2001/2002 aiaa undergraduate team aircraft designmason/mason_f/swghost.pdf · 2002 slob works, a...

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Bomber of the 21st Century and Beyond

2001/2002 AIAA Undergraduate Team Aircraft Design

2002 SLOB Works, a Virginia Tech AOE Design Team 1

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


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.


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


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


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


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


Index of Abbreviations AEW AGM AIM AMRAAM CAD cg GBU GCI JDAM JSOW LDGP MAC nm RDTE RFP RCS SLOB SW TLFC TOP USAF

Anti-Electronic Warfare Air-to-Ground Missiles Air Intercept Missiles Advanced Medium-Range Air-to-Air Missile Computer Aided Design Center of Gravity Guided Bomb Unit Ground Control Intercept Joint-Direct Attack Munition Joint-Stand Off Weapon Laser Designated General Purpose Mean Aerodynamic Chord nautical miles Research, Development, Test, & Evaluation Request for Proposal Radar Cross Section Supercruising, Low-Observable Bomber SLOB Works Thermal Laminar Flow Control Take-off Parameter United States Air Force

Index of Symbols A C CD0 CL CLmax CLTO D E e g Κ L n Ps q R S SLanding T TTO W We Wf Wo Wp WTO V ρ σ

Apect Ratio Specific fuel consumption Zero lift, drag coefficient Lift coefficient Maximum lift coefficient Lift coefficient at take-off Drag Endurance, Loiter time Oswald efficiency factor Gravity Induced drag coefficient Lift g-loading Specific power Dynamic pressure Range Wing area Landing distance Thrust Thrust at take-off Weight Empty weight Fuel weight Take-off gross weight, TOGW Weight of payload Weight at take-off Velocity Density Density Ratio


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


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.


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

Take-Off

Warm-Up Climb

Best Cruise Altitude1000 nmM = 1.6

Dash and Store Drop

50,000 ft1500 nmM = 1.6

Dash-in Dash-out

180° Turn

Internal Store Drop

Best Cruise Altitude1000 nmM = 1.6

Landing

Loiter30 min

Descent

No Distance Credit No Distance or Fuel Credit

3,500 nm


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.


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

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.

FIGURE 1.2: Concept SW-1


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

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



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.


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.

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.



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


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.



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.


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.


FIGURE 2.2: XB-70 CD0 versus Mach Number Comparison

FIGURE 2.3: XB-70 CL vs. CD Comparison at Mach 1.6

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18

CD

CL

XB-70 AeroDYNAMIC Data

XB-70 Flight Data

0

0.005

0.01

0.015

0.02

0.025

0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40 1.60 1.80 2.00

Mach

Cdo


XB-70 Flight Data


FIGURE 2.4: XB- 70 Coefficient of Lift Curve Comparison at Mach 1.6

0

0.005

0.01

0.015

0.02

0.00 0.50 1.00 1.50 2.00

Mach Number

CD

0

XB-70SW-1SW-2SW-3SW-4

FIGURE 2.5: Adjusted CD0 vs. Mach Number for Concepts

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

-2 3 8 13 18

Alpha (deg)

CL


XB-70 Flight Data


-0.2

0

0.2

0.4

0.6

0.8

1

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35

CD

CL


FIGURE 2.6: CL vs. CD for Concepts at Mach 1.6

FIGURE 2.7: Lift Curves for Concepts at Mach 1.6

-0.1

0.1

0.3

0.5

0.7

0.9

-3 2 7 12 17

α (degrees)

CL


CL


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

edition5. 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.

99.097.00

1 −=WW (warm-up, taxi, & takeoff) (2.1)

2

1

2 01.0007.0991.0 MMWW

−−= (climb & accelerate) (2.2)

)(

2

3 DLV

RC

eWW

−

= (cruise) (2.3)

)(

3

4 DLV

RC

eWW −

= (dash in) (2.4)

4

4

4

5 )8700(W

WWW −

= (weapons drop) (2.5)

)(

5

6 DLV

RC

eWW

−

= (dash out) (2.6)

)(

6

7 DLV

RC

eWW −

= (cruise) (2.7)

)(

7

8 DL

EC

eWW −

= (loiter) (2.8)

