.
This new plot, figure 10.1, indicates a new required T/W of approximately 1.2, and a reduced wing loading of 70 lb/ft2.
All specified performance parameters were important in the design of the aircraft. The aircraft was required to have a top speed of Mach 2.2, an instantaneous turn rate of 18 deg/sec, and certain specific excess power requirements. The engine was sized for the speed requirement and hence the tremendous amount of excess power. Driving the instantaneous turn rate was the CL max. Using this we were able to reach a maximum instantaneous turn rate at a speed of Mach 0.9 and an altitude of 35000 ft of 21.2 deg/sec with half the internal fuel and half of the weapons deployed.
The maximum speed requirement of Mach 2.2 coupled with the 4 hour loiter specification required the consumption of massive amounts of fuel forcing the use of removable external fuel tanks. The CAP mission requires the use of the fuel tanks to sustain the 4 hour loiter without the use of in flight refueling. The tanks however increase the total drag on the aircraft and reduce the top speed characteristics of the aircraft.
Deciding that the loiter period of the aircraft for this particular mission was more important than the speed,
48 Figure 10.1: Sizing Chart for Conceptual Design
49 accordingly the aircraft for the CAP mission still reaches a top speed of Mach 1.7, the tanks were placed on the aircraft to provide for the required loiter. The design of the tanks allowed for them to be removable and for the remaining missions can be stored in the hanger since all required fuel can be stored on the aircraft and no penalty to the performance of the aircraft arises. All other performance requirements were met and exceeded by the aircraft for all missions.
The aircraft is required to operate from all existing NATO airfields and has a take off and landing parameter of 8000 feet. Using equations 10.2 through 10.5 the balanced field length can be determined.
The shortest takeoff length for the aircraft with all high lift devices deployed was found to be 907 feet, the landing distance was found to be 903 feet.
T A = g 0 − µ (10.2) W
g 1 B = ρS (C − C µ + a) (10.3) W 2 D Lg
T0 ! W w − µ STO = ln (10.4) g (ρS (C − µC ) + a) T0 1 2 D Lg w − µ − 2W (ρS (CD − µCLg) + a) V
1 B S = ln 1 − V 2 (10.5) Landing 2B A c
Although the top speed requirement is not met for the CAP mission it was determined that the loiter requirement was more important and the top speed was then sacrificed (see figure 1.1). The excess power is far exceeded for all missions due to over sizing the engine for the CAP mission 10.4. The maximum sustained load factor was required to be at least 5 g’s. The use of an UCAV allowed the maximum sustained g-loading to be increased to facilitate the high speed turning requirements. The Gavial was designed to withstand a sustained loading of 9 g’s. This aided in both the wing loading and CL requirements. The instantaneous turn rate was found to be 21.2 deg/s since the aircraft was over sized for the loiter, however the maneuver weight of the aircraft is 64% of the original weight. This allows for better maneuverability at the specified altitudes.
To verify the accuracy of the aircrafts size a program entitled ACS[21] and supplied by NASA Langley and AVID, LLC was used. The initial aircraft sizing was completed by hand and then verified using the ACS program. All initial calculations were valildated. The ACS program was also able to show certain design
50 criterion in a comparative graphical form. The instantaneous turn rate, the excess power, and the maximum sustainable g loading could all be calculated by the ACS program. It should be noted that the aircraft does conform to the specified requirements set forth by the RFP[1]. This can be seen in the figures generated using the data drawn from ACS (Figures 10.2 - 10.4).
Figure 10.2 illustrates the maximum instantaneous turn rate attainable at various altitudes and Mach numbers. The plot showa that the Gavial interceptor attains the required maximum instantaneous turn rate at 35,000 ft.
Figure 10.2: Turn Rate vs Mach Number from ACS
Figure 10.3 depicts the maximum sustained load factor the Gavial can produce in maneuver. The chart is capped at 9 g’s because the airplane will be limited by the control systems to not exceed maximum safe loading.
The excess power curves shown in figure 10.4 illustrate the capability of the Gavial´s propulsion system.
The excess power requirements delineated in the RFP [1] are easily exceeded as shown.
51 Figure 10.3: Maximum Sustained Load Factor
Figure 10.4: Excess Power Available
52 Chapter 11
Propulsion Design
The first thing done with the propulsion section was to look at the missions and analyze them and how each aspect of the mission would affect the propulsion. The driving factors for engine selection were the specific fuel consumption (SFC) and the thrust. The engine deck data that was given was looked at and at first glance, the group decided that the SFC was too high to be an effective engine for the missions that airplane must complete. Other engines were researched to see if they had better SFCs and thrusts.
