AIAA UNDERGRADUATE INDIVIDUAL DESING COMPETITION: DESIGN AND ANALYSIS OF THE PEGASUS JET TRAINER

Transcription

1 AIAA UNDERGRADUATE INDIVIDUAL DESING COMPETITION: DESIGN AND ANALYSIS OF THE PEGASUS JET TRAINER Instructor: Dr. Ron Barrett Department of Aerospace Engineering May 10, 2014

2 Designer: AIAA Member ID: Faculty Advisor: Dr. Ron Barrett AIAA Member ID: Aerospace Engineering Department ii

3 ACKNOWLEDGEMENTS The author would like to thank the following people for their help and guidance throughout this process: Dr. Ron Barrett, Dr. Jan Roskam, Dr. Anemaat, and the wonderful employees at Textron Aviation. TABLE OF CONTENTS ACKNOWLEDGEMENTS... iii TABLE OF CONTENTS... iii LIST OF FIGURES... iv LIST OF TABLES... v LIST OF SYMBOLS... v 1 Mission Specification and Profile Historical Overview Statistical Time and Market Predictive Engineering Design Vector Analysis Statistical Time and Market Predictive Engineering Design Weight Sizing Statistical Time and Market Predictive Engineering Design for Wing Sizing and Installed Power Class I Configuration Selection Class I Design for Down Selected Configurations Class I Cockpit and Fuselage Propulsion Integration Class I Wing and High Lift Devices Class I Empennage Class I Undercarriage Class I Weight and Balance Final Down Selection Final Configuration Cockpit and Fuselage Final Configuration Propulsion Integration Class I Propulsion Integration Class II Propulsion Integration Class II Propulsion Performance Wing and High Lift Empennage Undercarriage Weight and Balance Stability and Control Drag Buildups and Drag Polars Performance Verification Systems and Structures Advanced Technologies Cost Estimation Aerospace Engineering Department iii

4 21 Characteristics REFERENCES LIST OF FIGURES Figure 1.1 T-38 Replacement Mission Profile... 9 Figure 1.2: BAE Systems Hawk (Ref. 2) Figure 2.1: T-38 Aircraft Figure 2.2: F-16 Fighter Figure 3.1: Design Philosophy (Reference 5) Figure 3.2: Market Share Philosophy (Reference 6) Figure 4.1: STAMPED Analysis Weight Fraction Projection Figure 5.1: STAMPED Wing Loading Results Figure 5.2: STAMPED Analysis Thrust Results Figure 5.3: Pegasus Jet Trainer Sizing Chart Figure 7.1: Fuselage 1 Isometric View and Side View Figure 7.2: Fuselage 2 Isometric View and Side View Figure 7.3: Configurations 1 and 2 Engine Placement Figure 7.4: Design 1 Forward Swept Wing, Design 2 Hershey Bar Wing Figure 7.5: Empennage Integration (Configuration 1, Configuration 2) Figure 7.6: Lateral and Directional Tip Over Criteria for Designs 1 and Figure 7.7: C.G. Excursion Diagrams (Configurations 1 and 2) Figure 8.1: Air Force Fuel Expenditure (Ref. 10) Figure 9.1: A10 Ladder Fully Extended (Ref. 11) Figure 10.1: CAD Engine Figure 10.2: AI F Turbofan Engine With Afterburner (Reference 12) Figure 10.3: Forces and Moments Acting on Aircraft (Ref. 13) Figure 10.4: Integrated Bihrle-Weissman Chart for the Aircraft (Ref. 13) Figure 10.5: Uninstalled Dry Thrust Engine Performance (Non-Afterburner) Figure 10.6: Installed Dry Thrust Figure 11.1: Ref 17 (Tail Buffet Allevation Wing Strake) Fillet Geometry Figure 11.2: Pressure Coefficient Distribution (Ref 17) Figure 11.3: High AOA operational envelope for two models (Ref. 18) Figure 11.4: Comparison Between Vortex Interaction with On and Off Fuselage Figure 11.5: Final Pegasus Wing Configuration Figure 11.6: Tornado Wings (Diamond Fillet - Baseline Configuration) Figure 12.1: Vertical and Horizontal Tails Figure 13.1: Longitudinal Clearance Angle Figure 13.2: Lateral Ground Clearance Figure 14.1: Weight Distribution Figure 14.2: C.G. Component Locations Figure 14.3: Class II Center of Gravity Excursion Figure 15.1: Directional X-Plot Figure 15.2: Pegasus Jet Trainer AAA Configuration Figure 15.3: Trim Diagram Aerospace Engineering Department iv

5 Figure 16.1: Fuselage Area Figure 16.2: Aircraft Cross Sectional Perimeter Figure 16.3: Drag Polar of Pegasus Jet Trainer Figure 16.4: Class II Component Drag Buildup Figure 17.1: Pegasus Jet Trainer V-n Diagram (Ref. 21) Figure 17.2: L/D vs Mach at feet Figure 17.3: 1-g Maximum Thrust Specific Excess Power Envelope (Non Afterburner) Figure 17.4: 5-g Maximum Thrust Specific Excess Power Envelope (Afterburner) Figure 17.5: 9-g Maximum Thrust Specific Excess Power Envelope Figure 18.1: Frames and Ribs of the Pegasus Jet Trainer (Side View) Figure 18.2: Frames and Ribs of the Pegasus Jet Trainer (Top View) Figure 18.3: Frames and Ribs of the Pegasus Jet Trainer (Isometric View) Figure 18.4: Aircraft Components for Assembly Figure 20.1: Cost of Aircraft as a Function of Total Built Figure 21.1: Final Aircraft 3-View LIST OF TABLES Table 4.1: STAMPED Analysis Aircraft Weights Table 4.2: Preliminary Weight Estimation for 2 Engine Configurations Table 5.1: Aircraft Weights, Wing Area, Thrust Table 7.1: Wing Characteristics Table 7.2: Class I Component Weight Distribution Table 11.1: Pegasus Aircraft Wing Characteristics Table 12.1: Empennage Characteristics Table 14.1: Class II Weight Determination Table 20.1: Predicted Aircraft Cost Table 21.1: Salient Characteristics LIST OF SYMBOLS Symbol Description Units A,B... Linear Regression Coefficients... - A new... Linear Regression Coefficient With New Materials Accounted For... - AR... Aspect Ratio... - C p... Specific Fuel Consumption... C, D... Sensitivity Constants...-, lbs C D... Drag Coefficient Zero Lift Drag Coefficient... - Aerospace Engineering Department v

6 CGR... Climb Gradient... rad c f... Friction Coefficient... - C L... Lift Coefficient... - D... Drag... - E... Endurance... hrs e... Oswald s Efficiency Factor... - F... Weight Sensitivity... lbs f... Parasite Area... ft 2 h... Altitude... ft I p... Power Index... L... Lift... - L/D... Lift to Drag Ratio... - M ff... Mission Fuel Fraction... M tfo... Trapped Fuel and Oil Fraction... - M res... Reserve Fuel Fraction... - N... Number of Engines Dynamic Pressure... psf R... Mission Range... nmi S... Wing Area... ft 2 s... Distance (Used in Landing and Takeoff Equations)... ft S wet... Wetted Wing Area... ft 2 TOP FAR 25 Takeoff Parameter... psf T/W... Thrust to Weight Ratio... - V... Velocity... ft/s, knots, mph W... Weight... lbs W/S... Wing Loading... psf Greek Symbols Description Units... Mathematical Partial... - η... New Material to Old Material Coefficient... - Aerospace Engineering Department vi

