TEAM Four Critical Design Review. Kai Jian Cheong Richard B. Choroszucha* Lynn Lau Mathew Marcucci Jasmine Sadler Sapan Shah Chongyu Brian Wang

TEAM Four Critical Design Review Kai Jian Cheong Richard B. Choroszucha* Lynn Lau Mathew Marcucci Jasmine Sadler Sapan Shah Chongyu Brian Wang 03.XII.2008

0.1 Abstract The purpose of this report is to present the final configuration and design of the cargo plane designed for use with the Federal Emergency Management Agency (FEMA) for relief and transoceanic missions. In addition, we will quickly recall the preliminary configurations and explain the downselect process with key changes to our final design. FEMA is in need of a long range cargo jet to transport emergency relief supplies with the capability of short take off and landing (STOL). The plane is required to fly at cruise speeds of about Mach 0.80 while carrying 60,000 pounds (lbs) of cargo for a 500 nautical mile (nm) relief mission. It should fly without the cargo for a transoceanic mission of at least 3,200 nm. A very important constricting factor is the ability to take off and land within a 2,000 to 3,000 feet (ft) critical field length on wet surfaces which may or may not be paved. The jet s fuselage must be able to support and hold a cargo container that is 560 inches (in) long by 128 in wide by 114 in high. 0.2 Executive Summary Spinning Cube Aviation designed three individual configurations for use with FEMA and USAF during the preliminary design stage. The design chosen was configured for FEMA due to a higher profit margin of 135 million dollars. The overall design of the aerocraft consists of a standard circular cross-section fuselage, twin engine propulsion system, T-tail empennage design, rear cargo bay door, and quadricycle landing gear system. The key design additions from the preliminary designs to the final included the use of a quadricycle landing gear, both fuel tanks contained inside the fuselage, a longer and more slender nose, and an all moving T-tail system. Our final FEMA cargo jet fulfills and surpasses all requirements set by the Federal Aviation Administration (FAA): One Engine Inoperative (OEI), lateral and longitudinal tip over, and controllable static margin. Some of the key risk factors of the aerocraft include the all-moving tail and quadricycle landing gear system due to its complexity and maintenance issues. In addition to standard maintenance problems that could arise from the intricate system for both, there is also a higher cost and complexity in the manufacturing and production processes of the aerocraft. Spinning Cube Aviation has taken these issues into consideration while designing our final configuration for FEMA and decided it is still the most beneficial and efficient design for the aerocraft even though it is slightly more expensive. The all-moving tail will drastically improve the aerocraft s control and stability during flight, along with the increased ability to counteract high crosswinds. The added landing gear weight is also justified because this will help keep the aeroplane in full contact with the ground and increase lateral stability while on wet and uneven landing and takeoff surfaces. 1

Contents 0.1 Abstract.......................................... 1 0.2 Executive Summary.................................... 1 List of Figures.......................................... 3 List of Tables........................................... 4 List of Symbols.......................................... 5 1 Introduction and Mission 6 1.1 Mission Requirements................................... 6 1.2 Introduction......................................... 7 2 Candidate Configurations 9 2.1 Design Drivers....................................... 9 3 Down Selection 10 4 Design Comparison 11 5 Introduction to Final Design 12 5.1 Introduction to Final Design............................... 12 5.2 Aeroplane Specifications.................................. 12 6 Layout and Design 14 6.1 All-Moving Tail....................................... 14 6.2 Cockpit Layout....................................... 15 6.3 Cargo Bay Layout..................................... 15 6.4 Cargo Bay Door...................................... 16 7 Aerocraft Sizing and Sizing Trade Studies 19 7.1 Aerocraft Sizing Plot.................................... 19 7.2 Sizing Trade Study..................................... 19 8 Weight Estimates 22 8.1 Initial Weight Estimates.................................. 22 8.2 Sensitivity Studies..................................... 23 8.3 Refined Mission Fuel Fraction............................... 23 8.4 Refined Weight Estimates................................. 24 9 Center of Gravity Location and Excursion 25 9.1 CG Excursion During Mission............................... 25 9.2 Loading CG Excursion................................... 26 2

10 Wing Design and High Lift Systems 28 10.1 Aerofoil Selection...................................... 28 10.2 Wing Design Characteristics................................ 28 10.3 High Lift Systems..................................... 29 11 Empennage and Control Surfaces 31 11.1 Empennage Characteristics................................ 31 11.2 Control Surfaces...................................... 32 11.3 One Engine Inoperative Requirements.......................... 33 12 Landing Gear Design 34 13 Propulsion 36 14 Performance Analysis 38 14.1 Stall Speed......................................... 38 14.2 Takeoff and Landing.................................... 38 14.3 Cruise, Climb and Flight Ceiling............................. 39 14.4 Trimmed Lift and Drag.................................. 39 14.5 Stability........................................... 39 15 Aero Loads 41 15.1 Maneuver and Gust Loads................................. 41 15.2 Lifting Surface Loads.................................... 42 16 Load Path 44 17 Cost Analysis 46 18 Conclusions 47 Bibliography 48 A Configuration Comparison 49 B Mission Fuel Fraction 50 C Centers of Gravity 51 D Engineering Drawings 53 E One Engine Inoperative 59 F Distributed Aerodynamic, Inertia and Gravity Loading Plots 60 G Wing Aerofoil Trade Study 63 H MATLAB Code 80 H.1 Powell s Parasite Drag Coefficient Matlab Code..................... 80 H.2 Weights Sensitivity Matlab Code............................. 80 H.3 Preliminary Sizing Matlab Code............................. 83 H.4 Preliminary Static Margin Matlab Code......................... 84 H.5 Preliminary Weights Estimate Matlab Code....................... 85 3

H.6 Powell s Inviscid Flapped Wing Loading Code...................... 86 I Calculations 92 I.1 Calculations for Stall Speed................................ 92 I.2 Calculations for Climb Rate................................ 92 4

List of Figures 1.1 Relief Mission Profile.................................... 6 1.2 Transoceanic Mission Profile................................ 7 1.3 Isometric View of Final Design: FEMA Aerocraft.................... 8 6.1 Isometric View of the All-Moving Tail.......................... 14 6.2 View of the All-Moving Tail Deflected.......................... 15 6.3 View of the General Layout of the Cockpit....................... 16 6.4 Isometric View of Cargo in Cargo Bay.......................... 16 6.5 Engineering Drawing of Cargo in Cargo Bay....................... 17 6.6 Example of Cargo Bay Doors Opened.......................... 18 7.1 Sizing Plot......................................... 20 7.2 Trade Studies Matrix.................................... 20 7.3 Carpet Plot......................................... 21 9.1 CG Excursion During Mission............................... 25 9.2 CG Excursion At Loading................................. 26 10.1 NLF0215F Aerofoil View................................. 28 10.2 Detailed View of Wing Full Span............................. 29 10.3 General View of High Lift Systems............................ 30 10.4 Sectional Lift Coefficient According to Normalized Wing Span Location....... 30 11.1 Detailed CAD Top View of Control Surface Locations................. 32 12.1 Position of Landing Gear with Respect to Fuselage................... 34 12.2 Landing Gear Retraction Mechanism........................... 35 13.1 General Electric (GE) CF6-80E1A3 Engine....................... 36 14.1 Trimmed Lift and Drag Data from AVL......................... 40 15.1 V-n Diagram for Minimum Cruise Weight........................ 41 15.2 V-n Diagram for Maximum Cruise Weight........................ 42 15.3 Aero Loads on Wings................................... 43 16.1 Torsion Moment About the y-axis for Critical Points With Maximum Weight.... 44 D.1 Engineering Drawing of Assembled Plane........................ 53 D.2 Engineering Drawing of Fuselage............................. 54 D.3 Engineering Drawing of Nose............................... 55 D.4 Engineering Drawing of Tail................................ 56 D.5 Engineering Drawing of Wing............................... 57 5