After looking at all the data gathered from researching other engines, the engine deck that was supplied by AIAA was decided on by the group to be used because of a lack of available engine decks for the other engines. The engine deck that was supplied had a thrust of 22,500 lbs, a diameter of 35 inches, a length of
148 inches, and a dry weight of 2,651 lbs. From the research, it was decided that this was a variant of the
F-414 engine used on the F-18 Super Hornet.
Figure 11.1: GE F-414 Engine
Table 11.1: Engine Selection Engine Diameter [in] Weight [lbs] Thrust [lb] Cruise SFC Max SFC F-110 46.5 4,050 32,000 0.64 2.09 F-404 34.8 2,282 17,700 0.84 1.74 F-414 35 – 22,000 0.8 1.7 F-119 – – 35,000 – – AL-31 50.28 – 27,557 – – AL-37 36.7 3,660 31,967 0.677 – NK-22 59 7253 44,090 – 1.95
53 11.1 Single vs Multi Engine
Since the plane needs to be as light and cheap as possible, a single engine was chosen instead of dual engines.
The single engine needs to have enough thrust to overcome the drage at Mach 2.2 and have a good SFC.
The single engine concept provides more internal volume for fuel or payload. Also, having one engine will reduce the time and cost of maintenance.
11.2 Engine Scaling
The group then decided that it was best to try to scale the engine deck to whichever size thrust we would need to satisfy the missions. We estimated the initial weight of the plane to be around 40,000 lb and the
AIAA engine deck could not provide the thrust required. The desired thrust for the mission was to initially be 35,000 lbs. Using the following equations from Raymers book:
T SF = req (11.1) Tactual
L = 0.4LactualSF (11.2)
D = 0.5DactualSF (11.3)
W = 1.1WactualSF (11.4)
The scale factor used was 1.57. Then the net thrust was multiplied by the scale factor. The SFC was calculated for the original engine deck by the equation:
F uelF low SFC = (11.5) NetT hrust
The SFC was decreased by 5% for the scaled engine deck because of technology improvements. Then the fuel flow for the scaled engine was calculated by the scaled net thrust multiplied by the decreased SFC. The values for the new engine is that of a scaled up F-414 and has a max thrust of 35,000 lbs, a length of 178 inches, a max diameter of 44 inches, and a weight of 4,364 lbs. Graphs were made of the SFC vs. net thrust
54 for varying altitudes at Mach numbers of 0.9 and 2.2. Also, the net thrust and SFC were graphed against the Mach number for varying altitudes.
Figure 11.2: Thrust vs SFC for Cruise Mach Number of 0.9
On the underside of the engine is the Full Authority Digital Engine Control (FADEC). It digitally calculates and precisely controls the fuel flow rate to the engines giving precise thrust. In addition to the fuel metering function, the FADEC performs numerous other control and monitoring functions such as
Variable Stator Vanes (VSV’s) and Variable Bleed Valves (VBV’s) control, cabin bleeds and power off-takes control, control of starting and re-starting, turbine blade and vane cooling and blade tip clearance control, thrust reversers control, engine health monitoring, oil debris monitoring and vibration monitoring.[22]
11.3 Inlet Design
The inlets are an integral part of the propulsion system; they must be the right size so that the engine gets enough air flow. Also, the inlet must have a ramp as to slow down the supersonic flow by using oblique shocks instead of normal shocks to keep the pressure recovery to a maximum. Since the front-face diameter is not known, Raymer[12] says that it can be estimated to be 80% of the maximum diameter. Since the max diameter is 44 inches, the front face diameter is 35.2 inches. The mass flow of the engine is then estimated
55 Figure 11.3: Mach Number vs Thrust for Varying Altitudes as 26 times the square of the engine front face diameter in feet. The mass flow is calculated to be 348 lb/sec.
Using the mass flow rate equation:
m˙ = ρAV (11.6)
The capture area required for a design Mach number of 2.2 at 35,000 ft was calculated to be 7.134 ft2.
Since there will be two inlets on the underbelly of the plane, the required area for each inlet is around 3.6 ft2.
11.4 Shock Ramps
Each inlet has a set of shock ramps to decelerate inlet flow before it reaches the compressor fan of the engine.