995.0990.08

9 −=WW (descend) (2.9)


Concept SW-1 Concept SW-2 Concept SW-3 Concept SW-4

1,204,758,987$ 1,924,266,243$ 1,259,118,957$ 1,708,459,452$

402,369,134$ 674,639,444$ 422,452,782$ 591,637,639$

2,058,404,632$ 3,241,701,052$ 2,132,842,432$ 2,867,835,179$

343,741,283$ 683,071,335$ 366,724,106$ 573,728,015$

6,912,541,441$ 11,247,720,816$ 7,208,859,096$ 9,899,414,283$ CRDTE

Cftor

Flight Test AirplanesCftar

Flight Test Operations

Airframe Engr & Design Caedr

Develop Support & Testing Cdstr

SegmentWarm up/Takeoff 0.970

Climb – Accelerate 0.954Cruise 0.859Dash in 0.891

Drop Weapons 0.904Dash out 0.891Cruise 0.859Loiter 0.941

Descend 0.993Land 0.995

Wdescend/Wloiter

Wland/Wdescend

Wdrop/Wdashin

Wdashout/Wdrop

Wcruise2/Wdashout

Wloiter/Wcruise2

Wwarmup/takeoff/Winitial

Wclimb/Wwarmup/takeoff

Wcruise1/Wclimb

Wdashin/Wcruise1

997.0992.09

10 −=WW (land) (2.10)

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

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)



13,935,401,171$ 15,199,508,655$ 14,017,063,908$ 14,804,392,227$

1,393,540,117$ 1,519,950,865$ 1,401,706,391$ 1,480,439,223$

18,034,048,574$ 19,669,952,377$ 18,139,729,763$ 19,158,625,235$

ManufacturingCMAN

Profit by ManufacturerCPRO

CACQ

C o n c e p t S W - 1 C o n c e p t S W - 2 C o n c e p t S W - 3 C o n c e p t S W - 4

3 , 5 6 0$ 3 , 5 6 0$ 3 , 5 6 0$ 3 , 5 6 0$

3 , 1 6 9 , 9 5 7 , 5 0 0$ 3 , 1 6 9 , 9 5 7 , 5 0 0$ 3 , 1 6 9 , 9 5 7 , 5 0 0$ 3 , 1 6 9 , 9 5 7 , 5 0 0$

1 , 3 4 4 , 9 2 7 , 1 9 9$ 8 4 1 , 9 4 6 , 2 9 5$ 8 4 1 , 9 4 6 , 2 9 5$ 8 4 1 , 9 4 6 , 2 9 5$

2 8 2 , 0 1 8 , 7 5 0$ 2 8 2 , 0 1 8 , 7 5 0$ 2 8 2 , 0 1 8 , 7 5 0$ 2 8 2 , 0 1 8 , 7 5 0$

1 , 9 7 2 , 5 5 9 , 8 9 1$ 1 , 2 3 4 , 8 5 4 , 5 6 6$ 1 , 2 3 4 , 8 5 4 , 5 6 6$ 1 , 2 3 4 , 8 5 4 , 5 6 6$

1 , 8 8 2 , 8 9 8 , 0 7 8$ 1 , 1 7 8 , 7 2 4 , 8 1 3$ 1 , 1 7 8 , 7 2 4 , 8 1 3$ 1 , 1 7 8 , 7 2 4 , 8 1 3$

3 1 3 , 8 1 6 , 3 4 6$ 1 9 6 , 4 5 4 , 1 3 6$ 1 9 6 , 4 5 4 , 1 3 6$ 1 9 6 , 4 5 4 , 1 3 6$

8 , 9 6 6 , 1 8 1 , 3 2 5$ 5 , 6 1 2 , 9 7 5 , 3 0 1$ 5 , 6 1 2 , 9 7 5 , 3 0 1$ 5 , 6 1 2 , 9 7 5 , 3 0 1$