D.6 Engineering Drawing of Fuselage Tail........................... 58 F.1 Normal Force Load Distribution............................. 60 F.2 Chordwise Longitudinal Load Distribution........................ 61 F.3 Spanwise Pitch Moment Distribution........................... 61 F.4 Inertia Load Distribution................................. 62 G.1 Trefftz Plot for α = 10................................. 64 G.2 Trefftz Plot for α = 09................................. 64 G.3 Trefftz Plot for α = 08................................. 65 G.4 Trefftz Plot for α = 07................................. 65 G.5 Trefftz Plot for α = 06................................. 66 G.6 Trefftz Plot for α = 05................................. 66 G.7 Trefftz Plot for α = 04................................. 67 G.8 Trefftz Plot for α = 03................................. 67 G.9 Trefftz Plot for α = 02................................. 68 G.10 Trefftz Plot for α = 01................................. 68 G.11 Trefftz Plot for α = 00.................................. 69 G.12 Trefftz Plot for α = 01.................................. 70 G.13 Trefftz Plot for α = 02.................................. 70 G.14 Trefftz Plot for α = 03.................................. 71 G.15 Trefftz Plot for α = 04.................................. 71 G.16 Trefftz Plot for α = 05.................................. 72 G.17 Trefftz Plot for α = 06.................................. 72 G.18 Trefftz Plot for α = 07.................................. 73 G.19 Trefftz Plot for α = 08.................................. 73 G.20 Trefftz Plot for α = 09.................................. 74 G.21 Trefftz Plot for α = 10.................................. 74 G.22 Trefftz Plot for α = 11.................................. 75 G.23 Trefftz Plot for α = 12.................................. 75 G.24 Trefftz Plot for α = 13.................................. 76 G.25 Trefftz Plot for α = 14.................................. 76 G.26 Trefftz Plot for α = 15.................................. 77 G.27 Trefftz Plot for α = 16.................................. 77 G.28 Trefftz Plot for α = 17.................................. 78 G.29 Trefftz Plot for α = 18.................................. 78 G.30 Trefftz Plot for α = 19.................................. 79 G.31 Trefftz Plot for α = 20.................................. 79 6

List of Tables 1 Table of Symbols...................................... 5 1.1 General Aerocraft Characteristics............................. 7 3.1 Table of Candidate Comparisons............................. 10 4.1 Comparison Aeroplanes for Final Blue Configuration.................. 11 5.1 Aeroplane Specifications.................................. 13 8.1 Mission Fuel Fractions for a Transport Mission..................... 22 8.2 Mission Fuel Fractions for Diversion........................... 22 8.3 Initial Weight Estimate.................................. 22 8.4 Sensitivity Studies..................................... 23 8.5 Refined Mission Fuel Fraction (Segment I)........................ 23 8.6 Refined Mission Fuel Fraction (Segment II)....................... 23 8.7 Refined Mission Fuel Fraction (Segment III)....................... 23 8.8 Refined Weights Estimates................................. 24 10.1 Wing Geometry Characteristics.............................. 29 11.1 Horizontal and Vertical Tail Characteristics....................... 31 11.2 Control Surface Locations................................. 33 12.1 Tire Dimensions...................................... 34 13.1 Engine Specifications.................................... 37 14.1 Takeoff and Landing Parameters............................. 38 14.2 Stability Derivatives due to Angle of Attack and Sideslip Angle............ 40 14.3 Stability Derivatives due to Roll Rate, Pitch Rate and Yaw Rate........... 40 15.1 Critical Loading Angles of Attack............................. 42 17.1 Program Costs....................................... 46 17.2 Market Pricing Strategy.................................. 46 A.1 Comparison Aeroplanes for Red Configuration..................... 49 A.2 Comparison Aeroplanes for White Configuration.................... 49 B.1 Mission Fuel Fraction at Various Stages......................... 50 C.1 CG Analysis......................................... 52 E.1 One Engine Inoperative.................................. 59 7

List of Symbols F: Degrees Fahrenheit AOA, α: Angle of Attack AR: Aspect Ratio AVL: Athena Vortex Lattice (computer program) CAD: Computer Aided Design C D : Drag Coefficient C D0 : Parasite Drag Coefficient CG: Center of Gravity C Lmax : Maximum Lift Coefficient DAPCA IV: Development and Production Cost for Aircraft FAA: Federal Aviation Administration FAR: Federal Aviation Regulation FEMA: Federal Emergency Management Agency ft: Feet GE: General Electric L/D: Lift to Drag lb: Pound-Mass MAC: Mean Aerodynamic Chord MATLAB: Matrix Laboratory (computer program) n: Load Factor nm: Nautical Miles OEI: One Engine Inoperative psf: pounds per square foot RDT&E: Research, Development, Test and Evaluation SFC: Specific Fuel Consumption S L : Landing Distance S T O : Takeoff Distance STOL: Short Takeoff and Landing T/W: Thrust to Weight Ratio TOP: Takeoff Parameter USAF: United States Air Force V-n: Velocity versus Load Factor W/S: Wing Loading W E : Empty Weight W P : Payload Weight W T O : Takeoff Weight Table 1: Table of Symbols 8

Chapter 1 Introduction and Mission 1.1 Mission Requirements The overall requirements for the cargo aerocraft given to us by FEMA for relief and transoceanic include the ability to fly at a Mach number of 0.80 or higher, the ability to takeoff on a balanced field length of 2,500 ft on a 95 F day, and the capability to warm up and taxi for 8 minutes. The crew includes two pilots and one loadmaster. There are also various requirements that are specific to the transoceanic and relief type missions. Relief missions must have a flight range of at least 500 nm, a descent to 1,000 ft for 100 nm at a Mach number of 0.6, landing distance of 3,000 ft with the ability to handle 25 knot crosswinds, and enough reserve fuel to accommodate an extra 350 nm missed approach along with a 30 minute hold at 5,000 ft. Transoceanic flights must have a flight range of 3,200 nm which can include an aerial refuel, a 2,500 ft landing distance or less, a descent to sea level, and enough reserve fuel to accommodate a missed approach of 150 nm and a 45 minute hold at 5,000 ft. In addition, the relief mission must be able to carry 60,000 lbs of payload including support equipment, and the transoceanic mission must be able to carry 10,000 lbs of bulk cargo with density of 20 lb/ft 3. The relief payload dimensions are 46.7 ft long by 10.7 ft wide by 9.5 ft high with a 1 ft clearance around the top and both sides of the payload. The relief and transoceanic mission profile can be seen below in Figure 1.1 and Figure 1.2. Figure 1.1: Relief Mission Profile A military version of this cargo plane was considered for the United States Air Force (USAF) that would have the capability to carry 120,000 lbs of payload with two pilots, a loadmaster, and 25 paratroopers. The takeoff distance must not exceed 5,500 ft on a 95 F day, the range must be 5,000 ft at Mach 0.80 or above, a landing distance must be less than 3,500 ft, and must hold enough reserve fuel to accommodate a 200 nm diversion with a 30 minute hold at 5,000 ft. This mission design was not carried out in final stage so no further information will be presented for this 9

