AAE 451 System Design Review March 30, 2010

Team 1 Alex Mondal Beth Grilliot Brien Piersol Heath Cheung Jason Liu Jeff Cohen Jeremy Wightman Kit Fransen Lauren Hansen Nick Walls Ryan Foley Tim Fechner

TABLE OF CONTENTS

EXECUTIVE SUMMARY ...... 3 MISSION STATEMENT ...... 4

DESIGN MISSION ...... 4 TYPICAL OPERATING MISSION ...... 5 MAJOR DESIGN REQUIREMENTS ...... 5 AIRCRAFT CONCEPT SELECTION ...... 6

ENGINEERING LOGIC ...... 6 OUTCOME AND LAYOUTS ...... 10 COMPLIANCE MATRIX ...... 15 ADVANCED AND UNCONVENTIONAL TECHNOLOGIES UNDER CONSIDERATION ...... 16

CANARDS ...... 17 FORWARD‐SWEPT WINGS ...... 18 BLENDED‐WING‐BODY ...... 19 PROPFAN ENGINES ...... 21 LAMINAR FLOW CONTROL ...... 21 CONSTRAINT ANALYSIS AND DIAGRAMS ...... 23

BACK SWEPT WING AIRCRAFT CONCEPT ...... 24 FORWARD SWEPT WING AIRCRAFT CONCEPT ...... 25 BLENDED WING BODY AIRCRAFT CONCEPT ...... 26 TRADE STUDIES ...... 28 SIZING STUDIES ...... 30

CURRENT APPROACH ...... 30 DRAG PREDICTION ...... 35 ENGINE MODEL ...... 35 ENGINE AND PROPULSION SELECTION ...... 37

APPROACH TO MODEL ENGINE ...... 44 STABILITY ...... 45

INITIAL VERTICAL TAIL SIZING ...... 48 SUMMARY AND NEXT STEPS ...... 50 WORKS CITED ...... 52 APPENDIX A: HOUSE OF QUALITY ...... 55 APPENDIX B: MERIT POOL ...... 56

EXECUTIVE SUMMARY

From market research, Team 1 determined that a long range business jet, capable of a still‐air mission of 7100 nmi would be most marketable in the year 2020. To complete these requirements, Team 1 sketched over 20 aircraft designs and assessed them using engineering logic. Three designs were selected: a back swept, forward swept, and blended wing aircraft. We researched the unconventional technologies of canards, forward swept, and blended wing bodies, as well as propfans. From this research, canards demonstrated improved lift characteristics at the cost of difficult placement. Forward swept aircraft have improved lift characteristics, allowing for decreased takeoff and landing distances. However, their design also necessitates improved composite materials used in the wing. Blended wing bodies demonstrate increased lift performance. However, aesthetically it may be less desirable due to the lack of windows in the cabin. From the constructed sizing code, we found initial empty weights. For the back swept configuration, its empty weight was 95,000 lbs. The forward swept’s empty weight was 99,000 lbs, and the blended wing body’s was 88,000 lbs. With these determined weights, and estimates for the passengers, fuel, and engines, we calculated stability. We found all three designs to be stable. The back swept aircraft had a 5.1% static margin, the forward swept aircraft had a 13.4% static margin and the blended wing body aircraft had a 10.2% static margin. Overall, the designs demonstrated viability. Further sizing refinement will better quantify the performance of these designs. We will be integrating our component weight build up into the sizing code in order to create a better iterative process. Improving our drag code will also provide better analysis for the generated values.

MISSION STATEMENT

To engineer a conceptual business aircraft solution capable of transporting esteemed passengers, in luxury, while adhering to NASA’s N+2 environmental goals. In order to abide by NASA’s N+2 environmental standards, our concept will provide reduced NOx emissions, reduced noise pollution, increased fuel efficiency, and an increased percentage of recyclable materials used in construction. These key topics will help address the primary concerns of environmentally conscious groups.

