Systems Definition Review

Transcription

1 Super Sonix Inc. Systems Definition Review An analysis of performance, concepts, and sizing Akshay Ashok, Nithin Kolencherry, Steve Skare, Michael McPeake, Muhammad Azmi, Richard Wang, Mintae Kim, Dodiet Wiraatmaja, Nixon Lange 3/12/2009

2 Table of Contents TABLE OF CONTENTS... 2 TABLE OF FIGURES... 4 LIST OF TABLES... 5 INTRODUCTION... 6 Review of SRR... 6 Opportunity Description... 6 Mission Statement... 7 Market Analysis and Concept of Operations... 7 Initial Sizing... 7 CONSTRAINT ANALYSIS g Supersonic Cruise g subsonic Maneuver Takeoff and Landing Ground roll Second Segment Climb Gradient CONCEPT SELECTION Brainstorming Pugh s matrix Hybrid Concepts Hybrid concept 1: Hybrid concept Hybrid concept Aft arrow wing concept ADVANCED TECHNOLOGIES Boom shaping technologies Nose Design Blunt Nose Nose Keel Dihedral Angle Effective Area distribution Aerodynamic Cowling Engine Nacelle Placement Canards Efficient supersonic cruise Active Flow Management Engine Technology Modifications for Samara NK-321 Engine CABIN LAYOUT SIZING STUDIES Sizing Approach Page 2 of 61

3 Component Weights Center of Gravity Shock Prediction N-wave Plateau wave SUMMARY Compliance Matrix Walkaround and CAD Model Next Steps References APPENDIX Progress Update Chart Concept Selection Process Initial Concept Drawings Detailed Design Criteria Design Mission Profile Historical Seating Data Combined Mass fraction Averages Initial Sizing results Requirements Compliance Matrix Next Step Flowchart Page 3 of 61

4 Table of Figures Figure 1: Constraint Diagram Figure 2: Concepts used for the Pugh s method analysis Figure 3: Hybrid concept Figure 4: Hybrid concept Figure 5: Hybrid concept Figure 6: Aft Arrow wing concept [12] Figure 7: 3-view and isometric drawing of final design Figure 8: Blunt nose effect on boom overpressure [23] Figure 9: Dihedral angle design Figure 10: Non-Axisymmetrical fuselage shape [22] Figure 11: Engine Database Figure 12: Thrust-to-weight ratios for viable engines Figure 13: Proposed engine exhaust configurations [14] Figure 14: Directivity of Overall sound pressure level [14] Figure 15: Perceived Noise Level [14] Figure 16: Passenger Seating Options (pitch) Figure 17: Passenger Seating Options (width) Figure 18: Passenger Seating vs. Overall a/c Length Figure 19: Seating Layout with Galleys and Restrooms Figure 20: Seating Dimensions, Coach (L) First Class (R) Figure 21: Sizing Approach flowchart Figure 22: Component weight mass fraction Figure 23: Center of Gravity Calculator Figure 24: Shape factor and pressure amplification factor charts Figure 25: Walkaround chart Figure 26: 3-View of Super Sonix Aircraft Page 4 of 61

5 List of Tables Table 1: Flight Conditions and Aircraft Parameters for 1-g steady level flight Table 2: Parameters for subsonic 2-g maneuver constraint Table 3: Landing and Takeoff parameters for DXB airport Table 4: Takeoff and Landing parameters for JFK Table 5: Second segment climb constraint parameters Table 6: Assumptions Sensitivity Table Table 7: First run of Pugh s matrix Table 8: Second run of Pugh s matrix Table 9: Viable supersonic engines [5] Table 10: Current Sizing Attributes Table 11: Definition of variables for overpressure calculation Table 12: Input Parameters for N-wave overpressure calculation Table 13: Input Parameters for Plateau wave overpressure calculation Page 5 of 61

6 Introduction Over the last few years, air traffic has seen a high increase in capacity, range and efficiency. In addition to these, the need for a faster aircraft is imperative at the present time. Transporting people over longer distances within a short span of time will be the objective of the airline industry in the near future. This report addresses the systems definitions for a small supersonic airliner with Initial Operating Capability (IOC) in A progress update chart is presented in the appendix, and summarized tasks that were performed earlier in the Systems Requirements Review (SRR), while illustrating the steps that are taken as part of the Systems Definition Review (SDR) process. The SRR included an initial sizing exercise based on historical data, to estimate the aircraft s takeoff gross weight. Market analysis was performed to determine the focus markets and passenger loads, operating costs and aircraft utilization. Customer requirements and engineering design characteristics were detailed in the House of Quality as part of the Quality Function Deployment (QFD) phase. To better define the aircraft system, as part of the SDR, additional work is done to identify performance constraints using a constraint diagram. The process drives the need for a comprehensive engine database that is used to select an engine, or a concept, that fits the initial performance requirements of the Super Sonix aircraft. Concept generation is an important phase of the SDR which allows the aircraft to be visualized by taking on a geometrical form. The brainstorming of ideas and the Pugh s selection matrix exercises leads to further research in advanced technological concepts. These advanced considerations allow a validation of shortlisted concepts, and enables a preliminary calculation of the sonic boom overpressure. A component weight database is built to aid in determining component weight fractions, and leads to a better resolved takeoff gross weight prediction. Along with the cabin layout, these weights are used to calculate the center of mass of the aircraft system. The next section serves as a brief review of the detail of the analysis done in the SRR. Review of SRR Opportunity Description NASA ARMD University Competition provides a framework that is used to guide the conceptual design of a supersonic aircraft. The aircraft is expected to meet a set of goals specified by the competition. These include, but are not limited to: Mach cruise speed of Design Range of 4000 nautical miles Accommodate passengers (preferably in a mixed class configuration) Fuel efficiency of 3 passenger-miles per pound of fuel or better Takeoff field length < 10,000 ft Page 6 of 61

7 Additionally, the aircraft is expected to achieve supersonic cruise efficiency, have a low sonic boom (<70 PldB) and high lift for takeoff or landing all while making a reasonable profit for the company. Mission Statement The Super Sonix mission statement is as follows: A cost-effective, advanced, high-speed commercial air transport that connects major worldwide hubs Super Sonix has set out several key design goals that is expected to be met and exceeded through the course of the conceptual development: Supersonic flights over land (boom overpressure less than 0.3 lb/ ft 2 ) Capture significant market of Supersonic Business Jets and supersonic transports Initial Operational Capability (IOC)in 2020 Manufacturing capabilities exist for the aircraft Payload of 60 passengers in a twin class configuration Still-air ground range of 4000nm Market Analysis and Concept of Operations Three focus markets emerge from the market analyses: Trans-Continental, Trans-Atlantic and Inter-Asia. Through detailed studies on market viability and cost models, potential worldwide hubs are selected to set forth a viable business proposition for customer airlines. These are: Los Angeles International Airport (LAX), John F. Kennedy International Airport (JFK), London Heathrow International Airport (LHR), Dubai International Airport (DXB) and Beijing International Airport (PEK). A hub and spoke structure is to be employed, with the reliance on the growing number of Low Cost Carriers to provide regional connections. Initial Sizing The initial sizing process is performed using an iterative approach. First, a database of similar aircraft is compiled. Next, a least squares regression is used on the database to find the coefficients of the basis function. The design mission is then used to determine W f /W 0. The sizing equation was then iterated to find the final weight of the aircraft. The results of this initial sizing are plotted against the database aircraft to validate the results. These results are summarized in the appendix. In the following section, constraint studies are performed to determine the limits that are placed upon the Super Sonix Aircraft during the operations to and from hub airports. Page 7 of 61

