n
Bomber of the 21st Century and Beyond
2001/2002 AIAA Undergraduate Team Aircraft Design
SLOB Works
(Supercruise Low-Observable Bomber Works)
May 2, 2002
Team Roster
Member AIAA Number Signature
Steve D’Adamo 177918
Derek Geiger 188822
Scott Henderson 214947
Andy Krohn 203787
Brian Shepard 204729
Matt Stephan 000000
Zach Sherman 000000
Obie Woods 000000
Arthur Jarjisian 000000
Harsh Vasavada 215016
Faculty Advisor
Dr. W. H. Mason
2002 SLOB Works, a Virginia Tech AOE Design Team 1
Executive Summary
The SLOB Works Group presents the SW-Ghost as a solution to the 2001-2002 AIAA Undergraduate Team Aircraft
Design Competition Request for Proposal (RFP) for an Advanced Deep Interdiction Aircraft.
The main drivers for this proposal were cost, stealth, supercruising capabilities, medium payload with multiple configurations, and range. An extensive aircraft comparative study was performed to evaluate past aircraft capabilities. With the knowledge gained from this study, and keeping the drivers in mind, it was possible to develop four concepts, each meeting the requirements set forth by the RFP. To choose the best concept, three selection matrices were used to evaluate the characteristics of the aircraft. This concept evolved through the preliminary design phase leading to an optimized aircraft that meets and exceeds the requirements in the RFP.
The SW-Ghost is a blended wing body aircraft with canards and a split canted tail. It utilizes four 30,000 lbs engines with afterburner capabilities. A diamond shaped wing is blended into the fuselage and area ruling was used throughout the aircraft to improve the stealth and aerodynamic characteristics. The cranked outer portions of the diamond wing provide better stability and control characteristics at low subsonic speeds. The weapon and engine pods are integrated into the fuselage and wing. The pod locations are close to the ground, which provide easy maintainability to the aircraft and reduce the turn around time. The weapon pods were sized to meet the largest payload and can be arranged with all combat configurations. The RFP engines are used for their better performance over other engines. By introducing all of these characteristics into the aircraft, the Ghost will be able to fly efficiently at subsonic and supersonic speeds.
To meet the requirements in the RFP, it was necessary to concentrate on the structures, materials, and systems of the aircraft. The bomber will have an integrated configuration of aluminum, steel, titanium, and magnesium materials for an optimal combination of strength, weight and cost. The aircraft will be lighter and become stronger by using Sine wave technology in the spars. For the aircraft’s systems, the majority of the items will be the government furnished equipment to keep the cost down. Even though most of the equipment used is government furnished the aircraft will incorporate the top-of- the-line avionics, flight control and propulsion systems.
2002 SLOB Works, a Virginia Tech AOE Design Team 2
Table of Contents
Executive Summary...... 2 Index of Tables ...... 5 Index of Figures...... 5 Index of Abbreviations ...... 7 Index of Symbols...... 7 1. Aircraft Requirements and Proposed Concept Designs ...... 8 1.1. Introduction ...... 8 1.2. Analysis of Request for Proposal ...... 8 1.3. Aircraft Comparative Study...... 10 1.4. Concepts ...... 11 1.4.1 Concept SW-1...... 11 1.4.2 Concept SW-2...... 13 1.4.3 Concept SW-3...... 15 1.4.4 Concept SW-4...... 16 2. Concept Analysis and Selection Process...... 18 2.1. Concept Design Tools ...... 18 2.1.1 Nicolai’s Aircraft Sizing Program ...... 18 2.1.2 AeroDYNAMIC Program...... 18 2.1.3 Development of SW Excel Sizing Program...... 23 2.1.4 Cost Analysis Program ...... 24 2.2. Generation of Carpet Plots ...... 26 2.3. Concept Selection Process...... 30 2.3.1. Concept Design Matrix...... 30 2.3.2. Risk Management Matrix ...... 32 2.3.3. Cost Analysis Matrix ...... 34 2.4. Final Analysis & Selection Process...... 35 2.5. Aircraft Design & Layout...... 35 3. Aerodynamics ...... 41 3.1. Planform and Airfoil Selection...... 41 3.2. Lift Analysis ...... 42 3.3. Drag Analysis...... 43 3.4. Aircraft Geometry ...... 46 3.5. High Lift Devices ...... 47 4. Structures and Materials...... 48 4.1. Materials...... 48 4.2. Structures...... 49 5. Stability and Control ...... 53 5.1. Method of Analysis ...... 53 5.2. Static Stability ...... 54 5.3. Engine Out...... 55 5.4. Dynamics and Flight Qualities ...... 55 6. Systems and Payloads ...... 57 6.1. Basic Layout...... 57 6.2. Fire Control and Defensive Systems ...... 57 6.3. Radar Cross Section (RCS) Prediction/Evaluation...... 58 6.4. Cockpit ...... 58 6.5. Electrical System...... 60 6.6. Flight Controls...... 61 6.7. Digital Flight Controller and Engine Control System ...... 62 6.8. Landing Gear...... 62 6.9. Fuel System...... 64 6.10. Environmental Control System...... 65 6.11. Anti-Icing Equipment ...... 65 6.12. Aircraft Lighting...... 65 2002 SLOB Works, a Virginia Tech AOE Design Team 3
6.13. Weapons ...... 65 6.14. Bomb and Missile Bays ...... 65 6.15. Defensive Systems...... 67 7. Propulsion Systems...... 69 7.1. Propulsion system comparative study ...... 69 7.2. Thrust Requirements...... 70 7.3. Propulsion System Selected...... 70 7.4. Inlet Geometry...... 71 8. Performance ...... 73 8.1. Performance Parameters...... 73 8.2. Maneuvering Performance Diagram...... 79 8.3. Sustain Load Factor Envelope...... 81 8.4. Conclusions on Performance Analysis ...... 82 9. Weight Analysis...... 83 9.1. Weight Breakdown...... 83 9.2. Center of Gravity...... 85 9.3. Weights and C.G. Conclusion ...... 86 10. Cost Analysis ...... 87 10.1. Introduction...... 87 10.2. Cost Analysis Method...... 87 10.3. Aircraft Life Cycle Cost ...... 87 10.4. Unit Cost, Fly-Away Costs and Cost per Pound...... 90 10.5. Cost Trade Study ...... 91 10.6. Cost Conclusion...... 92 11. Manufacturing and Maintenance...... 94 12. References...... 97
2002 SLOB Works, a Virginia Tech AOE Design Team 4
Index of Tables
1.1 Results from Aircraft Comparative Study………………………………………………………………………….11 2.1 Example of Results…...... …………………………………………………………………………….…………….24 2.2 Research Test Evaluation and Development cost ………………………………………………………………… 24 2.3 Acquisition Program Cost…………………………………………………………………………………………. 25 2.4 Operating Program Cost………………………………………………………………………………...... 25 2.5 Disposal Program Cost…………………………………………………………………………………………….. 25 2.6 Life Cycle Cost……………………………………………………………………………………………………..25 2.7 Concept Design Matrix……………………………………………………………………………………………. 32 2.8 Risk Management Matrix…………………………………………………………………………………………. 32 2.9 Cost Analysis……………………………………………………………………………………………………….35 2.10 Final Decision Selection…………………………………………………………………………………………..35 3.1 Key Aerodynamic Parameters for Mission Segments……………………………………………………………...42 4.1 Material Properties………………………………………………………………………………………………… 48 5.1 Stability Derivatives of SW-Ghost at Supercruise (Mach 1.6) ……………………………………………….. 53 5.2 Stability & Control Derivatives for SW-Ghost at Takeoff (Mach 0.3)………...…………………………………..54 5.3 Engine Out Data for SW-Ghost ……………………………………………………………..……………………. 55 5.4 Comparison Chart for SW-Ghost with the MIL-F-8785 B………………………………………………………... 56 7.1 Engine comparative study for resized engines using equations from RFP..………..……………………………... 70 7.2 Thrust required and Thrust available for the RF P engine at given conditions…………………………………….70 7.3 Base engine specs vs. sized engine specs…………………………………………………………………………..71 8.1 Take-off distances for the three different surfaces…………………………………………………...……………. 73 8.2 Evaluation of each altitude’s (L/D)/SFC…………………….……………………………………………………..73 8.3 Fuel Burned during 1000 nm Supercruise out……………….……………………………………………………. 73 8.4 Landing distances for the three different surfaces………………………………………………………………….78 8.5 Summary of each segment giving the important performance parameters………..…………………...... 82 9.1 Structural Weights Group………………………...………………………………………………………………...83 9.2 Propulsion System Weights Group……………………………………………………………………...... 83 9.3 System Weights Group……………………………………………………………………………………………..84 9.4 Ordinance Weights Group………………………………………………………………………………………….84 9.5 Fuel and Crew Weights Group…………………………………………………………………………………….. 84 9.6 Weight Summaries and Inertias of SW-Ghost…………………………………………………………………….. 85 9.7 SW-Ghost Ratios…………………………………………………………………………………………………... 85 10.1 Aircraft Inputs……………………………………………………………………………………………………. 88 10.2 Adjustment Factors………………………………………………………………………………………………. 88 10.3 Rates…………………………………………………………………………………………………...... 88 10.4 RDT&E Cost Breakdown …………………………………………………………..…………………………… 88 10.5 Acquisition Cost Breakdown…………………………………………………………………………………….. 89 10.6 Operating Cost Breakdown………………………………………………………………………………………. 90
Index of Figures
1.1 RFP Mission ………………………………………………………………………………………………………10 1.2 Concept SW-1………………………………………………………………………………………..…………….12 1.3 Concept SW-2……………………………………………………………………………………………………...13 1.4 Concept SW-3…………………………………………………………………………………………………….. 15 1.5 Concept SW-4……………………………………………………………………………………………………...17 2.1 Sample Input Screen for AeroDYNAMIC …………………………………………………………………………. 19 2.2 XB-70 CD0 versus Mach # Comparison… ……………………………………………………………...... 20 2.3 XB-70 CL vs. CD Comparison at Mach 1.6…………………………………………………………..…………….. 20 2.4 XB- 70 Coefficient of Lift Curve Comparison at Mach 1.6…………………………………………..…………….21 2.5 Adjusted CD0 vs. Mach Number for Concepts……………………………………………………….…………….. 21 2.6 CL vs. CD for Concepts at Mach 1.6…………………………………………………………………..…………… 22 2.7 Lift Curves for Concepts at Mach 1.6.……………..……………………………………………………………….22
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2.8 Carpet Plot of SW-1………………………………..……………………………………………………………….28 2.9 Carpet Plot of SW-2……………………………………………………………………………………………….. 29 2.10 Carpet Plot of SW-3……………………………………………………………………………………………… 29 2.11 Carpet Plot of SW-4……………………………………………………………………………………………… 30 2.12 Evolution of Aircraft…………………………………………………………………………………...... 36 2.13 Top View of SW-Ghost…………………………………………………………………………………………...37 2.14 Bottom View of SW-Ghost………………………………………………………………………………………. 38 2.15 Side View of SW-Ghost…………………………………………………………………………………………..39 2.16 Front View of SW-Ghost………………………………………………………………………………………… 40 3.1 SLOB Works Ghost semi-planform...………………………………………..……………………………………. 41 3.2 Parasite-drag buildup of SLOB Works concept at altitude 50,000 feet …………………………………………... 44 3.3 Drag polar of SLOB Works concept at cruise Mach 1.6………………………………………………………….. .45 3.4 Lift to drag ratio at cruise conditions……………………………………………………………………………… 46 3.5 Area Distribution at Mach 1.01……………………………………………………………………………………. 47 4.1 Material Breakdown of Aircraft………………………………………………………………………...... 49 4.2 Structures Top View/ Major Components………………………………………………………………………….50 4.3 Structures Side View/ Main Bulkheads…………………………………………………………………………….51 4.4 Structures Bottom View/ Removable Pods……………………………………………………………………….. 51 4.5 V-n Diagram………………………………………………………………………………………………………..52 4.6“Sine Wave” Spar Design………………………………………………………………………………………….. 52 5.1 Aerodynamic Center Shift with change in Mach #...... 55 6.1 Top view of the Fire Control Systems…………………………………………………………………………….. 58 6.2 View of the cockpit and instrumentation………………………………………………..………………………… 59 6.3 Effective envelope of the K-36D ejection seat……………………………………..…………………...... 60 6.4 Top view of the electrical system of the SLOB Works Ghost…………………………………………………….. 61 6.5 Top view of the Flight Controllers along with control lines and motors………………………………………….. 62 6.6 Side view of front landing gear……………………………………………………………………………………. 63 6.7 Side view of main landing gear……………………………………………………………………………………. 63 6.8 Top view of Fuel Tank Positions………………………………………………………………………………….. 64 6.9 Combat loads of the SW Ghost……………………………………………………………………………………. 66 6.10 Side and bottom views of weapons bays and their clearances………………………………………...... 66 6.11 Top view of defensive system locations…………………………………………………………………………..68 7.1 Reverse Thruster System Diagram……………………………………………………………………...... 71 7.2 Double –wedge intake geometry………………………………………………………………………………….. 72 7.3 S-Bend subsonic diffuser design for the Ghost……………………………………………………………………. 72 8.1 Climb Analysis at MTOGW, n=1, Mil. Thrust……………………………………………………………………. 75 8.2 Ps Plot for n=1, 50% fuel weight, and Mil. Thrust………………………………………………………………... 76 8.3 Ps Plot for n=2, 50% fuel weight, and Mil. Thrust………………………………………………………………... 76 8.4 Ps Plot for n=5, 50% fuel weight, and Mil. Thrust……………………………………………………...... 77 8.5 Ps Plot for n=1, 50% fuel weight, and afterburners……………………………………………………………….. 77 8.6 Sustained turn rate at 50,000 ft……..………………………………………………………………..……………. 79 8.7 Sustained turn rate at Sea Level…….………………………………………………………………..……………. 80 8.8 Sustained turn rate graph at 15,000 ft………………………………………………………………..……………. 80 8.9 Envelope for the Ghost at Ps=0 for a range of load factors……………………………………………………….. 81 9.1 Center of Gravity Movement……………………………………………………………………………………… 86 10.1 Breakdown of the Life Cycle Costs……………………………………………………………………………… 90 10.2 Unit and fly-away cost trade study in year 2000 dollars...……………………………………………………….. 92 10.3 LCC Breakdown Comparison between number of units produced……………………………………………….92 11.1 Weapon and engine pods………………………………………………………………………………………… 95 11.2 Engine pod removal…………………….……………………………………………….…………..…………….95 11.3 Weapon pod removal…………………………...…………………………………………………..……………. 95 11.4 Crew access to flight deck………………………………………………………………………..…...... 96
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Index of Abbreviations
AEW Anti-Electronic Warfare AGM Air-to-Ground Missiles AIM Air Intercept Missiles AMRAAM Advanced Medium-Range Air-to-Air Missile CAD Computer Aided Design cg Center of Gravity GBU Guided Bomb Unit GCI Ground Control Intercept JDAM Joint-Direct Attack Munition JSOW Joint-Stand Off Weapon LDGP Laser Designated General Purpose MAC Mean Aerodynamic Chord nm nautical miles RDTE Research, Development, Test, & Evaluation RFP Request for Proposal RCS Radar Cross Section SLOB Supercruising, Low-Observable Bomber SW SLOB Works TLFC Thermal Laminar Flow Control TOP Take-off Parameter USAF United States Air Force
Index of Symbols
A C Apect Ratio C Specific fuel consumption D0 Zero lift, drag coefficient CL C Lift coefficient Lmax Maximum lift coefficient CLTO D Lift coefficient at take-off E Drag e Endurance, Loiter time g Oswald efficiency factor Gravity Κ Induced drag coefficient L Lift n g-loading Ps Specific power q Dynamic pressure R Range S Wing area S Landing Landing distance T Thrust T TO Thrust at take-off W Weight W e Empty weight W f Fuel weight W o Take-off gross weight, TOGW W p Weight of payload W TO Weight at take-off V Velocity ρ Density σ Density Ratio 2002 SLOB Works, a Virginia Tech AOE Design Team 7
1. Aircraft Requirements and Proposed Concept Designs
1.1. Introduction
In September of 2001, SLOB Works was given a request for proposal (RFP) to design an advanced, supercruising, deep interdiction bomber1. After review of the proposal, it was decided that an extensive comparative study of previous related aircraft was required. This would enable SLOB Works to analyze the characteristics of relevant aircraft. These characteristics were used to evaluate the performance of any proposed concept aircraft. SLOB Works was able to develop criteria upon which the concepts were based through these steps. It was briefly determined that the aircraft needed to have substantial fuel weight, a generous amount of thrust, a large wing planform, etc. Once this was accomplished the initial designing began.
Four different aircraft were designed to address the requirements in the RFP, each having its own unique shape.
Analysis was conducted on each of these aircraft to determine their performance characteristics. The main goal was to make sure that each of these aircraft could meet the specifications of the RFP. After this analysis was conducted, selection of the final design started. SLOB Works used the comparative study matrix, a risk management matrix, and a cost analysis matrix to do this. Through research, calculations, and sound engineering judgment the best aircraft was selected.
1.2. Analysis of Request for Proposal
The United States needs to develop new aircraft. When the United States Air Force retired the F-111 in 1996 the
U.S. had already accounted for the plane going out of service by introducing new aircraft. The F-15E was the main successor to the F-111, but a few other aircraft were available for this role. These airplanes were the F-117, B-1, and the B-2. It is expected that these four planes will reach the end of their service lives by the year 2020. There must be a plane ready to fulfill the same type of requirements that the planes cited above meet once these aircraft are retired.
According to the RFP, the aircraft must be capable of delivering precision-guided weapons from long range without an extended preparation time. It must be able to accomplish its mission without the support of other aircraft (fighter, reconnaissance, refueling, etc). There is also the need for the plane to supercruise (cruise at supersonic speed without afterburners). This requirement is important because the aircraft would be able to cut travel times in half, making it much quicker in responding to crises around the world.