C D E P O T S

M i s c e l l a n e o u sC M I S C

C O P S

C C O N M A T

S p a r e sC S P A R E S

D e p o t s

C D I R P E R SP r o g r a m I n d i r e c t P e r s o n n e l

C I N D P E R S

C o n s u m a b l e M a t e r i a l

P r o g r a m F u e l , O i l , & L u b r i c a n t sC P O L

P r o g r a m D i r e c t P e r s o n n e l


COPS 342,553,246$ 368,996,449$ 312,743,072$ 350,212,271$


6,912,541,441$ 11,247,720,816$ 7,208,859,096$ 9,899,414,283$

18,034,048,574$ 19,669,952,377$ 18,139,729,763$ 19,158,625,235$

8,966,181,325$ 5,612,975,301$ 5,612,975,301$ 5,612,975,301$

342,553,246$ 368,996,449$ 312,743,072$ 350,212,271$

34,255,324,587$ 36,899,644,942$ 31,274,307,232$ 35,021,227,090$

121,099,952$ 150,085,792$ 123,051,402$ 141,058,444$

CDISP

Aircraft Estimated Price (AEP)

Life Cycle Cost (LCC)

RDTE CostCRDTE

Acuqisition CostCACQ

Operational CostCOPS

Disposal Cost

TABLE 2.3: Acquisition Program Cost (CACQ)

TABLE 2.4: Operating Program Cost (COPS)

TABLE 2.5: Disposal Program Cost (CDISP)

TABLE 2.6: Life Cycle Cost (LCC)


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.

vsC

Co

C

o

CCo

o

e KMS

WWTAbWa

WW

+= 5

max

4321 (2.11)

Then using equation 2.125 values for take-off gross weight were found.

oo

efuelloaddroppedpayadfixedpaylocrewo W

WWWWWWW

++++= (2.12)

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.

Equation 2.135 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.

aL

landing SCS

WS +

=

max

180σ

(2.13)

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.

( )

( )WTCS

WTOP

TOLσ= (2.14)


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.

−−

=

SW

qKn

SWqC

WTVPs oD 2 (2.15)

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.

V

SWqCg L 1))//(( 2 −=ϕ (2.16)

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.

( ) ( )AeqS

W

SW

qCWTD

L

oD

π+

==11

(2.17)

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.

=

takeoff

erest

erest

takeoff

erstofpotakeoff WW

TT

WT

WT int

intintint (2.18)

eresterest

takeoff

takeoff SW

WW

SW

intint

=

(2.19)

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.


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.

162000

164000

166000

168000

170000

172000

174000

176000

TOG

W (l

bs.)

T/W = .35

2g Ps = 0 ft/s Constraint

T/W = .6

T/W = .55

T/W = .4

T/W = .45

T/W = .5

W /S = 60

W /S = 70

W /S = 80

W /S = 90

W /S = 100

W /S = 50

Instantaneous Turn Constraint

Takeoff Constraint

Landing Constraint

Dash Constraint

1 g Ps = 200 ft/s Constraint

FIGURE 2.8: Carpet Plot of SW-1


250000

252000

254000

256000

258000

260000

262000

264000

266000

268000

TOG

W (l

bs.) T/W = .35

T/W = .4T/W = .45

T/W = .5

T/W = .55

T/W = .6 W /S = 50

W /S = 60

W /S = 70

W /S = 80

W /S = 90

W /S = 100


Takeoff Constraint

Landing Constraint


Dash Constraint



170000

172000

174000

176000

178000

180000

182000

184000

186000

TOG

W (l

bs.)

T/W = .35

T/W = .4

T/W = .45

T/W = .5

T/W = .55

T/W = .6 W /S = 50

W /S = 60

W /S = 70

W /S = 80

W /S = 90

W /S = 100


Takeoff Constraint

Landing Constraint


Dash Constraint




192000

194000

196000

198000

200000

202000

204000

206000

208000

210000

TOG

W (l

bs.) T/W = .35

T/W = .4

T/W = .45T/W = .5

T/W = .55

T/W = .6W /S = 50

W /S = 60

W /S = 70

W /S = 80

W /S = 90

W /S = 100


Takeoff Constraint

Landing Constraint



Dash Constraint


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.