7 η p... Propeller Efficiency... - π or Product Function... - ρ... Air Density... σ... Air Density Ratio... - Subscripts Description A... Approach... cl... Climb... cr... Cruise... crew... Crew... E... Empty... f... flaps... F... Fuel... FL... Field Length... guess... Guessed Value... L... Lift or Landing... LO... Liftoff... ltr... Loiter... max... Maximum... OE... Operating Empty... pass... Passenger... PL... Payload... RC... Rate of Climb... res... Reserve... TO... Takeoff... TOFL... Takeoff Field Length... tent... Tentative... tfo... Trapped Fuel and Oil... used... Used Fuel... wet... Wetted... Aerospace Engineering Department vii

8 Acronyms Description Units AAA... Advanced Aircraft Analysis... - AIAA... American Institute of Aeronautics and Astronautics... - AEO... All Engines Operating... - ARALL... Aramid-Aluminum... - FAR... Federal Air Regulation... OEI... One Engine Inoperative... - Aerospace Engineering Department viii

9 1 Mission Specification and Profile The Request for Proposal set forth by the American Institute of Aeronautics and Astronautics, (AIAA), for the Undergraduate Individual Design Competition calls for an advanced pilot training aircraft to replace the retiring T-38 trainer (reference 1). This proposed aircraft is to be more cost efficient at the acquisition and maintenance cost than the current costs for the T-38. The proposed aircraft must also have significantly better flying qualities than the T-38. The following figure demonstrates the profile that the aircraft will adhere to, as specified by the Request for Proposal (RFP). Figure 1.1 T-38 Replacement Mission Profile The following is a description of each phase leg, as shown on Figure 1,1 and specified by the RFP: Phase one consists of starting, warming up the aircraft, and taxiing. Phase two consists of take-off and acceleration for the climb which takes place during phase 3. Phase four is a 150 nautical mile cruise at the best cruise Mach and altitude which are presented further down in the report. Phase five is a 100 nautical mile, 300 knot airspeed rendezvous with a tanker at feet above sea level. Phase six indicates a simulated or actual refueling of the trainer which is to last 20 minutes while maintaining 200 knot airspeed. A climb from feet to the best cruise altitude takes place during phase 7 and a cruise to practice area of 100 nautical miles takes place during phase 8. Once at the practice location the trainer must descend to feet and perform air combat maneuvering training which is to take twenty minutes, while seeing a nine g-force. This combat maneuvering takes place as phases nine and ten respectively. Phase eleven is a climb back to the best cruise altitude, and the cruise back of 150 nautical miles takes place as phase twelve. Descending to sea level and landing take place during phases thirteen and fifth-teen respectively. Phase fourteen indicates a reserve fuel cruise of thirty minutes in the case of having to divert to another airport or being placed on a hold pattern before landing. Aerospace Engineering Department 9

10 These phases were set forth by AIAA with the intention of yielding aircraft designs that meet the current necessity of the United States Military. Due to the age of the T-38 program, most of the T-38 aircraft are currently flying over their intended service life (Ref. 1). As a result the T-38 is becoming too costly to maintain and/or to upgrade. With the resulting technological advances that have become available since the introduction of the T-38 aircraft, a need for a more recent military trainer has emerged. This is to properly train pilots for the new generation fighters and their maneuvering characteristics (Ref. 1). The RFP specifically states that the proposed jet trainer must be able to support modern aircraft such as the F-22 and be able to prepare the pilots for high-g operations, leading to the 9-g maneuvering requirement. When compared to other military trainers, the RFP has higher objective requirements than those of the newest trainers, such as the BAE Systems Hawk which has a lower Mach number and g capability (Ref. 2). Figure 1.2: BAE Systems Hawk (Ref. 2) Aerospace Engineering Department 10

11 2 Historical Overview From the request for proposal it is very clear that the goal of the Pegasus should be to outclass the older T-38C fleet that is currently being used to train pilots. The T-38C aircraft is not the only aircraft that the Pegasus should Figure 2.1: T-38 Aircraft outclass; the proposed trainer should have characteristics that match those of the F-16 fighter as well. Although it is classified as a fighter, the 2 seat versions of the F-16 are used to help train pilots for high-g maneuvers, air-to-air interception, and refueling, as the original T-38 was not capable of training the pilots for these missions (Ref. 1). This is one of the reasons why the Pegasus Jet Trainer presented in this report will adhere strictly to being able to achieve the 9 g-loading asked for in the RFP, although it will not outclass the F-16 in terms of velocity. This section provides a short historical overview of the T-38 series and the F-16 aircraft that the Pegasus intends to replace for military training. Nearly 1200 T-38 Talons were produced between 1962 and 1972 and more than U.S. Air Force pilots have trained using a T-38 aircraft during the lifetime of the program. The T-38 Talon was designed and produced by the Northrop Grumman Corporation. Currently over 500 Talons are currently in operation with either the Air Force for military training or NASA for testing. Due to the popularity of the T-38 Talon, Northrop Grumman designed and produced a replacement wing for the T-38 as necessity called for the extension of the program. This wing extends the life of the aircraft created in the 1960s until approximately the year The wings have aluminum outer layers with an internal honeycomb design to improve usage and life of the wing. The final T-38 that the Air Force received from Northrop Grumman was in 1972 which is one of the reasons why the Air Force is currently seeking a replacement for the aircraft. The T-38C variant of the T-38 Talon had its lifespan extended and received an avionics upgrade as well. All of the previous information can be found on reference 3. The F-16 produced by Lockheed Martin had many advantages over the F-15 which was its predecessor. These advantages at the time included fly by wire, it could withstand high g-loads which was accomplished by reclining the seat of the pilot to reduce the vertical distance between the heart and head of the pilot, allowing the pilot to stay conscious during the maneuvers. This is one of the reasons why the g-loading of the F-16 is much higher than that seen on other aircraft before it, and why aircraft such as the T-38 Talon have a g-loading of approximately 2/3 that of the F-16 Fighter. The canopy of the F-16 was design in such a way that allowed the pilot to see more of the field. Aerospace Engineering Department 11

12 These are just a few of the advantages that the F-16 offered when it was first released making it one the best fighters during its era. There have been over 4500 F-16 s produced in the lifespan of the program with many nations using the aircraft in their respective military. Some versions of the F-16 have a two pilot seat configuration making it ideal for use by the United States Air Force for training. The previous information can be found on reference 4. Figure 2.2: F-16 Fighter Aerospace Engineering Department 12

13 3 Statistical Time and Market Predictive Engineering Design Vector Analysis To complete the initial sizing process, the designer looked into market trends and used the Statistical Time and Market Predictive Engineering Design Vector Analysis method to project the expected variables of the proposed jet trainer aircraft for the near future. This chapter serves the purpose of explaining the fundamentals behind using the Statistical Time and Market Predictive Engineering Design Vector Analysis process, how a design philosophy was established for the proposed jet fighter, and explains how this vector analysis can be successfully applied to the preliminary aircraft design. Figure 3.1: Design Philosophy (Reference 5) The Statistical Time and Market Predictive Engineering Design Vector Analysis process gives engineers a tool with which they can track and predict any vector of engineering design variables at any given point in time with the use of gathered data. The following is the procedure that should be followed to correctly determine variables using the Statistical Time and Market Predictive Engineering Design Vector Analysis. First a product timeline must be developed using data pertaining to aircraft of the same class as the one being designed; in this case it will be the military jet trainer class. The major dates in the life of the aircraft program are very important to this process, as they help establish a design philosophy. Once all of the dates have been gathered, the designer is to generate a Bell-distribution of each date, in terms of the date of first flight of the aircraft. So the design fix date will be x amount of years before the first flight of the aircraft and the final aircraft retirement date will be x amount of years after the first flight date. This difference is what is used to generate the Bell curve. Figure 3.1 shows the Bell curve and how it would be divided for each important date. That is from the mean, every year after or before is graphed.. The peak of the design fix date is to be labeled as year zero. The next major step to complete is to establish the market vectors for the aircraft of similar class. Now the yearly production rate of each Aerospace Engineering Department 13