Figure 1.2: Transoceanic Mission Profile aerocraft. 1.2 Introduction Spinning Cube Aviation has developed a final design to be used for FEMA for its purposes of relief mission cargo transport. We have arrived at a final design after down selecting from three previous preliminary designs that included: a military version of an aerocraft meant to carry extra payload and paratroopers, a twin engine configuration with T-tail and high wing, and a configuration with one engine on the vertical fin of the T-tail. Our design was centered on the mission requirements and profile provided by FEMA for transoceanic and relief mission flights. After cost analysis for both the FEMA and USAF mission aerocrafts, Spinning Cube came to the conclusion that it was beneficial for us to pursue the FEMA design for a higher profit margin compared to USAF aerocraft. Important characteristics of the FEMA aerocraft final design are presented in Table 1.1 below. More detailed explanations of all parameters listed in Table 1.1 as well as discussion of all aerocraft subsystems will be presented in the following chapters. Takeoff Weight (lbs) 243,500 Empty Weight (lbs) 131,557 Mach 0.85 Cruise Ceiling (ft) 45,000 Type of Engine General Electric CF6-80E1A3 Takeoff Thrust (lbs) 63,121 Wing Span (ft) 143 Aspect Ratio (AR) 9 Fuselage Length (ft) 123 Fuselage width (ft) 17.5 Range (nm) 1,924 Cost per plane (Millions) (2008 USD) 308 Table 1.1: General Aerocraft Characteristics This report discusses the final configuration of the design used for the FEMA mission aerocraft. The cargo jet consists of a high wing structure with two General Electric (GE) CF6-80E1A3 turbojet engines and a T-tail empennage design. The layout and design of the plane was meant to take the effective characteristics of historically successful planes while making more efficient changes to fulfill the special Short Take Off and Landing (STOL) requirements set forth by FEMA. An isometric view of the final design of the aerocraft can be seen in Figure 1.3 below. 10

Figure 1.3: Isometric View of Final Design: FEMA Aerocraft 11

Chapter 2 Candidate Configurations 2.1 Design Drivers To meet FEMA s need for a long range cargo jet to transport emergency relief supplies with STOL capability, Spinning Cube Aviation developed three different designs (Blue, Red, and White), all of which met these basic mission requirements. All three designs used a high wing structure and T-tail. The Blue design used a conventional design with two wing-mounted GE turbofan engines. This plane adopted the effective design characteristics from successful historical planes while still modifying details to fulfill the STOL requirements. The Red design was unique because it had a single engine mounted on the vertical fin in a pusher configuration. This design was to understand the effects from a high aspect ratio. The White design considered the USAF s need for a transport plane with STOL requirements. This design incorporated four wing-mounted engines and an expanded fuselage to carry additional payload weight from the combat vehicles and paratroopers. It was substantially larger, more expensive to produce, and the only one to be limited by the 200 ft wing span constraint. 12

Chapter 3 Down Selection The Blue plane was chosen as the final design for continued work and analysis based on configuration, cost, and feasibility. The White design did not fulfill the OEI requirement of the feasibility study and thus, was not considered for the final design. Both the Blue and Red designs met all of the criteria set by mission requirements and feasibility studies. The Red design was discounted due to the higher risk associated with the single-engine design during the transoceanic mission. Also, preliminary cost estimations showed the Blue design was less expensive than the Red design. An overall comparison between the three designs is shown below in Table 3.1. All margins are fulfilled for all designs. Blue Red White Final Config Mission (FEMA or Military) FEMA FEMA Military FEMA W E (lbs) 88949 99160 249660 88949 W T O (lbs) 198310 194462 550255 198310 Cruise Mach number 0.85 0.85 0.85 0.85 Price (Millions) (2008 USD) 197.5 223 330.9 197.5 RDT&E + Flyaway Cost (Billions) (2008 USD) 8.9 7.8 33.1 8.9 Takeoff Distance Margin () 4.5 7.2 6 4.5 Landing Distance Margin (w/o / with thrust rev. ) 0/+32 +11/+46 0/+31 0/+32 Static Margin () 8.8 7 0.7 8.8 One-Engine Inoperative Test Passed Yes N/A No Yes Tipover Test Passed Yes Yes Yes Yes Table 3.1: Table of Candidate Comparisons 13

Chapter 4 Design Comparison This downselected design compares closely with three other aerocrafts: Lockheed Martin C- 130J, Tupolev Tu-330, and the BAE Nimrod. The values that vary dramatically are primarily due to the effects of having a shorter takeoff and landing distance. However, from the comparison chart in Table 4.1 below, the historical assumptions were somewhat poor, but were a good beginning step. Our design has changed dramatically from the preliminary stage to this design review. It remains to be fairly consistent with performance parameters of existing aerocrafts. The weight has increased about 50,000 lbs. The takeoff and landing constraints were optimized resulting in increased wing loading and thrust to weight ratio. Also, with this improved design, we are able to increase the landing distance and decrease the takeoff distance so the aerocraft effectively uses the allotted runway. Downselected Final Blue Lockheed Martin Tupolev BAE Nimrod Blue Config Config C-130 J Tu-330 W T O (lbs) 198,310 243,500 164,000 228,175 234,165 W E (lbs) 88,949 131,557 75,562 151,013 112,765 T/W 0.369 0.56 0.09-0.59 W/S (psf) 72 108 94 108.4 92.26 S L (ft) 2,250 3,200 2,550 7,220 - S T O (ft) 2,613 2,280 4,700 7,220 - AR 6.633 9 10.1 9.7 6.4 Wing Span (ft) 135 143 133 143 127 Table 4.1: Comparison Aeroplanes for Final Blue Configuration Comparison data for the Red and White design can be found in Appendix A. 14

Chapter 5 Introduction to Final Design 5.1 Introduction to Final Design The downselected blue configuration was refined and further analyzed in order to produce an optimized design for the mission requirements. The primary modifications were made in the areas of tail configuration, nose geometry, aspect ratio, and landing gear arrangement. These changes, along with the other subsystem designs are discussed in detail in the following chapters. An overall summary of the aeroplane s specifications is first presented in the following section. 5.2 Aeroplane Specifications This aeroplane is designed to fulfill FEMA s need for a disaster relief cargo aerocraft capable of long range transport with austere STOL field capability. It has a gross takeoff weight of 243,500 lbs., with an empty weight of 131,557 lbs. and payload weight of 60,000 lbs. The aeroplane requires a thrust-to-weight of 0.56 and a wing loading of 108. It cruises at Mach 0.85 at a ceiling of 45,000 ft. with a range of 1,924 nm. Its high lift system, consisting of slats and triple slotted flaps, helps produce a max coefficient of lift of 3.2. Two General Electric CF6-80E1A3 s produce a max thrust of 72,000 lbs. each with a specific fuel consumption of 0.344. The aeroplane has a wingspan of 143 ft with an AR of 9 (nine). It is a high wing configuration with a supercritical aerofoil and an all moving T-tail. The total length of the fuselage is 123 ft, while its maximum height is 17.5 ft. The landing gear is quadricycle with independent steering to make taxiing and ground operations feasible. It is capable of landing within 3,200 ft without the use of thrust reversers and can takeoff in 2,800 ft. The total number of aeroplanes to be produced is 35, each with a planned price of $308 million. A summary of these specifications is given below in Table 5.1. 15