Through analysis of the RFP, SLOB Works decided on the following major concept design drivers: cost, stealth, supercruise ability, payload and range capabilities. The RFP requires that 200 aircraft be produced at a maximum fly-away cost of $150 million each. Radar, infrared, visual, acoustical, and electromagnetic signatures must be reduced to minimum 2002 SLOB Works, a Virginia Tech AOE Design Team 8
levels and balanced with each other so that no one signal is more detectable than the others. A frontal radar cross-section against 1-10 GHz GCI, acquisition, and tracking radar of less than 0.05 m2 is necessary. The RFP dictates the ability to supercruise at Mach 1.6 while carrying various payload configurations, with a maximum weight of 8,700 lbs,
(2) AIM-120 + (4) GBU-27
(2) AIM-120 + (4) 2,000 LB JDAM (standard configuration)
(2) AIM-120 + (4) AGM-154 JSOW
(2) AIM-120 + (4) Mk-84 LDGP
(16) 250 LB Small Smart Bomb
The design mission requires the configuration carry two AIM-120s and four 2,000 lb JDAMs. A total mission range of 3,500 nm is set by the RFP.
Through further analysis of the RFP, the following important yet secondary concept design drivers were taken into account: operation, maintenance and crew requirements. The aircraft must operate in all weather conditions from existing
NATO runways of 8,000 ft. Maintenance requires easy access to primary elements of all major systems. Although the cockpit is designed around a crew of two, due to the mission length, the cockpit must be fully operational for one pilot control.
The RFP specifies that all competitors must complete a high altitude, supercruising, interdiction mission consisting of eleven phases, as seen in figure 1.1. Phase number one is made up of two sub-phases, engine warm-up and acceleration to climb speed. Next, the aircraft must climb from sea level to optimum supercruise altitude. Phase three requires the aircraft to supercruise out 1,000 nm at Mach 1.6 and optimum altitude (Note: there is no distance or credit for take-off and climb). At the end of this segment it is instructed to climb above 50,000 ft. Once this altitude is reached our aircraft must dash out 750 nm at Mach 1.6. Upon completion of the dash out, the aircraft must perform one 180-degree turn at 50,000 ft and Mach 1.6.
At the end of this turn the air-to-surface weapons must be dropped while retaining racks, pylons, and air-to-air missiles.
After this is accomplished, the aircraft must dash back 750 nm at Mach 1.6 and above 50,000 ft. It then descends back to optimum cruising altitude and supercruises back 1,000 nm at Mach 1.6.
The last phase of our mission is to descend to sea level and land (Note: there is no distance or fuel credit for descent and landing). There is a fuel reserve criteria that states there must be excess fuel to fly for thirty minutes at sea level at speed for maximum endurance.
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Dash and Store Drop 50,000 ft Descent 1500 nm M = 1.6 Take-Off Best Cruise Best Cruise Loiter Altitude Altitude 30 min 1000 nm 180° Turn 1000 nm M = 1.6 M = 1.6 Warm-Up Climb Dash-in Dash-out Landing
Internal Store Drop
No Distance Credit No Distance or Fuel Credit
3,500 nm FIGURE 1.1: RFP Mission
1.3. Aircraft Comparative Study
After analyzing the RFP and determining the requirements that were set forth, SLOB Works conducted a comparative study of relevant aircraft. It was possible to gain an idea of what has already been accomplished and where more development is needed through this study. Eleven aircraft were analyzed in this study. These planes could not meet the
RFP, however, they had special characteristics that could satisfy one or more of the RFP requirements.
Some of the aircraft characteristics that were examined were wing planforms, sizes, weights, and flight regimes.
The types of wing configurations that were included in this study were forward-swept, variable sweep, delta wings, flying wings, and canards. Aircraft of many different sizes and weights were also included. Each of the aircraft in this study had a different mission requirement, for example, fighters, commercial, reconnaissance, and experimental were included alongside the bombers. Since there are few supercruising aircraft, other high subsonic and supersonic aircraft were also examined.
International aircraft from Russia and Europe were included, not to limit the study to purely domestic aircraft.
As a result, it was possible to narrow the search to the following eleven aircraft:
1. B-1B Lancer 5. F-22 Raptor 9. X-29
2. B-2 Sprit 6. F-117 Nighthawk 10. TU-22 2002 SLOB Works, a Virginia Tech AOE Design Team 10
3. B-58 Hustler 7. XB-70 Valkyrie 11. TU-160
4. SR-71 Blackbird 8. Concorde
TABLE 1.1: Results from Aircraft Comparative Study
Parameter Value (Wp+Wf) / Wt 0.606
Wf / Wt 0.499
Wp / Wt 0.231 T/W 0.46 W/S 90 The technical characteristics of these aircraft were analyzed in a comparative matrix, whose structure will be described in section 2.3.1. Shown in table 1.1 are some results of the study. The results from this study were used as a basis for determining the characteristics of the four concepts described in detail below.
1.4. Concepts
Using the comparative study information, four concepts were developed based on the aircraft examined in the comparative study. These aircraft were sized and evaluated using carpet plots and a combination of in-house and commercial programs. Each of the four concept designs will be described in detail in the following sections. The sizing, analysis, selection programs, and technical data will be presented in chapter 2.
1.4.1 Concept SW-1
The SW-1 (figure 1.2) was designed with the goal of carrying out the long-range mission. The overall weight will be directly related to the final cost of any of the concepts. By minimizing the weight of this concept, it ideally resulted in an aircraft design that falls beneath the RFP requirement of $150 million fly-away cost.
Three other aircraft were analyzed with the approach of combining several key features of each plane. The B-58, B-
1B, and F-117 were the three main drivers for this concept. The B-58 can fly well over Mach 1.6 thus its physical design was in the target range for this concept. With a length of just over 100 feet, SW-1’s size was determined to be ideal when using just two engines; giving a total thrust of 52,000 pounds. For stealthiness, the F-117’s idea of upper wing mounted engines was used. Grills over the inlets of the engines also help with radar evasiveness. Other stealth technologies, like radar absorbent materials, were also carefully examined. For additional long-range bomber features, the B-1B came into play.
Although its size was much larger than the SW-1 concept, it provided a good source for internal weapons carriage and payload configuration.
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SW-1 combines these three plane concepts in a way that meets some of the requirements of performance and stealth.
The simple fuselage/conventional wing design will drive down aerodynamic difficulties when supercruising at a speed of at least Mach 1.6. The plane was designed to work with only two full-size engines (each 35 ft. long), based on the RFP engines.
By mounting the engines on the top of the wing it allows for a flat bottom wing. Also, they do not interfere when payloads are being dropped. With a maximum thrust of just over 52,000 pounds, it is necessary that the overall weight be low. Fewer
FIGURE 1.2: Concept SW-1 engines means less fuel, and that results in a lower overall cost. A lower weight also will provide a higher thrust-to-weight ratio. This will, like the aerodynamics, aid in the ability of the plane to supercruise. The fuselage fineness ratio of this aircraft was designed to be large, and with a fuselage length of 102 ft, and a width of 12 feet, the fuselage fineness ratio was
8.5. Large, split elevons are used for control. The conventional wing planform has a 55-degree sweep, lower than the 60- degree that will help meet a NATO take-off runway requirement of 8,000 feet. It was designed to be a simple concept.
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1.4.2 Concept SW-2
SLOB Works concept SW-2 (figure 1.3) was based on the design of the XB-70, the Concorde, the F-22, and the F-
117. The XB-70’s delta wing with front canard configuration was the main inspiration for this design. The difference was the decision to use a blended wing body with diamond shape wings and canards. The concept used the V-shaped tail, borrowed from the F-117, to help reduce the radar cross section. This shape proved to be successful in the F-117, which enhanced the stealthiness of the aircraft. The F-22 was also incorporated in concept SW-2 because of its blended wing body
FIGURE 1.3: Concept SW-2 and capability of supercruising at Mach 1.5. The last aircraft analyzed was the Concorde. The Concorde was primarily used for its long range, high payload, and supersonic flying capabilities.
Before designing SW-2, SLOB Works had to figure out an effective way to combine all the characteristics of the mentioned aircraft into a reasonable and unique design. Through research, it was found that the blended wing body design, along with the V-shaped tail, helped in reducing the radar cross section. Canards were used in the design of the XB-70 to provide control for the aircraft, while also sharing the lifting loads. The diamond wing shape was designed as a delta wing
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with a trailing edge sweep. The delta swept wings helped in the ability to fly supersonically, as well as in the ability to supercruise.
One of the main purposes of the design of the XB-70 was to create a compression lift aerodynamic design.
Compression lift was a successful aerodynamic concept that could improve the lift-to-drag ratio at supersonic speeds. Any body shape will create shock waves at supersonic speeds, forming at the nose and at any other place where the cross-section area is increasing. These shocks trail back at the determined Mach angle. In the XB-70 design, the inlet duct was faired back into a wide nacelle, with a steadily widening cross-sectional area until a maximum was reached. The engines and payload were also carried in this nacelle, which created a strong shock on either side with greatly increased static pressures behind the shocks. By placing the wings above the shocks, the increased pressure beneath the wing provided free lift, roughly 30% of the total lift required. Since SW-2 was close to the design of the XB-70, SLOB Works will be able to make use of this aerodynamic concept. One of the concerns for SW-2 was that the compression lift concept worked on the XB-70 cruising at
Mach 3, but SW-2 would only be cruising at a maximum speed of Mach 1.8. Even though the concept might not have the total output of lift due to compression lift as the XB-70, SW-2 should still benefit from this aerodynamic concept.
The next step in creating SW-2 was applying these characteristics to a physical design. First, the wing was integrated into the fuselage, which created the blended wing body. Next the canards were applied to the design with the expectations of adding control to the design. Then the trailing edge sweep was added. The trailing edge sweep is -17 degrees and has a leading edge sweep of 60 degrees. The concept is going to have a total of four engines producing approximately
105,000 lbs of thrust. This helped increase the thrust-to-weight ratio, which improves its likelihood to supercruise. To allow for enough room for fuel and payload, the overall length was 143 ft with a wingspan of 90.25 ft. This is the largest of SLOB
Works proposed designs, with an estimated total weight of approximately 227,000 lbs. Some key advantages in this concept are the blended wing body design that provides for better aerodynamics and stealth capabilities. The diamond wing shape insures the ability to conduct a supersonic flight.
There are a few disadvantages for this concept. The blended wing body design is complex and difficult to manufacture. This is the biggest of the four proposed concepts, which could increase its radar cross-section. The large size of the aircraft increases the amount of fuel burned, which in-turn increases the overall cost. Finally, the compression lift design will need more research to establish its applicability to this concept.
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1.4.3 Concept SW-3
This concept was based primarily on the design of the F-117A Nighthawk. The design of SW-3 (figure 1.4) concentrated on creating a stealthy design. The most unique aspect of this aircraft is the use of faceting for the design of the fuselage. This feature serves to reduce the radar cross section (RCS) by reflecting the radar waves away from the receiver.
The engines of this aircraft exhaust through a “platypus” type exhaust. This system vents the engine exhaust over a series of heat absorbent tiles to help reduce the infrared signature of the plane.
FIGURE 1.4: Concept SW-3
At 128,000 pounds TOGW, the plane is among the lightest in weight relative to the other three concepts. The wing shape is primarily a delta wing with a trailing edge sweep. This creates a near diamond, the optimal shape for both supercruising and low-speed flight. The wing of this concept is swept at 60 degrees. The wingspan is 97 feet and the wing area is 1,620 square feet. The aspect ratio of this aircraft is 5.81. The length of the aircraft is 107 feet. This aircraft is equipped with two full size engines according to the RFP. Using the data from the RFP the total thrust of this aircraft was found to be 52,800 pounds.
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1.4.4 Concept SW-4
SLOB Works concept SW-4 (figure 1.5) is based off of the designs of the Tu-160, X-29, and the Concorde. The
Tu-160’s blended wing and body is the first characteristic that was incorporated into concept four. Examined next was the
X-29’s idea of using canards while also having a forward swept MAC. The last aircraft analyzed was the Concorde. This was done purely based on its long range and supersonic flight capability. The characteristics of these aircraft were incorporated into concept SW-4.
Before approaching the design aspect of SW-4, there was a need to determine how to combine all the characteristics of the mentioned planes. Through research it was determined that a blended wing and body design was helpful in reducing the radar cross-section, which was one of the major design drivers. Canards are used in the design of the X-29 to provide additional control for the aircraft while sharing the lifting loads as well. The forward swept MAC was the only way to have a trailing edge sweep angle larger then the leading edge sweep angle without having a forward swept wing. This concept needs to have a large volume due to massive fuel and payload requirements.
The next step in creating the concept SW-4 was applying these characteristics to a physical design. First the wing was integrated smoothly into the fuselage, which creates the blended wing body. Next, canards were applied to the aircraft with expectations of added control. To maintain the forward swept MAC, a leading-edge sweep of 2 degrees and a trailing- edge sweep of -30 degrees were incorporated into the design of the wing. This concept has three engines that are aimed at increasing the thrust-to-weight ratio and ensuring the ability to supercruise. To allow for ample storage of fuel, payload, and systems, the overall length reached 117 feet. Vertical tails angled at 66 degrees from the horizontal and four separate ailerons were used as SW-4’s rolling surfaces.
The concept SW-4 has some key advantages that are incorporated into the design. It has the blended wing body design that provides for better aerodynamics and stealth capabilities. The small leading edge sweep angle eliminates cross flow instability. This concept also aims at utilizing thermal laminar flow control (TLFC) over the wing. TLFC’s purpose is to create a long stretch of laminar flow over as much of the wing as possible. If accomplished, the total drag could be reduced and would decrease the total fuel weight. The thrust-to-weight ratio is comparable to the Concorde that giving the indication of the ability to supercruise is possible for this design.
The problem with a low sweep angle is that when flying through Mach 1 CDO rises dramatically, requiring larger amounts of thrust to overcome this additional drag. Another disadvantage with a small leading edge sweep angle is the possibility of aerodynamic divergence and the increase in frontal RCS. At this time TLFC is in the experimental phase and
2002 SLOB Works, a Virginia Tech AOE Design Team 16
has not yet been applied to previous aircraft. This proves that it will be a significant challenge. Finally, since a BWB is complex it will prove difficult to manufacture.
FIGURE 1.5: Concept SW-4
2002 SLOB Works, a Virginia Tech AOE Design Team 17
2. Concept Analysis and Selection Process
2.1. Concept Design Tools
To find the weight, size, thrust-to-weight ratio and wing loading for these aircraft, a variety of analysis tools were used. These programs, for example, determined design points for constraints, weights and aerodynamic characteristics. The next sections describe these design tools.
2.1.1 Nicolai’s Aircraft Sizing Program
Nicolai’s sizing program was used2. This program requires 27 inputs and finds an estimate of the weight of fuel used, the empty weight, and the TOGW. This program was used initially to evaluate various estimations of weight for the individual concept aircraft.
After using this program initially, it was decided by the SW group that Nicolai’s program was too sensitive to some input parameters, thus Nicolai’s sizing program was used mostly as a guide, as opposed to providing the primary results.
2.1.2 AeroDYNAMIC Program
Having a sizing tool that will completely analyze an aircraft accurately is extremely difficult. From the resources that were available, it was decided that AeroDYNAMIC v1.00.023 would best satisfy the sizing requirements. Getting an approximate aircraft weight along with the basic aerodynamic characteristics was the goal. AeroDYNAMIC was used because all that was required for inputs were an initial design and a mission.
The initial design definition requires that the user input geometric shapes to form a wing and fuselage combination.
From there, high lift devices, vertical tails, engines, landing gears, and systems can be incorporated into the analysis. These all come from the initial design and RFP requirements and are individually entered into the program as geometric shapes.
Once the aircraft design is completed, a mission can be set up so that the aerodynamic features can be analyzed. Each leg of the mission is defined with all the details entered by the user. Figure 2.1 shows an example of the AeroDYNAMIC program.
2002 SLOB Works, a Virginia Tech AOE Design Team 18
FIGURE 2.1: Sample Input Screen for AeroDYNAMIC
Under the analysis there are three main analysis categories. The first is the aerodynamic analysis, from which it was possible to obtain estimates of the drag polar, lift curve, and lift over drag versus lift coefficient data. A major advantage is the graphical output. Plots of CD0 vs. Mach number, lift coefficient versus angle of attack, thrust and drag versus Mach number, and cross sectional area versus length of aircraft show directly how different designs perform. The next analysis is performance. From this analysis we get specific excess power, constraint diagrams, V-n diagrams, and a mission analysis.
The mission analysis gives the amount of fuel used during the desired mission. Finally, a weight and stability analysis is given and the code provides the take-off weight and wing loading.
It was important to validate the program so that the level of uncertainty was established. The validation was done using the actual data for the XB-704 and comparing that data with the default prediction for the XB-70 in AeroDYNAMIC.
The XB-70 was already available as a test case, and after an analysis run it was possible to compare the results to the actual data. Figure 2.2 shows CD0 versus Mach number comparison between the AeroDYNAMIC output and flight data for the XB-
70. This plot was used to determine a percentage error of the AeroDYNAMIC results with the flight data. As a result, the CDO versus Mach number plot for the concepts were adjusted according this percentage difference between the two sets of XB-70 data. Since the other XB-70 AeroDYNAMIC results were adequate in comparison to the flight, there was no adjustment made to this data. Refer to figures 2.2 through 2.4 for the validation of the AeroDYNAMIC and figures 2.5 through 2.7 are the aerodynamic data for the four concepts. 2002 SLOB Works, a Virginia Tech AOE Design Team 19
0.025
XB-70 AeroDYNAMIC 0.02 Data
0.015
Cdo
0.01 XB-70 Flight Data
0.005
0 0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40 1.60 1.80 2.00 Mach
FIGURE 2.2: XB-70 CD0 versus Mach Number Comparison
0.7
0.6
0.5 XB-70 Flight Data
0.4
0.3 XB-70 AeroDYNAMIC Data 0.2 CL
0.1
0 0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18
-0.1
-0.2
-0.3 CD
FIGURE 2.3: XB-70 CL vs. CD Comparison at Mach 1.6
2002 SLOB Works, a Virginia Tech AOE Design Team 20
0.6
0.5 XB-70 AeroDYNAMIC Data
0.4
0.3
CL
0.2
XB-70 Flight Data 0.1
0 -2 3 8 13 18
-0.1 Alpha (deg)
FIGURE 2.4: XB- 70 Coefficient of Lift Curve Comparison at Mach 1.6
0.02
0.015
XB-70 SW-1
D0 SW-2 C 0.01 SW-3 SW-4
0.005
0 0.00 0.50 1.00 1.50 2.00 Mach Number
FIGURE 2.5: Adjusted CD0 vs. Mach Number for Concepts
2002 SLOB Works, a Virginia Tech AOE Design Team 21
1
0.8
0.6
XB-70 SW-1
L 0.4 C SW-2 SW-3 SW-4
0.2
0 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35
-0.2
CD
FIGURE 2.6: CL vs. CD for Concepts at Mach 1.6
0.9
0.7
0.5
L C L C
0.3
XB-70 SW-1 SW-2 SW-3 SW-4 0.1
-3 2 7 12 17
-0.1 α (degrees)
FIGURE 2.7: Lift Curves for Concepts at Mach 1.6
2002 SLOB Works, a Virginia Tech AOE Design Team 22
With all this data, the next step was to evaluate the concepts with AeroDYNAMIC. From the analysis it was then possible to compare each concept with one another. The major components that were compared between the four concepts were the wing loading values, take-off weights, center of gravity locations, and the amount of fuel burned.