To minimize pressure losses across the shocks, a series of oblique shocks are used. To accomadate multiple design cruise points, specifically Mach 1.2 and 2.2, the ramps have a variable second stage. The variable ramp can also control the amount of flow during subsonic flight. This arrangement gives and inlet flow speed of Mach 0.674, with a total pressure recovery of 88.37% for an initial Mach number of 2.2. Once the engine is installed into the plane, some the thrust will be lost because of pressure differences and the shape of the
56 Figure 11.4: Propulsion Unit - Inlets, Ramps, Engine and Power Systems inlet. Using the equation from Raymer[12]:
P 1 P 1 %thrustlost = Cram[ − ]x(100) (11.7) P 0 ref P 0 actual
where Cram is the ram recovery correction fact and for subsonic flight, it is estimated to be 1.35. The thrust lost due to installation is 5.4% due to the curvature of the inlet. The final thrust of the installed engine is 33110 lbs.
Figure 11.5: Shock Ramp Schematic for Mach 2.2
57 11.5 Nozzle
The nozzle needed to help the plane achieve good subsonice and supersonic speeds. The nozzle that was chosen was a converging-diverging nozzle. The area of the nozzle during subsonic flight was calculated to be
4.304 ft2. The subsonic part was found by multiplying the capture area by a range of 0.5-0.7. The maximum area of the diverging nozzle was calculated to be 10.044 ft2. The max area was found by multiplying the the capture area by the range 1.3-1.5. To find the length of the nozzle required, an angle of 15 degrees was estimated between the end of the engine and the maximum area. The length of the nozzle was then calculated to be 14.4 inches.
Figure 11.6: Nozzle Geometry, Supersonic (left) and Subsonic (right)
58 Chapter 12
Structural Design
12.1 Materials Selection
Table 12.1 outlines the chosen materials and locations in the aircraft where they are used. Due to the
18/second turn rate, the Gavial must withstand 9g’s sustained load. Therefore, strong materials must be used to allow for these high loads. Secondarily, weight is a driving factor with any aircraft, so Youngs
Modulus to density ratios were taken into consideration, the higher obviously being more desirable. Since the RFP requires a design service life of 12,000 hours, chosen materials must also have high resistance to environmental effects, such as corrosion. All four of these are definitely listed under this category. As a
final consideration, cost was taken into account, although not nearly as much as other aspects of material choice. The graphite epoxy and aluminum are fairly expensive, but high strength, low weight, and good environmental resistance more than compensate for their price.
Table 12.1: Material Selection Material Components Titanium IMI551 Landing Gear Struts Aluminum Carbide Matrix (40% Al) Wing Ribs Wing Spars Titanium Carbide Matrix Fuselage Stringers Fuselage Ribs Fuselage Bulkheads Landing Gear Bays Carbon Reinforced IM6/3501-6 Composite Fuselage Skin Vertical Tail Access Doors Landing Gear Covers Canards Rudder Flaps Ailerons
59 12.2 Fuselage Structure
The fuselage will be constructed of equally spaced bulkheads and ribs for upholding larger systems, fuel tanks, wing load carry-through, and engine weight. These ribs reinforce the shape of the fuselage skin, which in turn bears a large portion of the bending loads on the aircraft. To enhance structural stability, four stringers, spaced every 90 degrees around the fuselage. These stringers support the skin against buckling and transfer strain into the ribs and bulkheads.
Upholding the structure of the Gavial is five ribs, six bulkheads, and three half-bulkheads. The half- bulkheads uphold the engine, and allow for ease of maintenance and removal, a requirement of the RFP. All of the bulkheads and ribs are spaced 40.1 inches apart, allowing for minimum weight while still maintaining structural stability. The six bulkheads are located at heavy systems, such as the FLIR pod and Radome, at landing gear locations, and also at two primary spar locations, due to high lifting loads. The five ribs are located wherever thickened bulkheads were not necessary, primarily at smaller systems and the remaining wing spars. Wherever any system did not coincide with a structural element, stringers were placed to connect that system to the two closest ribs/bulkheads. The shapes of the ribs and bulkheads were adapted to fit the cross-section of the fuselage at each location, ranging from one to two inches thick, and 40 to 100 pounds apiece.
The entire fuselage will be composed of a titanium composite, Titanium Carbide Matrix (80% TiC). The composite is very strong, with a Young´sModulus of 60,000,000 psi and a yield strength of 2,900,755 psi [23].
It has a moderate density of approximately 0.197 pounds per cubic inch [23]. This is much better than steel, but not nearly as light as aluminum composites. Titanium is known to be incredible at withstanding high loads and temperatures, and TiC is no exception. This is the main reason for the composite comprising the half-bulkheads upholding the engine; large amounts heat will be created from the Gavials high-powered engine.
The skin of the aircraft will be 0.3 inches thick everywhere except on the wing. Towards the leading edge and the root of the wing, higher thicknesses are required to withstand the higher loads encountered.