W E (lbs) 131,557 W T O (lbs) 243,500 W P (lbs) 60,000 T/W 0.56 W/S (psf) 108 Cruise Speed (Mach) 0.85 Cruise Altitude (ft) 45,000 Range (max payload) (nm) 1,924 C Lmax 3.2 Thrust Required at T/O (lbs) 136,000 SFC (lb/lbf-hr) 0.33 Static Margin () 7.9 S L (ft) 3,200 S T O (ft) 2,280 Wingspan (ft) 143 Aspect Ratio 9 Fuselage length (ft) 123 Fuselage height (ft) 17.5 Landing Gear quadricycle Tail Configuration T-tail Quantity Produced 35 Price (2008 USD) 285 million Table 5.1: Aeroplane Specifications 16

Chapter 6 Layout and Design The Blue Design is a conservative design with the exception of the all-moving tail and quadricycle landing gear. All engineering drawings can be found in Appendix D. All engineering drawings are done in inches. 6.1 All-Moving Tail To reduce the size of the tail empennage, we have chosen to use an all-moving tail. Upon further research, this has proven very effective for aeroplanes that fly in the transonic and supersonic regime. The all-moving tail provides an interesting challenge from a controls perspective, but has been successfully implemented on a number of planes in different fashions first with the North American F-100 Super Sabre to the Russian MIG-21, with varying type of purposes; from fighters to transports. [2, p.78] The all-moving tail will be operated by three independent hydraulic actuating system. Figure 6.1 shows an isometric view of our all-moving tail and Figure 6.2 gives a small demonstration of the all-moving tail deflected relative to the tail. An engineering drawing can be seen in Figure D.4. Figure 6.1: Isometric View of the All-Moving Tail 17

Figure 6.2: View of the All-Moving Tail Deflected 6.2 Cockpit Layout The cockpit will have a classical layout. It will have three chairs, two for the pilot and co-pilot and one directly behind the pilot for the load master. There is also room for a galley and lavatory in the left corner, or possibly on a separate floor in the nose. The cockpit will be pressurized behind a locked door to the cabin. The shape of the nose section is designed to resemble a slender cone. Considerations such as the pilots vision angle are taken into account while deciding how slender the nose should be. The smallest angle between the pilot s line of vision and the cockpit windscreen is known as the transparency grazing angle. A minimum grazing angle of 30 is required. This is because should the angle fall below 30, the transparency of the glass will be greatly reduced to the extent that under adverse lighting conditions the pilot will only see the reflection of the instrument panel instead of what is outside the aerocraft. In Figure 6.3, an isometric view of the general layout of the cockpit can be seen. 6.3 Cargo Bay Layout The cargo bay is rear loaded through the cargo bay doors, as seen in Section 6.4. A key feature of the cargo bay is the two exit doors; one on the port side close to the nose and another aft on the starboard side. Although Federal Aviation Regulation (FAR) dictates the need for only one door, in the interest of safety due to the placement of cargo, we recommend using the two door design. Given the placement of the engines, the crew should be able to exit without injury under normal operating conditions, as dictated by the FAR documents section 783. This can be seen in the engineering drawing Figure D.1. Drawings of the cargo bay can be seen in the isometric view, Figure 6.4, and an engineering drawing of the cargo inside the cargo bay in Figure 6.5. For an engineering drawing of the cargo bay, please see Figure D.2. 18

Figure 6.3: View of the General Layout of the Cockpit The fuel tanks will be placed on the walls of fuselage anywhere there is space. Figure 6.5 also shows the position relative to the beginning of the fuselage. Figure 6.4: Isometric View of Cargo in Cargo Bay 6.4 Cargo Bay Door The cargo bay doors will open sideways from a hinge on the tail, a ramp will then extend from the fuselage to ground for deployment. As designed, the opening of the cargo bay doors allows for 1 ft of clearance around the box while being loaded and unloaded. Figure 6.6 shows a solid model 19

Figure 6.5: Engineering Drawing of Cargo in Cargo Bay of the tail with the cargo bay doors opening at a 90 angle. For the engineering drawing please refer to Figure D.6. 20

Figure 6.6: Example of Cargo Bay Doors Opened 21

Chapter 7 Aerocraft Sizing and Sizing Trade Studies Given our mission profile and requirements, we proceeded to determine what characteristics our aerocraft should have, namely what restrictions there are on the thrust to weight ratio (T/W) and the wing loading to surface area (W/S) of our aerocraft. Subsequently, we performed a sizing trade study to see if we should change our design for a lighter design. 7.1 Aerocraft Sizing Plot There is little restriction on the maximum T/W and the minimum W/S. A high T/W corresponds to an oversized engine, and a low W/S corresponds to an oversized wing, both of which are undesirable. Thus, it is desirable to have the highest possible W/S while having the lowest possible T/W. Captured in Figure 7.1 are the requirements that the aeroplane needs: 1. to take off with a TOP of 66.7 which fulfills our balanced field length requirement. 2. to landing on a landing field of 3250 ft. 3. Climb gradient of 0.12 (7 climb). 4. Load factor of 2.5.s For our initial design, a safe sizing was chosen to give a large margin on the landing and takeoff distance. This is changed later during the trade-studies as we find that we can get a much lighter plane with a higher W/S and T/W, and stay in the acceptable bounds of the sizing plots. 7.2 Sizing Trade Study The trade study is basically a perturbation of the two parameters, T/W, as well as W/S. For our trade study, we did a 20 perturbation on our baseline design using our tools for refined weights estimation. We produced the following findings shown in Figure 7.2. From Figure 7.2, the weight estimations, landing distance, and the TOP values for each configuration can be seen. Designs 1,4,7 and 8 fulfill both landing and takeoff requirements, but are oversized. Designs 2,3 and 6 fail the takeoff requirement. The only viable options are 5 and 9, and design 9 has a lower weight. We can see the lines of constant weight and the configurations they correspond to in Figure 7.3. Thus, we understand from our trade studies that it would be best if we used design 9. Subsequent work would be done on this configuration, as reflected in the refined weights estimate in Section 8.4. 22

Figure 7.1: Sizing Plot Figure 7.2: Trade Studies Matrix 23

Figure 7.3: Carpet Plot 24

Chapter 8 Weight Estimates 8.1 Initial Weight Estimates An initial weights estimate based on the mission profile was required so that the design process could be started. Roskam s regression data for the Jet Military cargo class of aerocraft was used to determine the relationship between the empty weight of the aerocraft to the takeoff weight of the aerocraft. Once this relation was established, an initial mission fuel fraction was calculated based on the mission profile. Using Breguet s equations for range and loiter, as well as the historical data that Roskam has for the other portions of the mission, the following mission fuel fractions in Table 8.1 and 8.2 were generated. Mission segment Warmup Taxi Take-off Climb Cruise Descent Landing Fuel fraction 0.99 0.99 0.995 0.9789 0.9565 0.99 0.992 Table 8.1: Mission Fuel Fractions for a Transport Mission Mission segment (Divert) Cruise Loiter Fuel Fraction 0.9694 0.9691 Table 8.2: Mission Fuel Fractions for Diversion Using the mission fuel fractions computed, as well as the relation that Roskam derived between the empty weight and the takeoff weight of the military cargo transport plane, we then proceeded to an iterative process in which we calculated the following results for our cargo plane: Item Weight (lbs) Empty Aerocraft 120,000 Available Fuel 54,800 Trapped Fuel 3,300 Crew 615 Payload 60,000 Gross Take-off Weight 239,000 Table 8.3: Initial Weight Estimate 25