2.1.3 Development of SW Excel Sizing Program
To help determine the optimum sizes for the chosen concepts, a sizing program was developed using Microsoft
Excel. The program takes into account an aircraft’s geometry and aerodynamic performance. These values were obtained from both the AeroDYNAMIC program as well as the original concept drawings.
The basis for this sizing program was obtained using Raymer’s Aircraft Design: A Conceptual Approach, 3rd
5 edition . Equations 2.1 through 2.10 detail the various weight ratios that were obtained in order to define the mission. Wo is the initial weight in these equations.
W 1 = 0.97 − 0.99 (warm-up, taxi, & takeoff) (2.1) W0
W 2 = 0.991− 0.007M − 0.01M 2 (climb & accelerate) (2.2) W1
−RC W V (L ) 3 = e D (cruise) (2.3) W2
−RC W V (L ) 4 = e D (dash in) (2.4) W3
W (W − 8700) 5 = 4 (weapons drop) (2.5) W4 W4
−RC W V (L ) 6 = e D (dash out) (2.6) W5
−RC W V (L ) 7 = e D (cruise) (2.7) W6
−EC W (L ) 8 = e D (loiter) (2.8) W7
W 9 = 0.990 − 0.995 (descend) (2.9) W8
2002 SLOB Works, a Virginia Tech AOE Design Team 23
W 10 = 0.992 − 0.997 (land) (2.10) W9
Table 2.1 is an example of the results obtained from these equations using data from one of the concept aircraft.
Using these ratios, it was possible determine the weight of each of the concepts.
TABLE 2.1: Example of Results
Segment
Warm up/TakeoffWwarmup/takeoff/Winitial 0.970
Climb – AccelerateWclimb/Wwarmup/takeoff 0.954 CruiseW /W 0.859 cruise1 climb Dash inWdashin/Wcruise1 0.891 Drop WeaponsWdrop/Wdashin 0.904 Dash outWdashout/Wdrop 0.891 CruiseWcruise2/Wdashout 0.859 LoiterWloiter/Wcruise2 0.941
DescendWdescend/Wloiter 0.993
LandWland/Wdescend 0.995
2.1.4 Cost Analysis Program
The cost analysis program was designed for the four final concepts that SLOB Works chose to compare. The foundation of the program was from Dr. Jan Roskam’s Airplane Design Series, called Part VIII: Airplane Cost Estimation:
Design, Development, Manufacturing, and Operating6. The purpose of this program was to obtain an accurate comparison in cost estimation, not necessarily a precise, final estimation. Refer to section 10 for the details of cost estimation method used to the determine the costs of the concepts. Tables 2.2 through 2.6 show the results of the initial cost analysis.
TABLE 2.2: Research, Development, Test, & Evaluation Cost (CRDTE)
Concept SW-1 Concept SW-2 Concept SW-3 Concept SW-4 Airframe Engr & Design
Caedr $ 1,204,758,987 $ 1,924,266,243 $ 1,259,118,957 $ 1,708,459,452 Develop Support & Testing Cdstr $ 402,369,134 $ 674,639,444 $ 422,452,782 $ 591,637,639 Flight Test Airplanes
Cftar $ 2,058,404,632 $ 3,241,701,052 $ 2,132,842,432 $ 2,867,835,179 Flight Test Operations C $ 343,741,283 $ 683,071,335 $ 366,724,106 $ 573,728,015 ftor
CRDTE $ 6,912,541,441 $ 11,247,720,816 $ 7,208,859,096 $ 9,899,414,283
2002 SLOB Works, a Virginia Tech AOE Design Team 24
TABLE 2.3: Acquisition Program Cost (CACQ) Concept SW-1 Concept SW-2 Concept SW-3 Concept SW-4 Manufacturing
CMAN $ 13,935,401,171 $ 15,199,508,655 $ 14,017,063,908 $ 14,804,392,227 Profit by Manufacturer
CPRO $ 1,393,540,117 $ 1,519,950,865 $ 1,401,706,391 $ 1,480,439,223
CACQ $ 18,034,048,574 $ 19,669,952,377 $ 18,139,729,763 $ 19,158,625,235
TABLE 2.4: Operating Program Cost (COPS)
Concept SW-1 Concept SW-2 Concept SW-3 Concept SW-4
COPS $ 342,553,246 $ 368,996,449 $ 312,743,072 $ 350,212,271
TABLE 2.5: Disposal Program Cost (CDISP)
Concept SW -1 Concept SW -2 Concept SW -3 Concept SW -4 P ro gram F u el, O il, & L u b rican ts
C POL $ 3,560 $ 3 ,5 6 0 $ 3 ,5 6 0 $ 3 ,5 6 0 Program D irect Personnel
C DIRPERS $ 3,169,957,500 $ 3,169,957,500 $ 3,169,957,500 $ 3,169,957,500 Program Indirect Personnel
C INDPERS $ 1,344,927,199 $ 841,946,295 $ 841,946,295 $ 841,946,295 C o n su m ab le M aterial
C CONMAT $ 282,018,750 $ 282,018,750 $ 282,018,750 $ 282,018,750 Spares
C SPARES $ 1,972,559,891 $ 1,234,854,566 $ 1,234,854,566 $ 1,234,854,566 Depots
C DEPOTS $ 1,882,898,078 $ 1,178,724,813 $ 1,178,724,813 $ 1,178,724,813 M iscellaneous
C MISC $ 313,816,346 $ 196,454,136 $ 196,454,136 $ 196,454,136
C OPS $ 8,966,181,325 $ 5,612,975,301 $ 5,612,975,301 $ 5,612,975,301 TABLE 2.6: Life Cycle Cost (LCC)
Concept SW-1 Concept SW-2 Concept SW-3 Concept SW-4 RDTE Cost
CRDTE $ 6,912,541,441 $ 11,247,720,816 $ 7,208,859,096 $ 9,899,414,283 Acuqisition Cost CACQ $ 18,034,048,574 $ 19,669,952,377 $ 18,139,729,763 $ 19,158,625,235 Operational Cost
COPS $ 8,966,181,325 $ 5,612,975,301 $ 5,612,975,301 $ 5,612,975,301 Disposal Cost CDISP $ 342,553,246 $ 368,996,449 $ 312,743,072 $ 350,212,271
Life Cycle Cost (LCC) $ 34,255,324,587 $ 36,899,644,942 $ 31,274,307,232 $ 35,021,227,090
Aircraft Estimated Price (AEP) $ 121,099,952 $ 150,085,792 $ 123,051,402 $ 141,058,444
2002 SLOB Works, a Virginia Tech AOE Design Team 25
2.2. Generation of Carpet Plots
To determine an optimized design of the concepts, carpet plots were created. For these plots, wing loadings between
50 lbs/ft2 and 100 lbs/ft2 using 10 lbs/ft2 increments were used. In addition, thrust-to-weight ratios between 0.35 and 0.6 in increments of 0.05 were used.
Equation 2.115 was used to find the empty weight fraction at these individual wing loadings and thrust to weight ratios.
C3 C4 We C1 C2 T Wo C5 = a + bW A M K (2.11) W o W S max vs o o
Then using equation 2.125 values for take-off gross weight were found.
We Wo = Wcrew +W fixedpayload +Wdroppedpayload +W fuel + Wo (2.12) Wo
The results of these two equations were used to establish the base for the carpet plots. It should be noted that these equations require an initial weight approximation. These approximations were obtained from AeroDYNAMIC and the SW
Excel program.
To obtain useful data from these carpet plots, constraints were established and plotted. The constraints chosen for examination were landing distance, takeoff distance, two specific power requirements, instantaneous turn rate, and cruise/dash requirements.
5 Equation 2.13 shows how wing loading was determined given a landing distance. For this equation Sa is the obstacle clearance distance and was assumed to be fifty feet. The landing distance requirement was set at 8,000 feet, but in order to represent this constraint on the plots, a landing distance of approximately 6,700 feet was used. Note that this equation is independent of thrust-to-weight ratio.
W 1 S = 80 + S (2.13) landing S σC a Lmax
To obtain the take-off distance constraint, Figure 5.4 of Raymer’s book was used5. From this plot, a take-off parameter (TOP) was chosen and applied to equation 2.145. For this constraint, a take-off distance of 8,000 feet was used. (W ) TOP = S (2.14) σC T LTO ()W
2002 SLOB Works, a Virginia Tech AOE Design Team 26
Two specific power requirements were examined as set forth in the RFP1. One was evaluated as Ps = 0 at 50,000 ft.,
Mach = 1.6, 2-g’s loading. The other was evaluated as Ps = 200 ft/s at 50,000 ft., Mach = 1.6, 1-g loading. Equation 2.155 relates Ps, wing loading and thrust-to-weight ratios.
qC T Do 2 K W Ps = V − − n (2.15) W W q S S
The instantaneous turn requirement set forth in the RFP was a turn rate of 8 degrees/second at Mach = 1.9, at 15,000 feet. Equation 2.165 shows how wing loading was determined given this requirement. Note that this constraint is independent of the thrust-to-weight ratio.
2 g (qCL /(W / S)) −1 ϕ = (2.16) V
Finally, the cruise/dash constraint was related to the thrust-to-weight ratio and wing loading by using equation 2.175.
Since the aircraft would be in level flight conditions at this constraint, thrust is equal to drag and lift is equal to weight.
Therefore, lift-to-drag ratio (L/D) is equal to the inverse of the thrust-to-weight ratio. Note that the cruise conditions were calculated at an altitude of 38,000 feet. This is the optimum altitude at which the aircraft will cruise. Using these conditions, the cruise constraint does not appear on these plots due to the plot area.
L 1 1 = = (2.17) T qC D Do W ()W ()W + S S qπAe
To place these constraints properly, the values of wing loading and thrust-to-weight ratios had to be normalized with respect to the take-off thrust and weight conditions. Equations 2.18 and 2.195 detail how this was accomplished.
T T Ttakeoff W = int erest (2.18) W takeoff W po int of int erst Tint erest Wtakeoff
W Wtakeoff W = (2.19) S takeoff Winterest S interest
On these plots, the blue lines signify the base carpet plot. The red lines signify a constraint set forth by the RFP. The green dot represents the point to which the concept needs to be sized.
2002 SLOB Works, a Virginia Tech AOE Design Team 27
When the constraints were added to each concept’s base carpet, it was determined that the specific requirement of
200 ft/s at 1-g, as well as the instantaneous turn requirement were the major constraints. At this stage of the design process, it was decided that all concepts could be augmented later with afterburners in order to meet the specific power requirement. As the afterburners would not be used on a typical design mission, they would not factor into the cruising and dashing fuel consumptions.
Using this analysis, it is seen that none of the four proposed concepts meet the constraints set forth by these carpet plots. SW-2 was the closest to fulfilling the design requirements. When the errors from AeroDYNAMIC are taken into account, all of the concepts begin to approach the desired design area. The thrust-to-weight ratios of these concepts are still low. SW-2 has the greatest capability of engine upgrading. It is therefore felt that this concept has the best chance of succeeding and thus increased its chance of selection. The following plots show the carpet plots for each concept and their design points.
176000 2g Ps = 0 ft/s Constraint
T/W = .6 W/S = 50
T/W = .55 174000 T/W = .5
T/W = .45 W/S = 60 T/W = .4 172000 1 g Ps = 200 ft/s Constraint T/W = .35 W/S = 70
170000 W/S = 80
Instantaneous Turn Constraint W/S = 90 168000 W/S = 100 Takeoff Constraint TOGW (lbs.)
166000
164000 Landing Constraint
Dash Constraint
162000
FIGURE 2.8: Carpet Plot of SW-1
2002 SLOB Works, a Virginia Tech AOE Design Team 28
268000
T/W = .6 266000 W/S = 50 T/W = .55 2g Ps = 0 ft/s Constraint 1 g Ps = 200 ft/s Constraint T/W = .5 264000 T/W = .45 W/S = 60 T/W = .4 262000 T/W = .35 W/S = 70 260000
W/S = 80 258000 Instantaneous Turn Constraint W/S = 90 TOGW (lbs.)
256000 W/S = 100 Takeoff Constraint 254000
252000
Dash Constraint Landing Constraint 250000
FIGURE 2.9: Carpet Plot of SW-2
186000 T/W = .6 W/S = 50 T/W = .55 2g Ps = 0 ft/s Constraint 184000 T/W = .5
T/W = .45 W/S = 60 182000 T/W = .4 1 g Ps = 200 ft/s Constraint T/W = .35 W/S = 70 180000
W/S = 80 178000 W/S = 90
W /S = 100
TOGW (lbs.) 176000 Takeoff Constraint 174000 Instantaneous Turn Constraint
172000 Dash Constraint Landing Constraint
170000
FIGURE 2.10: Carpet Plot of SW-3
2002 SLOB Works, a Virginia Tech AOE Design Team 29
210000
T/W = .6 2g Ps = 0 ft/s Constraint W/S = 50 208000 T/W = .55 T/W = .5 206000 T/W = .45 W/S = 60 T/W = .4 1 g Ps = 200 ft/s Constraint 204000 T/W = .35 W/S = 70
202000 W/S = 80 Instantaneous Turn Constraint
200000 W/S = 90
TOGW (lbs.) W/S = 100 198000 Takeoff Constraint
196000 Dash Constraint Landing Constraint 194000
192000
FIGURE 2.11: Carpet Plot of SW-4
2.3. Concept Selection Process
Prior to finalizing all concepts, SLOB Works decided upon three selection processes for determining the final aircraft design. The first process was constructed using the foundations of the original comparative study. The second process used information generated from the cost analysis program. Finally, the third process selection was founded off the
“stop-light” chart. Below is a detailed description of each process and its importance in determining the final concept design.
The three processes were determined from the RFP: total performance, technical data, and cost analysis.
2.3.1. Concept Design Matrix
During the aircraft comparative study, a comparison chart was created to analyze general characteristics that were driving factors in the RFP. From the created matrix, SLOB Works created a design comparative study to analyze the same characteristics for all four preliminary concepts. From interpreting the aircraft comparative study, SLOB Works established approximations for some design drivers, these are found under the column heading “Approximations”. The values listed, in the other columns on the next page in table 2.7, were calculated and determined from either the computer programs described in the above sections, or from the equations found in Raymer’s5 and Roskam’s8 textbooks. The importance of this matrix was to compare technical data that could be achieved by each preliminary aircraft.
2002 SLOB Works, a Virginia Tech AOE Design Team 30
TABLE 2.7: Concept Design Matrix
SW-1 SW-2 SW-3 SW-4 Rating of Approximations Data Ranking Rating Data Ranking Rating Data Ranking Rating Data Ranking Rating Importance Cr ui se Ma ch 1.6 -- -- 1.6 -- -- 1.6 -- -- 1.6 -- -- Mach 1.6 -- Cruising Altitdue (ft) 54,258 -- -- 64,036 -- -- 67,948 -- -- 86,045 -- -- 38,000 -- Dash Mach 1.6 -- -- 1.6 -- -- 1.6 -- -- 1.6 -- -- Mach 1.6 -- Dash Altitude (ft) 50,000 -- -- 50,000 -- -- 50,000 -- -- 50,000 -- -- 50,000 -- Range (nm) 3,500 -- -- 3,500 -- -- 3,500 -- -- 3,500 -- -- 3,500 -- Take Off Gross Wt (lbs) 119,091 1 5 226,916 4 20 128,000 2 10 200,000 3 15 180,000 5 Fuel Weight 69,438 2 10 113,922 4 20 58,900 1 5 98,606 3 15 75,000 5 Payload Wt (lbs) 8,700 -- -- 8,700 -- -- 8,700 -- -- 8,700 -- -- 8,700 -- Wing Span (ft) 91.23 -- -- 90.65 -- -- 92.54 -- -- 108 -- -- 80 -- Wing Area (ft 2 ) 2,291 -- -- 4,650 -- -- 1,620 -- -- 2,565 -- -- 2,000 -- Wing Sweep (deg) 55 -- -- 60 -- -- 67 -- -- 2 -- -- 60 -- Wing Thickness (ft) Aspect Ratio 3.536 -- -- 2.15 -- -- 1.626 -- -- 4.547 -- -- 3.2 -- Wing Planform Type Propulsion System Total Installed Thrust (lbs) 52,760 -- -- 105,520 -- -- 52,760 -- -- 79,140 -- -- 60,000 -- Control System High Lift System 2 Bays 2 Bays 2 Bays 2 Bays Weapon Bay Configuration (Missle - -- -- (Missle - -- -- (Missle - -- -- (Missle ------Bomb Config) Bomb Config) Bo mb Confi g) Bomb Config) Aircraft Length (ft) 108.02 -- -- 143.91 -- -- 107 -- -- 117 -- -- 130 -- Take Off Distance 3,790 -- -- 3,700 -- -- 6,724 -- -- 7,287 -- -- 8,000 -- Fineness Ratio 0.845 3 3 0.630 1 1 0.865 2 2 0.923 4 4 1 L / D max @ M = 1.5 6.55 1 4 5.86 3 12 6.34 2 8 5.18 4 16 4
C d0 @ M= 0.8 8 0.0114 4 16 0.0104 2 8 0.0100 1 4 0.0113 3 12 4
C d0 @ M= 1.5 0.0168 2 8 0.0204 3 12 0.0159 1 4 0.0273 4 16 4 W / S @ MTOGW 51.982 3 9 48.799 2 6 79.012 1 3 77.973 4 12 90 3 T / W @ MTOGW 0.443 3 9 0.465 4 12 0.412 2 6 0.396 1 3 0.3 3 Wf / Wt 0.583 -- -- 0.502 -- -- 0.460 -- -- 0.493 -- -- 0.436 -- Wp / Wf 0.125 -- -- 0.076 -- -- 0.148 -- -- 0.088 -- -- 0.379 -- (Wp+Wf) / Wt 0.656 -- -- 0.540 -- -- 0.528 -- -- 0.537 -- -- 0.537 -- Special Operational System Co st ($ ) 121.1M 1 5 150.1M 4 20 123.1M 2 10 141.1M 3 15 150M 5 Average 8 11.38 5.25 11.63
As one can see, there are three different columns for each concept. The first column, “Data” is a detailed, technical description of each airplane in the preliminary design phase. It should be noted that not all areas are filled in for the second column, “Ranking”. This datum, from the first column, is only there to compare and understand some of the requirements needed to complete the specified mission found in the RFP or to better understand the concept. The areas that have a number are there to represent the best concept for that specific component of the aircraft. The numbers range from one to four, one being the best design and four being the worst. Finally, the third column, “Rating”, is a multiple of the “Ranking” for each design times the “Rating of Importance” for each column. This is the column that was added and then averaged to find the best design for technical data (lowest score was the best).