8 Constraint Analysis During the operation of the Super Sonix aircraft, there are several performance constraints that the aircraft needs to meet in order to successfully carry out its mission, perhaps even to be airworthy in the first place. Several operational conditions are considered in this analysis: 1-g supersonic steady level flight, 2-g subsonic maneuver, takeoff and landing ground roll during a hot day and short runway, and a 3% second-segment climb gradient. Several values are based on the design mission (which is included in the appendix). The constraint diagrams aid in identifying a viable range of operating wing loading and thrust-to-weight ratios. 1-g Supersonic Cruise The ability to maintain supersonic cruise is vital to the success of the Super Sonix aircraft. The derivation of the equation that relates thrust to weight T SL /W 0 to wing loading W 0 /S begins with the specific excess power relation: P s = V(T D) W = d dt + V dv g dt Equation 1 V is the aircraft velocity, T is the thrust at altitude, D is the drag force, and W is the current aircraft weight. We note that the thrust of a turbofan engine decreases with altitude by the following definition: T = αt SL Equation 2 Where the thrust lapse rate α, is approximated as: α = ρ altitude ρ SL Equation 3 For subsonic conditions, the density at altitude ρ altitude can be found from standard atmospheric tables. However, for supersonic flight, the lapse rate is approximated using an effective density ρ effective. This effective density is obtained using normal shock jump relations, assuming a normal shock in front of the body. This approximation will be justified in a later part of this report, where it is in fact seen that a bow shock is preferred for boom mitigation. It is seen that this approximation yields the effective density as seen by the engine inlet, and as such will be compatible with the estimate obtained for subsonic conditions. Page 8 of 61

9 The normal shock jump relation for density yields: ρ effective ρ altitude = γ + 1 M 2 γ 1 M Equation 4 For subsonic speed, drag at any point in flight is given by D = C D0 + C L 2 πare + C D w qs Equation 5 where e is Oswald Efficiency Factor typically between 0.7 and For supersonic flight, drag value is estimated using the equation D = C D0 + KC L 2 + C Dw qs Equation 6 where K = 1 πar 1 + (M2 1) πar 4 2 Equation 7 C Dw = E WD M πλ 0.77 LE 100 9π A max 2 ι 1 S Equation 8 C L = nw qs Equation 9 C D0 is approximated to be for a clean jet aircraft as given by Raymer [1]. The current weight of the aircraft is obtained by multiplying the takeoff gross weight W 0 by the fuel fraction β. The load factor n is set to 1 for this flight condition. Combining the equations and setting the specific excess power to zero, we obtain the following relationship between T SL W 0 and W 0 S : T SL W 0 = β α q β C D0 + C Dw W 0 S + 1 πare nβ q 2 W 0 S + 1 d V dt + 1 dv g dt Equation 10 The variable q is the dynamic pressure set by the cruise velocity and density: q = 1 2 ρv2. Table 1 shows flight conditions and aircraft parameters that are used for the 1-g steady level flight. Page 9 of 61

10 1g steady, level flight, M h=50000ft Altitude ft slug/ft M 1.8 V ft/s C D, AR 2.1 K E WD 1.9 LLE 60 deg d max 13 ft Length 180 ft Wing Area 3335 ft 2 C D,W q lb/ft 2 n 1 dh/dt 1.67 ft/s g ft/s 2 dv/dt 0 ft/s 2 Table 1: Flight Conditions and Aircraft Parameters for 1-g steady level flight Flight condition such as Mach number, altitude and aircraft parameters such as aspect ratio (AR), leading edge sweep (LLE), maximum diameter (D max ), length and wing area are determined from sizing exercises and the mission profile. The value of C D0 is increased slightly to to provide a conservative estimate. The value of is chosen based on the design mission of the Super Sonix aircraft, specifically the fuel fraction at the start of cruise. 2-g subsonic Maneuver The subsonic 2-g maneuver is of utmost importance especially in the design mission for Super Sonix. Inherent in the mission of hub to hub service lies the fact that the aircraft needs to operate in and out of high-traffic airports. A 2-g maneuver may be required to initiate hard turns and/or evasive maneuvers to clear traffic, or clear the airspace quickly about these airports. For the 2-g subsonic maneuver, the equation for supersonic flight is used, without any wave drag contribution: that is, C Dw =0. The load factor is set to 2, for a 2-g maneuver. The Table 2 below shows flight conditions and aircraft parameters that are used for the 2-g subsonic maneuver. Page 10 of 61

11 subsonic 2g maneuver, V h=10000ft Altitude ft slug/ft V 422 ft/s C D, e 0.8 AR 2.1 q lb/ft 2 n 2 dh/dt 0 ft/s g ft/s 2 dv/dt 0 ft/s 2 Table 2: Parameters for subsonic 2-g maneuver constraint An altitude of ft is chosen to represent the maximum altitude at which such a maneuver may be required. Flight speed is set at 250 kts, which relates to 422 ft/s True Air Speed (TAS). Aircraft parameters are, as before, determined based on sizing, and the fuel fraction is derived using the takeoff and climb portions of the design mission. An important aspect to be aware of is the maximum lift coefficient required to perform the maneuver during subsonic flight. As wing loading increases, higher lift is required to perform the 2-g maneuver. The wing loading can only increase until a point where the CL max of the lifting surface is reached, and any further increase in wing loading would lead to the stalling of the wing. A CL max, subsonic is assumed to be 1.2, a number that is chosen to be lower than the maximum lift coefficient at takeoff CL max,to while being within reasonable limits. This places an upper limit on the achievable wing loading, manifesting itself on the constraint diagram as a vertical line. Takeoff and Landing Ground roll Takeoff constraints are implemented through the following equation: T SL W 0 = β 2 W 0 αgρc Lmax S TO S Equation 11 Where S TO is the takeoff ground roll distance. The equation is obtained assuming that thrust is of a much larger magnitude than aerodynamic and ground friction drag. It also takes into account the performance with one engine inoperative during takeoff. Takeoff velocity V TO is assumed to be 10% higher than V stall. Page 11 of 61

12 Landing constraint is dependent on the availability of thrust reversers on the engines. For the present analysis, thrust reversers were used to aid in braking of the aircraft during its ground roll. T SL = β W 0 W 0 α rev ρc Lmax gs L S μ Equation 12 α rev is the ratio of the reversed thrust to the static sea level thrust, T SL. The friction coefficient, μ is typically set to be from 0.3 to 0.5 with the brakes being applied. S L, landing pavement length is mandated to be one third of the available runway pavement, the remaining two thirds for safety considerations. Takeoff and landing constraints are imposed at particular airports which Super Sonix serves. At present this analysis will include extreme conditions at hub airports, two of the most restrictive of which are Dubai (DXB) and New York (JFK). Tables 3(a) and 3(b) below show the parameters that are used for operations out of Dubai: landing ground roll h = 34ft,43 hot day[dxb] Altitude 34 ft ΔT 43 :R slug/ft rev CL max land 2 g ft/s 2 μ 0.3 S L 4374 ft Table 3(a): Landing constraint parameters for DXB airport takeoff ground roll h = 34ft,43 hot day[dxb] Altitude 34 ft ΔT 43 :R slug/ft CL max,to 1.5 g ft/s 2 S TO ft Table 3(b): Takeoff constraint parameters for DXB airport Table 3: Landing and Takeoff parameters for DXB airport Based on the typical data, rev was assumed to be 0.4. The fuel fraction was taken to be 1 for these maneuvers. In landing case, this implies that the aircraft can land immediately without burning any fuel. CL max,land and CL max,to were estimated to be 2 and 1.5 respectively [2]. The standard atmosphere temperature at sea level is 59:R; however Dubai presents a hot climate of 108:R due to its geographical location [9]. Since there is a significant variation of density at this temperature (and a corresponding degradation in engine performance), a ΔT from standard temperature of 43:R is incorporated into this performance constraint. Page 12 of 61