14 aircraft being used in the process is necessary. Once this data has been gathered a Bell-distribution graph is generated for each aircraft characteristic that the designer wishes to track, per year. The likely hood of each characteristic is determined by applying the production rate of the aircraft as a fraction of the total production rate for all of the aircraft being used. From this graph the aircraft that had the most sales any given year, or dominated the market, will represent the highest peak of the graph. This makes the most wanted characteristic, that of the aircraft with the most sales any given year. The design philosophy, or degree of risk, refers to how far away from the mean the designer wants to deviate. A low or zero risk philosophy signifies that the designer wants to follow the highest point for every year, leading the designer to believe that this correlates to having the most desired characteristic. All of the points with the same degree of risk will be connected and a trend will be formed from year to year. An equation will need to be found that accurately models the previously determined trend to be able to successfully project the design variable into future years. This procedure is repeated for all of the variables that are wanted. It is generally a good option to pick as a design philosophy to follow the trend corresponding to the biggest market share, which means riding the peaks of the plots similar to what is shown in Figure 3.2. Other design philosophies, such as needing a short time between design fix date and date of first flight to meet a predetermined deadline are acceptable as well as long as there is proper reasoning behind the decision to go to a different philosophy. A conservative design philosophy will be to the left of the mean on every Bell curve while an aggressive philosophy will be located to the right of the mean on the plot. A designer will employ this analysis because it provides him/her with an accurate estimation of where the market is headed as we move Figure 3.2: Market Share Philosophy (Reference 6) forward. This process helps the designer estimate future critical dates in the life cycle of the aircraft while letting the designer estimate if deadlines will be met with the desired design philosophy. All of the previous information can be found on reference 5. Aerospace Engineering Department 14

15 Weight Fraction, W_E/W_TO 4 Statistical Time and Market Predictive Engineering Design Weight Sizing The Statistical Time and Market Predictive Engineering Design Vector analysis procedure introduced in Chapter 3 was applied by the author to determine the major characteristics of the proposed aircraft. This process and major results are presented in this Chapter as well as Chapter 5. The major characteristics include: the aircraft s weights, Table 4.1: STAMPED Analysis Aircraft Weights wing size, and necessary power for this design. By gathering information on different aircraft it became possible to project the empty weight over maximum takeoff weight of the aircraft through time and into the future. Table 4.1 lists the different aircraft used for this analysis as well as their maximum takeoff weights and empty weights. The breakdown of the analysis is yearly, starting with the production of the aircraft in After calculating the probability of each weight fraction occurring at each year l the highest probable weight fraction was plotted against the year at which this fraction was the most produced. This can be seen on Figure 4.1 and the trend line shown was used to estimate the weight fractions for the year Year Figure 4.1: STAMPED Analysis Weight Fraction Projection Aerospace Engineering Department 15

16 Following this procedure resulted in an empty weight over takeoff weight ratio of approximately 57 percent for the year 2020, which is the expected introduction date of the aircraft being designed. Now applying the procedures listed by Dr. Jan Roskam in Part I of his Airplane Design series, reference 6, the fuel weight was calculated. More specifically the fuel fractions from Table 2.1 page 12 of Dr. Roskam s Part I, reference 6, were used for takeoff, climb, and landing fuel consumption calculation. Equations 4.1, 4.2, and 4.3 were used to determine the fuel fraction for all of the cruise phases of the mission profile. ( ) ( ) ( ) (Eq. 2.10, Ref. 6) Equation 4.1 ( ) (Eq. 2.12, Ref. 6) Equation 4.2 Equation 4.3 The equations were not applied if the specific fuel consumption and duration of phase were specified in the Request for Proposal. The spreadsheet Pegasus Calculations spreadsheet was created to calculate all of the fuel consumed during each phase and subtract the fuel, pilot, and payload weight from the maximum takeoff weight (Ref. 7). This sheet was created to then allow the designer the freedom to alter the initial maximum takeoff weight until the empty weight was at 58 percent of the takeoff weight. Table 4.2 shows at what values this percentage occurs for two different engine configurations. Some of the values estimated for this preliminary process are the following: C j =.66 lbs/lbs/hr and L/D = 10. These values were estimated using the values given on Table 2.2 on page 14 of Reference 6. For the preliminary sizing of the aircraft the fuel consumption was estimated for two engines and one engine, as the author had not decided on the amount of engines to use in the configuration up to this point. The fuel fractions were determined for an aircraft mission range of 1500 nautical miles. Aerospace Engineering Department 16

17 Table 4.2: Calculated Weights for One Engine For 1 Engine Fraction Desired 0.57 total fuel Takeoff Guess Phase C_j (lbs/lbs/hr) time (m) Per engine *x enginesfuel Fraction lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs Pilots and Equpt = Payload 550 lbs 1000 lbs W_E = lbs W_E/W_TO = W_F = lbs Table 4.2: Preliminary Weight Estimation for 2 Engine Configurations For 2 Engines Fraction Desired 0.58 total fuel Takeoff Guess Phase C_j (lbs/lbs/hr) time (m) Per engine *x enginesfuel Fraction lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs lbs Pilots and Equpt = Payload 550 lbs 1000 lbs W_E = lbs W_E/W_TO = W_F = lbs Aerospace Engineering Department 17

18 5 Statistical Time and Market Predictive Engineering Design for Wing Sizing and Installed Power Using the same procedure as that described in Chapter 4, the aircraft wing size and power needed were estimated. The different aircraft wing sizes were gathered and arranged together. The fraction, wing surface area divided by takeoff weight was determined for each of the aircraft and applied in a similar fashion as how it was performed in Chapter 4. Figure 5.1 shows the compilation of the different probabilities of encountering any given ratio for any Table 5.1: Aircraft Weights, Wing Area, Thrust given year. The STAMPED Procedure yielded better results for the power plant Aircraft Takeoff Weight (lbf) S (ft^2) T (lbf) sizing as the power needed for the aircraft Aero L-39 Hongdu L-15 T-38 T-45A Dassault Alpha Jet KAI T-50 Daussal Mirage II Aermacchi MB-339A is comparable with the power needed for the different trainers and fighters that were used for the analysis. Table 5.2 shows the different aircraft used for the analysis as F-5A well as the power installed on each of the F aircraft. Figure 5.2 shows the trend through-out the years. Figure 5.3 shows the design point plotted against the trends seen on Figures 5.1 and 5.2. The wing loading used is that of 80 ft^2/lbf. It is noticed that the maneuvering line dictates the necessary maximum thrust that the aircraft will need at any point during the flight. Following the trend the thrust to weight ratio was determined to be equal to.44. This is the ratio that the sizing chart will demonstrate for the normal flight conditions. This point is higher for maneuvering and thus an afterburning engine will be necessary. For preliminary purposes the wetted area was estimated using reference 8, a value of.029 was determined for the of the aircraft. The original preliminary sizing chart gave a value of.86 for the thrust to weight ratio for maneuvering at a wing loading of 80 square feet per pound.. This however was updated with the actual wetted area that was determined in the Drag Buildups Chapter of this report. This reduced the thrust to weight ratio for maneuvering and is the final sizing chart presented as Figure 5.3. Aerospace Engineering Department 18

19 Thrust to Weight, T/W_TO Wing Loading, W/S,, lbf/ft^ Year Figure 5.1: STAMPED Wing Loading Results Year Figure 5.2: STAMPED Analysis Thrust Results Aerospace Engineering Department 19

20 Thrust-to-Weight, T/W Takeoff_Distance_C_L=.5 Takeoff_Distance_C_L=.7 Takeoff_Distance_C_L=.9 Landing_Distance_CL =.4 Landing_Distance_CL=.7 Landing_Distance_CL=1 Landing_Distance_CL=1.3 Maneuvering_requirement s Design Point Wing Loadig, W_TO/S, lb/ft^2 Figure 5.3: Pegasus Jet Trainer Sizing Chart Aerospace Engineering Department 20