The “Rating of Importance” was weighed according to what SLOB Works thought was the most crucial design element. The most critical element received a rating of importance of “5”, while the lowest rating of importance received a rating of “1”. The others not rated were not important characteristics for evaluating and comparing for the final selection.
There are some key elements listed in the design comparative study that must be noted. Since the weights were calculated through the program AeroDYNAMIC, the weights are relatively inaccurate. When comparing the design results to 2002 SLOB Works, a Virginia Tech AOE Design Team 31
the calculated XB-70 results, all weights were low by a factor of approximately 14%. Thus, all weights will increase and all wing loadings will increase as well. However, since this is the preliminary phase of the design, SLOB Works was more concerned about relative and comparative data than exact data. For the final concept chosen, these errors will be fixed and the appropriate measures will be taken.
After averaging the concepts’ ratings, the conclusion on the technical and performance data obtained so far, was that
Concept SW-3 was the winner, followed by SW-1, SW-2, and then finally SW-4. This selection process was not as important as the risk management since it only accounted for a small amount of the total aircraft design.
2.3.2. Risk Management Matrix
The Risk Management Matrix, found on the next page in table 2.8, was designed as a total comparison of all four preliminary aircraft. The categories include: aerodynamics, stability & control, structures, propulsion systems, stealth characteristics, maintenance, manufacturability, and design. All eight categories will be explained in great detail below. The rating system included a “stop-light” chart, where red was considered difficult or hard to accomplish, while green was considered easy or can be accomplished with great ease. Anywhere that yellow appears implies a cautious category that could suggest a problem.
Most of the aerodynamics associated with the TABLE 2.8: Risk Management Matrix preliminary design phase came from the programs AIRCRAFT CONCEPTS SW-1 SW-2 SW-3 SW-4 mentioned above. With a lot of emphasis placed on the Rank Rank Rank Rank Aerodynamics 1 231 aerodynamic and performance characteristics through the Stability & Control 2 1 2 1
RFP, a more detailed version of the technical data is Structures 1 2 2 3 Propulsion Systems mentioned and described in the next section. However, Inlets 3 1 2 1 with a total aircraft comparison, aerodynamics was Supercruising 1 113 analyzed and rated. Concept SW-1 was rated as one of Stealth 2 113 the highest due to its low weight and swept wings. The Maintenance 3 222 Manufacturability 1 2 1 2 swept wings were determined at an angle of 55 degrees so Design 2221 that it would meet both the takeoff and landing Average 1.778 1.556 2.000 1.667 Ranking 3142 requirements while also trying to maintain low coefficients of lift at zero drag at higher Mach numbers. Concept SW-4 also received an excellent rating due to the possibility of the benefits of laminar flow control. Both of these concepts are relatively easy to analyze due to the simplicity 2002 SLOB Works, a Virginia Tech AOE Design Team 32
of design with lifting surfaces and drag. The others were rated relative to the best design in this category, thus Concept SW-3 was the poorest.
The second category in the management matrix was stability and control. For all four preliminary designs, the c.g. locations were acceptable before and after the completion of the mission. Next, since canards may help reduce aerodynamic center shift at supersonic flight, a higher rating was given to those with canards. With standard configurations, vortices disturb and decrease the effectiveness of conventional elevators. However, using these wing devices there is little downwash over the wing and the disturbance becomes smaller. For this reason, SLOB Works preliminary approximation for stability & control had Concept SW-2 and Concept SW-4 as the best designs. The other two designs closely followed with an equal rating.
Next, SLOB Works approximated the difficulty with the structural design of each aircraft. With Concept SW-1 having the simplest design, and the swept wings, the structures were, by far, the simplest. Concept SW-2 and Concept SW-3 were more difficult due a blended-wing body and a faceted body, respectively. Finally, the poorest rating was given to
Concept SW-4 because of its very small leading edge sweep. Although there is a 2-degree sweep, at supersonic flight, the divergence and aero-elasticity will be too difficult to overcome.
Within the propulsion category, SLOB Works felt it necessary to have two separate sub-categories. The first deals with the inlet geometry for each design. Concept SW-2 and Concept SW-4 were rated superior due to the ability of the inlet to change the flow from supersonic to subsonic. Concept SW-2’s inlet was created from the XB-70’s mixed compression inlet design. This gives a high efficiency over a wide Mach number range. Concept SW-4 used a diffuser length for optimum efficiency of eight times the fan-face diameter. Longer lengths have internal friction loss, as well as, a weight penalty. Shorter lengths, less than four times the diameter, produce some flow separation. Concept SW-3 was a little more difficult due to the inlet geometry while Concept SW-1 was the most difficult because of the lack of asymmetrical spiked inlets. These inlets are normally used for short diffuser lengths at high Mach numbers.
The second sub-category deals with supercruising. All aircraft would meet the RFP requirements mentioned in
Chapter 1 except for Concept SW-3. After researching comparable flying wings in the aircraft comparative study, it would be difficult to have Concept SW-3 cruise supersonically without afterburners because of high drag from the poor fineness ratio.
The RFP gives a maximum RCS of 0.5 m2 for the front of the aircraft. After reanalyzing Concept SW-4, the required stealth capabilities will be extremely difficult to accomplish. The wing causes large radar reflection due to the
2002 SLOB Works, a Virginia Tech AOE Design Team 33
almost perpendicular leading edge sweep. The blended body and faceted body of Concepts SW-2 and SW-3, respectively, as well as, the canted horizontal tails on both designs decrease the front RCS below that of Concept SW-1.
Another category of interest for the customer accounts for maintenance. Maintenance includes engine location, height of payloads, and even hatch locations. Concept SW-1 is considered the worst due to the engine location. With the engines located on top of the wing, any maintenance will be more difficult than the others. Concepts SW-2, SW-3, and SW-4 are relatively equal in maintenance capabilities; however, all concepts still lack simplicity. Thus, these three were rated as mediocrity.
Manufacturability includes the simplicity of design and the compatibility of interchangeable and removable parts.
Most of our analysis was geared toward the simplicity of the CAD drawings. The two smaller aircraft, namely Concept SW-
1 and SW-3 were simple with respect to the other designs. Concepts SW-2 and SW-4 were more difficult due to the blended wing design. However, SW-2 has a removable pod, which is very close to the ground to make interchangeable parts a swift operation per aircraft. The simple wing design, as well as, the low fuselage gives an edge to the fourth concept, which keeps all aircraft above a “most difficult” rating.
Finally, the design category was created for each aircraft’s originality. Since the RFP has a credited point value for originality, SLOB Works thought it necessary to compare each design. Since Concept SW-4 was the only “cutting-edge” design, it was rated the highest. The other concepts were not as original, so an equal, lower rating was assigned to each.
With all ratings completed, the final analysis provided Concept SW-2 the winner, followed by Concept SW-4, then
SW-1, and finally SW-3. Since the matrix compared all aspects of the aircraft, SLOB Works made this selection process the most important comparison of the three.
2.3.3. Cost Analysis Matrix
The cost analysis matrix was selected to stand alone in the selection process due to the importance of government budgeting. SLOB Works wanted to rate the cost analysis as much as the design comparative study, however, with all aircraft matching or bettering the RFP, SLOB Works decided to weigh this selection process the least. The two smallest aircraft were estimated to cost the least with the two larger designs were estimated to cost about 20 to 30 million dollars more. Table 2.9 shows the final analysis and ranking of the four aircraft in the selection process.
TABLE 2.9 Cost Analysis Aircraft Estimated Price (AEP) $ 121,099,952 $ 150,085,792 $ 123,051,402 $ 141,058,444
Ranking 1423
2002 SLOB Works, a Virginia Tech AOE Design Team 34
2.4. Final Analysis & Selection Process
After analyzing the three selection processes, SLOB Works had to determine how to rate the importance level of each process. Since the Risk Management Matrix measured close to all aspects of each design, the outcome of that matrix were given the highest rating of “3”. Then, the design comparative study was given the next highest rating of “2”. Finally, as mentioned above, originally, it was determined to also give the cost analysis outcome a rating of “2”, however, with all aircraft meeting the RFP, the rating dropped to “1”. Thus, the final decision matrix is shown in table 2.10 and it should be noted that Concept SW-2 was determined to be the best aircraft design. Concept SW-1 was next, followed by SW-3, and finally SW-4. TABLE 2.10: Final Decision Selection SLOB Works final design concept, AIRCRAFT CONCEPTS SW-1 SW-2 SW-3 SW-4 Concept SW-2, will know be known as SW- Rank Rank Rank Rank
Ghost. This aircraft was further investigated Trade Study/Technical Data 2314 in the preliminary design phase for more Cost Analysis 1423 precise data, in all aspects of the design. Risk Analysis 3142 (3R+2T+C) Average = / 6 2.333 2.167 2.667 2.833
Final Rankings 2 1 3 4
2.5. Aircraft Design & Layout
After the conceptual design phase, concept SW-2 was selected to continue into the preliminary design phase. As problems were encountered during the preliminary design phase, the aircraft was optimized. Figure 2.12 is a chart showing the different stages of the aircrafts design with a short description of what was changed.
The final version of the aircraft is a blended wing body design, with engines and weapons bays located under the main fuselage section of the aircraft. The concept is controlled by canards, ailerons, and ruddevators. Leading edge slats and roughly quarter span flaps provide the aircraft with high lift. Figures 2.13 through 2.16 show the top, side, front, and bottom views of the selected concept
2002 SLOB Works, a Virginia Tech AOE Design Team 35
FIGURE 2.12: Evolution of Aircraft
Concept SW-2 Refinement 1 - Overall aircraft refinement
Refinement 3 – Inlets & engines Refinement 2 – Wing root extended, detailed, landing gear finalized V-tails split, aircraft area ruled
Final, SW-Ghost
2002 SLOB Works, a Virginia Tech AOE Design Team 36
Figure 2.13 Top view of SW-Ghost
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Figure 2.14 Bottom view of SW-Ghost
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Figure 2.15 Side view of SW-Ghost
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Figure 2.16 Front view of SW-Ghost
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3. Aerodynamics
Predicting the aerodynamics for SLOB Works concept encompasses a variety of tasks. These tasks include planform selection, airfoil selection, area distributions and lift and drag analysis. A large part of the planform and airfoil selection was based on current aircraft characteristics which came close to meeting the RFP’s requirements of supercruising and range, in particular the concord and XB-70.
3.1. Planform and Airfoil Selection
The planform selected can be classified as a delta wing with modifications (fig. 3.1). The first deviation from a conventional delta wing is that the tips of the wings are clipped. In doing this, a portion of the wing that contributes a negligible amount to the lift is eliminated while also decreasing the span. Next the outboard section of the wing was unswept. This is recommended to reduce the aerodynamic center shift between subsonic and supersonic flight. The last modification to a classic delta wing is that area has been added to fill in the inboard trailing edge of the wing. The most influential benefit of this modification is that the trailing edge flaps become more effective. Other advantages of filling the inboard trailing edge include helping the plane with subsonic pitch-up, making the wing more efficient structurally and an increase in the planform area, which reduces wing loading7.
FIGURE 3.1: SLOB Works Ghost semi-planform
After the planform shape was chosen, the details of the wing needed to be specified. The leading edge sweep is varied from 60 degrees on the inboard portion of the wing to 54 degrees on the outboard portion. This sweep is relatively high for a cruise Mach number of 1.6. Historically a sweep of 50 degrees is used for a design Mach number in this range5.
However, SLOB Works found that to reduce shock formation and increase roll stability the sweep needed to be greater than
50 degrees. This improvement in stability is due to a natural dihedral effect caused by greater sweep. To prevent excess
2002 SLOB Works, a Virginia Tech AOE Design Team 41
stability from such a high sweep, this wing has a zero dihedral angle. The outboard section the wing sweep was changed simply to increase the wing area. In this phase of the design the wings have no twist to keep the drag at the cruise Mach number as low as possible. Lastly, SLOB Works chose to have sharp wing tips to reduce the induced drag as much as possible. This reduction of induced drag is caused by a difficulty in the flow rolling around the tips of the wings during a positive angle of attack.
The airfoils that were selected to use in the SLOB Works concept were NACA 6 series supersonic airfoils. The airfoils chosen are symmetrical and have low thickness to cord ratios. A NACA 64-004 model is being used for the planform. The decision to use a 4% thick airfoil was made from both a historical trend line5 and fuel storing needs. Since a large percentage of fuel is being stored in the wings this was as thin of an airfoil which could be used. Using the thinnest airfoil possible helped reduce the wave drag of the aircraft. Both the canards and vertical tail use NACA 64-006 airfoils.
This is a 6% thick airfoil and was chosen over the NACA 64-004 for structural purposes. The additional thickness will allow for larger struts to attach to the body of the aircraft.
3.2. Lift Analysis
To complete the subsonic lift analysis, SLOB Works used the program Tornado8 and equations from Raymer5.
Tornado is a MATLAB code which uses vortex panel method to determine lift coefficients. Only the planform, canards and vertical tail were entered when running this program. The slope of the lift coefficient line is 2.698 per radian. These results were compared with the lift curve slope of 2.531 per radian obtain by equations found in Raymer5. The lift curve slope values used in all aerodynamic calculations were acquired through Raymer. The CLmax for the SLOB Works concept is 0.884 without the use of high lift devices. Table 3.1 displays various aerodynamic parameters at the different mission segments.
TABLE 3.1: Key aerodynamic parameters for missions segments
SEGMENT Mach # Cd Cdo K Cl L/D (max) Take off 0.255 0.0428 0.009 0.1845 0.428 12.346 Climb 0.8 0.0190 0.00934 0.1817 0.23 12.136 Cruise out 1.6 0.0230 0.0136 0.345 0.165 7.299 Dash out 1.6 0.0311 0.0141 0.345 0.222 7.168 Dash in 1.6 0.0258 0.0141 0.345 0.184 7.168 Cruise in 1.6 0.0228 0.0141 0.345 0.159 7.168 Loiter 0.4 0.0136 0.00847 0.1841 0.168 12.658 Landing 0.278 0.04269 0.00889 0.1845 0.428 12.346
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3.3. Drag Analysis
The next step in completing the aerodynamics of SLOB Works concept was to determine the drag built up at various mach numbers. Skin Friction/Form Factor Drag2 and AWAVE2 programs were used to obtain the different types of drag, which compose the parasite drag. The Skin Friction/Form Factor Drag program found the skin friction and form factor coefficients. This software calculates these coefficients based on the wetted areas of the aircraft. SLOB Works concept was broken in to six pieces to achieve an accurate estimation of the total wetted area of the plane. The six pieces were the nose cone, fuselage section 1, fuselage section 2, wing planform, the canards and the canted tails that yielded a wetted area of
9,070 square feet. At a cruise Mach number of 1.6 and an altitude of 50k feet the drag coefficient due to friction and form drag is 0.0063. The next piece of the total drag to analyze was the wave drag, which is due to the formation of shockwaves.
The program AWAVE was used in this analysis. It generates its solutions based on the volume distribution the aircraft occupies. At the specified cruise Mach 1.6 this value was found to be 0.00589. This drag count is very similar when compared to the value of 0.00601 produced by equation 3.1.
4.5π A C = ( max )2 E (0.74 + 0.37 cos Λ )[1 − 0.3 M − M ] Dwave S l WD LE CDo max (3.1)
1 where, M CDo max = 0.2 (3.2) cos Λ LE
and EWD ≈ 1.5 for a blended-delta-wing aircraft with a smooth area distribution. The wave drag coefficient used in the aerodynamic analysis was obtained by AWAVE since the whole geometry of the plane is inputted instead of only Amax, l and
ΛLE as in equation 3.1.
The parasite-drag buildup is also composed of miscellaneous drag and drag due to leaks and protuberance. The miscellaneous drag takes into account any antennas, doors, lights, etc. A value of 0.0007 was added to the total drag of the aircraft for the range of mach numbers it encounters. Historically, drag due to leaks and protuberance are estimated to be between 2-5% of the total parasite drag for bombers. SLOB Works estimated this percentage to be 3% since the aircraft uses a blended wing-body design, thus eliminating many corners and edges where plates meet. Figure 3.2 is a graph of the parasite drag buildup for SLOB Works concept at an altitude of 50,000 feet. The parasite drag value at a Mach number 1.6 was found to be 0.0136.