13 The parameters that are used for JFK are as follows: landing ground roll h = 13ft,25 hot day[jfk] Altitude 13 ft ΔT 25 :R sl/ft 3 1 rev CL max land 2 g ft/s 2 μ 0.3 S L 2800 ft Table 4(a): Landing constraint parameters for JFK takeoff ground roll h = 13ft,25 hot day[jfk] Altitude 13 ft ΔT 25 :R slug/ft CL max,to 1.5 g ft/s 2 S TO 8000 ft Table 4(b): Takeoff constraint parameters for JFK Table 4: Takeoff and Landing parameters for JFK The constraint in this case is the short takeoff and landing ground roll. This represents the shortest runway at JFK airport [8,10], and meeting this constraint will allow us to utilize all of the runways not only at JFK but all other hub airports as well. Second Segment Climb Gradient Second segment climb gradient is a Federal Aviation Authority (FAA) requirement specified in the FAR [11]. It dictates the rate of climb for an aircraft from takeoff at 35 feet to clear 400 foot high obstacles along the climb out flight path as a function of the number of engines it possesses. The second segment climb gradient is defined by: CGR = T W 1 L D Equation 13 Where T is the thrust at takeoff, W is the aircraft weight, and L/D is the lift-to-drag coefficient. T SL W 0 = β α N 1 N CGR + 1 L D Equation 14 Page 13 of 61

14 Substituting L/D=C L /C D and their respective expressions, we arrive at the following relationship of thrust-to-weight ratio and climb gradient. T SL W 0 = β α N 1 N CGR + C D 0 + ΔC D0 C LTO + C L TO πare TO Equation 15 Equation 15 assumes one engine is inoperative during the climb (N being total number of engines). It is noted that ΔC D0 represents the additional parasite drag due to deployment of flaps during takeoff. Typical value for this variable is for plain flaps and for slotted flaps. The parameters used in the analysis are as follows: second segment climb gradient of 3% above 34ft,43 hot day Altitude 34 ft ΔT 43 :R slug/ft CL max TO 1.5 g ft/s 2 N 4 CGR 3 % C D, AR 2.1 e TO 0.6 Δ C D, Table 5: Second segment climb constraint parameters The DXB airport is a constraint here due to its hot climate. As the Super Sonix aircraft has 4 engines, the CGR is mandated to be 3% [11]. The determination of the number of engines will be discussed in a later part of this report that deals with engine selection analysis. The Oswald efficiency factor is reduced from the earlier value of 0.8 to 0.6 to account for the deployment of flaps. The following figure shows the constraint diagram that is obtained from the above-mentioned analyses. Page 14 of 61

15 Thrust -to-weight Super Sonix Inc. 12 March Wing Loading (lb/ft 2 ) 1g steady, level flight, M h=50000ft subsonic 2g manuever, V h=10000ft takeoff ground roll h = 34ft,43 hot day[dxb] landing ground roll h = 34ft,43 hot day[dxb] second segment climb gradient of 3% above 34ft,43 hot day takeoff ground roll h = 13ft,25 hot day[jfk] landing ground roll h = 13ft,25 hot day[jfk] Figure 1: Constraint Diagram The performance of the Super Sonix aircraft is primarily constrained by the supersonic 1-g cruise, second segment climb and subsonic 2-g maneuver. The current design point lies at or above T SL /W 0 =0.45 and W 0 /S=70 lb/ft 2. Compared with historically averaged data of T SL /W 0 =0.38 and W 0 /S=85 lb/ft 2 (from SRR initial sizing analysis), it is observed that these values are within acceptable limits. Page 15 of 61

16 Table 6: Assumptions Sensitivity Table The table above organizes certain key assumptions that are used in the constraint analysis in order of the sensitivity of the solution on the values. Notably, small variations in C D,0 and α cause significant changes in the constraint plots. The nominal cruise climb capability, on the other hand, has little to no appreciable effect on the behavior of the graphs. Further analysis is required, for example the generation of a carpet plot, using the predicted values of T SL /W 0 and W 0 /S as starting points to determine actual aircraft performance. Concept Selection The concept generation and selection process is a very important part in the design process. The objective of this process is to develop a concept that meets the customer requirements and well as the needs of the design mission. In order to create the best concept for the specific mission, team Super Sonix went through a comprehensive concept selection process outlined by the flowchart in the Appendix. Each of the steps is discussed in detail. Brainstorming To come up with a comprehensive database of design concepts, each member in team Super Sonix generated ideas and sketches of what they considered to be a good concept for the mission. The concepts generated from the brainstorming session are included in the appendix. Various design aspects such as engine type/placement, wing and fuselage shapes, number/size/ location of control surfaces and the cabin layout were described. Page 16 of 61

17 Pugh s matrix In order to make a qualitative comparison between various concepts, Pugh s method was used. This method is an iterative process where the different concepts were compared with a predefined datum or baseline concept, for a set of design criteria. For the selection of the Super Sonix concept, two runs of the Pugh s matrix had to be made. In order to generate a good concept, a good set of design criteria needed to be chosen. The list of design criteria used by team Super Sonix was generated using the Quality Function Deployment and the mission of the aircraft. A list of ten design criteria was used for the 1 st run of the Pugh s matrix. Using the various ideas generated from the brainstorming session, 8 different concepts were generated and these were compared to the Concorde. These concepts are shown in Figure 2. Figure 2: Concepts used for the Pugh s method analysis For the 1 st run of the Pugh matrix analysis, the Concorde was chosen as the datum because it is the best commercial supersonic transport aircraft that ever flew. Each of the concepts was compared to the datum for each design criteria and was evaluated on a three level scale positive(+), negative(-) and same(s). After the entire matrix had been traversed, the total number of positives, negatives and same s were compiled. Based on the evaluation, three baseline concepts were selected to undergo the next iteration of the Pugh s method. These Page 17 of 61

18 were concept 1, concept 5 and concept 6. Table 7 shows the first run and the compiled results of the Pugh s matrix. Hybrid Concepts Table 7: First run of Pugh s matrix From the 1 st run of Pugh s matrix, the positive attributes of the various concepts were incorporated into the three baseline model to create hybrids. The negative attributes in these concepts were reduced or eliminated. A high performance aft arrow wing concept retrieved from [12] was included in this list because of its predicted high performance supersonic characteristics. Different technologies were incorporated into each of the different concepts so that these technologies could be compared. A very qualitative assessment of various technologies was made at this point. Viability and actual performance of the concepts need to be tested. Page 18 of 61

19 Hybrid concept 1: Figure 3: Hybrid concept 1 Hybrid concept 1 looked like a scaled down version of the Concorde. The droop nose was missing and was replaced by a hybrid nose design, which reduced the number of moving parts, and hence the weight of the aircraft. It also decreased complexity and increased safety of the aircraft. The sonic boom mitigation was taken care by the nose design which created a series of weak compression shocks similar to the engine inlet of the SR-71. Canards and horizontal stabilizers are missing on the aircraft. However, like on the Concorde, inner, middle and outer elevons are used to control roll, pitch and yaw motion for the aircraft. The concept has a delta wing with the wing root at mid fuselage. It also has an anhedral angle and droops along the chord, from leading edge to the trailing edge. The engine location is aft of the delta wing and is placed under the wing. Page 19 of 61

20 Hybrid concept 2 Figure 4: Hybrid concept 2 Hybrid concept 2 has an over wing engine which should help reduce the effect of engine noise. This is located on the aft end of the delta wing. The wing has no anhedral or dihedral angle on this concept. Nose shaping, similar to the one on the F-5 shaped sonic boom demonstrator [13], would be used to reduce the sonic boom overpressure. Canards with a dihedral angle would be present on the upper section of the fuselage. Page 20 of 61

21 Hybrid concept 3 Figure 5: Hybrid concept 3 The 3 rd hybrid concept made use of the Gulfstream / NASA Quiet Spike technology for sonic boom mitigation. A major difference with this concept is that it has under wing inlets for the engine and over wing outlets, which is similar to the YF-23. Canards with an anhedral angle are its primary control surfaces. This concept has a delta wing mounted on the bottom section of the fuselage. The main wing has a dihedral angle. Page 21 of 61