21 6 Class I Configuration Selection This chapter presents very rough sketches of the different designs that the author took into consideration for the design of this aircraft. From the considered aircraft configurations, the author narrowed the configurations to three. From this point on, the author performed the Class I design process as presented in Dr. Jan Roskam s Airplane Design Series (Ref. 6). The author has a design philosophy of trying to sell the trainer at the least price possible, while maintaining high standards and performance. As a result the top three designs are meant to be simple which should translate to a lower acquisition cost. By simple the author means that there is nothing in terms of configuration that the market has not seen before. Although this might appear as un-imaginative the author realizes that in the real world, efficiency is the most important factor. That is why a lot of fighters appear to have the same configurations, one engine and a type of delta wing or normally swept wing. As a result having aircraft that deviate from these typical configurations did not appear to be the wisest choice, as top engineers have been designing the fighters and trainers,,and all appear to agree on the best configuration type. The top three designs are shown on the next page, with a black border around them and are labeled as configurations 7,8, and 9. The first design picked was a conventional design that is powered by two buried aft engines. It has a forward sweep and a twin tail configuration. The twin tail is to try to better prepare the pilots for the F-18 and F-35 configuration as they possess this empennage configuration. Configuration 8 represents a typical, one engine configuration, although the most common and used, is very effective and is still being applied, as seen on the recently unveiled Cessna Scorpion. One vertical tail would suffice under this configuration as only one engine is being used and is located on the center of the fuselage. The aim of this type of configuration will be to minimize the acquisition cost, as the author believes that this design would be the most cost effective out of all the ones presented. The final configuration selected from the list is that of a forward swept wing, with two engines. Forward swept wings have been proven to be beneficial; however they have not been applied to fighters due to elastic stability issues seen at the tips. The author believes that these issues can be overcome with the implementation of new technology and thus selected this configuration fore preliminary design. Aerospace Engineering Department 21

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23 7 Class I Design for Down Selected Configurations This Chapter contains the cockpit and fuselage designs for the three aircraft, as well as show more detailed figures of the final selected configuration. It should be noted that this chapter shows the results for two of the final three configurations selected, as the other configuration is the final one and all of the figures for that configuration can be found starting in Chapter Class I Cockpit and Fuselage Most of the designs presented in Chapter 6 have a similar fuselage as there is not a lot that the aircraft must accommodate. The major difference between most of the designs is the location and configuration of the wing and the configuration and location of the empennage. For the most part, the aircraft is being made to just carry the student and instructor, engine, fuel, and avionics in the fuselage. The following two fuselages will be applied in the designs. Each general fuselage shape will stay the same for Class I; however the dimensions vary depending on the final location and size of: the wing, empennage, landing gear, fuel tank, and payload. The following figures show the fuselages for the configurations that were contemplated. The overall length of each of these designs was 34 feet, which is short when compared to other jet fighters on the market, which tend to have lengths between feet. However the author felt that this was all of the room that would be needed on the fuselage, for the systems and fuel. Some of the fuel would be stored on the wings, with the majority of it being on the fuselage on a tank right behind the second pilot for the first configuration. The second configuration has a tail that shifts upwards at the end of the fuselage. This was done because the engines would be located underneath the wing and to reduce the landing gear length which is sized primarily by the longitudinal rotation angle. The inlet locations for the two configurations would be on the sides. The inlets for the first configuration would run along the side of the fuselage, and then merge into the fuselage. This would generate a large flop of air that the two aft engines would receive. For the second configuration the inlets would run on the side of the fuselage directly into the engines. One inlet for each engine since the engines are located on the sides of the fuselage underneath the wing it is not necessary to have the inlets enter the fuselage at any point. The fuel for the second configuration would be stored mostly on the wing. 24

24 Figure 7.1: Fuselage 1 Isometric View and Side View Figure 7.2: Fuselage 2 Isometric View and Side View 7.2 Propulsion Integration Following the configurations determined in Chapter 6, the two finalists that were selected were the twin engine configurations. As a result both of the CAD s presented in this Chapter, demonstrate where the engines would be located on the fuselage. It can be seen that the engine location was at the back of the fuselage; however for two engines this proposed a center of gravity issue when determining the maximum shift. To solve this the author decided to move the two engines forward to be nested under the wing, although this helped solve the center of gravity issue presented for the first configuration the author decided that the implementation of two engines was not necessary. Figure 7.3: Configurations 1 and 2 Engine Placement 25

25 7.3 Class I Wing and High Lift Devices Table 7.1: Wing Characteristics Wing 1 Wing 2 S (ft^2) AR b (ft) C_r (ft) C_t (ft) L.E. Sweep (deg) Forward 35 0 Taper Ratio From the preliminary sizing it was determined that the design would not need a high lift coefficient for takeoff, as a result the wings would not need high lift devices. Only ailerons will be needed for maneuvering. As a result the following are the wings for the first two configurations. The following figures show the contemplated wings for the first two configurations. Figure 7.4: Design 1 Forward Swept Wing, Design 2 Hershey Bar Wing 7.4 Class I Empennage Class I empennage was completed in accordance with the method presented by Dr. Jan Roskam s Aircraft Design Part III, reference 9. This preliminary method of design is performed before determining the center of gravity of the aircraft. Once the center of gravity the x-plots are generated to verify the sizing of the empennage from Class I. The empennages were incorporated into the designs and are shown in the following figures. Figure 7.5: Empennage Integration (Configuration 1, Configuration 2) 26

26 7.5 Class I Undercarriage Using statics and dynamics the author determined the force of the weight that each of the landing gear locations would see as a downward force, this in turn is the force that the tire would be able to withstand. The actual extension of the landing gear strut was then estimated by applying the later and longitudinal tip over criteria. This is explained more in depth in the following chapters. The following are the Class I locations of the tires of the landing gear for the first two configurations. The lines indicate a 5 degree angle for the front view and a 15 degree angle for the side view. Figure 7.6: Lateral and Directional Tip Over Criteria for Designs 1 and Class I Weight and Balance The center of gravity excursion was determined for each of the two configurations by approximating the weight component of each component of the aircraft. This was done using the methods presented in reference 8. The C.G. excursion was specified by the RFP to be less than 7% as a fraction of the mean geometric chord of the wing. The following table contains the weights of the different components as estimated by using the equations provided in reference 8. The C.G. graphs are also given below. 27

27 Table 7.2: Class I Component Weight Distribution WEIGHT BREAKDOWN Fraction Weight MAX T.O Empty Crew Payload Fuel Total Non Colored Fractions Estimated Using Appendix A in Part V Further Breakdown is Necessary of Empty Weight Fuselage Wing Engines Avionics Avionics Vehicle Management System Electrical System Auxiliary Power Unit Ejection Seat Ejection Seat Onboard Oxygen Generation System Onboard Inert Gas Generation System Empennage Group Landing Gear Group Total Empty Weight *F124-GA- 200 Engine * RFP Specified Figure 7.7: C.G. Excursion Diagrams (Configurations 1 and 2) 28