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Parasite-Drag Buildup @ 50k feet
0.02 Leaks and 0.018 protuberances 0.016
0.014 Miscellaneous
0.012 Wave drag
0.01 Cdo
0.008
0.006 Form drag 0.004 Skin friction drag
0.002
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Mach Number
FIGURE 3.2: Parasite-drag buildup of SLOB Works concept at altitude 50,000 feet
The next step in finding the total drag of the concept was to obtain values of induced drag for different lift coefficients at the designated Mach cruise number. To obtain this the minimum total drag the parasite drag is added to the induced drag. The induced drag for an uncambered airfoil is found using equation 3.3.
C = C + KC 2 (3.3) D DO L
In equation 3.3 the K for subsonic flight was found using equation 3.4,
K = SK100 + (1 − S)K0 (3.4) where S is the leading edge suction5. SLOB Works used a leading edge suction of 89% for Mach numbers of 0.2 to the drag divergence Mach number MDD of 0.92. The K used in supersonic flight was found using a supersonic aerodynamics code
Arrow5. This software takes into account the leading edge sweep angle and notch ratio of the wing at a user specified Mach number. It outputs two K values, K0 and K100. K0 is the value of K with 0% leading edge suction and K100 is the value with
100% leading edge suction. These two values were then averaged (50% LE suction) to give K equal to 0.345. Table 3.1 gives K values at different segments of the mission. Figure 3.3 displays the polar drag for a Mach number of 1.6.
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Drag Polar @ M=1.6 and 50k ft
0.4
0.3 L/Dmax = 7.30
0.2 Lift Coefficient
0.1
0 0 0.01 0.02 0.03 0.04 0.05 0.06 Drag Coefficient
FIGURE 3.3: Drag polar of SLOB Works concept at cruise Mach 1.6
The next characteristic of the aircraft to be determined was the maximum lift to drag ratio, L/Dmax, labeled on figure
3.3. This value occurs at the point on the curve in which a tangent line can be drawn to the origin of the x-y axis. L/Dmax for cruise is calculated to be 7.30 for the SLOB Works concept. Figure 3.4 shows L/D values at various lift coefficients. Of particular interest is the cruise lift to drag ratio. The L/D at which the aircraft flies varies to achieve the greatest range. The values of L/D that are used for the supersonic flight vary from 7.29 to 6.92. Table 3.1 displays L/D ratios at different mission segments.
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Lift / Drag @ Mach 1.6
8
7
6
5 L/Dmax 4 Lift / Drag 3
2
1
0 0 0.05 0.1 0.15 0.2 0.25 0.3 Lift Coefficient
FIGURE 3.4: Lift to drag ratio at cruise conditions
3.4. Aircraft Geometry
The area distribution of SLOB Works concepts was analyzed using the AWAVE software program mentioned earlier in this chapter. Figure 3.5 shows the area due to each piece of the aircraft as well as the total area distribution. The area distribution was taken at a Mach number of 1.01 in order to non-dimensionalize the theta cuts. From this graph the
2 maximum area Amax, and the corresponding location can be found. Amax is 157.8 ft at a location of 64% of the total length of the aircraft. The finesse ratio can also be obtained from figure 3.5 using equation 3.55.
l f = (3.5) 4Amax π
This parameter is a ratio of the aircraft length to maximum area and provides a value for the sleekness of the aircraft. The fineness ratio for SLOB Works concept is 10.66, which is comparable to the Concorde’s finesse ratio of 12.8. The total volume of the aircraft was calculated by Awave to be 11,200 ft3.
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Area Distribution @ M=1.01
180
160 Total Fuselage 140 Wing Pod 120 V. Tail Canard 100
80 Area (ft^2)
60
40
20
0 0 20 40 60 80 100 120 140 160 x Location (ft)
FIGURE 3.5: Area distribution at Mach 1.01
3.5. High Lift Devices
The final aspect of the aerodynamic analysis is the use of high lift devices. To ensure the SLOB Works concept obtains the greatest lift coefficient possible slotted flaps and slotted leading edge flaps (slat) are utilized. Slotted flaps were chosen because they help to increase lift while reducing the drag. This drag reduction is accomplished by allowing the high- pressure air to exit through the slot between the wing and flap which reduces separation. By using slotted flaps in their
O maximum deflection configuration of 40 degrees, CLmax is increased by 0.157 and ∆αOL is –2.5 . Slats are being used in our design to give the wing camber which increases the lift it can produce. The CLmax value can be increased 0.031 and the zero
O lift angle of attack can be decreased by –2.2 . After the addition of the high lift devices the CLmax is 1.07 and the zero lift angle of attack is –4.7O instead of zero degrees.
2002 SLOB Works, a Virginia Tech AOE Design Team 47
4. Structures and Materials
4.1. Materials
The direct link between the structural performance and the cost of the aircraft is also integrated with the material properties of every structural member. When dealing with supersonic aircraft it is important to pay close attention to the following:
-High altitude operations -Temperatures from –58OF to 320OF for main structure -Thermal cycling under moisture and radiation impact -12,000 hour service life on all parts -High engine temperatures
The aircraft will have an integrated configuration of aluminum, steel, titanium, and magnesium materials9. Table 4.1 shows some specific materials that were chosen because of their performance characteristics. Along with these, some simple and complex composites were utilized.
TABLE 4.1: Material Properties
Material General Density Bulk Modulus Endurance Limit Fatigue Strength Modulus of Elasticity Thermal Expansion Designation lb/in^3 10^6 psi ksi ksi 10^6 psi 10^-6/F Al alloy: 2024-TO 0.01 9.86-10.88 5.67-6.24 5.33-6.61 10.59-11.17 12.5-13.17 S steel:AISI 410 0.28 20.31-23.64 37.27-48.01 28.5-59.11 27.56-30.46 5.0-6.11 MMC: Cerme-Ti 0.16 17.4-18.13 45.69-48.73 37.96-57.83 15.95-17.4 4.37-4.71 Mg Alloy: AZ31 0.064 5.22-5.94 15.23-16.68 14.94-16.96 6.38-6.67 14.44-14.50
Aluminum has always proven to be a good resource in aircraft production. It was chosen because of its reasonable yield and ultimate tensile strength, as well as its good machinability and surface finish. In addition, the density helps lower weight9. Although the density of steel is the highest, its large tensile and ultimate yield stresses are important for areas of large loads. Sacrificing a higher weight is necessary for structural integrity. A large portion of the plane will be made from titanium and titanium composites. Titanium holds the highest yield and ultimate tensile stresses even with a very low density. This enables the plane to be both strong and light. Magnesium is also used for its extremely low density. However, the ultimate and yield tensile stresses are among the lowest of the chosen materials so placement of this material is important to avoid high loads.
Titanium was not used throughout the entire plane, despite its superior characteristics, because of the cost. Titanium is harder and takes more time to machine11. As with all materials, the cost is a driving factor and certain materials were placed in certain locations because of the relationship between cost and structural ability. Figure 4.1 shows a diagram of our design showing where each material is going.
2002 SLOB Works, a Virginia Tech AOE Design Team 48
FIGURE 4.1: Material Breakdown of Aircraft
The main bulkheads in the SW-Ghost will be made of titanium castings. As in the F-22 Raptor, the castings will be welded together to eliminate mechanical joints and ultimately be stronger and more cost effective12. Most skin of the aircraft will be an aluminum lithium composite. Areas that are RCS critical will have sheets of special radar absorbent material.
4.2. Structures
Every airplane’s structural and material components will experience heating, cooling, bending, twisting, shaking, tearing, and breaking. It is important to design an integrated system that maximizes the performance of the structural and material composition and location.
The primary structure consists of 8 longerons running from the rear of the cockpit to the midsection of the canted tails, as shown in figure 4.2. Five longerons form a skeleton on the upper half of the plane, and the other three provide support below the wings of the aircraft. Bulkheads are positioned in high load areas5. All the main bulkheads, along with their location and purpose, are shown in figure 4.3. In addition, there are bulkheads that join to the spars in the wings. All the
2002 SLOB Works, a Virginia Tech AOE Design Team 49
bulkheads connect to longerons to help distribute loads. All bulkheads are not the same size; they are sized according to their location and purpose. The all-moveable canard has smaller structural elements, as shown in figure 4.2. At the wing root the chord is about 64 feet. The first spar is located at 14% of the root chord and the last spar is at 87% of the chord. Typically, the last spar is located at 65-75% of the chord, but because of the delta wing design this shifts the location of the last spar back8. Spars are placed about 7 feet apart in the wing because of the extreme length at the root and are sized accordingly.
This spacing was determined by comparisons with the Concorde. The root chord of the Concorde is larger than the proposed aircraft and has spars spaced even further apart. The spars in the last quarter of the wing have a negative sweep to help accommodate the diamond wing design. In the belly of the aircraft is an arrangement of structural elements that provide support to the bomb/missile bay pod, inlets, landing gear, and engine pods, as seen in figure 4.4. The missiles and bombs are in a pod that can be easily removed from the aircraft. The engines are in dual pods that allow the removal of two or all four engines without losing structural integrity. Extra support is placed around the four engines to help alleviate stress due to torque. A boom extends off the back of the aircraft to help support the weight and forces seen by the canted tails10.
The V-n diagram, seen in figure 4.5, shows some of the boundary limits put on the aircraft. The plane has been designed to withstand g-loadings ranging from +10.5 to –4.5. With a factor of safety of 1.5 on design ultimate loads, the design limit load factors range from +7.0 to –3.0 g’s. Also shown are the corner point, dive, and cruise velocities. The left boundaries of the graph are the stall lines. The right boundary is the maximum dive speed that can be reached without structural damage to the aircraft. With such high limits to the g-loading wind gusts are not a factor. The wind gust lines are a function of many things including gust speeds, equivalent airspeeds, and wing loading5.
FIGURE 4.2: Structures Top View/ Major Components
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FIGURE 4.3: Structures Side View/ Main Bulkheads
FIGURE 4.4: Structures Bottom View/ Removable Pods
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FIGURE 4.5: V-n Diagram
As used in the F-22 Raptor, the wings have a “sine-wave” design that makes them stronger and lighter than the traditional I-beam12. Figure 4.6 shows a sketch of this design. Holes are drilled in the ribs and spars, helping to reduce weight while not affecting the structural integrity.
FIGURE 4.6: “Sine Wave” Spar Design
2002 SLOB Works, a Virginia Tech AOE Design Team 52
5. Stability and Control The control surfaces on the SW-Ghost are very unique because they include an all-moveable canard and ruddervators on both of the canted tails, and the wings have leading edge slats, single-slotted flaps, and ailerons. The all- moveable canard, which acts as a high lifting surface and also as a horizontal stabilizer, has a positive and negative deflection of 40 degrees. The canard has an area of 198 ft2. The tails are canted outward for RCS purposes, thus the tail was made larger to compensate for the loss of lateral control. Since each tail is deflected 42 degrees from the vertical, the ruddervators also affect pitch. The size of each ruddervator is approximately 45 ft2 and provide for yaw control and secondary pitch moment on maneuvering. The maximum deflection for each ruddervator is 30 degrees. Each trailing edge, single-slotted flap has a total moveable surface area of 51 ft2 with a maximum deflection of 40 degrees, while the ailerons each have an area of 13 ft2 with a maximum deflection of 30 degrees. The leading edge slats on both wings have a total moveable area of approximately 109 ft2. The leading edge slats are located across the entire span and help improve lift at high angles of attack, take-off, and landing.
5.1. Method of Analysis
The stability and control analysis of the SLOB Works team was, at the beginning, heavily dependent on Digital
DATCOM13. However, for this concept, the DATCOM program was not reliable due to the “unconventional features” of the bomber. The program could not compute lateral-directional derivatives at supersonic speeds. There was also either no method or approximations for handling nacelles of such shape and size, three-surface configurations (canard-wing-vertical tail), twin vertical tails, a blended wing body, or double delta wings14. With all these problems, the DATCOM output was not dependable. Thus, very little supersonic analysis was completed.
TABLE 5.1: Stability derivatives for SW-Ghost at Supercruise (Mach 1.6) Supersonic Derivatives
CLα 2.1949
Cmα 0.0564
In the subsonic region, both longitudinal and lateral derivatives were found using methods in JKayVLM15 and
Tornado16. JKayVLM and Tornado both used the vortex lattice method for calculations. A separate code was used for engine out requirements17. Due to the aircraft’s similarities, all data obtained was validated to the XB-70, which was examined by Razgonyaev and Mason18.
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TABLE 5.2: Stability & control derivatives for SW-Ghost at Takeoff (Mach 0.3)
CL Derivatives CD Derivatives CY Derivatives
CLα 2.52 CDα -0.000019 CYα 0.0
CLβ 0.0 CDβ 0.0 CYβ -0.1407
CLp 0.0 CDp 0.0 CYp -0.0021
CLq 2.0037 CDq 0.0 CYq 0.0
CLr 0.0 CDr 0.0 CYr -0.2325
Cl Derivatives Cm Derivatives Cn Derivatives
Clα 0.0 Cmα 0.3625 Cnα 0.0
Clβ -0.0267 Cmβ 0.0 Cnβ 0.0977
Clp -0.2872 Cmp 0.0 Cnp -0.00094
Clq 0.0 Cmq -1.3365 Cnq 0.0
Clr 0.0449 Cmr 0.0 Cnr -0.1650
Flaps Ailerons Canards V-Tails
CLδ 0.4112 0.1144 0.046826 0.10933
CDδ 0.0 0.0 0.0000017 0.0
Dyδ 0.0 0.0 0.0 0.0
Clδ 0.1046 0.0427 0.0 0.0
Cmδ -0.1885 -0.0566 0.084759 -0.12724
Cnδ 0.0194 0.0019 0.0 0.0
5.2. Static Stability
Calculations were made by hand19, 5 and validated from the longitudinal stability code to find the neutral point of the aircraft. The subsonic neutral point was found to be 32% of the Mean Aerodynamic Chord (MAC) for takeoff conditions.
For this condition, the subsonic static stability of the aircraft is –14.4%. The c.g. location of the aircraft stays at 46.4% MAC with the proper fuel transfer. The supersonic neutral point of the aircraft is 43.9% MAC. Thus, the supersonic static stability is –2.57%. The AC shift (figure 5.1) ranges from 37.4% MAC at takeoff to 49.2% MAC at cruising speed. With an increase in speed and angle of attack, the neutral point shifts aft, creating a more stable aircraft.
2002 SLOB Works, a Virginia Tech AOE Design Team 54
50%
48%
46%
44%
42% % MAC
40%
38%
36%
34% 0.4 0.6 0.8 1 1.2 1.4 1.6 Mach #
FIGURE 5.1 Aerodynamic Center Shift with change in Mach Number
5.3. Engine Out
In the event of an engine failure, the Ghost must be able to maintain controlled flight. Since the takeoff is the worst case for this condition (fully loaded), the LDStab code was used for analysis at this takeoff. The requirements include full rudder deflection and a 5o bank angle. The LDStab code outputs sideslip angle and other control deflections to allow for straight and level flight.
TABLE 5.3: Engine out data for SW-Ghost
β 2.49 φ 5.0
δa 3.07 δr 30.0 Cn avail -0.0023
5.4. Dynamics and Flight Qualities
For the dynamic stability characteristics of this aircraft, a program written by Dr. Frederick Lutze was used20.
Derivatives obtained by hand from Etkin & Reid and Raymer were used in the stability program. The twin-canted tails,
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along with the canards, contribute to the dynamic stability performance as displayed below in Table 5.4. Due to the instability of the SW-Ghost, the aircraft is controlled by a Flight-By-Light system. This system functions similarly to a Fly-
By-Wire system but operates much faster. The Flight-By-Light system offers a high degree of response accuracy in the controls and also eliminates the need for excessive hydraulic controls. In the next chapter, sections 6.6 and 6.7 provide more detail for the aircraft’s stability and control systems, as well as, all other aircraft flight systems.
TABLE 5.4: Comparison Chart for the SW-Ghost with the MIL-F-8785 B
MIL-F-8785 B SW-Ghost Requirements (Class II, Cat. B, Level 1) Subsonic Supersonic
Damping ξsp > 0.15 no short period: controlled Short Period rad rad with fly-by-light system Natural Frequency 0.1 /s < ωsp < 2.0 /s
Phugoid Damping Τ2 > 55 sec 55.46 sec 213.75 sec
Damping ξd > 0.02 0.0473 0.222 Dutch Roll rad rad rad Natural Frequency ωnd > 0.4 /s 0.800 /s 2.830 /s Minimum time to Spiral Roll 4 sec 4600 sec 5950 sec double amplitude Rolling Maximum time 1.4 sec 0.897 sec 0.097 sec Convergence constant
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6. Systems and Payloads
6.1. Basic Layout
The layout of this aircraft is a blended wing body type. The weapons will be carried internally in two main bays; a missile bay and a bomb bay. These weapons bays are conformal with the aircraft body and feature quick opening bay doors.
6.2. Fire Control and Defensive Systems
The aircraft will have numerous means of detecting threats and targets, both surface and airborne. The primary device is the Raytheon21 AN/APG – 70 active array RADAR provided in the RFP. This device is a multi-mode air-to-surface and air-to-ground radar system currently in service with the F-15E strike eagle. It is mounted in the nose cone of the aircraft and allows active scanning and ranging of airborne threats. This radar array also provides the crew with a means of obtaining high-resolution radar maps of their target areas on the ground. It will be mounted behind a selective bandpass radome, allowing only the radiation from the AN/APG-70 to travel through.
The aircraft is also equipped with a LANTIRN targeting pod. This pod contains a laser range-finder/designator beam for precision-guided weapons. It also incorporates a Forward Looking Infrared (FLIR) camera. The pod is mounted vertically on the centerline of the aircraft behind the cockpit. It is fully retractable serving to decrease the drag and radar signature of the aircraft en route to its target. The LANTIRN targeting pod allows this aircraft to incorporate precision munitions.
A LANTIRN navigation pod is mounted to the port side of the targeting pod. This device provides a means of low light navigation for the aircraft. It is also fully retractable.
A High Speed Anti-Radiation Missile (HARM) targeting pod provides an extra measure of engagement capability.