22 Aft arrow wing concept Figure 6: Aft Arrow wing concept [12] The aft arrow wing concept [12] was included in the list of hybrid concepts for the next run of the Pugh s matrix because of its expected performance. The aircraft geometry helps to reduce the sonic signature of the aircraft. Having swept canards also helps bring the sonic signature from an N-wave to a Plateau wave form. Another major difference in the aircraft geometry is the inclusion of two vertical tails on either wing. The concept has a highly swept delta wing, with a dihedral angle, on the bottom of the fuselage. Although it helped to increase options for the final concept, choosing between various technologies on a qualitative basis was difficult. For this reason, further research was done into each of the design criteria used for comparison of aircraft concepts. This included making a detailed description of each design criteria and the systems it related to. A table of the design criteria along with its detailed description is in the appendix. Page 22 of 61

23 D A T U M Super Sonix Inc. 12 March 2008 The 2 nd run of the Pugh s matrix used the aft arrow wing concept as the datum. The three hybrid concepts were compared to each of the design criteria (using the detailed description table). Like the previous run, the positives, negatives and the same s were compiled. 2nd run AFT ARROW WING CONCEPT HYBRID CONCEPT 1 HYBRID CONCEPT 2 HYBRID CONCEPT 3 SONIC BOOM SUBSONIC NOISE s + s CONTROL SURFACES - s s TURN AROUND TIME s + + AIRPORT COMPATIBLE SAFETY s + - EASE OF MANUFACTURE - s - EMPTY WEIGHT + + s COST Table 8: Second run of Pugh s matrix s From the 2 nd run of the Pugh s matrix, it was possible to see how the hybrid concept 2 had better design characteristics for the mission when compared to the other three concepts. For the final design, the positive attributes of this concept were enhanced and the negatives were reduced. Using this information and the positive aspects of the other concepts, the final hybrid conceptual design was created. Page 23 of 61

24 Figure 7: 3-view and isometric drawing of final design Figure 7 shows the 3-view and isometric drawing of the final design concept. The final concept selected from the Pugh s matrix analysis had a blunt nose design. This creates a bow shock in front of the aircraft, which will eliminate shock effects on the wings of the aircraft and also reduces the effect of coalescent shocks along the bottom of the aircraft. The aircraft concept had a double delta wing with aft mounted engines. The concept also had swept canards like the aft arrow wing concept. Certain aircraft characteristics, such as the engine placement (over/under wing) and vertical stabilizers- one on fuselage or two on wings had not been decided. Further research needed to be done before a definitive choice was made. The effects of dihedral or anhedral angles on the wings and canards needed to be examined as well. Door locations on the aircraft are yet to be decided. Page 24 of 61

25 Advanced Technologies The Super Sonix aircraft will need to balance three major design qualities to be successful. These qualities are: low sonic Boom, efficient supersonic cruise, and advance engine design to reduce noise and pollution. There are many ways that each design quality can be met. It is important to note that not all design qualities can be met to 100%, this is primarily due to mission requirements and a coupled relationship between the different design solutions. What follows is a description of the different ways that each design quality will be met. Boom shaping technologies A modern supersonic aircraft will need to not only be able to fly fast, but also be able to fly fast quietly. Quiet supersonic flight is a challenge because of the many different technologies that must be used in tandem to reduce an aircraft s sonic boom signature. There are two different types of sonic boom signatures, an N-wave and a plateau wave. An N-wave sonic boom signature is created when aircraft develops lift rapidly. This can be in the form of high aspect ratio wings with low sweep. A plateau wave is created when lift is developed gradually and or a blunted nose configuration is used. The plateau wave is more desirable because the sonic boom energy is spread out over a longer period of time, and thus there is a lower overpressure and a smaller sound signature. Thus, it can be concluded that the objective of the Super Sonix aircraft is to produce plateau wave sonic boom signatures with significantly lower overpressures. The methods in which this objective will be achieved are presented below for each design configuration. Nose Design Blunt Nose The implementation of a blunt nose on the Super Sonix aircraft will result in a bow shock in front of the aircraft. Normally this would not be desired; however the objective of producing this bow shockwave is to reduce coalescence shockwaves produced under the aircraft. These coalescence shockwaves are stronger and produce the N-wave pressure signatures seen on earlier supersonic aircraft. Using a blunt nose configuration results in a plateau wave signature, thus lowering the sonic boom effects. The effects of this technology on the overpressure produced by the aircraft are presented below in Figure 8. Page 25 of 61

26 Figure 8: Blunt nose effect on boom overpressure [23] It is important to note that implementing a blunt nose configuration on the Super Sonix aircraft will result in a large increase in wave drag. The wave drag penalty is one of the main reasons why this has not been implemented in previous designs. Thus, another big challenge of the Super Sonix aircraft is to reduce wave drag in other parts of the aircraft by careful fuselage and tail design. These concepts are presented in more detail later. Nose Keel Another nose design that may reduce sonic boom sound signature includes installing a keel at the front of the aircraft. The keel has a channel down the centerline of the device that allows air to flow in-between two panels. Then high voltage electricity is arched across the air to ionize the incoming air. This process produces a plasma out in front of the aircraft that when implemented will increase the effective length of the aircraft [20]. While this nose design results in a significant overpressure drop across the aircraft, there are many environmental issues associated with this design. Creating plasma at high altitudes results in the production ozone which is a greenhouse gas. Given the altitude that this gas is produced at, it will have much more severe environmental consequences than if the same device was implemented on sea level. Furthermore, there are issues associated with how one might Page 26 of 61

27 Dihedral Angle Using dihedral angle on any of the lifting surfaces on the aircraft can have a profound impact on the sonic boom signature [21]. Tilting the lifting surface upward a set angle results in lift produce that is on multiple planes. This has the effect of increasing the effective length of the aircraft, thus smoothing out the effective area distribution. An example of dihedral angle is shown in Figure 9. Figure 9: Dihedral angle design In Figure 9, the angle gamma is the amount that the wings are tilted upward. Full implementation of this ideal over an entire lifting surface may not be feasible; however, sections of the main wing can have dihedral angle. This will help with increasing the effective length of the aircraft while slowly building a lift producing surface. While using a dihedral angle on the wings may have several advantages for supersonic aircraft, there are several issues associated with using it in the subsonic regime. During subsonic flight, wings with dihedral angle will experience wash out due to lateral flow instabilities in the airflow. Furthermore, increasing dihedral angle reduces planform area of the wing and introduces structural issues in wingbox design. Effective Area distribution Aerodynamic Cowling Carefully shaping the area distribution is a primary design consideration for the Super Sonix aircraft. Research has found that for sonic boom minimization it is important to have a rapid buildup of volume at the nose of the aircraft [17]. This idea was introduced earlier in the blunt nose design section. A nose design that involves a rapid buildup of volume at the nose of the aircraft will produce significant wave drag, and thus the rest of the aircraft must compensate by Page 27 of 61

28 having a low wave drag design. These ideas are implemented by using a varying diameter fuselage, which if designed improperly can lead to severe weight penalties. The solution to this problem is to use a classic cylindrical fuselage design, while varying the effective diameter of the body by using cowling held in place by light-weight trusses. The spaces between the fuselage cylinder and the outer cowling truss will be filled with avionics equipment, fuel, insulation, or some form of payload. When the fuselage travels over the wing structure, the fuselage cowling will shrink to fit the fuselage cylinder, thus smoothing out the area distribution of the aircraft. An example of this is presented below in Figure 10. Figure 10: Non-Axisymmetrical fuselage shape [22] From Figure 10 it can be seen that the supersonic aircraft concept has a carefully shaped fuselage to reduce the drag and sonic boom signature. Engine Nacelle Placement The placement of the engine nacelles can contribute significantly to the overall area of the aircraft. To help smooth the area distribution with the engine nacelles, one can add carefully shaped cowlings to the engines. If the engines are placed underneath the fuselage and a part of the tail of the engine is stuck out behind the wing, the cowling on the engine could be larger close to the end of the wing, to help smooth the effective area of the aircraft. Canards Another component of the Super Sonix aircraft is the canards located at the nose of the aircraft. These lifting devices help add longitudinal stability to the aircraft and are the primary elevator control surfaces of the aircraft. For these devices to be efficient on the Super Sonix aircraft, they will need to have large dihedral angle, such that the lifting area contribution is gradual and the sonic boom signature is carefully crafted. Page 28 of 61