TABLE 2.7: Concept Design Matrix

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

Data Ranking Rating Data Ranking Rating Data Ranking Rating Data Ranking RatingCruise Mach 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 5Fuel 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 TypePropulsion System

Total Installed Thrust (lbs) 52,760 -- -- 105,520 -- -- 52,760 -- -- 79,140 -- -- 60,000 -- Control SystemHigh Lift System

Weapon Bay Configuration2 Bays

(Missle - Bomb Config)

-- -- 2 Bays


-- -- 2 Bays


-- -- 2 Bays


-- -- -- --

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 1L/ 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=0.88 0.0114 4 16 0.0104 2 8 0.0100 1 4 0.0113 3 12 4C 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 3T/ 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

Cost ($) 121.1M 1 5 150.1M 4 20 123.1M 2 10 141.1M 3 15 150M 5Average 8 11.38 5.25 11.63

Rating of Importance

SW-4Approximations

SW-1 SW-2 SW-3


SW-1 SW-2 SW-3 SW-4Rank Rank Rank Rank

Propulsion Systems

Average 1.778 1.556 2.000 1.667Ranking 3 1 4 2

1

1 1

2 2 2

2Manufacturability

Maintenance

Stealth

AIRCRAFT CONCEPTS

Design

2

2

3

1

3

2 2 2

1

Aerodynamics

Stability & Control

1

2

1

2 3 1

1

2

2

2

1

3

Inlets

Supercruising

3

1

1

1 1

12

3

Structures

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

preliminary design phase came from the programs

mentioned above. With a lot of emphasis placed on the

aerodynamic and performance characteristics through the

RFP, a more detailed version of the technical data is

mentioned and described in the next section. However,

with a total aircraft comparison, aerodynamics was

analyzed and rated. Concept SW-1 was rated as one of

the highest due to its low weight and swept wings. The

swept wings were determined at an angle of 55 degrees so

that it would meet both the takeoff and landing

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

TABLE 2.8: Risk Management Matrix


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


121,099,952$ 150,085,792$ 123,051,402$ 141,058,444$

1 4 2 3

Aircraft Estimated Price (AEP)

Ranking

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


SW-1 SW-2 SW-3 SW-4Rank Rank Rank Rank

Average = (3R+2T+C) / 6 2.333 2.167 2.667 2.833

Final Rankings 2 1 3 4

AIRCRAFT CONCEPTS

Trade Study/Technical Data 2 3 1 4

3

Risk Analysis 3 1 4 2

Cost Analysis 1 4 2

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.

SLOB Works final design concept,

Concept SW-2, will know be known as SW-

Ghost. This aircraft was further investigated

in the preliminary design phase for more

precise data, in all aspects of the design.

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

TABLE 2.10: Final Decision Selection


FIGURE 2.12: Evolution of Aircraft

Concept SW-2 Refinement 1 - Overall aircraft refinement

Refinement 3 – Inlets & engines detailed, landing gear finalized

Final, SW-Ghost

Refinement 2 – Wing root extended, V-tails split, aircraft area ruled


Figure 2.13 Top view of SW-Ghost


Figure 2.14 Bottom view of SW-Ghost


Figure 2.15 Side view of SW-Ghost


Figure 2.16 Front view of SW-Ghost


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


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 12.346Climb 0.8 0.0190 0.00934 0.1817 0.23 12.136Cruise out 1.6 0.0230 0.0136 0.345 0.165 7.299Dash out 1.6 0.0311 0.0141 0.345 0.222 7.168Dash in 1.6 0.0258 0.0141 0.345 0.184 7.168Cruise in 1.6 0.0228 0.0141 0.345 0.159 7.168Loiter 0.4 0.0136 0.00847 0.1841 0.168 12.658Landing 0.278 0.04269 0.00889 0.1845 0.428 12.346


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.

]3.01)[cos37.074.0()(5.4

max2max

CDoLEWDDwave MMEl

AS

C −−Λ+=π

(3.1)

where, LE

CDoMΛ

= 2.0max cos1

(3.2)

and 5.1≈WDE 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.


Parasite-Drag Buildup @ 50k feet

0

0.002

0.004

0.006

0.008

0.01

0.012

0.014

0.016

0.018

0.02

0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

Mach Number

Cdo

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.

2LDD KCCC

O+= (3.3)

In equation 3.3 the K for subsonic flight was found using equation 3.4,

0100 )1( KSSKK −+= (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.