This system gives the aircraft a limited suppression-of-enemy-air-defenses (SEAD) ability. The HARM pod detects enemy radar systems and accurately determines their range and type. The HARM pod is mounted in the same manner as the
LANTIRN pods; it is internally stowed aft of the cockpit. Figure 6.1 shows the location of these systems as a top view of the nose section. The M61A cannon can also be seen in this figure.
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FIGURE 6.1: Top view of the fire control systems
6.3. Radar Cross Section (RCS) Prediction/Evaluation
The RFP states that the aircraft must have a maximum frontal RCS of 0.05 m2 against 1-10 GHz Ground Control
Intercept (GCI), acquisition, and tracking radars. As of this time, we have no means of numerically evaluating the size of the radar cross-section of the SLOB Works Ghost.
To improve the RCS of the aircraft, the weapons are carried internally. Also, the engines are “buried”, meaning that an enemy cannot radiate straight down the inlet and illuminate the spinning fan of the engine. The wings, canard, and vertical tails are all swept to deflect incoming radar waves away from the originating source. Both the landing gear doors and the weapons bay doors are “saw-toothed” to further deflect incoming radar energy. Finally, the radome on the nose consists of single band-pass material allowing the free travel of radar waves from the APG-70 radar system but not the incoming threat radar frequency ranges.
6.4. Cockpit
The cockpit of this aircraft is designed for two pilots, but a single pilot can fly it due to the long nature of the mission. Pilot controls are input using a stick and throttle in a “hands on throttle and stick (HOTAS)” concept. To improve the pilot-aircraft interaction, head’s up displays (HUD) are provided for both pilot positions. These features allow the pilot to fly the aircraft with little distraction and maximum efficiency.
The instrument layout of the cockpit has been designed with the idea that in some emergency cases there may be a need for single pilot flight. Because of the side-by-side seating configuration, it is possible to give both the pilot and copilot the necessary instruments for flying the aircraft alone. Both pilots have a heads-up display, throttle control, access to the weapons launch panel, and multifunction displays.
For the multifunction displays, flat screen LCD active matrix screens are used to ensure that both pilots can view all of the displays without image distortion, and to minimize any potential glare. The weapons launch control panel is a touch
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screen and is placed between and in front of the pilot and co-pilot for ease of use and optimum view. Also, because both pilot and co-pilot can easily see and use the panel, no weapon can be launched until both pilots concur on timing and targeting, thus increasing accuracy.
The monochrome displays, which can be used as a backup display for miscellaneous warnings and tasks, can be set according to the pilot’s preferences.
Analog instruments are included in case of main systems failure. These instruments include, altimeter, airspeed, false horizon, compass, and directional gyro.
Figure 6.2 shows the view from the cockpit along with instrument placement.
1. 8 inch heads up display 2. 12 inch active matrix LCD multifunction display 3. 8 inch active matrix LCD multifunction display 4. 14.1 inch active matrix LCD multifunction display 5. 6 inch monochrome multifunction display 6. 6 inch LCD touch screen (Weapons launch) 7. Analog altimeter, airspeed, artificial horizon, and directional gyro 8. Compass 9. Engine start and status toolbar 10. Misc. warning lights 11. Throttle 12. Flaps, landing gear, landing sequence initiation 13. HUD adjustment 14. Flight control stick (with autopilot) 15. Radio Selector and control 16. Cabin pressure, temperature and lighting control
FIGURE 6.2: View of the cockpit and instrumentation
The pilots are seated on K-36D model ejection seats22. This seat is designed by the Zvezda Design Bureau in
Tomilino, Russia. The seat features the greatest available ejection envelope for the pilots. It has a zero-zero capability as well as providing ejection at high supersonic speeds. Figure 6.3 shows the operating envelope of this seat as compared to other premier ejection seats.
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22 FIGURE 6.3: Effective envelope of the K-36D ejection sea (red field)
Upon ejection the seat uses belts attached to the pilot to reel in the head, waist, and extremities. It then deploys a windblast deflection shield and opens stabilizing booms to facilitate safe removal from the aircraft. The aircraft cabin features a roof that is jettisoned by explosive bolts just prior to seat ignition.
6.5. Electrical System
The systems of this aircraft have large electric power requirements. To provide the necessary power, there are four separate electrical systems.
The Auxiliary Power Unit (APU) is located in the tail. It is used to power the systems and avionics before and during engine start. The four turbofan engines provide the main electric power to the aircraft using turbine generators which produce 90 kVa per generator5. The generators of engines 1 and 2 (the port side engines) are tied into the primary power harness (bus) and the generators of engines 3 and 4 (the starboard side engines) are tied into the secondary power harness
(bus). These power busses are run separately through the aircraft to ensure redundancy. The aircraft is also equipped with two sealed lead-acid batteries underneath the cockpit to provide interior lighting, instrumentation, and power for APU start.
These batteries can also be utilized in the event of an electrical failure for a short period of time.
Finally, two ram air turbines (RATs) are installed under the bases of the vertical tails. When activated by the pilot, explosive charges opens the intakes and allows the turbines to generate power enabling the pilot to run instruments and controls long enough to facilitate a safe ejection. The details of the electric wiring system can be seen in figure 6.4.
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FIGURE 6.4: Top view of the electrical system of the SLOB Works Ghost
6.6. Flight Controls
The aircraft is controlled by a Fly-By-Light system. This system functions similarly to a Fly-By-Wire system but operates faster. Electro-Hydrostatic motors are used to operate all control surfaces. These motors run off of the plane’s electrical system. The plane is physically controlled by canards, ailerons, and ruddervators. The high lift system consists of leading edge slats and quarter span flaps.
This system offers a high degree of response accuracy in the controls and also eliminates the need for excessive hydraulic controls. The main drawback to a Fly-By-Light control system is the relatively large power requirement.
However, with four engines, the aircraft should generate sufficient power (approximately 90 kVa per generator). Figure 6.5 shows the control line scheme and flight controller locations. The hydraulics reservoir shown (in pink) handles landing gear retraction and steering.
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FIGURE 6.5: Top view of the flight controllers along with control lines and motors
6.7. Digital Flight Controller and Engine Control System
As mentioned above, the aircraft is controlled by means of a Fly-By-Light system. There is a dual redundancy capability obtained by using two entirely separate control systems. If one system malfunctions or is disabled, the other system will take over.
When the pilot applies control input, the digital flight computer decides if the input is correct and then moves the appropriate surface to the required deflection to obtain what the pilot desires. This is done nearly instantaneously thanks to the fiber optics. This allows the aircraft to be controlled similar to a fighter and keeps control inputs within required limits.
The flight controls will also determine maximum control deflection without overstressing the airframe (a g-limiter). There will, however, be a switch to disable the g-limiter. When held down, this button will allow the pilot to overstress the aircraft to a point possibly bending the airframe but not destroying it.
6.8. Landing Gear
The Front Landing Gear is modeled after a McDonnell Douglass MD-10 aircraft. The front landing gear consists primarily of an oil/air shock absorber (Oleo). The Oleo forms the main component of the main cylinder. A piston is attached to the main cylinder, driving the landing gear. The landing gear folds in to the nose. The landing gear is attached to aircraft at a height of 9.87 feet above the ground.
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FIGURE 6.6: Side view of front landing gear
The Rear Landing Gear of Ghost is modeled on the landing gear of a B-2 bomber. The system of landing gear used is quadricycle. The main feature of the landing gear is the sensing wheel. The purpose of the sensing wheel is to help the aircraft slow down in case of a hard landing. The landing gear rotates on its axis and folds back in to the plane. The landing has thick shock absorbers, to allow hard landings of the aircraft. A piston in front of the landing gear pulls the gear system back into the wheel well. The landing gear is 7.33 feet long with a wheel radius of 1.705 feet. The gear is retracted back and rotated about the leg into the aircraft.
FIGURE 6.7: Side view of main landing gear
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6.9. Fuel System
The JP-8 fuel will be carried in bladder style tanks concentrated inside the rear fuselage and wings of the aircraft.
The tanks will be self-sealing to prevent leaks if the tanks are damaged during a mission. The overall fuel capacity of this design is 139,000 lbs, but the design mission only requires 129,000 lbs. Due to the self-sealing nature of the tanks, only 85% of the volume is usable for fuel.
The main fuselage tanks are divided into four separate bladders. Each wing consists of four total bladders covering the length of the wing and 20% chord to 70% chord. This configuration allows the c.g. of the fuel to be concentrated at 86 feet aft of the nose.
There is a re-fueling receptacle located on the top of the aircraft, aft of the cockpit. This device will add greater utility to the design and also allow the pilots to use the afterburners with the knowledge that they can top off their fuel tanks after combat.
FIGURE 6.8: Top view of the fuel tank positions.
Due to landing requirements of 8,000 feet, the aircraft has a maximum landing weight of 120,000 lbs. This will require a system of fuel dumps for lightening the aircraft in the event of an early turn around and landing. The main fuel dumps are located at each wing tip of the aircraft. An electric auxiliary fuel boost pump from Hydro-Aire Inc powers each fuel dump23. This allows a combined fuel dump rate of 80,000 lbs/hr. In the event of a worst-case scenario involving immediate turn around for landing, 120,000 lbs. of fuel must be disposed of in a time of 1.5 hrs.
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6.10. Environmental Control System
The cockpit is climate controlled using bleed air from the four engines. This allows for cockpit pressurization along with heating and cooling of the ambient air. Air vents are located beside each pilot and temperature and cabin pressure is controlled on the overhead console.
6.11. Anti-Icing Equipment
Although many military aircraft do not have de-icing systems, the RFP states that this aircraft must fly and fight in all conditions. The de-icing system of this aircraft consists of heating elements embedded in the leading edges of the wings, canards, and vertical tails. De-icing boots are not used due to the supersonic nature of this aircraft and the difficulty of fitting rubber boots to all those surfaces. It would also require various pumps and hoses, further cluttering the interior of the aircraft. A fluid based de-icing system was not used because of the lack of space for a reservoir of alcohol and the piping and pumps required. This method would also not last as long as needed (there is a finite amount of alcohol that could be carried).
6.12. Aircraft Lighting
The aircraft has minimal lighting due to its stealthy nature. Due to its requirement to operate alone, there are no formation lights. There are standard navigation lights which can be turned off in a combat environment. These consist of a green light on the starboard wing and vertical surface and a red light on the port wing and vertical surface. The nose landing gear has a landing light attached for taxiing and landing at night.
6.13. Weapons
The weapons24 specified by the RFP include the AIM-120 AMRAAM, the Joint Deployed Attack Munition (JDAM)
(the aircraft can carry the 1000 lb. and 2000 lb. variety), the GBU-27, the Laser Guided Mk-84, AGM-154 Joint Standoff
Weapon, and the 250 lb. small smart bomb (currently the miniaturized munition technology demonstration).
6.14. Bomb and Missile Bays
The weapons of this aircraft are carried internally. This improves the Radar Cross Section of the airplane. Since the
RFP states that this plane must perform its mission with minimal help from other assets, it is equipped to carry air to air missiles in addition to its air to ground payload. The various required loads are shown in figure 6.9.
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FIGURE 6.9: Combat loads of the SW Ghost
FIGURE 6.10: Side and bottom views of weapons bays and their clearances
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The air-to-ground (AG) munitions are mounted 7 feet forward of the center of gravity of the aircraft. All of the weapons have the required 10 degree clearance between bay walls and at least 3 inches between each weapon.
Due to flow interference at high speeds, munitions experience a force that tries to pull them back into the bay of the dropping aircraft. To alleviate this problem the bombs are mounted on lugs that are in turn attached to hydraulic pistons.
When the pilot elects to employ a weapon, the bay doors open, the pistons extend, and a blast of inert gas, generated by the on-board inert gas generation system (OBIGGS), is fired to give the weapon enough separation. The pistons then retract and the bay doors close. This entire operation should take no more than 5 seconds, keeping the exposure time to enemy threat radars to a minimum.
The AIM – 120 type AMRAAM missiles are housed in a bay separate and forward of the bomb bay. When an air target is engaged a missile can be launched from the bay with priority given to the port missile. This bay is 22 feet forward of the CG. Due to the relative lightweight of the missiles, the CG shift is small each time an AMRAAM is employed. The release system operates in the same manner as the bomb release.
There is also an M-61A Vulcan cannon required by the RFP. This is mounted under and two and a half feet to port of the cockpit. This positioning allows the gun to fire and not upset the radar. There is an electrically operated door allowing smooth flow around the nose when the gun is not in use. An ammo drum containing 500 rounds is located just aft of the cannon.
6.15. Defensive Systems
This aircraft must perform its mission alone with minimal help from other assets. As such, there are numerous defensive systems incorporated in the design. For the most part these systems are all located in the tail structure.
There is an AN/ALE-50 towed decoy developed by Raytheon21. This device strings out a decoy attached to a wire.
The decoy emits signals which attempt to attract incoming missiles and away from the actual aircraft.
In the event of an engagement, the aircraft is also equipped with an AN/ALQ-161A Integrated Electronic Warfare
System (INEWS) and an Infrared Missile Warning System (IRMWS). These systems will also be obtained from Raytheon.
The purpose of these systems is to detect incoming missiles, both radar targeted and infrared targeted. The systems will alert the pilot and automatically or at the pilots’ discretion release chaff and/or flares. Figure 6.11 show the locations of these defensive systems.
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FIGURE 6.11: Top view of defensive system locations
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7. Propulsion Systems
A propulsion system was needed that could fulfill the RFP’s requirement for the ability to supercruise. In this chapter, the propulsion system comparative study, thrust requirements, the selected propulsion system, and its characteristics and inlet geometry will be discussed.
7.1. Propulsion system comparative study
A comparative study of aircraft powerplants was used to select which powerplant would be used for the SLOB Works
Ghost. The initial take-off gross weight (TOGW) estimate for the Ghost was 240,000 lbs. Through performance analysis it was determined that the preliminary thrust to weight ratio (T/W) of the Ghost should be 0.5. To obtain this, a propulsion system was required that produced 30,000 lbs of thrust (Note that the concept requires the use of four engines). After further analysis using the carpet plot it was determined that a new T/W ratio of 0.45 and a new TOGW of 238,000 lbs was needed.
At this T/W, the engine is required to produce 26,775 lbs of thrust. For the study however, each engine was sized to 30,000 lbs of thrust for comparison purposes. Each engine was sized based on the baseline engine using the following equations1.
0.4 LengthNEW = LengthOLD * (TREQ /TBASE) (7.1)
0.5 DiameterNEW = DiameterOLD * (TREQ /TBASE) (7.2)
1.0 WeightNEW = WeightOLD * (TREQ /TBASE) (7.3)
Six engines were selected for the comparative study (figure 3.1). The Pratt & Whittney PW F100-232, PW F119-
100, and PW J58, the General Electric F101-102 and F110-132, and two configurations (with and without afterburners) of the engine given in the RFP were the propulsion systems that were compared. The PW F119-100, the F-22 powerplant, was eliminated from selection due to lack of availability of data. Specific fuel consumption (SFC) and size were the major engine selection criteria. The GE F110-132 and the PW J58 had high SFCs of 2.09 and 2.174 respectively. While the RFP engine configuration with afterburners had an SFC of 2.618, the highest of all the systems, it had to be taken in account that this was the SFC at max thrust with afterburners engaged. Afterburners increase the thrust of an engine by 60% and SFC is increased by 120% while the afterburners are in operation5. At Military thrust the SFC was 1.19, while at any thrust setting higher
(afterburner will be in operation) the SFC changed to 2.618.
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TABLE 7.1: Engine comparative study for resized engines using equations from RFP.
Weight(lbs) Length(in) Max Diameter(in) SFC PW F100-232 3,785 185 45 1.91 PW J58 5,501 200 52 2.174 GE F101-102 4,289 179 54 N/A GE F110-132 3,744 177 45 2.09 RFP Engine w/ AB 5,122 271 55 2.618 RFP Engine w/o AB 8,195 326 69 1.19
7.2. Thrust Requirements
The thrust required at take-off, loiter, cruise, and dash were very important in the selection process. Thrust required
(TREQ) is a function of density, velocity, wing area, and drag coefficient. The equation is:
2 TREQ = ½ * ρ * V * S * CD (7.4)
The thrust required divided by the number of engines is the total thrust required per engine at a given altitude. The thrust required for the SLOB works Ghost at the required altitudes and Mach numbers is tabulated in the table 7.2. Thrust is low at loiter due because the aircraft is flying at max endurance speed which slightly above the stall limit. Since thrust is low, the fuel burn at loiter is low.
7.3. Propulsion System Selected
Initially the RFP engine, without afterburners (type I), was selected for the propulsion system of the SLOB works
Ghost. The SFC for this engine was significantly low compared to the others. Also, the thrust requirements based on max range were well met by this engine. The thrust required is compared to thrust available in the table 7.2.
TABLE 7.2: Thrust required and Thrust available for the RF P engine at given conditions.
Condition Altitude (ft) Mach No. TREQ (lbs) TAVAIL(lbs) Take-off 0 (sea level) 0.3 30,000 30000 Dash 52k 1.6 3,625 8213 Cruise 52k 1.6 3,625 8213 Loiter 36k 0.6 785 6339
The supercruise low-bypass turbofan was sized up to 30,000 lbs of thrust (using the equations from section 3.1).
The relationship between the original dimensions and the increased dimensions for this propulsion system is in table 7.3.
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TABLE 7.3: Base engine specs vs. sized engine specs
Characteristics Base Engine Sized Engine Weight (lbs) 7,200 8,195 Length (ft) 25.83 27.21 Max Diameter (ft) 5.42 5.78 Fan Face Diameter (ft) 4.17 4.45 SFC 1.19 1.19
After further analysis, it was determined that an afterburner must be incorporated so that Maneuvering requirements could be fulfilled. This produced the RFP engine Type II. This upgrade increased the max installed thrust to 48,000 lbs at a throttle setting of afterburner. While the max military thrust is 30,000 lbs at a throttle setting of 100%. A reverse thruster system (figure 7.1) was also installed so that landing requirements could be met.