29 Efficient supersonic cruise The majority of a supersonic aircraft s design mission is spent in the supersonic flight configuration. This means that the Super Sonix aircraft needs to be efficient in the supersonic flight regime. The key design features that enable an aircraft to perform efficiently in the supersonic flight regime are: high aspect ratio low sweep wings resulting in natural laminar supersonic flow. It is unlikely that a natural laminar supersonic flow wing can be created, however getting the flow over the wing to be close to laminar will result in significant drag reduction. This wing design is in direct contradiction to what is required for low sonic boom signature. For low sonic boom sound signature the Super Sonix aircraft needs to develop lift more gradually; thus highly swept wings and dihedral angle are major design requirements. The wing design for efficient supersonic flight and low sonic boom signature are fundamentally different, and thus a balance needs to be reached where the Super Sonix aircraft can fly quietly at supersonic speeds. Active Flow Management Because the Super Sonix aircraft will have a hybrid wing design that balances supersonic flight efficiency and quiet supersonic flight operations, some design solutions may not be efficient in during the entire operating mission of the aircraft. For that reason, active flow control will be used to help keep the flow over the wings attached to the body, and thus improve lift. An example of where this technology might be needed is during takeoff and landing. Engine Technology Correctly pairing an engine to an aircraft is one of the most important steps in the conceptualization process. For this purpose, a database of over 200 engines was created; outlining details about each engine and the aircraft that it was used in. The engines were then filtered into two categories, subsonic and supersonic engines. While some engines that operate in the subsonic regime can fly supersonic, the majority of the engines put into the subsonic category were high-bypass turbo fans. This is primarily due to the size of the engine and design of the engine inlet. Therefore, all of the engines considered for the Super Sonix aircraft were medium bypass turbofans, low bypass turbofans, and turbojets. The engines from this set were then filtered for the proper thrust profile that fit the preliminary sizing studies. Figure 11 shows the flow chart detailing the engine selection process. The engine results from this process are listed in Table 9. Page 29 of 61

30 Figure 11: Engine Database Manufacturer Designation Dry Thrust Weight, dry (lb) Aviadvigatel (Solovyev) D-30F ,326 GE F ,460 GE/RR F ,300 P&W F ,750 P&W PW1000G ,885 Samara NK ,283 Samara NK ,588 Table 9: Viable supersonic engines [5] Out of the engines listed, the Samara NK-321 is thought to be the best engine that meets many of the requirements. The reason is due to the fact that the Samara NK-321 has supersonic capabilities, they have also been used in commercial supersonic transport with the TU-144. From the sizing code, it was determined that the aircraft will require 107,724 lbf of thrust. With the thrust rated at 30,843 lbf, the Super Sonix aircraft will have a 4-engine configuration. Apart from NK-321, another engine that is under consideration for Super Sonix aircraft is QSST engine. The engine is capable of producing 33,000 lbf of thurst and will be available by With the latest, state of the art technology and significant reduction of noise, this engine is a good option and fits perfectly into Super Sonix engine requirements. As of now, the Page 30 of 61

31 T/W0 Super Sonix Inc. 12 March 2008 development of the engine is contracted to General Electric, Pratt & Whitney, and Rolls-Royce and detail specifications of the engine are currently not available. There are a few aspects of note in regards to the Samara NK-321. Firstly, the Samara is a Russian engine and there have been problems in the past exporting TU-144 for the SST studies because the engines are export controlled. Secondly, the Samara engine is also very noisy. Furthermore Samara engine is quite old, entering service in Figure 12 below illustrates the Thrust to Weight ratio of the 7 engines against the manufactured year. It is noted that there are several engines that have yet to be produced (P&W 1000G, P&W F135 and the GE F136). Therefore, even though some of those engines are expected to produce compatible thrust, only the Samara NK-321 will be used as a baseline engine with additional technological improvement as of today. T/Wen vs. Manufactured Year 8 7 Samara NK-25 6 GE F Aviadvigatel D-30F6 Samara NK-321 P&W F GE F P&W 1000G T/We Year Figure 12: Thrust-to-weight ratios for viable engines Modifications for Samara NK-321 Engine The first technological concept in regards to reducing noise is employing the variable engine cycle [3, 4], an engine that is designed to operate efficiently under mixed flight conditions. One concept is the tandem fan engine where the engine has two fans, both mounted on the low Page 31 of 61

32 pressure shaft with an axial gap between the units. In normal flight, the engine is in 'series' mode, meaning that the flow leaving the front fan passes directly to the second fan. This works like a conventional turbofan. But during takeoff, climb out, final descent and approach, the front of the fan is allowed to discharge automatically through a nozzle on the underside of the nacelle. This will allow air to enter the rear fan and will flow to the rest of the engine. This is called the 'parallel' mode. In parallel mode, it increases airflow of the engine at thrust which leads to a lower jet velocity and quieter engine. One downside of the variable engine cycle is the complexity of it, which can lead to difficulty in manufacturing and also servicing. As has been discussed above, the Samara NK-321 engine is very noisy. Papamoschou and Debiasi [14] did a study in order to reduce engine noise and in order for the NK-321 to be a viable base engine, this study has to be taken into account. It was found that employing an eccentric exhaust configuration will be beneficial to reducing engine noise. They studied the effects of engine configurations have in regards to noise. Figure 13 shows the 4 configurations that they propose would help with engine noise. The first one, B03-MIX was the baseline engine, a military turbofan engine with BPR of 0.3 at mach 1.6 and height of 16,000 m. The notation MIX stands for mixed flow engine. The second configuration, B16-MIX, is a mixed flow turbofan. It was suggested an increased bypass ratio from 0.3 in the baseline engine to 1.6. This configuration will yield less noise but will not have an efficient supersonic performance. The third configuration, B16-SEP-COAX, is an engine with bypass ratio of 1.6. It has separate flow and has a coaxial configuration. A separate flow is normally better than mixed flow because the bypass exhaust design will reduce the mach wave emission from the core and will therefore produce less noise. This configuration is normally used in current subsonic aircrafts. The last configuration, also said to be the best solution to the noise versus performance tradeoff problem, according to Papamoschou and Debiasi, is the eccentric engine configuration. This configuration works by maximizing the theory of mach wave elimination discussed above. The eccentric configuration minimizes the mach wave and therefore produces less noise. The comparison between the four was based on the assumption that all engines produce the same supersonic cruise thrust, and have the same core characteristics. A solution that Papamopschou and Debiasi proposed to reduce noise was to have an intermediate bypass ratio in order to satisfy the need for a high bypass ratio on landing/takeoff, which would lead to quieter noise and also a low bypass ratio during supersonic cruise for efficient cruise. Page 32 of 61

33 Figure 13: Proposed engine exhaust configurations [14] The main source of the engine noise is Mach wave radiation. Mach waves are shock waves that are generated by supersonic motion of turbulent eddies relative to the air. It can be assumed that the velocity of the eddies will be close to 80% of the jet velocity engine. The problem with such a high eddy is that it will create a system of shocks that will produce noise that can be heard form a very far distance. A solution to this problem is known as the mach wave elimination method. This works by preventing the formation of shocks by making sure the eddies are subsonic. To make the eddies subsonic, a secondary flow was created to surround the jet so it will reduce the velocities of the eddies. Page 33 of 61