Therefore, 0.4 inches will be the standard thickness at these locations. The aircraft´sskin will be composed of a high-modulus graphite epoxy, IM6/3501-6. The composite has a fiber volume of 63.5%, giving an overall density of 0.0567 pounds per cubic inch [24]. As compared to the titanium, or even aluminum, this composite is extremely light. The canards and vertical tail, since they are not withstanding very large loads, will not have the typical rib and spar structure of a wing. Instead, they will be comprised completely of IM6/3501-6.
60 Figure 12.1: Fuselage Structural Layout
With a Youngs Modulus of 23,300,000 psi [24], the graphite epoxy is strong enough to withstand any and all loads encountered by the Gavial´s fuselage, wings, or control surfaces.
12.3 Wing Structure
A single wing on the Gavial is composed of six spars and five ribs. Three of the spars span completely to the wingtip, while one maintains structural stability towards the leading edge of the wing and the other combats the high loads encountered when flaps are deflected. The spars each connect to a bulkhead or rib on the fuselage. Wherever any two spars intersect, a wing rib is located, due to the possible high loading.
This ensures a large factor of safety in the case of, say, fully deflected flaps, or fully deflected leading-edge slats. The spars are two inches thick, and they span from the top to the bottom section of the wing for best possible support.
The wing ribs on the Gavial are located across the wing at high load areas, parallel to the fuselage. The closest rib is located very near fuselage due to high wing root twisting moments [25]. This largest rib resists these high twists. Another wing rib is placed between the two large flap segments, since their actuating systems need structural attachment points, and also since two different angled deflections (between the two
61 flaps) will cause a twisting moment between them. This wing rib, along with two others, as aforementioned, are also located at points where wing spars intersect/end. The smallest of the wing ribs is located very near the wingtip, and is placed there to ensure that the loads created by the ailerons do not cause the wingtip-aileron connection to fail or allow for high twists.
Figure 12.2: Wing Structural Layout
The entire wing, excepting the skin which is composed of the same graphite epoxy as the complete aircraft skin, of the Gavial is composed of an aluminum composite, fiber reinforced aluminum carbide (40%
Al). Originally, the wing spars were going to be made of the same titanium carbide matrix as in the fuselage, but weight savings were becoming more and more important. The aluminum composite is definitely more expensive, but at a density of approximately 0.081 pounds per cubic inch, is much lighter [23]. The composite is quite strong as well, having a Youngs Modulus of 35,000,000 psi [23]. Since material selection (ignoring price) is decided upon Youngs Modulus to density ratio, this aluminum carbide composite seems an obvious choice for aircraft wing ribs and spars.
62 12.4 Canard and Tail Structure
In order to keep weight and performance detriment to a minimum, the vertical tail and canards were formed of a single piece of high modulus graphite epoxy (HMGE). Numerous inexpensive methods of forming relatively simple structures exist for HMGE, which reduce both cost and structural weight. The actuation systems are buried directly into the HMGE mesh, saving room in the fuselage for more critical systems.
12.5 Landing Gear
The landing gear layout was chosen as the standard tricycle gear used on most current military vessels. Using methods outlined in [12], the landing gear was designed to meet airforce specifications and the rotation angles necessary for takeoff and landing. Table 12.2 lists the layout, and figure 12.3 shows the layout schematically.
Table 12.2: Landing Gear Specifications Tipover Angle 63 deg Tailscrape Angle 10.6 deg Wheel Track 17.0 ft Wheel Base 10.4 ft CG Height Above Deck 5.27-6 ft
Figure 12.3: Landing Gear Schematic View
The forward landing gear on the Gavial is composed of a single oil-type strut, 3.5 in diameter, with a spring constant of 2500 lb/in, retracted and extended by an actuator attached to its side. The actuator is
63 designed so the tire and gear retract forward. This has two major advantages; it allows for more space aft of the gear, where many systems are located, and upon landing approach, air drag helps the gear to extend, using less power from the aircraft [26]. Powering the actuator will be accomplished through a localized hydraulic system, since the rest of the aircraft is electrically actuated. The system will also power the anti- skid carbon brakes used to slow the plane during landing, and allow the tire to rotate slightly for runway turns. The brakes are quite expensive, but 40% lighter than conventional brakes [26]. The tire will be that of the F-16 Fighting Falcon, the Michelin AirX 18 x 6 [27]. This tire has been tested and proven more than satisfactory on aircraft over many years. IMI551 Titanium alloy will compose most of the forward gear, due to its high resistance to environmental effects and relatively low weight[28].