FIGURE 7.1 Reverse thruster system diagram
The nozzle exit area varies depending on subsonic or supersonic flight. The nozzle exit area in the subsonic regime varies from 9.4987 ft2 to 13.2981 ft2. While in the supersonic regime it varies from 22.7968 ft2 to 30.3957 ft2. These values were obtained using Raymer5.
7.4. Inlet Geometry
The inlet system consists of a three-shock intake and a subsonic diffuser. The intake is double-wedge external compression. The double-wedge intake utilizes two oblique shocks and a normal shock to slow the flow from Mach 1.6
(Mach at cruise) to Mach 0.7. Both of the wedge angles (δ) are two degrees (figure 7.2).
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FIGURE 7.2: Double–wedge intake geometry.
The intake capture area was calculated using Raymer’s equation5. The capture area per engine is 18.9973 ft2. The inlet system utilizes boundary layer suction to divert the boundary layer. A porous bleed is used for the throat bleed and secondary airflow.
The subsonic diffuser (figure 7.3) utilizes S-bend geometry to prevent the fan face from reflecting radar waves. This characteristic contributes to the low RCS of the Ghost. Geometry of the subsonic diffuser further reduces the Mach number to 0.55 entering the fan face. To manage the flow in the subsonic diffuser, boundary-layer suction and bleeds are used.
FIGURE 7.3: S-Bend subsonic Diffuser designed for the Ghost.
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8. Performance The performance was analyzed for the requirements given in the RFP. The RFP required a take-off, climb, supercruise out, dash-out, turn, dash back, supercruise back, and landing analysis. A 30-minute loiter requirement also needed to be met. After looking at Attachment 1 in the RFP, the performance analysis started with first obtaining take-off distances and then proceeded to obtaining Ps plots which will be seen later in this chapter. Most of the parameters needed to fulfill this analysis were obtained using the weights, aerodynamics, and propulsion data presented in these chapters.
8.1. Performance Parameters
The take-off analysis was completed using Raymer5 and Roskam25 as references. The SLOB Works Ghost was required to take-off on an 8,000 ft runway on dry, wet, and icy concrete at sea level and standard atmosphere conditions.
Using the following equations from Raymer5, assuming µ for dry concrete is 0.03, µ for wet concrete is 0.05, and µ for icy concrete is 0.02, CL= CLmax =1.2, the results are shown below:
2 1 KT + K AV f S = n (8.1) G 2gK l 2 A K T + K AVi
T KT = − µ (8.2) W
ρ 2 K A = (µC L − C DO − KCL ) (8.3) 2 W ()S
TABLE 8.1: Take-off distances for the three different surfaces
Total Ground Distance (dry concrete) brakes off (ft) 3400
Total Ground Distance (wet concrete) brakes off (ft) 3500
Total Ground Distance (icy concrete) brakes off (ft) 3370 Balanced Field Length (ft) 4500
The take-off speed is approximately 285 ft/sec and SW-Ghost will be taking-off at a 15 degree angle of attack. The transition to climb will change the angle of attack to 20 degrees.
To determine what altitude the SLOB Works Ghost will be supercruising out, the Breguet range equation was used, as seen below:
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V L W R = ln i (8.4) C D W f
To maximize the range and minimize the fuel burned, (L/D)/C needs to be maximized. After going through those calculations, the optimized altitude for supercruise out was determined to be 50,000 ft.
TABLE 8.2: Evaluation of each altitude’s (L/D)/SFC
Altitude (ft) L/Dcruise SFC (L/D) / SFC 50000 7.12 1.20 5.93 51000 7.19 1.23 5.85 52000 7.24 1.25 5.78 53000 7.24 1.28 5.68 54000 7.25 1.30 5.59 55000 7.23 1.30 5.55 56000 7.21 1.30 5.55 57000 7.14 1.28 5.56 58000 7.09 1.26 5.61 59000 7.02 1.23 5.68 60000 6.94 1.20 5.76
Below is a table that shows how much fuel is burned as SLOB Works Ghost supercruises through the 1000 nm range:
TABLE 8.3: Fuel Burned during 1000 nm Supercruise out
Range (nm) Wfuel (lbs) 100 3993 100 3927 100 3864 100 3803 100 3743 100 3686 100 3630 100 3577 100 3525 100 3474 1000 37222
SLOB Works Ghost needs to climb from sea level to 50,000 ft. A climb analysis was then performed and the plot shows the shortest time to climb to the supercruise altitude. Minimum time to climb was necessary to reduce the amount of
5 time spent in the areas of high drag. Using Raymer , the Ps plot containing lines of constant energy was created as seen below.
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60000 stall limit 140 Ps = 0 100 120 50000 80 60 50 40000 Ps = 100 40 Temp. limit = 150 F 30000 30 Ps = 200 Inlet Pressure Limit Altitude (ft)
20000 q limit = 2133 psf 20 Ps = 300 10000 10
0 5 0.0 0.5 1.0 1.5 2.0 Mach No.
FIGURE 8.1: Climb Analysis at MTOGW, n=1, Mil. Thrust
The climb line is tangent to both the Ps lines and to the lines of constant energy. The time of climb was calculated using the equation in Raymer5.
After supercruise out, SLOB Works Ghost will be dashing out, supercruising back, and dashing back at an altitude of 59,000 ft. The same analysis was used to determine the altitude for optimizing the range and minimizing the fuel burned.
Flying at 59,000 ft and M =1.6, the maximum value of (L/D)/SFC was achieved.
Below are the Ps plots for n=1, n=2, and n=5 at max military thrust and at 50% internal fuel weight. SLOB Works
Ghost will not be using afterburners during any part of the mission so maximum military thrust will be used in the performance analysis. The only time afterburners need to be used are to fulfill one constraint given in the RFP1. That plot will be seen later.
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16 70,000 Ps = 0 14 stall limit 12 10 60,000 Temp limit = 150 F 80 Ps = 50,000 60
Inlet pressue limit 5 40,000 Ps = 200 4
Altitude (ft) Altitude 30,000 q limit = 2133 3 Ps = 300 20,000 2
Ps = 400 10,000 1 Ps = 500 5 0 0.00 0.50 1.00 1.50 2.00 2.50 Mach No.
FIGURE 8.2: Ps plot for n=1, 50% fuel weight, and Mil. Thrust
Temp limit = 150 F 50000 stall limit Ps = 0
40000
Ps = 100 inlet pressure limit
30000
Ps = 200 Altitude (ft)
20000
Ps = 300 q limit = 2133 psf
10000
Ps = 400
Ps = 500 0 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Mach No.
FIGURE 8.3: Ps plot for n=2, 50% fuel weight, and Mil. Thrust
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28000
26000 inlet pressure limit
24000 stall limit Temp limit = 150 F 22000
20000
18000 q limit = 2133 psf Ps = 0 16000
14000
12000 Altitude (ft)
10000 Ps = 100 8000
6000
4000 Ps =
2000
0 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Mach No.
FIGURE 8.4: Ps plot for n=5, 50% fuel weight, and Mil. Thrust
Temp limit = 150 F 70000
60000 stall limit Ps = 200
50000 inlet pressure limit
40000 Altitude (ft) 30000 q limit = 2133 psf
20000
10000
0 0 0.5 1 1.5 2 2.5 3 Mach No.
FIGURE 8.5: Ps plot for n=1, 50% fuel weight, and afterburners
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At n=1, 50% fuel weight, SLOB Works Ghost has a max speed of M = 1.7 at 36,000 ft. The RFP1 had a constraint that at n=1, 50% fuel weight, the aircraft needs to be able to fly at M = 1.6 and at or above 50,000 ft for a Ps = 200 ft/sec.
This is where the use of the afterburner comes in. The plot showing this constraint is shown in figure 8.5 and is met. SLOB
Works Ghost is able to fly at 58,000 ft at M = 1.6.
The summary of the mission for the SW-Ghost will be supercruise at 50,000 ft at M=1.6, and after the 1000 nm cruise, the Ghost will be switching to the 750 nm Dash out at an altitude of 59,000 ft at M=1.6. At the end of the dash out, the Ghost will perform the 180-degree turn and at the end of the turn it will drop the payload. The next section of this chapter will go into more detail encompassing the turn analysis.
After the deployment of the weapons, a 750 nm dash back and the 1000 nm supercruise back are performed at
M=1.6 at an altitude of 59,000 ft. Finally a descent is performed leading up to the approach for landing. Once again, the plane must land on an 8,000 ft runway on dry, wet, and icy concretes. After careful calculations with the use of Raymer5, the
RFP requirements are met with the use of 10% reverse thrusters. The landing distances are shown below for the different types of conditions.
TABLE 8.4: Landing distances for the three different surfaces
Total Ground Distance (dry concrete) brakes on (ft) 5000
Total Ground Distance (wet concrete) brakes on (ft) 6200
Total Ground Distance (icy concrete) brakes on (ft) 7600
Loiter is the last segment of the mission before landing and must be 30 minutes long. The loiter equation was used to perform this analysis and is shown below:
L 1 W E = ln i (8.5) D C W f
To achieve max endurance conditions and minimum fuel burned, (L/D)/C must be maximized just as in the range analysis. After going through this analysis, SW-Ghost will be loitering at M = 0.4 at an altitude of 10,000 ft.
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8.2. Maneuvering Performance Diagram
Sustained g-turn rates are calculated for a particular altitude and give various turn rates we can have at different speeds for a given altitude. Turn rate is a function of load factor and speed. Figure 8.6 shows the sustained turn rate at an altitude of 50,000 ft.
15
14 n= 7 13 n= 6 12 n= 5 11 n= 4 Structural Limit 10 Stall Limit
9 n= 3 8
7 n= 2 6 Turn rate (deg/s)
5
4
3 Sustained Turn 2 Rate Envelope
1
0 450 550 650 750 850 950 1050 1150 1250 1350 1450 1550 1650 1750 1850 1950 Speed (ft/s)
FIGURE 8.6: Sustained turn rate graph at 50,000 ft
The structural limit for the aircraft is 7 g’s. On the graph we have the load factors plotted as contour lines. The intersection of the stall limit and structural limit gives us the corner speed of our aircraft. In a classic turning dogfight, a pilot would want to reach his corner speed as fast as possible. The RFP1 requires for us to calculate the time it takes to do a 180 degree turn at 50000 ft and at M =1.6. To plot the graph, different turn rates were calculated at different speeds and turn rates, plotted as a function of velocity. The figure shows that we can obtain a sustained turn rate under the structural limit of the aircraft. From our graph we calculated our required turn rate to be 2.45 deg/ sec at a speed of M 1.6. Thus the time it takes us to do a 180-degree turn is calculated to be 74 sec; the distance covered by the aircraft is about 19 miles. The radius of the turn was calculated to be about 3.45 miles. Sustained turn rate graphs for sea level and 15,000 ft were also required. They are shown below.
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40 38 36 34 32 30 Stall Limit 28 Structural Limit 26 n= 7 24 n= 6 22 n= 5 20 n= 4 18 n= 3
Turn rate (deg/s) 16 14 12 n= 2 10 Sustained Turn 8 Rate Envelope 6 4 2 0 200 250 300 350 400 450 500 550 600 650 700 750 800 850 900 950 100 105 110 115 120 0 0 0 0 0 Speed (ft/s)
FIGURE 8.7: Sustained turn rate at Sea level
30 29 28 27 26 25 Stall Limit 24 23 22 21 20 19 n= 7 18 n= 6 17 n= 5 Structural Limit 16 n= 4 15 14 13 n= 3
Turn rate (deg/s) 12 11 10 n= 2 9 8 7 6 Sustained Turn 5 4 Rate Envelope 3 2 1 0 300 350 400 450 500 550 600 650 700 750 800 850 900 950 100 105 110 115 120 125 130 0 0 0 0 0 0 0 Speed (ft/s)
FIGURE 8.8: Sustained turn rate at 15,000 ft
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8.3. Sustain Load Factor Envelope
The sustained load factor envelope shown below is at Ps=0 for a range of load factors from +1 to +7.
60000
50000
n = 1 40000 n = 3 n = 5 30000 n = 7 pressure limit
Altitude (h, ft) (h, Altitude q limit 20000 temp limit
10000
0 0.00 0.50 1.00 1.50 2.00 Mach (M, -)
FIGURE 8.9: Envelope for the Ghost at Ps=0 for a range of load factors
This plot shows that at n=1 the maximum speed the Ghost can fly at, at an altitude of 36,000 ft, is M=1.7. At a
Mach number of 1.07, the Ghost can fly around an absolute ceiling of 60,000 ft.
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8.4. Conclusions on Performance Analysis
Table 8.5 summarizes the mission duration, circumference of the 180-degree turn, the time for each segment, and the fuel burned for each mission segment.
TABLE 8.5: Summary of each segment giving the important performance parameters
MISSION Time to Distance Traveled Amount of fuel Total SEGMENT complete (nm) burned (lbs) Weight (hrs) (lbs) MTOGW - - - 240,000
Take off 0.004 0.560 (Dry) 7,500 232,500 Climb 0.45 17.7 7,500 225,000 Cruise out @ 50K 1.090 1,000 37,222 187,778 Dash out @ 59K 0.819 750 24,549 163,229 180 deg turn / 0.021 18.86 450 154,079 Ordnance Drop @ (circumference) 50K
Dash back @ 59K 0.819 750 19,950 134,129 Cruise back @ 59K 1.090 1,000 22,813 111,316 Landing 0.004 0.815 (Dry) 687 110,629 Loiter @ 10K 0.500 - 3,772 106,857 Reserve & Trapped - - 8,000 98,857 Total 4.8 3,542 132,443
All of the performance segments of the mission were calculated using a maneuver weight of 50% internal fuel,
170,000 lbs of fuel, for the air-to-ground design mission loadings.
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9. Weight Analysis
9.1. Weight Breakdown
An in-depth weight and center of gravity analysis was done for the SLOB Works Ghost aircraft. To determine an
5 initial value empty weight, Wo, of the aircraft an approximation was done using Raymer’s approximate weights methods .
From the initial calculations, the aircraft’s empty weight was determined to be 99,389 lbs. This approximate method was validated using the Concorde aircraft and the approximate Wo was within 10% of the actual value.
After determining the initial empty weight, a detailed analysis of the weights was done. To determine the weights of the structural components, the algorithms listed in Roskam’s26 airplane design series were used. These methods used statistical algorithms based upon sophisticated regression analysis of aircraft. The weights of the other components were obtained from the RFP as well as from the systems & payload, propulsion systems, and performance groups. Tables 9.1 through 9.5 provide a breakdown of all of the weights of the components in the aircraft. Table 9.6 summarizes the aircraft weight’s subgroups and lists the inertias for theses groups, and table 9.7 shows the aircraft’s ratios.
TABLE 9.1: Structural Weights Group
Weight (lbs) xcg (ft) zcg (ft) x-mom. (ft-lb) z-mom. (ft-lb) Fuselage 38057 78.025 0 2969359 0 Wings 9838 102.378 0 1007147 0 Canards 1049 43.46 3.77 45570 3953 Vertical Tails 1929 126.102 4.19 243298 8084 Front Landing Gear 621 44.386 -4.45 27573 -2764 Rear Landing Gear 5591 95.537 -5.59 534136 -31253
TABLE 9.2: Propulsion System Weights Group
Weight (lbs) xcg (ft) zcg (ft) x-mom. (ft-lb) z-mom. (ft-lb) Engines w/ Afterburner (4) 32800 113.375 -3.681 3718700 -120737 Inlets 202 89.556 -3.681 18078 -743 Engine Controls 172 113.375 -3.681 19529 -634 Engine Starter 269 113.375 -3.681 30457 -989 Thrust Reversers 1476 113.375 -3.681 167342 -5433
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TABLE 9.3: System Weights Group
Weight (lbs) xcg (ft) zcg (ft) x-mom. (ft-lb) z-mom. (ft-lb) AN/APG 70 Radar 450 6.95 0 3128 0 M61A Cannon 275 22.7 -2.01 6243 -553 Ammo drum 300 27.6 -1.45 8280 -435 OBOGS 35 35.1 5.34 1229 187 OBIGGS 35 37.2 5.34 1302 187 Battery 1 20 24.76 2 495 40 Battery 2 20 24.76 2 495 40 Flight Controller 1 50 31.38 5.81 1569 291 Flight Controller 2 50 31.38 5.81 1569 291 LANTIRN Nav Pod 350 33.05 -3 11568 -1050 LANTIRN Targeting Pod 350 33.05 0.58 11568 203 HARM Targeting Pod 150 33.05 -2 4958 -300 Engine Generator 1 300 116.24 0 34872 0 Engine Generator 2 300 116.24 0 34872 0 RAT 1 50 131.23 0.5 6562 25 RAT 2 50 131.23 0.5 6562 25 APU 100 134.91 0.2 13491 20 INEWS 100 138.66 0.47 13866 47 IRMWS 28 140 0.66 3920 18 AN/ALE 50 80 142.08 0.29 11366 23 AC, Press, De-ice 432 25 5 10806 2161 Ejection Seats 320 22.503 5.862 7201 1876 C.G. Controller 246 55 3 13546 739 In-flight Refueling 50 16.527 2 826 100 Hydraulics 702 55 0 38598 0 Paint 720 78.025 0 56178 0
TABLE 9.4: Ordinance Weights Group
Weight (lbs) xcg (ft) zcg (ft) x-mom. (ft-lb) z-mom. (ft-lb) AIM - 120 AMRAAM (2) 700 67.1 -4.09 46970 -2863 JDAMS (4) 8000 82.238 -4.09 657904 -32720
TABLE 9.5: Fuel & Crew Weights Group
Weight (lbs) xcg (ft) zcg (ft) x-mom. (ft-lb) z-mom. (ft-lb) Pilots (2) 500 22.503 5.862 11252 2931 Fuel Used 125000 85 -1 10625000 -125000 Fuel Reserve 6667 85 -1 566695 -6667 Fuel Trapped 1333 85 -1 113305 -1333 Miscellaneous 234 78.025 0 18249 0
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TABLE 9.6: Weight Summary and Inertias of SW-Ghost
Ixx Iyy Izz Izx 2 2 2 2 Weight (lbs) xcg (ft) zcg (ft) (slugs-ft ) (slugs-ft ) (slugs-ft ) (slugs-ft ) TOGW 240000 87.9 -1.3 - - - -
W0 97566 93 -2 123 81547 81424 -3165 Structures 57084 84.6 -0.4 1488 21835 20347 -5503 Propulsion 34919 113.2 -3.7 6148 700288 694140 -65325 Systems 5563 54.8 0.7 697 190279 189582 -11498 Weapons 8700 81.0 -4.1 2103 15080 12976 5224 Fuel & Crew 133734 84.8 -1.0 448 42833 42385 -4359 Note: Aircraft is symmetric and all components are placed for ycg = 0, therefore Ixy and Iyz are 0
TABLE 9.7: SW-Ghost ratios
Symbol Value Wing Loading W/S 86.6 lbs/ft2 Thrust-to-Weight (mil.) Tmil/Wt 0.50 lbs-st/lbs Thrust-to-Weight (max.) Tmax/Wt 0.80 lbs-st/lbs Fuel Ratio Wf/Wt 0.55 Payload Ratio Wp/Wt 0.04
9.2. Center of Gravity
Moments about the nose of the aircraft were taken to determine the center of gravity of the entire plane. The movement of the center of gravity was calculated by determining the weight change at different stages of the flight mission.