34 Figure 14: Directivity of Overall sound pressure level [14] Figure 14 shows the difference of the Overall Sound Pressure Level for the baseline engine with the three derivatives. It is seen that the sound levels are decreased appreciably with the new derivatives. Figure 15 below is a comparison of the Perceived Noise Level (PNL) time histories between the different configurations on a given flight path. All engines produced the same thrust of 126 KN. As can be seen from the plot, the eccentric exhaust configuration will be 13dB quieter than the baseline engine, while the mixed and coaxial configurations are 5 and 6 db quieter than that of the baseline. Page 34 of 61

35 Figure 15: Perceived Noise Level [14] There are two ways to go about selecting the engine. The first will be to use the Samara NK-321 as a base engine and improve that engine with the technology discussed above. The result will be the Samara engine with less noise and that will be good given that the Samara has been tried and tested in a supersonic commercial aircraft. Another route would be to consider an engine still being developed today that will be available in The QSST engine discussed earlier would be a good option assuming it will produce a similar Thrust to what is required. Details for the QSST engine are unavailable at the moment so it is not possible to make a selection basing on so little information today. Page 35 of 61

36 Seat Pitch (in) Super Sonix Inc. 12 March 2008 Cabin Layout Several factors contribute to the cabin layout of the Super Sonix aircraft concept. Passengers desire a comfortable, spacious seating environment, but it is important to meet aircraft design constraints and FAA requirements as well. Passengers will be paying a hefty sum of money to fly supersonic on this aircraft, but it is debatable to what extent a passenger will hold seating layout a priority in their travel with such short flight times. Business analysis thus far has shown that the target customer base values time a bit more than comfort, but at the same time people who can afford to pay for time savings like this generally may expect more from their seating. There is a delicate balance between passenger comfort, safety, efficiency, and aircraft size and shape. 120 Seat Type - First Class and Business Class Flat Bed Lie Flat Recline Suite Figure 16: Passenger Seating Options (pitch) Page 36 of 61

37 Seat Width (in) Super Sonix Inc. 12 March Flat Bed Lie Flat Recline Suite Seat Type Figure 17: Passenger Seating Options (width) After analyzing historical data (Figures 15 and 16 above), it is apparent that typical subsonic transports run by the airlines have a maximum seat pitch in first class of about 40 inches, though it is generally much less. Larger specialty aircraft with long flight times may have seat pitches between 55 to even 100 inches or more with fully reclining sleeper seats or suites. Generally, supersonic flights will be no longer than 4 hours, so traditional reclining seats appear to be the most reasonable option. For coach, the maximum seat pitch is around 36 inches but the same idea still applies. These rough numbers give a general idea of how long the seating section would be. Given the overall aircraft length from sizing, the length seems practical. 180 ft 90 ft Figure 18: Passenger Seating vs. Overall a/c Length Page 37 of 61

38 As shown in Figure 17 above, the seats would take up 90 feet of the 180 feet of the aircraft length. Because of aerodynamics requirements the aircraft length was extended to 180 feet, giving a cabin that is just slightly more than half the length of our aircraft. With all this extra length in the fuselage, the positions of major aircraft components such as fuel tanks and landing gear can be adjusted to meet the c.g. requirements outlined in this report. Figure 19: Seating Layout with Galleys and Restrooms The current seating configuration gives first class passengers 7 rows of 2 seats, and gives coach passengers 15 rows of 3 seats for an overall total of 59 passengers. Two flight attendants will have folding jump seats available in the cabin between the two classes. Other parameters, such as aisle and seat width, are also obtained using the method described above. The first class aisle is 28 inches wide, while the coach aisle is 20 inches wide. First class and coach seats are also 28 and 20 inches wide respectively and also represent the upper limits for seating in a typical subsonic transport such as a Boeing 737 or Airbus 320. The seat width also takes into account the extra fuselage width available with only 2 seats per row in first class. Overhead compartments, though not yet modeled in CATIA, will fit easily into the 9 foot diameter inner-cabin cylinder. Placing the floor 2.5 feet from the bottom of the cylinder allows for a 6.5 foot aisle height. The inner cabin (inner fuselage) is designed as a basic cylinder for easy cabin pressurization. Though this inner section is cylindrical, the idea is that outer fuselage geometry can be built around this cylinder to allow for area rule compliance as well as aerodynamic shaping. Having this inner cylinder should allow more room for adjustability, and it is desirable from a structural standpoint. Also worth mention, the CATIA model of the cabin is easily adjustable for changes in any one of these parameters. A change in one number in the model will adjust the entire seating Page 38 of 61

39 arrangement for the aircraft. The seating is modeled separately so that it can be adjusted and imported into the main product file. Figure 20: Seating Dimensions, Coach (L) First Class (R) Page 39 of 61

40 Sizing Studies Sizing Approach The current sizing code being utilized is similar to the previous code, but now has more components integrated. For example, in addition to the iterative weight calculator, the sizing code now includes a center of gravity calculator and a boom overpressure calculator. An important consequence of this integration is the ability to quickly alter all of the estimations. For example, a change in the takeoff weight of the aircraft will result in a change in both the boom overpressure and the center of gravity. This is represented as a flowchart in Figure 20. This sizing code results in a slight increase in the estimated takeoff gross weight. It is now expected to be around 284,000 lb. The results from the current sizing are summarized in Table 10. Figure 21: Sizing Approach flowchart Page 40 of 61

41 Component Weights Table 10: Current Sizing Attributes Aircraft weight statement should list all possible weight within an aircraft. The sum of weight statement should be equal to the gross takeoff weight. The importance of estimating aircraft component weight cannot be understated as it is critical to the center of gravity calculation, structural design, and aerodynamic design. Well estimated aircraft component weights will help better placement of the wing, horizontal tail, canards, fuel, and engine. The baseline aircraft used as a comparison was the Concord. However because information could not be found on the Concord, a database of 16 small and large commercial aircraft component weights was created. List of aircrafts in the database includes small commercial aircraft of: Citation 500, MDAT-30, MDAT-50, F-28, MDAT-70, DC-9-10, , and For large aircraft database it includes: , , Dc-8-62, DC-10-10, L-1011, DC-10-40, , and SCAT-15. Total of 16 aircrafts, including mass fraction given: wing system, tail system, body system, alighting gear system, nacelle system, propulsion system (less dry Engine), flight control system, auxiliary power system, instrument system, hydraulic and pneumatic system, electrical system, avionics systems, furnishing and equipment system, air-conditioning system, anti-icing system, and load and handling system. Page 41 of 61

42 Mass Fraction Average Component Mass Fraction Component Mass Fraction Componenet Mass Fraction Super Sonix Inc. 12 March 2008 Small Commerical Large Commerical Combined (Small and Large) Commerical Average Mass Fraction for Aircraft Component Parts Small AVERAGE Large AVERAGE Figure 22: Component weight mass fraction Page 42 of 61

43 This aircraft component weight database was set to provide foresight on different component mass fraction of small versus large commercial aircraft, while providing a range that the component weight fraction should be within. Major concern was the compatibility of the aircraft within the database to the Super Sonix. None of the aircraft in the data base have the ability to fly at Mach The idea of the database is to create a general range with average and standard deviation of the component weight fraction with different aircraft. Figure 21 above show the individual component weight mass fraction in the form of average, standard deviation above and standard deviation below. This provides insight on the different range of component mass fraction for the small and large commercial aircraft. The combined average (third figure from top) for the mass fraction of the two aircraft size is also presented in the form of average, and plus, minus one STD. Specific values of combined averages can be found in the appendix. The average of the two is shown again on the very bottom of Figure 21. Using the database of 16 aircrafts, the combined average was chosen to be used as the component mass fraction. This provides an estimate for 87% of the aircraft weight. The table above shown a 95.8% estimate on component weight however, some components of the mass fraction will not be considered (ex. Horizontal tail) because the design does not include this component. The table above provides the estimate of each component weight of the Super Sonix; this is used to calculate the center of gravity and later used to provide insight on aircraft stability dynamic and control. Center of Gravity In order to determine the center of gravity, an excel spreadsheet was created and is shown in Figure 22. This spreadsheet utilizes the derived component weights of our aircraft. The weights can either be divided up and entered individually (passengers and seats), or the total weight of a part can be placed at the center of gravity for that specific part (engines). The current center of gravity is 97 feet from the tip of the nose. This is roughly 54% of the total length of the aircraft (180 ft). It should be noted that this estimation does not include all of the component weights. However, the spreadsheet is currently calculating the center of gravity using 87% of the gross takeoff weight. All of the major component weights such as wings, fuselage, fuel, canard, tail, engines and passengers are included in the 87%. Components such as baggage, air conditioning, and the anti-icing system are not included as of yet due to their placement not being accurate enough yet. The center of gravity is expected to change slightly when these components are added. However, it is not expected to change significantly because the location of most of the missing components is expected to be close to the center of gravity. Page 43 of 61