Skin friction drag

Wave drag

Form drag

Leaks and protuberances

Miscellaneous


Drag Polar @ M=1.6 and 50k ft

0

0.1

0.2

0.3

0.4

0 0.01 0.02 0.03 0.04 0.05 0.06

Drag Coefficient

Lift

Coe

ffic

ient

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.

L/Dmax = 7.30


Lift / Drag @ Mach 1.6

0

1

2

3

4

5

6

7

8

0 0.05 0.1 0.15 0.2 0.25 0.3

Lift Coefficient

Lift

/ D

rag

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

maximum area Amax, and the corresponding location can be found. Amax is 157.8 ft2 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.

πmax4A

lf = (3.5)

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.

L/Dmax


Area Distribution @ M=1.01

0

20

40

60

80

100

120

140

160

180

0 20 40 60 80 100 120 140 160

x Location (ft)

Are

a (f

t^2)

TotalFuselageWingPodV. TailCanard

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

maximum deflection configuration of 40 degrees, CLmax is increased by 0.157 and ∆αOL is –2.5O. 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

lift angle of attack can be decreased by –2.2O. 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.


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 ExpansionDesignation 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 12.5-13.17S steel:AISI 410 0.28 20.31-23.64 37.27-48.01 28.5-59.11 27.56-30.46 5.0-6.11MMC: Cerme-Ti 0.16 17.4-18.13 45.69-48.73 37.96-57.83 15.95-17.4 4.37-4.71Mg 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.


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


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


FIGURE 4.3: Structures Side View/ Main Bulkheads

FIGURE 4.4: Structures Bottom View/ Removable Pods


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


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.


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 -0.00094Clq 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.


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,

34%

36%

38%

40%

42%

44%

46%

48%

50%

0.4 0.6 0.8 1 1.2 1.4 1.6

Mach #

% M

AC


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

SW-Ghost

MIL-F-8785 B Requirements (Class II,

Cat. B, Level 1) Subsonic Supersonic Damping ξsp > 0.15

Short Period Natural Frequency 0.1 rad/s < ωsp < 2.0 rad/s

no short period: controlled with fly-by-light system

Phugoid Damping Τ2 > 55 sec 55.46 sec 213.75 sec Damping ξd > 0.02 0.0473 0.222

Dutch Roll Natural Frequency ωnd > 0.4 rad/s 0.800 rad/s 2.830 rad/s

Spiral Roll Minimum time to double amplitude 4 sec 4600 sec 5950 sec

Rolling Convergence

Maximum time constant 1.4 sec 0.897 sec 0.097 sec


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.


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


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.

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.

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 6.3: Effective envelope of the K-36D ejection sea22(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.


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.


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.


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


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.


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.


FIGURE 6.9: Combat loads of the SW Ghost

FIGURE 6.10: Side and bottom views of weapons bays and their clearances


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.


FIGURE 6.11: Top view of defensive system locations


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.

LengthNEW = LengthOLD * (TREQ /TBASE)0.4 (7.1)

DiameterNEW = DiameterOLD * (TREQ /TBASE)0.5 (7.2)

WeightNEW = WeightOLD * (TREQ /TBASE)1.0 (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.


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:

TREQ = ½ * ρ * V2 * 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.


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).


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.


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

2

21

iAT

fAT

AG VKK

VKKn

gKS l (8.1)

µ−

=WTKT (8.2)

( )( )2

2LDOLA KCCC

SW

K −−= µρ (8.3)

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:


=

f

i

WW

DL

CVR ln (8.4)

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) / SFC50000 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

time spent in the areas of high drag. Using Raymer5, the Ps plot containing lines of constant energy was created as seen

below.


0

10000

20000

30000

40000

50000

60000

0.0 0.5 1.0 1.5 2.0

Mach No.

Alti

tude

(ft)

5

10

20

30

40

5060

80100 120

140

Ps = 100

Ps = 0

Ps = 200

Ps = 300

q limit = 2133 psf

Temp. limit = 150 F

Inlet Pressure Limit

stall limit

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.


0

10,000

20,000

30,000

40,000

50,000

60,000

70,000

0.00 0.50 1.00 1.50 2.00 2.50

Mach No.