DESIGN MISSION To reach our mission statement goals, the idea of a long‐range business aircraft was chosen. By looking at long‐range business aircraft currently in production and choosing attributes that we believe we can improve, we created the design mission summarized below. ‐ 12 – 19 Passengers + 4 Crew ‐ Cruise Altitude > 40,000 ft ‐ Cruise Speed 0.85 Mach ‐ Still‐air Range of 7,100 nmi ‐ Takeoff Field Length 4,700 – 5,000 ft ‐ Landing Field Length 2,500 – 3,000 ft A high operating ceiling has many benefits. By choosing a cruise altitude of greater than 40,000 feet, our business jet will operate above the majority of air traffic allowing for higher speeds and a cruise‐climb method, increasing altitude as the aircraft becomes lighter from burning fuel. This method improves the overall efficiency of the engines and decreases fuel usage. Timely flights are a desirable characteristic that consumers desire in a business jet. High cruise speed directly correlates to the flight duration. Therefore, we chose the cruise speed of 0.85 Mach from historical data as it offers a high speed while maintaining fuel efficiency. A range of 7,100 nmi, a conservative distance from Los Angeles to Hong Kong with a 60 kts headwind, was the design mission range for the aircraft. Destination flexibility is also important for a desirable business jet solution. With a takeoff field length of 4,700 – 5,200 feet and a landing field length of 2,500 – 3,000 feet, these aircrafts will have access to many small 4

airports; this reduces the aircraft design’s reliance on larger and more congested terminals and, thereby, improves turnaround time and decreases wait times.

TYPICAL OPERATING MISSION It is not reasonable to expect the designed aircraft to operate at the full design mission at all times. Therefore, we designed the typical operating mission to carry 6 – 8 passengers, with 3 crew, over approximately 2,500 nmi. This mission allows for travel between many trans‐ continental cities. As a reference, a flight from New York to Los Angeles is 2,139 nmi. While this mission does not fully utilize the aircraft’s capabilities, the short takeoff and landing capacity will allow for more opportunities for shorter range flights in a given time frame.

MAJOR DESIGN REQUIREMENTS

From the definition of the target customer, we categorized the assumptions concerning desired characteristics for the business jet by performance, aesthetics, service, and extraneous attributes. After evaluating these characteristics, range, speed, and comfort are of highest priority, followed shortly by destination flexibility and environmental image. While range, speed, and comfort are regularly desired qualities for a transpacific personal business jet, destination flexibility and environmental image are developing concerns. In countries and locations where smaller airports are available, the ability for a business jet to avoid landing at busier airports gives the customer the ability to make deadlines without worry of external factors preventing prompt landings. To address these factors, we selected a variety of performance characteristics. The house of quality is a graphic tool that defines a relationship between customer desires and the business jet’s capabilities. From its results, we prioritized range velocity and fuel weight. Comparing the designed jet to the current competitors’, the range, price and speed are comparable, as seen in the House of Quality that is located in Appendix A. However, it is forecasted that a desire for technology which is less harmful to the environment will progress in the future. Using this House of Quality, we determined that the following properties were most important for meeting our design mission goals:

‐ Fuel Weight ‐ Empty Weight ‐ Cruise Speed ‐ Range ‐ Cabin Volume ‐ Take‐off Distance ‐ Cumulative Noise Level

AIRCRAFT CONCEPT SELECTION

With the properties above in mind, Team 1 produced over 20 aircraft concepts for consideration and debate. These concepts include many variations of conventional swept back wings with tails, forward swept wings, designs using canards and blended wing bodies. The numerous designs changed the position of the engines, wings, and tail to obtain many different possibilities to consider.

ENGINEERING LOGIC Because every member contributed an assortment of concepts, a large number of designs were categorically similar, if not identical. The generalizations became forward swept, back swept, and blended wing body configuration. Each of these had the possibility of H tail, T tail, V tail, tail mounted engines, wing mounted engines, engines above or below the wing, and canards. For discussion, the aircraft unable to be categorized under the above descriptions will be analyzed first. From Figure 1, concept A is a swing‐wing aircraft design. Team 1 discarded this design due to the structural problems such a design would encounter. B represents an aircraft with two sets of wings at different heights. Team 1 decided that such a design would suffer from the additional weight due to the extra structures, and the tradeoff with any benefits would not be worth the difficulties in overcoming the additional weight. C represents an engine running through the length of the plane. While interesting, the design would create excessive vibrations

6 and be disruptive to the passengers. Lastly, D represents a forward swept blended wing with canards, which was discarded due to the incompatibility of the three design features.