8.2 Sensitivity Studies The sensitivity studies are essential in uncovering the design drivers. The main goal of this exercise is to understand what factors drive the weight and ultimately the cost of the aerocraft down. This was done using the initial weight estimate codes we had, by perturbing the parameters individually and capturing their effect on the takeoff weight of the aerocraft. The results are captured in Table 8.4. Parameter varied Absolute change in WTO (lbs) Relative change in WTO () Payload Weight 2.61 0.836 Cruise L/D -3730-0.291 Range 76.9 0.206 Empty Weight 2.09 1.12 Fuel Consumption 81300 0.304 Flight Speed -98.8-0.264 Table 8.4: Sensitivity Studies As we can see from the relative change column, it is most desirable to have the lowest empty weight, as it reduces the takeoff weight most significantly. Also, a higher L/D, flight speed, and lower fuel consumption rate are desirable. Varying the flight speed and fuel consumption will depend heavily on the kind of turbo-fan engines we can find commercially, and that will be discussed in the Chapter 13. In order to have a higher L/D, the aerocraft needs to cruise at the optimum altitude. Experience tells us that the higher the plane cruises, the better the L/D. Thus, this will also depend on the engine performance as we will need to pick an engine that is able to give us a higher altitude. 8.3 Refined Mission Fuel Fraction The following refined fuel fractions were calculated to be used in the iteration for a refined weights estimate: Mission Segment(I) Warm-up & Taxi Take-off Climb Cruise Descent Landing Fuel Fraction 0.996 0.995 0.978 0.956 0.99 0.992 Table 8.5: Refined Mission Fuel Fraction (Segment I) Mission Segment(II) Warm-up & Taxi Take-off Climb Cruise Descent Fuel Fraction 0.996 0.995 0.978 0.952 0.99 Table 8.6: Refined Mission Fuel Fraction (Segment II) Mission Segment(III) Climb Cruise (Divert) Loiter Descent Landing Fuel Fraction 0.997 0.949 0.985 0.99 0.992 Table 8.7: Refined Mission Fuel Fraction (Segment III) The whole mission is analyzed in Tables 8.5, 8.6 and 8.7. Mission segment I corresponds to the portion of the mission where the payload is being brought to the relief zone. Mission Segment II 26

corresponds to the portion of the mission where the aerocraft returns from the relief zone. Mission Segment III corresponds to the requirement whereby the plane does a diversion to an alternate landing site. It should be noted that for both segments I and II, the cruise altitude is 45,000 ft to maximize the L/D and improve fuel efficiency. For segment III, the cruise attitude is 10,000 ft because the aerocraft does not have enough endurance for a climb to optimal cruise altitude. Also, the cruise speed is at Mach 0.30 and the loiter speed is at Mach 0.26 to improve efficiency as we are not required by mission specification to fly above Mach 0.80 for this portion of the mission. 8.4 Refined Weight Estimates After the sizing and the trade studies are done (as described in Chapter 7), a refined weights estimate was produced by analyzing the various weight groups of the aerocraft, as prescribed by Raymer. The refined weights estimation reflects the actual size of the wing according to the required W/S, as well as the actual weight of the engine selected that fulfills the T/W requirements. This estimation also makes use of the refined mission fuel fraction presented in the previous section. Item Weight(lbs) Item Weight(lbs) Fuselage 24,400 Wing 22,570 Horizontal tail 2,100 Vertical tail 1,400 Landing gear (nose) 3,400 Landing gear (main) 7,100 Installed Engines 29,200 All Else 41,400 Empty Weight 131,600 Trapped Fuel 2,900 Crew 615 Operating Empty Weight 135,200 Fuel Available 48,400 Payload 60,000 Gross Take Off Weight 243,500 Table 8.8: Refined Weights Estimates These numbers in Table 8.8 are based on the sizing determined after the trade studies, and are the weight estimation of the final design configuration. 27

Chapter 9 Center of Gravity Location and Excursion For stability issues, it is important to understand how the center of gravity (CG) shifts during the mission, so as to offset the effect using stability control systems, or shift the fuel tanks such that the CG excursion is minimal during flight. As such, an estimate of the CG location throughout the mission was calculated. The most forward and most aft locations are noted, to be used later in the landing gear sizing. There are two CG excursion plots we should look at, one of the mission profile, another of the loading of the plane. 9.1 CG Excursion During Mission Figure 9.1: CG Excursion During Mission As previously defined, (I) corresponds to the mission segment where the payload is transported to the relief zone, (II) corresponds to the aerocraft returning, and (III) corresponds to the diversion mission. As seen from Figure 9.1 the only anomaly occurs when the aerocraft has landed at the relief zone 28

and unloads the payload. This would manifest itself as a large forward shift in the CG location. However, the fuel is adjusted to pull the CG back to its original location. With a quadricycle landing gear, the aeroplane is more stable with a forward CG than an aft CG, so we opt to have the counter-balance fuel moved to the main fuel tank before removing the payload. Also, from Figure 9.1 we can see that the CG excursion during the flight is minimal. This is because the main fuel tank, where the aerocraft burns fuel from, is built at the CG of the plane. As fuel is burned, the CG of the aerocraft will not change much, keeping the CG excursion small during flight. Worthy of note is the fact that the CG location of segment I and segment II to III are very close. This is a result of the choice of a favorable location for the aft fuel tank that acts as a counter-balance to the payload. This keeps the overall excursion to within 0.1 feet, which is highly desirable. 9.2 Loading CG Excursion As mentioned in Section 9.1, there is a need for an aft tank configuration in the aerocraft in order to properly counter-balance the presence of the payload. Thus in the loading path, we have an additional concern that the aft tank should be loaded after the payload to prevent a longitudinal tip over. Figure 9.2: CG Excursion At Loading Figure 9.2 shows us the CG excursion at the different points during the loading process. The configuration that corresponds to each station number are as follows: 1. Operational Empty Weight 2. Operational Empty Weight + Crew 3. Operational Empty Weight + Crew + Main Tank 4. Operational Empty Weight + Crew + Main Tank + Payload 5. Operational Empty Weight + Crew + Main Tank + Payload + Aft Tank 29

From this study, we now know the most aft CG location to be 53 ft, and the most forward CG location to be 52 ft as measured from the nose of the aerocraft. 30

Chapter 10 Wing Design and High Lift Systems 10.1 Aerofoil Selection Our team of engineers performed a trade study of supercritical aerofoils to determine the most efficient lifting aerofoil for our Mach number of 0.85 at 45,000 ft. Using AVL, MATLAB, and a BATCH script, we were able to run a large selection of supercritical aerofoils found at angles of attack from -10 to +20. Using the Trefftz plots as well as aerodynamic wing loading vectors outputted by AVL, we determined the most beneficial aerofoil for our aerocraft wing and horizontal stabilizer was the NLF0215F. A sampling of various angle of attack executions for all aerofoils tested can be seen in Appendix G if needed. The aerofoil chosen can be seen below in a data point representation view in Figure 1. Figure 10.1: NLF0215F Aerofoil View 10.2 Wing Design Characteristics The wing geometry characteristics for the final design are shown below in Table 10.1. All parameters of the wing were sized according to an iterative process taking the initial takeoff weight, W/S, and static margins. The value of 9 for the aspect ratio was chosen in order to maximize the wing span according to the reference area calculated through sizing of the W/S. It can be proven through aerodynamics that as the aspect ratio of the wing increases, the lift forces across the span of the wing increases. Increased span-wise lift is highly desired for the FEMA conditions at takeoff and landing as well as reducing the amount of thrust needed for cruise. In addition, the wing span of 143 ft is significantly less than the 200 ft aerocraft width constraint set forth by FEMA, which is beneficial for our company because it will reduce excess weight and cost of materials. An important parameter of the wing that might attract attention is the high wing structure with an anhedral of 5. This design feature is not necessarily typical in most carrier or cargo aerocrafts but was thought to be a very important design to continue to incorporate into our final design. The reason for a high wing system is to allow for clearance between the ground and the engines as well as the undercarriage of the wing in harsh environments prevalent in relief zones. Additionally, since the supporting structure is a wing box configuration, a high wing would put the wing box 31