28 8 Final Down Selection The down selection to one configuration from three came down to the number of engines. All three of the configurations had similar fuselage layouts, with the exception of the locations where the engines would be placed. The author had originally intended on using two engines to better replicate the flying experience of modern jet fighters. This however was deemed to be a poor trade for the increased weight and fuel consumption that the two engine had over the one engine configuration. The author decided not to stray from his original design philosophy of producing the most cost effective aircraft, not only at the time of acquisition, but also during the life of the aircraft. The extra engine would significantly increase the fuel consumption of the aircraft, which amounts to approximately and extra 1500 pound of fuel per mission. That is something that the author sees as unacceptable because fuel prices are on the rise and might keep on rising. It must be understood that there is a business aspect to any type of design, as much as the designer would like to put more power break the specifications asked for, the least expensive method will be selected by the consumer most of the time, and as of right now the consumer is extremely concerned with the fuel consumption of the training aircraft. As stated by Major General Edward L. Bolton Jr. who is the Air Force s deputy assistant secretary for budge, the Air Force predicted the cost of the fuel used on training missions to be 1.3 Billion dollars less than it actually cost for the fiscal year of 2013 (Ref. 10) It is difficult to try to estimate the actual cost of fuel for the introduction year of 2020, but the author is convinced that the price can only increase from where is currently is. Making each pound of fuel saved per mission that much more valuable. For this reason the author decided to eliminate the two engine configurations, opting for the fuel and structural weight savings of the one engine configuration. Figure 8.1: Air Force Fuel Expenditure (Ref. 10) 29

29 9 Final Configuration Cockpit and Fuselage The final configuration selected was of one engine buried at the back of the fuselage. The inlets are placed on the sides and serve the purpose of supplying the engine with air as well as storing the landing gear mechanism. The landing gear mechanism is stored on the bottom of the inlet, and that area has been accounted for in the design and sizing of the inlet. The fuselage is thinner at the sections were the wing would be located to best accommodate the Area Rule. The overall length of the fuselage is 37.2 feet and this is sufficient to properly store the fuel tank in the fuselage as well as leave enough room for the engine inlets to converge and feed into the engine. This length is in accordance with the length of similar military trainers, such as the Yak-130, M346, T50 Golden Eagle, and BAE Hawk who have a fuselage length within the feet range (Ref. 2). It is necessary for a ladder to be stored within the aircraft for the pilots to have access to the aircraft as asked for by the RFP, so the author has decided to implement the same design that was used on the A-10 Thunderbolt II, which is of a retractable ladder located on the side of the fuselage. Due to the cockpit only opening on one side the author decided to have only one of retractable ladder on the side of the fuselage, this means that the pilots will Figure 9.1: A10 Ladder Fully Extended (Ref. 11) have to take turns when entering the aircraft. The following figures show the overall shape of the fuselage as well as the ladder fully extended and retracted. Consideration was made to creating a payload bay within the fuselage of the aircraft, however this was discarded as the small size payload asked by the RFP more than likely refers to combat missiles which are to be held by the wing. The aircraft cockpit is contains zero-zero ejections seats as requested by the RFP. It also contains heads-up displays (HUDs) and helmet-mounted displays (HMDs) and night-vision goggles. The front pilots line of sight is clear for the first 15 degrees as specified by Reference ##, while the aft pilot s line of sight is clear for the first 5 degrees as specified by the same reference. The line of sight 30

30 starts at the eye of the pilot. It should be noted that for both of the line of sight angles shown in the figures below have a small female pilot as specified by the RFP. Since the smaller female will have a line of sight that starts at a lower location than the 95 th percentile male, this became the line of sight to verify. The aircraft s cockpit is fitted for a 95 th percentile male as well as the small stature female. 10 Final Configuration Propulsion Integration Within this Chapter the reader will learn some of the reasons behind the selection of the engine to be used, as well as learn about the performance of the engine with varying altitude and Mach Class I Propulsion Integration Due to the Statistical Time and Market Predictive Engineering Design Vector Analysis results presented in Chapters 4 and 5, the design point was selected, such that the necessary power is 4752 pounds of thrust for normal flight. However this design point does not satisfy all of the requirements, as significantly more thrust is needed for the 9g maneuvering at feet above sea level. The total thrust necessary to Figure 10.1: CAD Engine accomplish the 9g maneuver is equal to 7560 pounds of force for the one engine configuration. The AI-222K-25F engine was selected for implementation on this design. This engine when equipped with an afterburner produces a thrust pounds of thrust which is more than is needed for the 9-g maneuver according to the sizing plot. Without the use of the afterburner the engine produces a thrust of 5520 pounds of force according to reference 11, which is more than needed according to the sizing chart for normal cruise condition. This Figure 10.2: AI F Turbofan Engine With Afterburner (Reference 12) engine has a bypass ratio of 1.19:1, length of 77 inches 31

31 which accounts for the afterburner portion of the engine, fan diameter of 25 inches, and weight of 1234 lbs. It should be noted that although the fan diameter is that of 25 inches, the widest part of the engine is 32 inches wide Class II Propulsion Integration The author decided to incorporate a thrustvectoring nozzle to the engine to help increase the maneuverability of the aircraft while during take-off, flight, stall, and landing. The AI F engine presents no issues when equipped with a thrust-vectoring nozzle as declared by the manufacturer. Figure 10.3: Forces and Moments Acting on Aircraft (Ref. 13) The incorporation of thrustvectoring emerges from various issues presented by the mission of the aircraft, which is to train pilots. The first major problem that the author considered was the high g-maneuver which can be a daunting task for a student pilot which has never completed it. In 2007, Mr. Atesglu and Mr. Figure 10.4: Integrated Bihrle-Weissman Chart for the Aircraft (Ref. 13) Ӧzgӧren published a journal article on the Journal of Guidance, Control, and Dynamic, from reference 13, describing the great advantages of a using a thrust-vectoring control on jet fighters and bomber aircraft. The gentlemen created a model of the aircraft dynamics during the thrustvectoring phase, by defining the Earth fixed reference frame and the aircraft body frame. For this the aerodynamic forces and moments are accounted for as disturbances under this model. Figure 10.3 shows the forces and moments acting on the aircraft as the model was generated. The aircraft aerodynamics were modeled for three conditions which are: high angle of attack behavior, departure indication parameters, and nonlinear modeling of the aircraft s aerodynamics. Figure 10.4 shows the Bihrle-Weissman chart for the aircraft 32

32 being modeled, it is to note that the region represented on the figure as region A is the safest region to be in and the angle of attack range that the aircraft sees in this range is -15 to 17 degrees (Ref. 13)The Bihrle-Weissman chart shows the departure and stall regions of the aircraft s angle of attack range. The gentlemen also modeled the engine, flight environment and thrust-vectoring paddles. The two controllers that modeled are what is being tested. The first controller is used to test the thrust-vectoring paddles by allowed the user to alter the aerodynamic effects on the thrust-vectoring paddles. Once the aerodynamic controller no longer has the capability to control the aircraft, the thrust-vector controller is engaged, this typically occurs at high angles of attack. The second controller was designed to control the angle of attack of the aircraft. From the journal article it was determined that the Cobra maneuver could only be reached if the thrust-vectoring control was on. The Herbst maneuver which deals with lateral as well as longitudinal movements by the aircraft cannot be initiated at moderate to high angles of attack, unless the thrust-vectoring control is on. Under the previous condition the aircraft is not yet in stall, however it cannot perform the maneuver. The study was intended to show the need for a controller which can automatically start the thrust-vectoring on the aircraft when there is indication of stall or departure. The author of this report expects the follow up studies asked for by the authors of reference 13 will be performed and a system generated. The incorporation of this system would help ease the stress of high angle of attack maneuvers on the aircraft and pilot. For now the author has decided to implement a thrust-vectoring nozzle to help with the maneuverability of the aircraft. No other modifications were made to the engine Class II Propulsion Performance The main concern with any engine is if it will be able to generate enough thrust to power the aircraft at various altitude and velocities. Starting from the uninstalled dry thrust from reference 14, the author scaled the available data to the engine selected as the data for the selected engine was not available to the author. The trends in thrust decrease as the flight Mach increases past Mach 1.5, however the author is not interested in that region so ti was omitted. From the installed dry thrust it was determined that the selected engine would have enough power as called for by the sizing chart at sea level. 33