Figure 12.4: Forward Landing Gear
The main landing gear is composed of four oil-type struts, angled out 32.3 degrees from the vertical for stability. The struts are 4” in diameter with spring constants of 6,198 lb/in. for the two forward struts, and 8,140 lb/in. for the two rear struts. The latter value is higher because the rear struts are angled more with the horizontal. In the main gear, the struts will act as both shock absorbers and actuators, as well as side braces and drag struts due to their angles, a revolutionary system which saves weight. The tires will retract forward very slightly, but mostly vertically. This allows for maximum volume around the main gear bay, which is an optimal place to store fuel. As in the nose gear, the struts will retract by a localized hydraulic system, which will also power its anti-skid carbon brakes, allowing for both braking and runway maneuverability [26]. The two tires used on the main gear will be larger than that of the nose, Michelin
AirX 26 x 6.6, [27] since they will be encountering much greater loads. The main gear will also be composed of nearly all IMI551 Titanium alloy like the nose gear.
In calculating the necessary strut diameters and spring constants for the landing gear, the worst possible
64 Figure 12.5: Main Landing Gear case scenario was applied. This occurs when the plane makes a landing and all its weight is channeled to one tire in contact with the runway. Therefore, the Gavial´s landing gear is designed for a high factor of safety, which unfortunately increases weight, but can save the aircraft from a catastrophic systems failure or if high winds/gusts cause difficulty in landing.
12.6 Structural Limits
The structural limits of the Gavial are presented in its V-n diagram, figure 12.6. The aircraft was designed to fly at an inverted g-limit of three, and a regular g-limit of nine. Typical values for regular g-limits for
fighter aircraft range around seven due to pilot limitations. With an unmanned aircraft, like the Gavial, this can be increased, which aids our 18 degree per second turn rate, a major driving force for design. As depicted, inverted and regular stall intersect with these minimum and maximum load limits between 900 and 100 feet per second. The constraint of maximum speed for the aircraft is deigned VD, or dive speed, and is calculated as 1.35 times the maximum straight and level flight speed, a typical factor used in structural design. Dotted lines indicate the gust portion of the V-n diagram. Since the gusts at the three calculated speeds are located within the maneuver diagram, the design V-n diagram is the maneuver V-n diagram.
65 Figure 12.6: V-n Diagram for Maneuver and Gust Loads
66 Chapter 13
Weapons/Fire Control System
13.1 Weapons
There are 3 weapons packages the HDI is required to accommodate.
1. M61-A1 + 500 rds, 2 AIM-120, 2 AIM-9[1]
2. M61-A1 + 500 rds, 4 AIM-120[1]
3. M61-A1 + 500 rds, 4 AIM-9[1]
Figure 13.1: Weapons Layout
The M61A1 is mounted above the forebody of the Gavial, in place of the pilot’s cockpit. The ammunition is stored directly below that to minimize the length and weight of the feed and shell-return system. The bullet path comes from the top barrel, and has a clearance of 14 degrees above the nose.
All missiles are launched from common LAU-129 missile rails. Each rail is placed to accomodate either an AIM-9 or AIM-120, so the only requirement for changing ordnance is keeping the load symmetrical.
67 13.2 Wide Angle Cameras
The IR 360 (Figure 13.2), available from the Sierra Pacific Corporation is combination thermal IR and CCD color camera system. It is a low cost, high resolution unit which can be mounted inside the fuselage behind a lens or outboard. The thermal IR and CCD modules can be mounted separately on rotating platforms and can be controlled remotely by computer for maximum range of visibility. To accomodate maneuver and threat identification, the CCD cameras are mounted flanking the M61A1, and the infrared camera is mounted directly underneath the barrel. The CCD’s provide full 180 degree forward vision, with +90/-15 degrees vertical. These views are trasmitted to the pilot in the remote cockpit and provide exceptional situational awareness. The IR 360 unit meets all necessary Mil Spec standards and is ideal for a UAV application.[4]
Figure 13.2: IR360-1 Camera.[4]
Figure 13.3: CCD Array in Canopy
68 13.3 FLIR
For targeting, navigation and laser tracking, the Raytheon AN/ASQ-228 ATFLIR was chosen. This Ad- vanced Tactical Forward-Looking Infrared system is a single-pod setup which integrates FLIR targeting, navigation, and laser spot tracking in one unit. The use of a multi-functional system such as this allows for two-thirds fewer parts than other systems. The pod should be mounted on the underside of the aircraft for maximum visibility. With a diameter of 13 inches, a length of 72 inches and an effective range of over
40 nautical miles, this unit is both compact and effective. Data and images from the ATFLIR pod will be transmitted from the UAV to the remote control station.[6]
Figure 13.4: AN/ASQ-228 ATFLIR.[5]
13.4 Radar
The AN/APG-79 AESA from Raytheon is a lightweight and highly effective radar system which is currently in use on the F/A-18E/F Super Hornet. The radar systems light and thin antenna is extremely reliable.