Reference Area 2257 ft 2 AR 9 Taper Ratio 0.4 Span 143 ft Root Chord 22.62 ft Leading Edge Sweep 30 Quarter Chord Sweep 27 Anhedral -5 Fixed Incidence 0 Table 10.1: Wing Geometry Characteristics mostly out of the way of the interior of the fuselage for the payload and loading mechanisms. Any other wing configuration would require the passing of the wing box directly through the fuselage which would require an increased length in aerocraft and added mass. Some other important parameters for the wing include a mean aerodynamic chord (MAC) of 16.8 ft, the y-direction distance of MAC of 30.54 ft, and distance of MAC from nose of 45.9 ft. A detailed CAD drawing of the full span wing can be seen below in Figure 10.2. Figure 10.2: Detailed View of Wing Full Span 10.3 High Lift Systems In order to take off and land in the short distance required by FEMA, we need a C Lmax of about 3 which can only be obtained with the help of the triple-slotted flap system and leading edge slats, increasing the lift at these conditions by 1.9 and 0.4 respectively. The flaps run across 68 of the wing half-span while the slats are about 83 of the wing half-span. The high lift triple-slotted flap system is displayed in a more general sense below in Figure 10.3 [7, p.336]. A more detailed explanation and CAD drawing of locations of the flaps will be presented below. In the preliminary design stage, the intent was to use a leading edge slat to increase the C Lmax 32

Figure 10.3: General View of High Lift Systems by about 0.4 and triple-slotted flaps to increase the C Lmax by about 1.9. With these two wing modifications intact, the wing can effectively have an increase of about 2.3 above the wing geometry C Lmax. In addition, the all moving horizontal stabilizer discussed earlier will slightly increase the lift performance of the aerocraft by increasing its ability to pitch the aerocraft up or down. We were able to determine the location of the flaps along the half span of the wing using a wing loading code provided by Professor Powell. After iterations of the code using our wing geometry we were able to determine the inboard and outboard span-wise location of the trailing edge flaps based on wing stalling at the particular C l values. Figure 10.4 shows the plot used to determine span-wise location of the triple-slotted flaps. Figure 10.4: Sectional Lift Coefficient According to Normalized Wing Span Location The leading edge slats are being placed across 83 of the half span wing because there are no other control surfaces placed on the leading edge to interfere with the slats. The slats will allow the aerocraft wing to have a higher angle of attack which increases lift without experiencing stall. 33

Chapter 11 Empennage and Control Surfaces 11.1 Empennage Characteristics The main feature of the aerocraft s empennage is the use of an all-moving T-tail. This design was chosen over a conventional tail so that the flow trailing the main wing would not create turbulent effects over the horizontal stabilizers leading to a greater induced drag and instability. The T- tail allows for more stable maneuvers and pitch control during flight. A key characteristic of the vertical fin in this configuration is the minimal taper of 0.9 with a higher leading edge sweep of 35 to allow for less incident drag. From the previous preliminary designs, the height of the vertical fin has been drastically reduced, and the span of the horizontal stabilizer has been reduced as well. Specific values for important vertical fin and horizontal stabilizer characteristics are summarized in Table 11.1 below. Horizontal Stabilizer Vertical Fin Reference Area (ft 2 ) 383.9 257.3 AR 4 1.9 Root Chord (ft) 10.89 12.93 Span (ft) 39.2 22.11 Taper Ratio 0.8 0.9 Leading Edge Sweep ( ) 35 35 Table 11.1: Horizontal and Vertical Tail Characteristics A very unique and innovative design feature of the T-tail design on our aerocraft is the use of an all-moving tail to increase the angle of attack that is induced by the vertical tail and stabilizer. There is a need for our aerocraft to be able to recover control in case of a one engine inoperative situation. This is usually accomplished through the use of a yawing moment caused by rudders. Instead, the plane utilizes an all-moving vertical fin that acts as one complete rudder to reduce the amount of deflection required. In addition, the all-moving horizontal stabilizer will also help improve the controllability of the aerocraft during cruise. The vertical fin will be able to move around its z-axis using a bar which runs through the entire fin to a hinge point in the middle of the horizontal stabilizer that will rotate and cause the fin to move. A 3-D view of the all moving mechanism using a rotating bar can be seen in Figure 6.2. Since the vertical fin is the only moving mechanism around the z-axis, the horizontal stabilizer will move through the same angle. Since the maximum angle deflected by the vertical fin will not be too large, the rotation of the horizontal stabilizer away from the neutral position will not have any real effects on the span wise coefficient of lift as well as the coefficient of drag. The need for the taper ratio to be so high for the vertical fin was due to the fact that there is not much drag penalty that results from flow around the aerofoil forming the vertical stabilizer. 34

Another reason was to allow for a larger root chord of the horizontal stabilizers to be used higher up on the vertical tail which is optimal for laminar flow around it. The T-tail design also allows for less interference between the stabilizers and the cargo bay door which would be a problem if a conventional tail was used. Figure D.4 illustrates the design of the T-tail as attached to the end of the fuselage on top of the cargo bay door. Other important characteristics of the empennage include the MAC, y-position of MAC, position of MAC from nose, and volume coefficient used to size. The horizontal stabilizer has a MAC of 9.84 ft, a y-position for MAC of 9.43 ft, distance of MAC from nose of 134.6 ft, and a volume coefficient of 0.75. The vertical fin has a MAC of 12.3 ft, a y-position for MAC of 5.43 ft, distance of MAC from nose of 116.95, and a volume coefficient of 0.062. These dimensions in relation to the entire aerocraft can be seen more visually in the detailed three view CAD drawing presented in the earlier sections. 11.2 Control Surfaces Control surfaces of this FEMA mission aerocraft play a vital role in the functionality of the plane. Due to the unusual takeoff and landing terrain as well as possible unsafe weather conditions, it is necessary to implement the most advanced system of control surfaces on the main wing as well as T-tail. The three main surfaces are ailerons, elevators, and rudders. Figure 11.1, shown below, illustrates where all surfaces are located in relation to the plane. Figure 11.1: Detailed CAD Top View of Control Surface Locations The ailerons are placed at about the last third of the wing (near the tip) on the trailing edge in order to allow for more control of the aerocraft s ability to roll. The pilot is able to operate the ailerons in order to roll for turn or to stabilize the aerocraft back to a natural state after a small perturbation disturbance. The same idea is applied to the rudder which in our aerocraft encompasses the entire area of the vertical fin. The rudder is used to control the yaw of the aerocraft so that if a disturbance impacts the aerocraft, the pilot can correct it using a minor deflection in the rudder to change the horizontal location of the nose. The elevator is located on the horizontal stabilizers to allow for more control of the pitch angle of the aerocraft so that the plane can either generate more lift when in need or less when not. Table 11.2 summarizes the locations of all control 35