Alti

tude

(ft)

60

5

4

3

2

1

5

80

1012

14

16

q limit = 2133

Inlet pressue limit

Ps = 0

Ps =

Ps = 200

Ps = 300

Ps = 400

Ps = 500

stall limit

Temp limit = 150 F

FIGURE 8.2: Ps plot for n=1, 50% fuel weight, and Mil. Thrust

0

10000

20000

30000

40000

50000

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

Mach No.

Alti

tude

(ft)

Temp limit = 150 F

q limit = 2133 psf

inlet pressure limit

Ps = 0

Ps = 100

Ps = 200

Ps = 500

Ps = 400

Ps = 300

stall limit



0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

22000

24000

26000

28000

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

Mach No.

Alti

tude

(ft)

Temp limit = 150 F


q limit = 2133 psfPs = 0

Ps = 100

Ps =

stall limit


0

10000

20000

30000

40000

50000

60000

70000

0 0.5 1 1.5 2 2.5 3

Mach No.

Alti

tude

(ft)

Ps = 200

Temp limit = 150 F

q limit = 2133 psf


stall limit

FIGURE 8.5: Ps plot for n=1, 50% fuel weight, and afterburners


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:

=

f

i

WW

CDLE ln1

(8.5)

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.


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.

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

450 550 650 750 850 950 1050 1150 1250 1350 1450 1550 1650 1750 1850 1950

Speed (ft/s)

Turn

rate

(deg

/s)

n= 7

n= 6

n= 5

n= 4

n= 3

n= 2

Stall LimitStructural Limit

Sustained TurnRate Envelope

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.


FIGURE 8.7: Sustained turn rate at Sea level

FIGURE 8.8: Sustained turn rate at 15,000 ft

0123456789

101112131415161718192021222324252627282930

300 350 400 450 500 550 600 650 700 750 800 850 900 950 1000

1050

1100

1150

1200

1250

1300

Speed (ft/s)

Turn

rate

(deg

/s)

n= 5n= 6

n= 7

n= 4

n= 3

n= 2

Structural Limit

Stall Limit

Sustained TurnRate Envelope

0

2

4

6

8

10

12

14

16

18

20

22

24

26

28

30

32

34

36

38

40

200 250 300 350 400 450 500 550 600 650 700 750 800 850 900 950 1000

1050

1100

1150

1200

Speed (ft/s)

Turn

rate

(deg

/s)

n= 3

n= 2

n= 4n= 5

n= 6

n= 7

Structural LimitStall Limit

Sustained Turn Rate Envelope


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.

0

10000

20000

30000

40000

50000

60000

0.00 0.50 1.00 1.50 2.00

Mach (M, -)

Alti

tude

(h, f

t)

n = 1

n = 3

n = 5

n = 7

pressure limit

q limit

temp limit

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.


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

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.

132,443 3,542 4.8 Total

98,857 8,000 - - Reserve & Trapped

106,857 3,772 - 0.500 Loiter @ 10K

110,629 687 0.815 (Dry) 0.004 Landing

111,316 22,813 1,000 1.090 Cruise back @ 59K

134,129 19,950 750 0.819 Dash back @ 59K

154,079 450 18.86 (circumference)

0.021 180 deg turn / Ordnance Drop @

50K

163,229 24,549 750 0.819 Dash out @ 59K

187,778 37,222 1,000 1.090 Cruise out @ 50K

225,000 7,500 17.7 0.45 Climb

232,500 7,500 0.560 (Dry) 0.004 Take off

240,000 - - - MTOGW

Total Weight

(lbs)

Amount of fuel burned (lbs)

Distance Traveled (nm)

Time to complete

(hrs)

MISSION SEGMENT


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

initial value empty weight, Wo, of the aircraft an approximation was done using Raymer’s approximate weights methods5.

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


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


TABLE 9.6: Weight Summary and Inertias of SW-Ghost

Weight (lbs) xcg (ft) zcg (ft) Ixx

(slugs-ft2) Iyy

(slugs-ft2) Izz

(slugs-ft2) Izx

(slugs-ft2) 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.