It weighs merely 95 pounds and should not require maintenance for 10 to 20 years. The AN/APG-79 is capable of 7 radar modes including: real beam mapping, synthetic aperture radar, air-to-air search, air-to-air track, passive, sea surface search, and ground moving target indication. It is capable of nearly simultaneous air-to-air and air-to-ground operation as well as multi-target and multi-missile tracking. [6] [6]
Figure 13.5: AN/APG-79.[6]
69 Chapter 14
Internal Systems
14.1 General Layout
A major concern when designing an unmanned aircraft are the avionics required for fulfill the mission.
The communication system necessary must have the ability to exchange real-time telemetry data as well as digital video and data from the targeting system. This data must be capable of being transmitted over a long distance without significant delay or loss of connection with the ground control station. This is essential in a combat situation where the ever changing nature of the environment is critical. Gavial will be equipped with a redundant, multi-link broadband communication system consisting of both satellite communication and direct line of sight links. A direct broadcast SATCOM antenna assembly as well as a line of sight radio frequency assembly and antenna with multiple UHF receiver/transmitters are utilized for air-to-ground communication. The feasibility of meeting the bandwidth requirements for such a communication system were confirmed by information published in MILSATCOM documentation.
”The key trend in communication systems is increasing data rates, primarily brought on by
migration towards higher RF frequencies and the emerging dominance of optical over RF sys-
tems. Optical systems are laser-based systems, which will offer data rates two to three orders of
magnitude greater than those of the best future RF systems. The advantages of optical communi-
cation were demonstrated in 1996 when a ground-based laser communications (lasercom) system
provided rates of 1.1 terabits/second (Tbps) at over 80 nm range. Airborne and spaceborne Tbps
lasercom systems will certainly be possible by 2025. Although lasercom will shortly surpass RF in
terms of data transfer rate, RF will continue to dominate at the lower altitudes for some time
into the future because of its better all-weather capability. Thus, both RF and optical technology
development will continue to progress out to 2025.
Data compression will remain relevant into the future as long as band-limited communica-
tions exist, but it is unlikely compression algorithms alone will solve the near term throughput
requirements of advanced sensors. A technology that intentionally discards information is not the
70 preferred technique. For now, compression is a concession to inadequate bandwidth.”[29]
While optical communication is preferred, the Air Force will soon have their Transformational Communi- cations Satellite System, (T-SAT) operational. This system is expected to be anywhere from 100-1000 times faster than current-day SATCOM systems. Figure 14.1 is from the Department of Defense Unmanned Aerial
Vehicles Roadmap and depicts projected bandwidth capabilities up to the year 2025. From this, it can be confirmed that the necessary bandwidth will be available by the time of possible deployment of Gavial .
Figure 14.1: Bandwidth Limitations
The communication system must also have anti-jam capability along with encryption and decryption of real-time telemetry data. The GPS and SATCOM dishes were placed just forward of the wing planform and internal fuel stores. This allowed for ample rotation space for the SATCOM dish without interference.
Active array radar was placed in the nose of the aircraft to allow maximum visibility. Several other systems were placed in the forward section of the aircraft just below the cannon. This would allow air vented from the cannon doors to be routed to critical systems for cooling. Vibration from the cannon was also considered when placing critical systems. This issue was remedied by using ample vibration damping where the cannon is mounted. Several situational awareness tools are employed on Gavial including radar, IR and CCD cameras, and forward looking infrared targeting. The forward looking infrared targeting and navigation pod allows for the engagement of multiple targets in any conditions. Several cameras were used to enhance situational awareness. Two CCD cameras mounted just forward of the cannon are capable of rotating and providing a wide field of view. There is also an IR camera mounted between the two CCD cameras which allows for greater visibility in low light situations.
71 Figure 14.2: Systems Layout
72 14.2 Cooling of the Avionics
An unexpected benefit of the unmanned approach was a simplification of the cooling system for the internal avionics in the nose. The systems cooling is accomplished by the opening of the forward bays to atmosphere.
There are 5 small holes in the underside of the forward fuselage, covered with simple filters to prevent water and particulates from entering the avionics bay. This air circulates inside, before being drawn by small fans through the OBIGGS and into vents that feed back into the main inlets. These cooling exhaust vents are visible in figure fig:propunit The total mass flow is 2% of the main inlet flow.