33 Installed Dry Thrust ~ T ~ thousand lbf Uninstalled Dry Thrust ~ T_tst/av ~ thousand lbf K feet 1K feet 2K feet 3K feet 4K feet 5K feet Mach Number ~ M Figure 10.5: Uninstalled Dry Thrust Engine Performance (Non-Afterburner) Mach Number ~ M Figure 10.6: Installed Dry Thrust 0K feet 1K feet 2K feet 3K feet 4K feet 5K feet 34

34 11 Wing and High Lift Several designs were considered for implementation on the Pegasus Jet Trainer, from forward sweep to a conventional aft sweep wing. The final wing selection was that of a double delta wing with a diamond fillet at the double delta juncture. This chapter outlines the preliminary though process, the reasoning behind the selection of the double delta wing, and details its implementation into the design. As discussed before, the sizing chart proved that there was no need for high lift devices as, take-off and landing lift coefficients were low. The landing coefficient from the sizing chart was that of 1, while the takeoff lift coefficient was equal to.65. As a result the aircraft clean coefficient would be able to accommodate the two other coefficients without much of an issue. The decided to implement a double-delta wing configuration as explained in the following paragraphs. Having decided on picking the best design out the three designs presented above, the author sought the guidance for the design of military trainers or fighters. The following quote was found by the author: Vertical tail buffeting at high angles of attack is a phenomenon associated with the impact of vertical flows generated by the aircraft on. This poses a serious problem for both single and twin-tail fighter aircraft from the point of view of combat maneuverability and structural integrity (Ref. 15). This same reference states the reason for a leading edge extension being used on the F-16 and F-18 is that leading edge extensions are great for generating nonlinear lift at high angles of attack. There is a major downfall to this advantage and that is that the vortices can dramatically reduce the lifespan of the vertical tail (Ref. 15). The damage created by the vortices can be historically traced to the F-18 and its problems with its vertical tails. Tail buffeting does not only occur with leading edge extensions, it is also encountered at different angles of attack or sideslip angles for conventional winged aircraft. Unlike the F-18, a single vertical tail configuration can also be significantly affected with the use LEX. Reference Figure 11.1: Ref 17 (Tail Buffet Allevation Wing Strake) Fillet Geometry 16, demonstrated that vortices formed along the 35

35 leading edge of delta-wings can account for as much as 30% of the total lift that the aircraft will see at moderate to high angles of attack (Ref. 16). Myose conducted an experiment to determine the best wing configuration for ideal vortex breakdown. From his results, the most efficient wing had a cropped or 90 degree wingtip, while maintaining a delta shape. All of the previous references and future references had similar conclusions, which were that to achieve better lift at moderate to high angle so attack the implementation of LEX and a cropped delta wing were essential. This will be referred to as a cropped double delta wing. Reference 17 was an experiment performed to determine if the LEX and delta wing juncture fillet shape had an impact on the performance of the aircraft or buffeting. Figure 11.1 demonstrates the different fillet shapes used in the experiment. The results showed that the diamond fillet increased the characteristic frequency off the wing and the buffet pressure amplitude to decrease. Both of these things lead the authors of reference 17 to deduce that the diamond fillet shape can be used to alleviate vertical tail buffet. Figure11.2 which was taken from reference 17 shows the average pressure coefficient for the four different fillets. The diamond fillet clearly has a lower average pressure coefficient than the other fillets at moderate to high angles of attack. Although the other fillets appear to be a downgrade from the baseline configuration, the diamond fillet improves the baseline configuration. 36

36 Figure 11.2: Pressure Coefficient Distribution (Ref 17) 37

37 Not only does the diamond fillet help with the buffeting on the vertical tail of the aircraft and helps augment the lift of the wing at moderate to high angles of attack, but it also improves the operational envelope of the wing Figure 11.3: High AOA operational envelope for two models (Ref. 18) when compared to the baseline model. This was the conclusion made by the authors of reference 18. Figure11.3 shows the results from that experiment, with the clearly defined high angle of attack operational envelope for the diamond fillet being better than that of the baseline configuration. This is a great thing for military trainers and fighters according to the reference because maneuvers at high AOA and appreciable sideslip angles provide more agility and are therefore desirable. According to the results, it is advantageous to have vortex flow control in this regime (pg. 7 Ref. 18). It might become apparent to the reader, if the references are consulted, that all of the previous experiments were performed on just the wing configuration. No fuselage was present on the wing for any of the previous tests and this might completely throw off the results obtained and conclusions made. The author of this report was skeptical about this as well, until reference 19 was encountered. This reference performs some the similar procedures in the other references to try and validate the results. The conclusions is the following: [The] double delta wing with the LEX in the present study maintains a stabilized and well-organized vortex system due to the strong interaction of the wing vertex and the LEX vortex, which subsequently results in the presence of the center body having minor effect on the flow pattern and the wing-upper-surface distribution (Pg. 7. Ref. 19). 38

38 Figure 11.4 demonstrates the similarities between the strake and wing vortex interactions for center-body fuselage on and off the wing. Figure 11.4: Comparison Between Vortex Interaction with On and Off Fuselage (Ref. 17) From all of the previous information that the author gathered and the fact that most jet fighters and newest jet trainers incorporate LEX and a delta wing, the author made the decision to alter the design of the wing to the following. The author had intended on generating a larger sweep angle for the second delta wing, however this was not possible due to the high aspect ratio for the low area wing. If Figure 11.5: Final Pegasus Wing Configuration the sweep was made any bigger the chord of the wing would have dramatically decreased. As seen in Figure 11.5, the author did decide to implement the 39

39 Table 11.1: Pegasus Aircraft Wing Characteristics Section b (ft) C_r (ft) C_t (ft) Equivalent Sweep (deg) 31.3 Total S (ft^2) 141 Total b (ft) 27.3 diamond fillet at the juncture of the LEX and delta wing as well. This is under the assumption that the 9-g maneuvering will include moderate to high angles of attack or significant sideslip angles. If this scenario is correct then the fillet gives the pilot an easier job of handling the aircraft at said angles of attack. This fillet is not detrimental to the design at low angles of attack either, however it is most effective at moderate to high. As stated in Chapter 9.1 of this report, the low lift coefficients for landing and takeoff prevent the aircraft from needing flaps for landing and takeoff. As a result they were not implemented. Ailerons however are and they span from 50% of the wing span to 90% of the wingspan, while having a chord of aileron to chord of wing of approximately 25%. The wing was split into four partitions to model in Tornado (Reference Tornado here), which is a software that uses the Vortex Lattice method for linear aerodynamic wing designs. This code is open to the public use and runs on Matlab. The author of this report made two wing configurations for the Pegasus. One with the diamond fillet at the juncture, while the other wing does not incorporate this characteristic. Figure 11.6: Tornado Wings (Diamond Fillet - Baseline Configuration) The baseline configuration was run first, and a lift coefficient was determined for a set flying condition. The second wing configuration as run and the lift was normalized by altering the flying condition until it matched the lift generated by the first wing. At this point the wing was swept as a function of angle of attack and the resulting drag polar was altered to more closely match that of the aircraft. This was done by increasing the C_D by the 40

40 C_Do which is not accounted for in Tornado. The results from tornado showed an increase to the lift coefficient for any given drag coefficient. Having the wing configuration selected, the author, sized the high lift devices to use the NACA 65A004 airfoil for both the tip and root of the wing. The reason for a non-cambered airfoil is that it makes no contribution to lift in tat supersonic airspeeds while producing a moment that is dependent on the camber of the airfoil and making a contribution to drag, which are things which are not desired (Ref 20). From reference 20 it is seen that the wave drag is proportional to the thickness of the airfoil squared. This means that the lower the thickness value, the less wave drag that will be generated. Thus the author of this report decided on selecting a thin airfoil with no camber since very little fuel is being stored on the wing. 41