As will be described in the systems chapter, fuel pumps will be used to control the movement of the center of gravity. The movement of the c.g. over the aircraft’s flight regime can be found in figure 9.8
Also, figure 9.1 compares movement of the c.g with and without fuel pumping. By using fuel pumping, it is possible to keep the c.g at 88 ft while the aircraft is cruising supersonically, which is a requirement set forth by stability and control. The c.g. moves a maximum of 26 inches throughout the entire cruise portion of the flight, compared to the 30 inches without the fuel pumping. If the final part of the supercruise back segment were to be excluded, the aircraft’s c.g. will only move by 1-3/4 inches with fuel pumping, while with a standard burn, the c.g. will move 5-5/8 inches.
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C.G. Movement M.A.C 0 % 10 % 20 % 30 % 40 % 50 % 60 % 70 % 80 % 90 % 100 % 240,000 at TOGW Take-off
220,000 Climb to 50,000 ft
Max. Foward C.G. Location 200,000
Supercruise at 50,000 ft Max. Aft C.G. Location 180,000 Fuel Control Standard Burn Dash-out at 59,5000
Weight (lbs) 160,000 180 deg turn; Bomb & Missle
140,000 Dash-back at 59,500
120,000 Land Supercruise at 59,500 ft Loiter at 10,000 ft
100,000 60.00 70.00 80.00 90.00 100.00 110.00 120.00 130.00 C.G. Location from Nose (ft)
FIGURE 9.1: Center of gravity movement
9.3. Weights and C.G. Conclusion
The weight, c.g. location, and c.g. movement of the SW-Ghost was determined. From the weight analysis, the aircraft was determined to have an empty weight of 97,566 lbs and a TOGW of 240,000 lbs. By using a fuel pump it is possible to keep the c.g. location at 88 ft, but since most of the fuel is burned by the end of the last supersonic stage, the c.g. shifts aft to
90 ft. This is a larger jump in the c.g. location than the Concorde, but is acceptable for the SW-Ghost since it remains in the forward and aft limits of the c.g. location.
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10. Cost Analysis
10.1. Introduction
One of the major constraints for this project was the cost requirement. A cost requirement was given in the RFP for
200 aircraft, each costing no more than $150 million in year 2000 constant dollars. Since this is one of the major design criteria, it was valued on the same level as the technical data. In the next few sections a breakdown of the costs will be introduced, as well as a trade study showing how the price of the aircraft varies with the amount built.
10.2. Cost Analysis Method
The primary method used to calculate these costs was the empirical formulas written by Dr. Jan Roskam6. It should be noted that these equations are based primarily on the correlations of subsonic aircraft. The reason for using this method is that there were not many supersonic transports, more specifically, supersonic bombers that have been produced. Since there aren’t any supersonic bombers to compare the following data with, a second cost algorithm created by J. Wayne Burns at
Vought Aircraft27 was used for comparison. Since the Vought Aircraft algorithms only determine the RDT&E and acquisitions costs, this code was only used to calculate the fly-away and unit costs. In the following sub-sections, the life cycle cost and the individual phase costs for 200 aircraft are shown.
10.3. Aircraft Life Cycle Cost
The life cycle costs of an aircraft were broken up into four major phases: research, development, testing & engineering (RDT&E); acquisition; operating; and disposal.
The first phase is the research, development, testing and evaluation (RDT&E) and encompasses all of the costs from research and development to the final detailed design drawing as well as the financing and profit. The next major phase is acquisition. Acquisition pertains to the costs necessary for manufacturing, production flight-testing, profit and financing.
The third phase is the operation. The costs calculated in this group are all of costs associated to operating the aircraft, for example, pilots, maintenance, and fuel. The final phase is the disposal phase and the only cost that is calculated is the cost required to dispose of the aircraft.
The research, development, testing & evaluation cost was calculated by determining the following costs:
• Airframe Engineering & Design • Development Support & Testing • Flight Test Airplanes • Flight Test Operations • RDTE Profit • RDTE Financing 2002 SLOB Works, a Virginia Tech AOE Design Team 87
These costs were calculated using algorithms so it was necessary to input the aircraft’s characteristics. It was also necessary to estimate some of the factors used in this algorithm. These factors were observability, design difficulty, CAD difficulty, material, finance, and profit. The values used were estimated based on other aircraft. The inputs for the algorithim are given in table 10.1 through 10.3 and the RDT&E cost breakdown for 200 aircraft can be found in table 10.4.
TABLE 10.1: Aircraft Inputs
Symbol Units Value Definition TOGW lbs 240,000 Take-off Gross Weight
Wf lbs 133,000 Fuel used
Wampr lbs 79,861 AMPR Weight
Vmax kts 1032.42 M 1.8 @ 36,089 ft
NRDTE - 6 Number of Test Aircraft
NST - 4 Number of Static Test Aircraft
NM - 200 Number of Production Aircraft
NProgram - 206 Total Program Production Aircraft
NRR units/month 0.33 Number of research aircraft produced per month
NRM units/month 5 Number of Program Aircraft produced per Month
NE - 4 Number of engines
TABLE 10.2: Adjustment Factors
Symbol Units Value Definition (Factor Rating)
Fdiff - 2 Difficulty Factor (Hardest)
Fcad - 0.8 CAD Difficulty Factor (CAD experts)
Fmat - 2 Material Factor (Standarad composite material)
Fobs - 3 Observability Factor (Stealthy)
Ftsf - 0 Test Facility Factor (No new facilities)
Fpror - 0.1 Profit Factor (10%)
Ffinr - 0.15 Finance Factor (15%)
Fftoh - 4 Overhead factor
Fol - 1.005 Oil & Lubricant Factor
fpersind - 0.14 Indirect Personnel Factor
fspares - 0.27 Spares Factor
fdepot - 0.22 Deport Factor
fmisc - 0.04 Miscellaneous Factor
TABLE 10.3: Rates Symbol Units Value Definition
RER $/hour 125 Engineering/ Research Rate 2000
RMR $/hour 69 Manufacturing Labor Rate Y:2000
RTR $/hour 88 Tooling Labor Rate Y:2000
RMMIL $/hour 61 Military Maintenance Rate Y:2000
RCONMAT $/hour 8.81 Average cost of consumable materials Y:2000
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TABLE 10.4: RDT&E Cost Breakdown (in millions $, 2000 constant)
RDT&E $12,601 Airframe Engineering & Design $2,970 Development Support & Testing $1,071 Flight Test Airplanes $4,935 Flight Test Operations $475 RDTE Profit $1,260 RDTE Financing $1,890
The second phase of the aircraft’s life cycle was the acquisition. Acquisition is broken down into the following cost sub- groups:
• Airframe Engineering & Design • Airplane Program Production • Production Flight Test Operation • Manufacturing Finance • Manufacturing Profit
The total cost for this phase of the project can be found in table 10.5. Also shown in the table below are the costs for
200 aircraft for each of the sub-groups.
TABLE 10.5: Acquisition Cost Breakdown (in millions $, 2000 constant)
Acquisition Cost $32,393 Airframe Engineering & Design $2,702 Airplane Program Production $22,076 Production Flight Test Operation $253 Manufacturing Finance $4,417 Manufacturing Profit $2,945
The third phase of the life cycle was the operational phase. The operating cost is defined by the summation of the following sub-groups:
• Fuel, Oil & Lubricants • Direct Personnel • Indirect Personnel • Consumable Materials • Spares • Depot • Miscellaneous Items
The operating cost over a period of 20 years of service, and 500 flight-hours per year for 200 aircraft, is given in table 10.6. This aircraft will require experience to fly, and the pilots will need to be a major or lieutenant commander in the
Air Force. The cost of each of the above-mentioned sub-groups is given in table 10.4
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TABLE 10.6: Operating cost breakdown (in millions $, 2000 constant)
Operating Cost $25,944 Fuel, Oil & Lubricants $4,356 Direct Personnel $3,844 Indirect Personnel $3,632 Consumable Materials $361 Spares $7,005 Depot $25,944 Miscellaneous Items $1,038
The final phase of the life cycle of the aircraft, the disposal phase, was calculated. The total disposal cost was calculated to be $717 million (2000 constant) for 200 aircraft.
The life cycle cost of the aircraft is defined as the summation of the costs for every phase of the aircraft’s life cycle.
By summing all of the costs for every phase, it was possible to determine that for 200 aircraft the life cycle cost will be approximately $71.7 billion. Figure 10.1 shows approximately how much of the total life cycle cost each phase encompasses. Disposal
RDT&E
Operating
Acquisition
FIGURE 10.1 Breakdown of the life cycle costs
10.4. Unit Cost, Fly-Away Costs and Cost per Pound
The data presented in section 10.3 are the costs for 200 aircraft. The total life cycle costs are not restricted in the
RFP, but there is a cost requirement of $150 million year 2000 constant dollars. In this section, the unit cost and the flyaway costs for the SW-Ghost will be presented. The unit cost is defined as the RDT&E and acquisition cost divided by the number
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of production aircraft, while the fly-away cost is defined as the cost required to produce one aircraft. Finally, the cost per pound of aircraft will be introduced.
Based on the definitions of the unit cost and flyaway cost, the unit cost was determined to be $226 million per aircraft and the fly-away cost was determined to be $145, which is $5 million below the requirement set forth in the RFP.
The cost per pound of the aircraft was determined by dividing the flyaway cost by the AMPR (Airframe unit weight) and the aircraft was determined to cost $1,800 /lbs. Therefore, no extra structural material or dead space is wanted.
10.5. Cost Trade Study
The effect of the number of aircraft on the unit and flyaway costs was analyzed in detail. From this simple cost trade study it was possible to determine how these costs vary in value for 100 to 1000 aircraft produced.
The unit and flyaway cost for different numbers of aircraft produced was determined using this study. This was accomplished by varying the number of aircraft produced from 100 to 1000. As a result, it was possible to see an important trend in the data. Figure 10.2 shows how the unit and flyaway costs vary with number of aircraft. In this figure it can be seen that the flyaway and unit costs decrease exponentially. Figure 10.3 shows what percentage of the total life cycle cost each phase encompasses versus the number of aircraft produced. As the number of aircraft increase, the sunken cost RDT&E decreases in total percentage, while the production dependent acquisition, operation, and disposal costs increase in total percentage with an increase in the number of aircraft produced.
From this study, it can be seen that 200 aircraft is the ultimate minimum that can be produced for the $150 million flyaway cost limit. But if it were possible to increase the number of aircraft produced it will be possible to reduce the unit and flyaway costs drastically.
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350E+06 300E+06 250E+06
200E+06 Unit Costs 150E+06 Fly-Away Costs Costs ($) 100E+06 050E+06 000E+00 0 200 400 600 800 1000 Units
FIGURE 10.2: Unit and fly-away cost trade study in Year 2000 dollars
100% 90% 80% 70% 60% Disposal Phases 50% 40% Operation 30% Aquistion 20% RDTE 10% 0% 100 300 500 700 900
Units
FIGURE 10.3: LCC breakdown comparison between number of units produced
10.6. Cost Conclusion
Through this analysis it was possible to determine the life cycle cost, unit cost, fly-away cost and cost per pound of aircraft. Also, it was possible to conduct a cost trade study, showing by increasing the amount of aircraft produced it is possible to decrease the cost per aircraft drastically.
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To reduce and keep the cost of the aircraft down, government furnished systems were used in the aircraft. Also, by using the technology and methods used in the F-22 and future SST aircraft for the SW-Ghost it will be possible to use and/or modify existing facilities for R&D, flight testing, manufacturing as well as others.
Since the algorithms do not take into account contractual work, it will be possible to reduce the $145 million price tag on the aircraft even further by sub-contracting labor intensive components to lower developed countries. It will also be possible to reduce the costs by spreading the R&D costs over more than one nation, thereby receiving national research funding for various high capital components.
On the other hand, since the aircraft is below the $150 million fly-away cost, the above-mentioned possibilities will be necessary if and only if costs need to be reduced even further.
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11. Manufacturing and Maintenance
The SW-Ghost concept must use the role of advanced methods of manufacturing to find a good balance between structural performance and cost that must be attained in order for the aircraft to be responsive to the RFP.
Manufacturing of the SW-Ghost is not only a main driver of this design, but also has an important effect on cost and weight. To reduce weight, titanium castings will be used within the structure of the aircraft, eliminating mechanical joints and reducing material costs along with machine time. The wing of the SW-Ghost will contain spars comparable to the F-22 that are produced by a method called “resin-transfer molding”. This advanced manufacturing process produces complex composite parts that reduces cost and improves quality and consistency. Also being incorporated on the spars is a “sine- wave” design, which provides more structural strength and is lighter than the original “I-beam design.” The wing spars will alternate materials between composites and titanium alloys for a more reinforced structure that will be more survivable in combat situations. For most parts, a welding method will be applied to hold the aircraft together instead of bolts and rivets, which tend to fail due to high vibrations. Welding methods are extremely weight efficient and reduce the use of traditional fasteners. Drilling techniques are more advanced with a laser-guided method, which reduces cost effectively by not needing expensive tools for manual drilling while ensuring more quality. However, a fastener method will be applied to the removable sections of the aircraft so that they can be easily detached and replaced with ease.
Another excellent feature of the SW-Ghost will entail the use of many interchangeable parts not only within itself, but also with the entire fleet of bombers.
The SW Ghost contains three removable sections called “pods” which contain the engines, and the weapons bays.
The purpose for the design of these pods is to expedite deployment and the repair process. If the aircraft returns from a mission and needs to receive a new payload, there can be an extra pod waiting to be switched in a matter of minutes, instead of each weapon being placed individually in the aircraft one at a time. For propulsion, if there is an engine problem, there will be an extra engine pod waiting to be switched in with full hookups to inlet and fuel connections. The aircraft contains a total of three pods, one for the armament, and two engine pods containing two engines each. The aircraft will now be fully operational, with minimal down time, ready for mission deployment. These pods can make maintenance and compatibility a swift, efficient way to save time and deploy the SW-Ghost’s in combat.
The aircraft is seven feet off the ground where the pods are contained, and therefore are easily accessible with minimal and simplistic equipment required. The weapon pod contains both the bomb bay and the missile bay, customizable to all combat configurations. With there being three smaller pods, as opposed to one large one, it will increase the ease of
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removal and decrease the cost of equipment needed to remove them. Because all of the pods are very similar in size, the same basic equipment can be used to remove all three pods. The dimensions on the engine pods are approximately 28 feet long by 13 feet wide by 7 feet high; similarly the weapon pod is 30.5 feet long by 13.5 feet wide by 3 feet high (figure 11.1).
FIGURE 11.1: Weapon and engine pods
The engine pods are removed by releasing the aft section onto the removal cart and lowering only that side of the removal cart. The cart then moves toward the rear of the aircraft, thus releasing the forward section of the detachable pod.
The entire pod is then lowered, and removed. The weapons pod is removed in a similar fashion except both the front and rear sections are released simultaneously, and the whole pod itself is lowered and transported out the rear of the aircraft. (figure
11.2, 11.3)
FIGURE 11.2: Engine pod removal
FIGURE 11.3: Weapon pod removal
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Another feature of the SW Ghost is a simple pilot access hatch. The pilots will enter through a hatch in the bottom of the aircraft, which is approximately 2.3 feet wide. A conventional ladder can be used for the pilot to climb 16 feet to the flight deck. This will provide an easy access for not only both pilots, but also for repair crewmembers to maintain and repair avionic instrumentation. (figure 11.4)
FIGURE 11.4: Crew access to flight deck
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12. References
1. American Institute of Aeronautics and Astronautics. 2001/2002 AIAA Foundation: Undergraduate Team Aircraft Design Competition. Request for Proposal. http://www.aiaa.org/membership/index.htm?mem=2
2. Mason, W. www.aoe.vt.edu/aoe/faculty/Mason_f/MRsoft.html [cited 11 February 2002].
3. AeroDYNAMIC
4. Arnaiz, Henry H. “Flight-Measured Lift and Drag Characteristics of a Large, Flexible, High Supersonic Cruise Airplane.” NASA TM X-3532, NASA, Washington DC, May, 1977.
5. Raymer, Daniel P. Aircraft Design: A conceptual Approach 3rd edition. American Institute of Aeronautics and Astronautics: Reston, Virginia. 1999.
6. Roskam, Dr. Jan. Airplane Design: PART VIII: Airplane Cost Estimation: Design, Development, Manufacturing, and Operating. Roskam Aviation and Engineering Corporation: Ottawa, Kansas. 1989.
7. Mason, W. www.aoe.vt.edu/aoe/faculty/Mason_f/ConfigAeroSupersonicNotes.pdf, February 2002.
8. Roskam, Jan Dr. Airplane Design Part III: Layout Design of Cockpit, Fuselage, and Empennage: Cutaways and Inboard Profile. Roskam Aviation and Engineering Corporation, 1989.
9. MatWeb, www.matweb.com/searchindex.htm, February 2002.
10. ATSE, www.atse.org.au/publications/focus/focus-polmear.htm, February 2002.
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