73 Chapter 15
Fuel and Electrical System
15.1 Fuel System
Figure 15.1: Fuel System (shown in green)
The fuel system was designed around the DCAP mission requirements, as this mission required the most fuel. The total fuel capacity of the Gavial is 30250 lbs. The internal stores can carry 17050 lb of JP-8. To make up the lack, three 660 gallon external tanks can be fitted. These tanks are retained for the duration of the mission, and are not equipped with automatic jettison systems. They can be removed by the ground crew. When the tanks are removed, there are smooth covers that fit flush to the skin to fill the connection slots.
The vast majority of the Gavial’s internal volume is given over to fuel storage, as can be seen in figure
15.1. On the DCAP mission, the fuel is nearly 75% of the takeoff gross weight. It is actually capable of carrying more fuel than is necessary for Point Defense Intercept mission. The Gavial would be able to have
74 a longer intercept range than is required, though not by a very large amount. This oversizing of the internal tankage did incur a penalty in wave drag, and created a fairly complex fuel system. The alternative, however, was a much larger set of external fuel tanks that caused an even greater drag penalty. With the current set of external tanks, the maximum attainable Mach number is 1.7. Using a larger design, two 1000 gallon tanks, reduced the maximum velocity to Mach 1.3 at altitude.
The internal fuel is segmented into several tanks, and a computerized fuel flow system ensures that the overall fuel CG stays relatively fixed. The section of the fuel tank directly over the engine core has heat shielding and a air cooling gap between the tank and the core to protect it.
15.2 Electrical System
Power for the aircraft is generated by a primary direct drive generator forward of the engine, and an APU housed underneath the combustor. The drive shaft for the generator runs forward from the compressor hub, through the inlet duct and into the generator.
Figure 15.2: Lateral Directional Control Planform
The power lines all run along the central spine ridge to the forward avionics area before running to individual systems. Those lines that power flap actuators run along the skin into the wings, then out along the forward most spar, and drop back along the rib nearest to the actuator it is needed to power.
75 Chapter 16
Servicing Plan
16.1 Forward Avionics Access
In order to facilitate servicing of the avionics and communication equipment, several doors have been built into the skin of the aircraft. The canopy covering the M61-A1 and camera array is removable. This allows access to the cannon, camera arrays and ammunition feed system. When the canopy is removed, the nose cone is freed to pivot starboard, revealing the radar. Similar to the F-18, the radar array and command computer systems aft of the dish slide out on rails. There is a removable hatch aft of the canopy that opens up the computer systems, and under those, the communications and OBIGGS system. There is also a hatch under the port canard that opens to allow reloading of the ammunition drum and removal of spent shell casings.
16.2 Aft Access
To access the engine and power train, there are two additional access panels in the aft of the airplane. The
first is the generator access panel, located between the inlet ducts on the underside of the fuselage. This allows servicing or replacement of the primary direct drive generator. The other panel is the engine access hatch. In this case, the entire lower half of the empennage is removable, to allow access to the APU and minor service access to the engine itself. In the case of replacement or full overhaul, then engine is unbolted from the half bulkhead engine supports, slid back 1/4 of its length, and lowered out.
76 Figure 16.1: Forward Access Panels
77 Figure 16.2: Aft Access Panels
78 Chapter 17
Weight Parameters
17.1 Takeoff Gross Weight
The primary objective in the weight discipline is to minimize the maximum takeoff gross weight (TOGW).
Roskam’s class I and II methods provided the basis for preliminary TOGW estimates. Despite the fact that weight minimization is directly related to cost increase; the gains in aircraft performance are immeasurable.
In order to improve the aircraft’s maneuverability, combat radius, loiter, etc. several key components were targeted for extensive weight reduction:
• Fuselage
• Wing
• Propulsion System
Sizing procedures, described in Section 3, gave initial component estimates based on weight fractions of the TOGW. Through the use of advanced metallic and composite materials in both primary and secondary structures drastic reductions in the weight of these components were achieved. The materials used in the
Gavial are all light weight composites widely used in modern aerospace applications, and were chosen for their high strength to density ratios. The chosen scaled F-414 propulsion system also utilizes modern advancements in materials and design to drive down weight and achieve high thrust output. The F-414, however, is heavier than the government furnished engine; due to the required maneuvers and long loiter periods specified in the RFP the weight gain was unavoidable. The Aircraft operational empty weight is shown in table 17.1, and the TOGW for each mission is shown in table 17.2.
17.2 Center of Gravity
The position of the center of gravity (CG) is an important parameter in the study of balance and performance.