Figure 1: Aircraft of Alternative Classification

From Figure 2, concept E is Back Swept Aircraft with a lifting control canards and a V‐ Tail. We discarded the V‐Tail because the extra pitch control was not needed because of the canard. Concept F represents a H‐Tail aircraft with a bubble canopy. While this design gives the passengers an extra view, such a configuration would be unable to construct with the number of passengers for our design mission. Concept G represents an anhedral configuration with propeller. This configuration is not feasible for our design mission because the propeller limits the required Mach. Concept H represents a tail‐mounted engine on a back swept wing. With the placement of the engine on top of the vertical tail, the pitch moment would be hard to overcome, deeming this configuration not feasible for the back swept wing design.

Figure 2: Aircraft of Conventional Classification

Concept I, seen in Figure 3, is a forward swept wing design with integrated engines and strakes. Due to the vibrations of the integrated engine, this design is not feasible for our mission. Concept J is an H‐tail forward swept wing design. This design was very promising due to the H‐tail’s ability to shield engine noise. Since the engines are mounted aft of the wings, the center of gravity will be pulled back to the aft of the aircraft making the aircraft hard to stabilize. Another issue arose with the location of control surfaces. With a center of gravity so far aft, the control surfaces on the wing and tail would be very close to this center of gravity. The short moment arm in both cases would necessitate larger controlling mechanisms and the potential benefits would not justify the added structural weight. Concept K is an engine mounted forward swept aircraft with a T‐tail. Due to the structural support of the T‐tail, the wings cannot be placed at the back of the aircraft. This causes the center of gravity to be shifted forward and the aircraft will be hard to stabilize. Concept L eventually became the baseline of our forward swept wing. However, in concept L, the wings are placed above the forward swept wing. When placing the engines above the wing, extra support is needed for the wing which makes the wings heavier.

Figure 3: Aircraft of the Forward Swept Classification

Concepts M and N, seen in Figure 4, are integrated engine Blended Wing Body aircraft with winglets. Team 1 discarded these two designs due to the vibrations of the integrated engines. Additionally, concept O created an unbalanced aircraft. By having the engines mounted on the tail, the BWB would resemble a back swept aircraft with a vertical tail. The separate tail nullifies the use of the Blended Wing Body. Concept P has canards with engines mounted in pods. Since the engines are fixed to the body, the vibrations will occur throughout the cabin. Also, the canards do not add any additional benefit since the Blended Wing Body can be treated as a lifting surface.

Figure 4: Aircraft of the Blended Wing Body Classification

OUTCOME AND LAYOUTS After these considerations, three designs were obtained: a back swept wing with canards, a forward swept wing with canards and a blended wing body. These three designs incorporated the aspects that were thought would be the most beneficial or had the best trade‐ offs when meeting the given design mission and environmental concerns. The outside of the aircraft can be shown using isometric and orthographic views. Figure 5 and Figure 6 represent a Back Swept Wing aircraft with lifting, control canards. Engines are placed further up, near the tail, in order to reduce the noise and vibrations observed by the passengers. This aircraft is the more conventional approach compared to the other two designs. The Back Swept Wing aircraft can be assessed with historical data since it is similar to current aircraft. Another advantage to this aircraft is the use of window seats and passenger placement inside the cabin. However, the main disadvantage is the stability of the aircraft. Due to the location of the canard and main wing, the stability is harder to improve. Since the center of gravity is shifted aft, the only measure to improve the static margin is to move fuel placement further forward into the fuselage.

Figure 5: Isometric View of the Back Swept Wing Design

Figure 6: Orthographic View of the Back Swept Wing Design

Figure 7 and Figure 8 represent a Forward Swept Wing aircraft with lifting, control canards. As seen in Figure 7, the engines are mounted on the wing instead of being mounted near the back tail. This engine placement was chosen due to the lack of structural support in the tail and also to improve the center of gravity. Since the wings are forward swept, the center of gravity is pushed forward. However, the wings are placed at the aft of the aircraft. Due to this placement, the engines must be placed on the wings to help the static margin and to make sure the aircraft is statically stable. This configuration does not maximum the noise 11

reduction. If the engines could be placed at the aft of the aircraft, the wings would provide shielding and noise reduction. Another disadvantage is the requirement of certain physical limitations. This will be discussed later in the document.

Figure 7: Isometric View of the Forward Swept Wing Design

Figure 8: Orthographic View of the Back Swept Wing Design

Figure 9 and Figure 10 represent a Blended Wing Body aircraft with engines mounted above the wing. Since the engines are mounted above the wing, noise will be shielded away from the fuselage. Also, the Blended Wing Body is considered a lifting surface. This lifting surface leads to a better fuel consumption. The aircraft allows for a large cabin volume which 12

translates to more comfort for the passengers. The main disadvantage is that there are no working prototypes of this jet transport. More disadvantages will be discussed later in this document.