44 Shock Prediction Figure 23: Center of Gravity Calculator The prediction of the sonic boom overpressure is of a great importance. It is vital for a sound argument as to the feasibility of allowing supersonic flights over land. The resolution of the boom overpressure prediction problem involves a fundamental understanding of sonic boom pressure signatures as well as the comprehension of boom prediction methods and implementations. Sonic boom signatures are discussed in an earlier section of the report: Predominant far-field signatures (i.e. the signatures seen on the ground) include N-wave and plateau shaped characteristics. As noted earlier, the plateau and N-wave signatures have very different magnitudes and are determined by specific geometry. In order to provide a full computation of the boom overpressure, two methods are considered to calculate the overpressure for the N-wave and plateau signatures respectively. N-wave N-wave boom overpressure prediction is based on the study performed by H.W. Carlson [16]. Much of the detail required for successful implementation of the simplified model is not available at this preliminary stage of design of the Super Sonix aircraft, therefore the super simplified model is used to calculate bow shock overpressure. In this model, charts and approximations are used in place of some of the more detailed equations to determine the shape factor K S and the pressure amplification factor K p. Page 44 of 61

45 The equation for the overpressure is given as: The variables are defined as follows: Equation 16 Variable M W l p v p g h h v K p K s Definition Cruise Mach number Aircraft weight (MTOW is used for this prediction) Aircraft Characteristic length (Fuselage length) Atmospheric pressure at altitude Ground atmospheric pressure (assumed to be Sea level pressure) Altitude of aircraft above ground (assumed to be Cruise altitude) Altitude of aircraft above sea level (assumed to be Cruise altitude) Pressure amplification factor Shape factor (assumed to be similar to Concorde) Table 11: Definition of variables for overpressure calculation To use the charts a lift parameter K L is calculated based on the following equation: Equation 17 Based on this, a shape factor is obtained from Figure 23. It is noted that the curve corresponding to the heavy bomber shape (highlighted on the figure) is assumed to be the closest planform shape to Super Sonix. It is used in conjunction with the cruise altitude to determine the pressure amplification factor. Figure 24: Shape factor and pressure amplification factor charts The super simplified analysis only depends on certain macro properties of the aircraft, which are determined from initial sizing studies conducted earlier. The aircraft weight is at the top of cruise, and is therefore the MTOW (W 0 ) multiplied by the fuel fractions of segments 1 and 2 in Page 45 of 61

46 the design mission profile. The following table displays the input parameter values as well as the final result: N-wave signature M 1.8 W lb h ft l 180 ft P g lb/ft 2 P v lb/ft 2 b 1.50 K l K s K r 1.9 K p 1.13 ΔP max 1.89 lb/ft 2 Table 12: Input Parameters for N-wave overpressure calculation It is seen that the overpressure is far above the target limit of 0.3 lb/ft 2, a fact that should not seem surprising as N-wave signatures do indeed result in high overpressures. To validate the model, the parameters for the Concorde are used and an overpressure of 2.02 lb/ft 2, and matches very closely with the actual overpressure of 2 lb/ft 2 [17]. Plateau wave Calculations for the plateau wave are based on work done by R. Seebass and A.R. George [15]. The formal derivation of the equations involves complex Whitham F-functions and only the resulting equations are presented as follows. The main equation for prediction shock overpressure p so is shown below: Equation 18 Page 46 of 61

47 It relies on several auxiliary equations that are computed as follows: Equation 19: Auxiliary Equations for Plateau wave overpressure calculation In these equations, the parameter H represents a characteristic altitude to act as a basis for computations. This height is taken to be H=25000ft, a value that has historically been found to provide accurate results [15]. characterizes the rate of advance of the signal as it propagates through a real atmosphere. This signal advance is highly nonlinear and therefore must be referenced to the advance in a homogeneous atmosphere, where it is well understood. symbolizes this ratio of real advance to homogeneous advance. The input parameters and computed overpressure are tabulated below: Plateau Boom Signature M 1.8 W lb h ft l 180 ft Pg 2116 lb/ft 2 beta 1.50 alfa 0.74 gamma 1.2 k 2.15 W~ 1.16 Pso 0.70 lb/ft 2 Table 13: Input Parameters for Plateau wave overpressure calculation For the same aircraft parameters and flight conditions, it is seen that the plateau wave overpressure is predicted to be almost 2.5 times lower, a fact that follows directly from the discussion earlier. Even though the overpressure of 0.70 lb/ft 2 is still higher than the requirement, it is re-assuring that the move toward a blunt nose profile leading to a plateau shock signature does indeed have the potential to reduce the boom overpressure considerably. It is important to mention that these prediction methods are still unrefined, and more accurate models and programs (such as NASA s PBOOM program) is to be used in future as the Super Sonix aircraft become more defined in its characteristics. Page 47 of 61

48 Summary Compliance Matrix The compliance matrix was prepared to check the design progress. Recent version of the matrix can be found in the appendix. Current concept generated has met several requirements including take off field length and still air range. However, current concept has not yet met the cabin volume per passenger, cruise efficiency and the sonic boom overpressure. The cabin is to have at least 50 pax/(ft 3 /pax); however our current design has This is due to having 59 people in such a small cabin. Rearranging the cabin layout slightly is considered to efficiently use all the volume from the fuselage. The cruise efficiency was roughly calculated by counting the number of passengers, total fuel weight and the total range. The threshold is 0.3 lb fuel/pax mi, but the current value came out to be 0.6, which is the twice of the threshold. We are studying to see if there is a way to improve efficiency. The threshold sonic boom overpressure is 0.3 lb/ft 2. The design needs to meet the threshold value however; the rough calculation from volume came out to be 0.70 lb/ft 2. This value is more than twice of the target value. Currently it seems that meeting the target value is very difficult. More trade off studies will be completed to reduce the overpressure produced by the Super Sonix aircraft. There are some requirements still unmet. Some requirements such as door height and stall speed have not been found. The level of details needed is higher than the current stage, more detail specifications studies will be conducted later in the project. Page 48 of 61

49 Walkaround and CAD Model Figure 25: Walkaround chart Figure 24 depicts a walkaround chart that is created for the Super Sonix aircraft. CATIA is used to design the airplane for visual aid and analysis. Previously discussed technologies and concepts were considered when designing this model. To start, the fuselage was designed to have different cross sectional area along its length. The fuselage part connected to the nose has diameter of 10 ft. The diameter gradually increases to 13 ft to about half way of the cabin. The fuselage then decreases to 9 ft on the back of the cabin section. Tail fuselage is about 60 ft long. The diameter then quickly decreases to 1 ft, forming concaved shape. Overall, this aircraft is designed to encompass most of its volume up front in order to reduce sonic boom over pressure. The wings consist of two parts, the strakes and main delta wings. Currently, the airfoil selection hasn t been done yet, therefore the wing and the strake is assumed to have symmetric airfoil. This will later be change when further analysis is conducted and the best airfoil for the aircraft has been selected. The wing has total span of 80 ft and mounted with dihedral angle for Page 49 of 61