Initial CG location estimates were calculated using Roskam’s methods, and were refined through stability
79 Table 17.1: Aircraft Dry Weight by Component Component Weight (lb) Component Weight (lb) Structure Avionics and Eauipment Fuselage Structure 1202.00 Generator 35.00 Wing Structure 1500.00 APU 100.00 Canard 38.31 Landing Gear Hydraulics 50.00 Vertical Tail 106.70 ANAPG-79 650.00 Fuselage Skin 292.99 CCD Array 30.00 Wing Skin 249.92 IR Camera 5.00 Spine Ridge and CFT Skin 44.20 Data Bus 10.00 TE Flap 69.63 INEWS 100.00 Aileron 60.00 FPCS 50.00 LE Flap 63.87 ICNIA 100.00 Main Landing Gear 633.81 SATCOM Antenna 8.00 Forward Landing Gear 258.45 GPS Antenna 6.00 Main Landing Gear Wells 61.66 OBIGGS 35.00 Forward Landing Gear Wells 30.00 ATFLIR 420.00 Inlets 48.93 Engine 4364.00
Weapons Fuel System Missile Hardpoints 100.00 Launch Rails 254.00 Missiles 1308.00 External Fuel Tanks 300.00 M61A1 275.00 Ammuniution System 300.00 Total 11352.47
and control analysis, Section 10.1, as well as mission performance procedures, Section 5. Once all components in the design were permanently placed, simple moment translation was used to calculate the Gavial’s exact
CG location at various payloads. The forward and aft limits of the CG are shown in figure 17.1, and the
TOGW CG is shown in table 17.3 for each mission.
17.3 Weight and Balance Analysis
The weight related performance of the Gavial is based on its CG location when fully loaded and its maximum
TOGW. With a high fuel to weight fraction, performance will greatly improve as payload is reduced, because of decreases in induced drag. Operational performance diagrams for all three missions are shown in Figure
17.2. Operational or empty weight consists of the airframe, all vehicle control systems, and the propulsion device. In order to calculate the forward and aft CG limits all possible flight conditions needed to be analyzed; analyses taking into account all various payload situations that may be encountered during flight.
The forward limit for CG location is 33.6 feet aft of the Gavial’s nose; which corresponds to a 40.39 percentage
80 Table 17.2: TOGW by Mission CAP Point Defense Intercept/Escort Structural Weight (lbs) 6169 6169 6169 Avionics/Sys. & Weight (lbs) 5963 5963 5963 Weapons Weight (lbs) 2237 2237 2237 Payload (lbs) 30560 16168 25433 TOGW 44929 30537 39802
Table 17.3: Center of Gravity Location by Mission CG Locations X (ft aft of nose) Y (ft right of centerline) Z (ft above runway) CAP Mission 34.41859 0.009939 5.27332499 Point Defense Mission 35.00322 0.014615 6.03625623 Intercept/Escort Mission 34.58532 0.011219 5.50426336
of the mean aerodynamic chord (MAC). The upper limit is 35.8 feet from the nose and 52.1 percent of the
MAC.
81 Figure 17.1: CG Limits
Figure 17.2: CG Excursion Chart: Defensive Counter Air Patrol
82 Figure 17.3: CG Excursion Chart: Point Defense Intercept
Figure 17.4: CG Excursion Chart: Intercept/Escort
83 Chapter 18
Cost Estimation
Along with designing the Gavial to be a long range, high speed, maneuverable, intercept fighter the RFP calls for a cost efficient aircraft. The given cap for the flyaway cost is a maximum of $15 million. The flyaway cost includes all values associated with the production of the airframe, propulsion system, and avionics. Price values for the systems and equipment used in the ground station and others not specified in the RFP were found through research, and communication with manufacturers. The estimated flyaway cost is given in
figure 18.1.
Figure 18.1: Airframe Fly Away Cost Breakdown
The RFP stressed the importance of expense, and so through out the design process attempts were made to keep the flyaway value as low as possible. Going with an unmanned approach led to increases in the needed avionics and systems, but after extensive research added costs were kept to a minimum through
84 the use of multi-functional equipment. To keep the Gavial light advanced metallic composites make up the airframe and delta wing structures. Although these materials are quite expensive when compared to traditional metals, their high strength capabilities reduced the number of needed spars and ribs. This not only kept our weight down, but also the associated bulk material cost.
With the standing flyaway value at roughly half of the $15 million cap, the remainder of the given amount will be used to pay for the associated manufacturing cost. This includes the manpower needed to construct the aircraft structure, install the avionics and systems, as well as test the finished product.
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87