C.G. Movement

100,000

120,000

140,000

160,000

180,000

200,000

220,000

240,000

60.00 70.00 80.00 90.00 100.00 110.00 120.00 130.00

C.G. Location from Nose (ft)

Wei

ght (

lbs)

Fuel ControlStandard Burn

0 % 10 % 20 % 30 % 40 % 50 % 60 % 70 % 80 % 90 % 100 %

M.A.C

at TOGWTake-off

Climb to 50,000 ft

Supercruise at 50,000 ft

Dash-out at 59,5000

180 deg turn; Bomb & Missle

Dash-back at 59,500

Supercruise at 59,500 ft Loiter at 10,000 ft

Land

Max. Aft C.G. Location

Max. Foward C.G. Location

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.


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


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

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

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 10.4: 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


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.

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

RDT&E

Acquisition

Operating

Disposal


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.


000E+00

050E+06

100E+06

150E+06

200E+06

250E+06

300E+06

350E+06

0 200 400 600 800 1000

Units

Cos

ts ($

)

Unit CostsFly-Away Costs

FIGURE 10.2: Unit and fly-away cost trade study in Year 2000 dollars

0%10%20%30%40%50%60%70%80%90%

100%

Phases

100 300 500 700 900

Units

DisposalOperationAquistionRDTE

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.


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.


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


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


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


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.

11. Titanium Industries, www.titanium.com/brochure/tech_manual/techl.cfm, February 2002.

12. Van Diermen, W. www.f-22raptor.com/index_ie.htm, February.

13. McDonnell Douglas Astronautics Company – St. Louis Division. “The USAF Stability and Control Digital

DATCOM: Volume 1, User’s Manual.” Air Force Wright Aeronautical Labs, Wright-Patterson AFB, OH, April, 1979.

14. Blake, W.B. “Prediction of Fighter Aircraft Dynamic Derivatives Using Digital DATCOM.” AIAA Paper 85-4070,

U.S. Air Force Wright Aeronautical Labs, Wright-Patterson AFB, OH, October, 1985. 15. Mason, Durham, Lutze, Benoliel. “Control Authority Issues in Aircraft Conceptual Design.” Department of

Aerospace and Ocean Engineering, Virginia Polytechnic Institute and State University, Blacksburg, VA. November 1996. http://www.aoe.vt.edu/aoe/faculty/Mason_f/VPI-Aero-200.pdf

16. Melin, Thomas. “Tornado: A Vortex Lattice Program.” The Royal Institute of Technology, KTH, Stockholm,

Sweden. http://www.flyg.kth.se/Tornado/htm/tornado.htm 17. Grasmeyer, Joel. “Stability and Control Derivative Estimation and Engine Out.” Department of Aerospace and

Ocean Engineering, Virginia Polytechnic Institute and State University, Blacksburg, VA. January, 1998. http://www.aoe.vt.edu/aoe/faculty/Mason_f/Ldstabdoc.pdf

18. Valery Razgonyaev and W.H. Mason. “An Evaluation of Aerodynamic Prediction Methods Applied to the XB-70 for

Use in High Speed Aircraft Stability and Control System Design.” AIAA Paper 95-0759, 33rd AIAA Aerospace Sciences Meeting And Exhibit, Reno, NV, January 12, 1995.

19. Etkin and Reid. Dynamics of Flight. John Wiley & Sons Inc., New York, 1996.


20. Lutze, Frederick H. “Dynstable Computer Code.” Department of Aerospace and Ocean Engineering, Virginia Polytechnic Institute and State University, Blacksburg, VA. http://www.aoe.vt.edu/~lutze/AOE3134/

21. Raytheon. http://www.raytheon.com/

22. http://www.ejectionseat.com/k-36tm1.html

23. HydroAire. http://www.hydroaire.com/products/Fuel.htm

24. Federation of American Scientists. http://www.fas.org/man/index.html

25. Roskam, Dr. Jan. Airplane Design: PART VII: Determination of Stability, Control and Performance Characteristics: FAR and Military Requirements. Roskam Aviation and Engineering Corporation: Ottawa, Kansas. 1989.

26. Roskam, Dr. Jan. Airplane Design: PART V: Component Weight Estimation. Roskam Aviation and Engineering

Corporation: Ottawa, Kansas. 1989.

27. Burns, J. W. “Aircraft Cost Estimation Methodology and Value of a Pound Derivation for Preliminary Design Development Applications”. SAWE Journal. Fall 1994

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