41 12 Empennage For the preliminary empennage sizing for the different designs, the methods presented in Chapter 8 of Dr. Jan Roskam s Part II, reference 8, of his Airplane Design sequence were used. Using equations 12.1 and 12.2 as well as Tables 8.8a and 8.8b on page 199 from the same reference, the preliminary surface area of the horizontal tail and vertical tail were approximated. (Eq. 8.1, Ref. 8) Eq (Eq. 8.2, Ref. 8) Eq Table 12.1: Empennage Characteristics Using this method the author generated a vertical tail with Horizontal Vertical S (ft^2) C_r (ft) C_t (ft) AR /4 chord Sweep ( i_w (deg) t/c (%) 10 0 an area of 24 square feet and a horizontal tail with an area of 19 square feet. This preliminary approximation was not deemed correct as is shown in the Stability and Control chapter of this report. As a result the empennage was resized to the values determine through the stability and control analysis. Figure 12.1: Vertical and Horizontal Tails 42

42 13 Undercarriage Using Dr. Roskam s Airplane Design Part II, reference 8, the initial landing gear location was estimated. Using ratios in Table 9.2, of reference 8 and basic statics and dynamics the loading was determined for each tire. With this using Table 9.2, the landing gear tire sizes were determined. For the nose gear the diameter is 17 inches and 4.4 inches wide. For the main gear the dimensions are 18 inches by 6.5 inches. The main gear has a pressure of 100psi while the nose gear is at 130psi. Type VII tires were chosen because they have a high load capacity, while keeping a narrow width. The tip over criteria were satisfied and shown in the following figures. The main gear will retract outwards, while the nose gear rotates downward. For the tip-over criterion, from Reference ##, there should be 5 degrees clear between the landing gear ground point and the rest of the aircraft. For the longitudinal this angle should be no less than 15 degrees. The minimums are shown in the figures below. Figure 13.1: Longitudinal Clearance Angle 43

43 Figure 13.2: Lateral Ground Clearance Using the following equations the strut sizing was determined for the aircraft landing gear. ( ( ) ) (Eq. 2.11, Ref. 8) Equation 13.1 (Eq. 2.13, Ref. 8) Equation 13.2 From these equations the following landing gear shock absorbers were determined. For the stroke of the main gear, 13 inches are needed while the diameter of the stroke should be equal to 2.6 inches. For the nose gear, the stroke of the shock absorber should be equal to 5.4 inches while the diameter of the shock absorber should be equal to 1.3 inches. 44

44 14 Weight and Balance Using the methods outlined in Chapter 10 of reference 8 the Class I Weight and Balance was created for the proposed aircraft. This preliminary C.G. excursion was performed to help in further the process, so it will not be presented in this section. The final aircraft weight and balance is determined through Class II weight and balance, so all of the information presented from here on refers to Class II weight and balance. Applying Torenbeek s equations to the various aircraft components the weight of said components of the aircraft were determined. Torenbeek method was preferred as it accounted for the g-loading that the aircraft would encounter, therefore becoming more accurate than the other equations for this procedure. It is to note that although the wing weight decreased by approximately a hundred pounds from what was originally estimated, the overall weight of the aircraft stayed relatively constant. The wing was not the only structure that had a change differential. With the new component weights the weight and balance was performed again and iterated until the design had a center of gravity shift less than 7% of the mean geometric chord of the wing. The following equations were used: ( ) ( ) [ ( ( ( ) ) ) ] ( ) ( ) (Eq. 5.5, Ref 21) Eq (Eq. 5.16, Ref 21) Eq ( ) ( ) (Eq. 5.27, Ref 21) Eq ( ) (Eq. 5.42, Ref 21) Eq

45 (Eq. 6.10, Ref 21) Eq ( ) ( ) (Eq. 6.17, Ref 21) Eq (Eq. 6.27, Ref 21) Eq ( ) (Eq. 6.34, Ref 21) Eq The following table summarizes the results and gives the C.G. excursion for the current aircraft configuration. It was determined that the change is weight was below 5% which indicates that a recalculation of any values up to this point is not necessary. The center of gravity maximum shift is under the 7% that was specified in the RFP, as it comes in at 6.6% of the mean geometric chord. The component number on Figure 14.1 corresponds to the number found on Table Figure 14.1: Weight Distribution 46

46 Table 14.1: Class II Weight Determination Object Weight X_CG W*X 1 Fuselage Wing Engines Avionics Avionics Vehicle Management S Electrical System Auxiliary Power Unite Seat Seat Oxygean System Gas System Horizontal Tail Vertical Tail main Gear Nose Gear Air Induction system Fuel System Air Conditioning, Pressurization Electrical System W_F Payload Pilot Pilot E 6109 X_CG_E OE 6659 X_CE_OE OE+Fuel 9748 X_CG_OE+Fuel TO X_CG_TO OE+Payload 7659 X_CG_OE+PAYLOAD CG_Shift inches C_bar= ft X_CG_Shift/C_bar = The parts of the table marked in yellow indicate the weights as specified by the RFP. 47

47 Figure 14.2: C.G. Component Locations Figure 14.3: Class II Center of Gravity Excursion 48

48 Yawing Moment Coefficient Due to Sideslip, C_n_beta, 1/deg 15 Stability and Control The previous equations correspond to equations 11.1 and 11.2 from reference 8. This analysis is meant to determine the validity of the horizontal and vertical tail area determined with the Class I sizing. It is to note that this analysis was performed before the completion of Chapter 12 of this report, therefore the values obtained here correspond the empennage configuration presented in said chapter. To determine the static longitudinal stability as well as the static directional stability, two x-plots were created. The x-plots were created following the steps laid out in Chapter 11 of reference 8 The calculations determined that the aircraft would need a horizontal tail area of approximately 16 square feet and a vertical tail area of approximately 29 square feet Horizontal Tail Area, S_v, ft^2 Figure 15.1: Directional X-Plot For the directional x-plot a yawing moment due to sideslip of.001/deg should be achieved for the aircraft to be stable, this can be found in reference 8. 49

49 AC and CG Location with Respect to mean geometric chrod, % X_bar_ac X_bar_cg Horizontal Tail Area, S_h (ft^2) 10 The longitudinal X-plot was generated to determine the horizontal tail area needed for the desired static margin. The static margin desired is that of -5% as requested by reference 8 This is done to allow the aircraft to be inherently stable, while making it maneuverable by the pilot. From the configuration entered in AAA the following results were determine for the aircraft s trim diagram. From this point forward the program Advanced Aircraft Analysis which is a program created and distributed by DARcorp was used to help determine and iterate the design as needed. The following figure demonstrates the aircraft configuration entered into AAA to help with the modeling of the Pegasus Jet Trainer. The name of the file used for this purpose is the Pegasus Jet Trainer AAA (Ref. 22). From the AAA file and using the lift equation for cruise condition it was determined that the optimum cruise altitude is at 37 thousand feet above sea level at a Mach of.9. The maximum Mach number is presented in Chapter

50 . Figure 15.2: Pegasus Jet Trainer AAA Configuration Figure 15.3: Trim Diagram 51

51 Static margin is equal to -1.95% according to AAA. After having all of the necessary inputs into AAA, the program returned a value of 3.2% as the static margin of the aircraft. It should be noted that the author had to give the horizontal tail an incidence angle of -2.5 degrees to get the current configuration to be at an elevator deflection angle of approximately zero. This was done because the current configuration refers to the aircraft during cruise conditions. 52