50 efficiency during supersonic cruise. The location where the strake and wing meets is arbitrarily selected since a lot of information about strakes and wings are currently not available. The strakes are added to produce more lift during subsonic flight and to move the center of gravity forward. The strakes are also mounted on dihedral angle. Future consideration for this design is to have all the connections between fuselage, strakes and wings to be smooth. It is important to have the entire cross sectional area smoothen so drag can be reduced. The canards are implemented on the airplane as control surface, and to help move the center of gravity forward. Currently, the canards are mounted on top of the fuselage due to problem with accessing the door when the gate is used in the airport. The location of canards and strakes on the fuselage caused the area around door to be small, risking possible collision between the airport gate and the canard. Further research about the flow effect from the canard to the wing and the engine inlet will be conducted. Vertical tail is very similar to Concorde s. Vertical wing also has symmetric airfoil. The tail is drawn so that it is swept. Further study is needed find the effect of having vertical tail during supersonic flight. There are few other concepts, such as having two vertical tails on the both side of the tip of the wing. Figure 26: 3-View of Super Sonix Aircraft Page 50 of 61

51 So far, airplane has been modeled to reflect researches currently done. For future work, the doors are going to be added. The location of door is already decided in the cabin layout section, so it will be drawn later. The landing gears are in consideration also. The location to put landing gears based on center of gravity analysis needs to be addressed. The nose is also being considered to redesign to have more flat bottom. This means, instead of having a symmetric elliptical cone, it will have asymmetric elliptical cone. This will reduce drag during flight. Finally, the number of engines and the location engines are in consideration. Further studies needs to be conducted to find the effect of having the engine on the wing and under the wing. Next Steps Given the current status of this project, the next steps will involve optimization. The system definition review has allowed the Super Sonix aircraft to have its primary characteristics defined. However at this point it is unknown what relative sizes of each surface or best mass position will result in the correct configuration to satisfy the design mission. Thus the next steps of the Super Sonix project is to iterate though sizing variables to reach a final system that satisfies the design mission. The manner in which this will be completed is outlined in the next steps flow chart present in the appendix. Some key design qualities that will need to be modified include: aerodynamic shaping of the aircraft, detailed dynamics and control calculations, and structural analysis and design. In conclusion, the Super Sonix aircraft has the key qualities that will allow it to satisfy the opportunity description. The seating arrangement is sufficient for airlines to make money, the aerodynamics has qualities that will allow the aircraft to travel supersonic over land, and the overall design of the aircraft is simple enough to allow for reasonable maintenance. While the system has been defined, it has not been perfected. In preparation for the system characteristics review the Super Sonix aircraft will undergo many controlled changes, in an attempt to find the most efficient design. Page 51 of 61

52 References [1]. Raymer, D. P., Aircraft design: A Conceptual Approach, 4 th Edition, AIAA Education Series 2006 [2]. Fun Aerospace Questions, [3]. Sweetman, B., USAF Funds Leap-Ahead Jet Technology ckscript=blogscript&plckelementid=blogdest&plckblogpage=blogviewpost&plckpos tid=blog%3a27ec4a53-dcc8-42d0-bd3a-01329aef79a7post%3a73d6b7cf-a2e8-4efb f7494c0747e5, 8/15/2007 [4]. Quiet Supersonic Platform (QSP) [5]. Meier, N., Jet Engine Specifications Database, [6]. Tupolev Tu-160 BlackJack, [7]. Fighter Gallery, [8]. TravelSmart Ltd, New York Airport Climate and Weather, [9]. TravelSmart Ltd, Dubai Climate and Weather, [10]. Jeppesen, Kennedy Intl Airport Chart, 20 May 2005 [11]. FAA regulations FAR 14: &idno=14#14: [12]. Carlson, H.W et al, Application of Sonic-boom Minimization Concepts in Supersonic Transport Design, NASA TP D-7218, Langley Research Center, Hampton, VA, June 1973 [13]. Morgenstern, J.M. et al, F-5 Shaped Sonic Boom Demonstrator s Persistence of Boom Shaping Reduction through Turbulence, Lockheed Martin Aeronautics Company, Palmdale, CA 93599; AIAA ; 43rd AIAA Aerospace Sciences Meeting and Exhibit - Reno, NV Page 52 of 61

53 [14]. Papamoschou, D. Debiasi, M., Conceptual Development of Quiet Turbofan Engines for Supersonic Aircraft, Journal of Propulsion and Power Vol. 19, No. 2, March April 2003 [15]. Seebass, R. et al, Sonic Boom Minimization Revisited, AIAA , 1998 [16]. Carlson, H., W., Simplified Sonic Boom Prediction, NASA TP 1122, Langley Research Center, Hampton, VA, 1978 [17]. Aronstein, D.,C. et al, Two Supersonic Business Aircraft Conceptual Designs, with and Without Sonic Boom Constraint, Journal of Aircraft, Vol. 42 No. 3, May-June 2005 [18]. Jane s All the World s Aircraft Engine Database, [19]. Sakata, K., Supersonic Experimental Airplane (NEXST) for Next Generation SST Technology AIAA ; 39 th Aerospace Sciences meeting & Exhibit - Reno, Nevada, 2001 [20]. Marconi, F., et al, Sonic Boom Alleviation using Keel Configuration AIAA ; 40 th Aerospace Sciences Meeting and Exhibit Reno, Nevada, 2002 [21]. Kroo, I., Sturdza, P., Tracy, R., Chase, J., Natural Laminar Flow for Quiet and Efficient Supersonic Aircraft, AIAA ; 40 th Aerospace Sciences Meeting and Exhibit Reno, Nevada, 2002 [22]. Makino, Y., Suzuki, K., Noguchi, M., Yoshida, K., Non-Axisymmetrical Fuselage Shape Modification for Drag Reduction of Low Sonic-Boom Airplane, AIAA ; 41 st Aerospace Sciences Meeting and Exhibit Reno, Nevada, 2003 [23]. Kroo, I. Unconventional Configurations for Efficient Supersonic Flight, VKI Lecture Series on Innovative Configurations and Advanced Concepts for Future Civil Aircraft, 2005 Page 53 of 61

54 Appendix Progress Update Chart Page 54 of 61

55 Concept Selection Process Page 55 of 61

56 Initial Concept Drawings Page 56 of 61

57 Detailed Design Criteria DESIGN CRITERIA DETAILED DESCRIPTION SONIC BOOM geometry altitude weight wetted area aspect ratio frontal area induced drag control surface effects SUBSONIC NOISE engine placement noise mitigation type of engine CONTROL SURFACES types of control surfaces static stability number location of control sufaces airport compatibility complexity size of control surfaces weight TURN AROUND TIME # of passenger doors # of service doors preflight checks AIRPORT COMPATIBLE location of control sufaces location of doors geometry fuel bay location engine placement SAFETY engine placement fuel bay location landing gear vertical stabilizer osciallations emergency exits stability debris preflight EASE OF MANUFACTURE moving parts materials used geometry costs EMPTY WEIGHT # of engines materials used # of landing gear wing size # of moving parts COST procurement operation manufacture maintenance crew Design Mission Profile Page 57 of 61

58 Historical Seating Data Combined Mass fraction Averages Combined Mass Fraction Average Aircraft Weight Wing System Tail System Body System Alighting Gear System nacelle System Propulsion System (less dry Engin Flight Control System Auxiliary Power System Instrument System Hydraulic and Pneumatic Sytem electrical System Avionic Systems Furnishing and Equipment System Air Conditioning System Anti-icing system Load and Handling System Empty Weight Dry Engine Weight Empy Weight (Maximum Empty) Page 58 of 61

59 Initial Sizing results Page 59 of 61

60 Requirements Compliance Matrix Page 60 of 61