surfaces including flaps and slats in relation the inboard and outboard span location and chord fraction. Flaps Slats Ailerons Elevator Rudder Inboard span location (normalized) 0.123 0.123 0.7 0.1 0 Outboard span location (normalized) 0.7 0.95 0.97 0.85 1 Inboard chord fraction 0.7 0 0.7 0.7 0 Outboard chord fraction 1 0.15 1 1 1 Table 11.2: Control Surface Locations 11.3 One Engine Inoperative Requirements While designing our T-tail and rudder system, we had to consider the requirement set by FEMA for OEI. This situation occurs if one of the two engines becomes inactive and does not function properly. The resulting yawing moment must be stabilized using a maximum rudder deflection of no more than 25. This requirement sized the length of the fuselage and geometric dimensions of the vertical fin. There were two options to satisfy the required moment needed to counteract the OEI during flight: make the vertical fin span longer and therefore the reference area larger, or increase the length of the fuselage to increase the moment arm of the yawing forces generated by the rudder. Our engineering team chose to increase the length of the fuselage because there was a worry that making the reference area of the vertical fin too large would make the span very large as well. A vertical fin that is too tall might have structural integrity issues. Therefore, we chose the safer option of having a longer fuselage. In conclusion, our current aerocraft configuration does satisfy the feasibility test set by OEI. The values used to analyze the OEI requirement can be seen in Appendix E. 36

Chapter 12 Landing Gear Design Our aeroplane will employ a quadricycle landing gear, whereby the nose gear will consist of two twin nose wheels, and the main gear will consist of two four-wheel bogeys. The quadricycle arrangement is ideal for our purpose as it allows for the cargo floor to be very low to the ground. This means that loading and unloading of the cargo will be easier. The nose wheels will be placed 25 ft from the nose, and the main wheels will be 66 ft from the nose. Figure 12.1 shows where the landing gear will be positioned with respect to the fuselage. Figure 12.1: Position of Landing Gear with Respect to Fuselage The tires to be used are the Michelin Air X Radial Tires. The tire dimensions are detailed in Table 12.1. Nose/Main Tire Diameter (ft) 3.08 Nose/Main Tire Width (ft) 0.98 Table 12.1: Tire Dimensions The main landing gear is located about 13 ft behind the most aft CG location which is located at 53.6 ft from the nose. At this distance, the vertical from the landing gear makes a 45 angle with the main gear line and thus satisfies the 15 angle required for longitudinal stability. The position of the main landing gear also takes into consideration that a certain amount of clearance is required for the tail section of the aerocraft to remain clear of ground during takeoff and landing. An angle of at least 15 between the landing gear and the tip of the tail is required for adequate 37

ground clearance. Calculations show that the aerocraft s landing gear makes a 20 angle with the tail tip of the aerocraft, thus providing sufficient ground clearance. The landing gear will be retracted into the fuselage when the aerocraft is not in takeoff or landing mode. The space containing the nose landing gear will be of length 4.7 ft, width 2.9 ft and height 3.6 ft. The space containing the main gears will be of length 7.8 ft, with width and height the same as that for the nose gears. The mechanism for the landing gear retraction is based on a four-bar linkage using three members that are connected by pivots. A diagonal arm called the drag brace breaks in the middle for retraction. The gear retracts backward while the drag brace retracts forward, as shown in Figure 12.2. The advantage of this mechanism is that in the event of a hydraulic failure, aero loads will blow the gear down and activate the landing gear. This mechanism will be used for both the nose and main gear. The retraction mechanism will be operated with the use of a hydraulics actuation system. Hydraulic fluid will flow through a valve and into an actuator when a command is made, which will subsequently move a piston that will directly cause the retraction or deployment of the landing gear. Figure 12.2: Landing Gear Retraction Mechanism 38

Chapter 13 Propulsion In order to power the aerocraft based on the iterative estimates for gross takeoff weight and calculated T/W as explained earlier, we chose a twin engine configuration; one turbo jet engine on each half-span of the wing. The total required thrust to takeoff is 136,000 lbs, therefore, each engine is only responsible for half that force in normal operable conditions. The General Electric (GE) CF6-80E1A3 turbofan engine, shown in Figure 13.1, was chosen because it had a maximum thrust output of 72,000 lbs. This engine provided the closest characteristics to what the aerocraft needs to operate with a safety cushion of about 7,300 lbs. The engine provides a specific fuel consumption (SFC) at sea level of 0.33 lb/(lbf-hour) with a dry weight of 11,000 lbs per engine. The diameter is 9.5 ft and the length is 14 ft. Figure 13.1: General Electric (GE) CF6-80E1A3 Engine The placement of the engines will have to be at least 9 ft from the fuselage since the engine diameter is 9.5 ft. An estimate of the center point of the engine location along the wing span is 11 ft from the fuselage or 0.27 inboard half-span fraction. The 6.25 ft space between one end of the engine to the main fuselage will allow for more laminar flow characteristics in that region in order to not create any circulation due to the colliding flow coming from the engine nacelle and fuselage. Table 13.1 displays some of the key features and dimensions of the GE CF6-80E1A3 turbofan engine. 39

Maximum Diameter (in) 114 Length (in) 168 Takeoff Thrust (lbf ) 72,000 Dry Weight (lbs) 11,225 Specific Fuel Consumption (lb/(lbf*hour) 0.33 Overall Pressure Ratio 32.4 Table 13.1: Engine Specifications Since our wings will be high, the two engines will be mounted on the underside of the wings. This allows for easy access from the ground, for maintenance purposes. The engines will be placed along the wing, away from the fuselage. The weight of the engines out along the wing provides a span-loading effect, which helps reduce wing weight. 40

Chapter 14 Performance Analysis 14.1 Stall Speed At takeoff, the aerocraft has a wing loading of 108 pounds per square foot (psf) and a C Lmax of 3.0 resulting in a takeoff stall speed of 174 ft/s. During landing, due to the amount of fuel that was burnt getting to the destination, the weight of the aerocraft decreases. Consequently, W/S of the aerocraft decreases to about 98 psf. At a C Lmax of 3.0, this will result in a stall speed of about 167 ft/s at landing configuration. 14.2 Takeoff and Landing At takeoff, the engines of the aerocraft are able to provide a maximum thrust of 144,000 lbs, giving a takeoff T/W of 0.591. This is more than sufficient to meet the minimum requirements for takeoff. With this amount of thrust and the use of high lift systems giving a C Lmax of 3.0, the aerocraft is able to achieve a takeoff distance of about 2280 ft at a W/S of 108 psf. During landing, the W/S decreases to about 98 psf due to the reduced weight of the aerocraft from the burning of the fuel. With the use of high lift systems giving a C Lmax of 3.0, the aerocraft is able to attain a landing distance of about 3,200 ft without the use of thrust reversers. This is within the landing distance requirement of 3,250 ft. Table 14.1 summarises the takeoff and landing parameters of the aerocraft. Takeoff Landing C Lmax 3.0 3.0 e 0.75 0.7 C DO 0.1255 0.1328 L/D 4.3 4.0 T/W 0.59 NA W/S 108 98 Distance (ft) 2,280 3,200 Margin (w/o thrust rev. ) 24 1.5 Margin (w/ thrust rev. ) NA 41 Table 14.1: Takeoff and Landing Parameters 41