52 Fuselage Perimeter, Inches 16 Drag Buildups and Drag Polars The following are the drag polars generated for the Pegasus Jet Trainer. These are based off the wetted area of the aircraft, which was determined following reference 9. The wetted area of the entire aircraft was calculated to be 760 ft^2. At a skin friction coefficient of.004, and equivalent parasite area of 2.9, the C_Do of the aircraft was determined to be.02. The fuselage area was determined by finding the perimeter at certain fuselage stations and integrating between the two fuselage stations. Figure 16.1 shows the fuselage perimeter as a function of fuselage station, with all of the area underneath the curve becoming the wetted area of the aircraft s fuselage. The wetted area of the wing, vertical, and horizontal tail were also calculated, this however was done using equation It became difficult for the author to meet the area rule with the double delta wing, as there is a sudden drop off in area, halfway through the fuselage. Eq Fuselage Station, F.S., Inches Figure 16.1: Fuselage Area 53

53 Lift Coefficient, C_L Figure 16.2: Aircraft Cross Sectional Perimeter Clean Configuration Takeoff Configuration Landing Configuration Drag Coefficient, C_D Figure 16.3: Drag Polar of Pegasus Jet Trainer The previous figure was generated using the following equation: 54

54 Eq It should be noted that there are only two major trends that the drag polars follow, this is because there are no high lift devices in the design, so the aircraft relies solely on the addition of the landing gear for both the takeoff and the landing cases. With the use of the configuration that was entered into AAA, the following component breakdown was determined for ach aircraft component. Figure 16.4: Class II Component Drag Buildup 55

55 Load Factor ~ n 17 Performance Verification The V_n diagram for this aircraft was created following the procedures laid out in reference 21and is the following: Speed ~ Mach ~ M Figure 17.1: Pegasus Jet Trainer V-n Diagram (Ref. 21) This can be compared with the V-n diagram of the T-38 although they differ in maximum load factor. The T-38 was however designed to reach a higher velocity than this aircraft, therefore the V-n diagrams vary in those aspects. With the help of the AAA program a lift to drag ratio comparison was made with Mach. As Mach increases, the aircraft L/D decreases. The following figure demonstrates this: 56

56 Lift to Drag Ratio, L/D, ~ Mach, M ~ Figure 17.2: L/D vs Mach at feet Based on these calculations it can be determined that the maximum Mach number at feet is Mach =.97. This value increases as altitude increases, so the maximum Mach number at feet is equal to Mach =.99 to maintain an L/D over 8. The reason why an L/D of 8 is desired to be kept is that when this lift to drag ratio will maintain the weights determined up to this point within 5%, which according to Dr. Roskam in reference 8 means that the calculations do not need to be iterated. 57

57 Specific Excess Power, ft/s, P_s Specific Excess Power, ft/s, P_s Mach, M Sea Level 5K feet 10K feet 15K feet 20K feet 25K feet 30K feet 35K feet 40K feet 45K feet 50K feet Figure 17.3: 1-g Maximum Thrust Specific Excess Power Envelope (Non Afterburner) Mach, M Sea Level 5K feet 10K feet 15K feet 20K feet 25K feet 30K feet 35K feet 40K feet 45K feet 50K feet Figure 17.4: 5-g Maximum Thrust Specific Excess Power Envelope (Afterburner) 58

58 Specific Excess Power, ft/s, P_s Mach, M Sea Level 5K feet 10K feet 15K feet 20K feet 25K feet 30K feet 35K feet 40K feet 45K feet Figure 17.5: 9-g Maximum Thrust Specific Excess Power Envelope From these figures it can be seen that the aircraft is capable of cruising at an altitude of feet which is what was desired for the design of the aircraft. The maximum Mach seen is of approximately 1.5 and the aircraft can perform the 9-g maneuver at feet altitude. 59

59 18 Systems and Structures The flight control system layout consists of the primary and secondary flight control systems. The primary flight control systems are the lateral controls to the ailerons, spoilers and differential stabilizer. Primary longitudinal controls include elevator, stabilizer, and canards if present. Primary directional controls only include the rudder. Secondary flight control systems include trim controls, high lift controls, and thrust controls. The aircraft will be fly-by wire with electromagnetic actuators. At this point the reader will be reminded that there are not flaps on this aircraft; therefore only the primary control surfaces must be connected to the cockpit. The following figure shows the airframes and expected ribs of the aircraft. Figure 18.1: Frames and Ribs of the Pegasus Jet Trainer (Side View) Figure 18.2: Frames and Ribs of the Pegasus Jet Trainer (Top View) 60

60 Figure 18.3: Frames and Ribs of the Pegasus Jet Trainer (Isometric View) The aircraft will be assembled by generating the fuselage first. Afterwards, the wing is attached by joining the two sides on the fuselage. Finally the horizontal tail is joined on the fuselage as well. The vertical tail is assembled on its own and is joined to the fuselage. The engine is pulled directly from behind the fuselage. Figure 18.4: Aircraft Components for Assembly 61

61 19 Advanced Technologies A lot of the things that the author is implementing have been covered in the previous chapters, as there is nothing that is completely new being used by the author. The LEX-wing combination has been used and the author deemed it necessary to explain the fillet juncture in the high lift device section. Thrust-vectoring is also not new to the industry so the author does not believe that it is essential to cover it under this chapter. 62

62 Cost of Individual Aircraft, $x10^6 20 Cost Estimation With the use of the available AAA program created for use in the previous sections, the various costs associated Table 20.1: Predicted Aircraft Cost with the aircraft were determined for various production rates. The For the Year 2013 Number of Aircraft Cost $ (x10^6) NRE Cost following figure shows the correlation between the number of aircraft built and the cost of each individual aircraft at that production rate. As specified in the RFP, the number of aircraft for which the cost should be observed is 100 to 700 aircraft in increments of 100 aircraft. The RFP requests that the aircraft not exceed a per unit cost of $20 million if valued in the year The objective is to have a design which is under $10 million if valued in The objective was met by the author for the production rates above 200 aircraft. The NRE cost was constructed partially by the RFP which called for 8 test aircraft. It called for one iron bird, so the author input this as 7 test aircraft and one static model. The NRE does not change depending on the number of aircraft being built, so the $340,400,000 expenditure is fixed. This however is lower than what the RFP called for as the objective set forth by the RFP was to be under $500,000,000. For sake of determining where the cost of the aircraft will try to even itself out, the author increased the production to a 1000 aircraft. At this point the cost per aircraft was still over $7 million, so it can be assumed that the aircraft price will be over $7 million at all production rates that are within reason Number of Aircraft Built, N_a Figure 20.1: Cost of Aircraft as a Function of Total Built 63

63 21 Characteristics The design of the Pegasus Jet Trainer was focused around saving fuel and lowering the acquisition cost. Those are the factors that drove the design to be of one engine with the implementation of the Ivachenko-Progress AI F engine as it has low fuel consumption when compared to other engines. The increase in maneuverability that the leading edge extension and thrust vectoring provide, are meant to help the student learn and recover from any minor mistakes he/she might make during the training process. Table 21.1: Salient Characteristics Wing Loading 80 psf Thrust to Weight Cruise 0.44 Thrust to Weight Maneuver 0.73 Engine Uninstalled Thrust 5520 lbf (dry) lbf (wet) S 141 ft^2 b 27.3 ft AR 5.2 Equivalent 1/4 Chord Sweep 31.3 deg Wing Airfoil NACA 65A004 S_h 29 ft^2 C_r_h 7 ft C_t_h 2.5 ft S_v 16 ft^2 C_r_v 1.76 ft C_t_v 1.25 ft Max C.G. Excursion % of MAC W_TO lbf W_OE 6659 lbf W_E 6109 lbf Cost per aircraft at 500 P.R $ Million Cruise Mach 0.9 Dash Mach 1.3 Optimum Cruise Altitude ft g-loading at ft MSL 9 Ceiling 50,000 ft Minimum Runway Length 6000 ft Payload 1000 lbf Range 1500 nm 64

64 Figure 21.1: Final Aircraft 3-View 65