14.3 Cruise, Climb and Flight Ceiling The use of the General Electric CF6-80E1A3 engines has allowed us to fly the aerocraft at a higher cruise altitude. We have chosen to fly the aerocraft at a cruising altitude of 45,000 ft. The amount of thrust required will be less and hence reduce the fuel consumed. Our calculations show that at 45,000 ft a thrust of about 12,000 lbs will be required for cruising flight. The engines are able to provide an estimated total thrust of 15,000 lbs at this altitude. Thus, there is 3,000 lbs of excess thrust which could be used for any maneuver corrections needed as the flight progresses. For the climb segment, we have chosen a climb angle of 7. For simplicity purposes, we assumed a constant acceleration in the direction of the thrust vector and that the aerocraft will climb along a straight line path until it reaches cruising altitude. Taking into considerations these assumptions and the climb angle, the aerocraft will achieve a climb rate of 3,520 ft/min and reach cruising altitude of 45,000 ft in about 13 minutes, covering a distance of about 60 nm along the ground. At takeoff, the engines are capable of producing an initial excess thrust of 55,000 lbs The flight ceiling of the aerocraft is determined primarily by the capabilities of the engine to provide enough thrust to keep the aerocraft at level flight. Data taken from Jane s Aerospace Engines show that the CF6-80 engines can cruise at a maximum altitude of 45,000 ft at a speed of Mach 1.0. Since our aerocraft is cruising at a lower speed, there is excess thrust to allow the aerocraft to climb to a higher cruise altitude. However, due to a lack of sufficient data regarding engine parameters, we can only estimate that the flight ceiling imposed by the engine capabilities would be slightly higher at about 48,000 ft. 14.4 Trimmed Lift and Drag Using the data acquired through our weights estimation and W/S sizing, we were able to size the wing appropriately and determine the max coefficient of lift (C Lmax ) of the aerocraft to be 0.5263 at 0o angle of attack. Using the step up method to determine a more precise value for the parasite drag coefficient (C D0 ), we were able to figure the coefficient of drag (C D ) at various stages in flight as well as a lift to drag ratio (L/D). The step up method involved using a more complex method of building up mission fuel fractions at each changing point in the mission profiles. Using the value for C D0 and a Mach number of 0.85 at cruise condition as inputs into AVL, we were able to output a Trefftz plot that displays the sectional coefficient of lift values across the wing and horizontal stabilizer span as well as C L and C D values for the aerocraft. The supercritical aerofoil used in AVL analysis was the NLF0215F. The plot of these span-wise values can be seen below in Figure 14.1 for cruise at an angle of attack of 0. The values given by AVL correspond to the condition whereby the control surfaces were trimmed in order to maintain the aerocraft s stability at the specified angle of attack and speed. The corresponding trimmed lift and drag coefficients are 0.5263 and 0.03709 respectively. In addition to giving the values for the induced and parasitic drag, the AVL plot also shows the lift distribution across the wings. 14.5 Stability From the CG analysis of the whole mission profile, we are able to derive the static margin changes through out the mission as well. The static margin is coupled to the CG excursion as the two have a very close relation. The fact that the CG excursion is small also means that the static margin stays within a tight range, and in our design we have our static margin to range from 6 to 7.5. This means that the aeroplane is generally stable throughout the flight without being too 42

Figure 14.1: Trimmed Lift and Drag Data from AVL difficult to manuever. For a more detailed presentation of the static margin analysis please refer to Appendix B. Besides having the ability to determine the lift and drag coefficients, AVL is also able to produce the stability derivatives of the wing for the various angles of attack. By observing the coefficients of the stability derivatives, we are able to tell if there exists static stability due to five different parameters: the angle of attack, sideslip angle, roll rate, pitch rate and yaw rate. Table 14.2 and 14.3 show the stability derivatives for the wing at zero angle of attack. Alpha Beta Z force 5.2139-4.4e-005 Y force 0.000783-0.27821 Rolling Moment, Cl -0.00026-0.085709 Pitching Moment, Cm 4.1704-9e-005 Yawing Moment, Cn -0.000438 0.091262 Table 14.2: Stability Derivatives due to Angle of Attack and Sideslip Angle Roll Rate, p Pitch Rate, q Yaw Rate, r Z force 0.000321-4.002-4e-006 Y force 0.20732-0.00158 0.156 Rolling Moment, Cl -0.479 0.000172 0.147 Pitching Moment, Cm 8e-005-27.51 4.1e-005 Yawing Moment, Cn -0.0109 0.001013-0.0673 Table 14.3: Stability Derivatives due to Roll Rate, Pitch Rate and Yaw Rate As can be seen from the tables, the stability derivatives are stable for all five parameters mentioned earlier. For example, the pitching moment coefficient due to the pitch rate is negative, implying stability. In conclusion, the aerocraft will exhibit static stability. 43

Chapter 15 Aero Loads An analysis of the aero loads exerted on the aeroplane was done in order to determine its basic strength and flight performance limits. The specific aero loads examined include maneuver, gust, and lifting surface loads. These are discussed in the following sections. 15.1 Maneuver and Gust Loads The generation of lift during high-g maneuvers typically accounts for the greatest aero loads on the aeroplane. At lower speeds the highest load factor an aeroplane may experience is limited by the maximum lift available. For higher speeds the maximum load factor is limited to the chosen value based upon the expected use of the aeroplane. The flight envelope, or V-n diagram, of the aeroplane describes these basic flight performance limits. A V-n diagram for the aeroplane s minimum cruise weight of 140,000 lbs is shown below in Figure 15.1. Similarly, a V-n diagram for the aeroplane s maximum cruise weight of 250,000 lbs is shown below in Figure 15.2. The flight conditions represented by the boundary of the flight envelope are the most severe structurally. Furthermore, the points located on the corners are the most critical, in practice, because they correspond to the highest loads produced on the aeroplane. Figure 15.1: V-n Diagram for Minimum Cruise Weight 44

Figure 15.2: V-n Diagram for Maximum Cruise Weight The first loading condition occurs at the highest possible angle of attack, which is obtained in a pull up. The second loading condition corresponds to the positive low angle of attack and maximum indicated airspeed at which the aeroplane will dive. At the most negative high angle of attack the third loading condition occurs, which is obtained in low speed maneuvers. The fourth critical loading condition corresponds to the negative low angle of attack, which occurs at high speed pitch down maneuvers. These four angles of attack specific to the design of this aeroplane are listed below in Table 15.1. Point AOA (α ) A 14 D 2 G -10 F -6 Table 15.1: Critical Loading Angles of Attack The flight envelopes in Figures 15.1 and 15.2 show the flight performance limits for various cases. These include flaps retracted and deployed for symmetric maneuvers, as well as gust loads. Gust loads were also considered because they can sometimes exceed the maneuver loads experienced by an aeroplane. These gust loads cause a change in the angle of attack, which affects the aeroplane s lift and ultimately the load factor. 15.2 Lifting Surface Loads An analysis of the aerodynamic, inertial, and gravity loads acting on the wing was done in order to aid in the structural arrangement of the wing components. A visualization of the aerodynamic loads can be seen below in Figure 15.3. Despite this information being considered when creating the initial structural arrangement, a more detailed analysis of the shear forces and moments will be required to correctly size the various structural wing components. Preliminary work can be found in Appendix F. 45

Figure 15.3: Aero Loads on Wings 46