Figure 9: Isometric View of the Blended Wing Body Design

Figure 10: Orthographic View of the Blended Wing Body Design

We created cabin layouts in order to properly visualize the inside of each aircraft. For the back swept and forward swept designs the same cabin layouts were used since both would use the same sized fuselage. Figure 11 and Figure 12 are 16 and 12 passenger cabin layouts respectively. Both layouts offer a galley, storage spaces, dual restrooms, and a crew rest for a reserve pilot. The crew rest is in place for trans‐Oceanic flights because the flight time exceeds the work day of a single pilot. The cockpit, though not specifically shown in either design, would be a “typical” layout

seen in most business jets. The tail section, to the right of the figures, would be largely empty, possibly housing a baggage compartment. Emergency exits would be located on the starboard side of the plane and out the aft pressure bulkhead of the cabin.

Figure 11: 16 Passenger Layout for the Back Swept and Forward Swept Concept

Figure 12: 12 Passenger Layout for the Back Swept and Forward Swept Concept

The cabinets and galley will provide ample storage for food, beverages, blankets, pillows, and other desired commodities. The single seats in the cabin will have the ability to swivel, which would facilitate conferencing and other tasks. The red rectangles and ellipses denote tables which can be stored inside the side walls when not in use. This will allow for increased personal space during other activities when desk space is not needed. To the rear of the passenger cabin is a pair of joined seats. These are intended to be used for high capacity flights to fit the last two passengers at the cost of some personal space. Overall, the designs are fairly typical of business jets in this class of operations. For the blended wing body concept, there are two cabin layouts being considered. Figure 13 is an 18 passenger layout which is designed like an extra wide conventional fuselage cabin. It includes the same accommodations as the back and forward swept wing cabin layouts including 2 restrooms, a galley, storage cabinets and a crew rest. While no tables are shown in

14 this layout, they would most likely swing up from the sides of the chairs, similar to a student desk in a lecture hall.

Figure 13: 18 Passenger Layout for the BWB Concept

Figure 14 is a 16 passenger layout that takes advantage of the blended wing body to allow for the cabin to flare out into the aft part of the wing. This layout features theater style seating and includes the previous amenities like a restroom and crew rest. This layout would also allow for a screen to be installed just aft of the cockpit doorway which would utilize the seating arrangement so that a presentation or in‐flight movie may be shown.

Figure 14: 16 Passenger Layout for the BWB Concept

COMPLIANCE MATRIX

Table 1 shows the compliance matrix that was complied for each concept using the performance characteristics obtained from the House of Quality.

Table 1: Compliance Matrix for the Three Designs

Currently historical data is used to obtain the amount of recyclable material and interior decibel level of the aircraft. The remaining values obtained are from a variety of methods. We found cabin height and volume from dimensioning our layouts. The Blended Wing Body cabin volume is larger due to the cabin being able to flare into the wing. The chosen engines determine the cruise speed, and the N+2 goals determine the cumulative certification noise level. We found Empty and fuel weight by using the sizing equations mentioned above. The Ground roll take‐off distance chosen accommodated the smaller airports we wish to be able to utilize. The payload for the aircraft is the total number of maximum passengers and crew multiplied by 225 lbs. Still‐air range is the distance from Los Angeles to Hong Kong in a 60 kts headwind. The threshold values were determined by looking at competitor business aircraft and making sure that we have at least the ability to perform on the same level of our competition.

ADVANCED AND UNCONVENTIONAL TECHNOLOGIES UNDER CONSIDERATION

The design team intends to meet or approach NASA’s N+2 goals for increased fuel efficiency, decreased emissions, and reduced noise generation. These are fairly lofty goals, and changes in aerodynamic design demonstrate the greatest potential for significant improvement in performance. As such, the team is evaluating several unconventional aerodynamic features

for use in the three concepts. These features include the use of canards, forward‐swept wings, and a blended wing body aircraft design.

CANARDS Although both are horizontal stabilizers, canards provide certain benefits over conventional horizontal tails. A horizontal tail experiences a nose‐up pitching moment due downwash from the main wing. This additional downward force must be countered by deflecting the elevator; this deflection adds trim drag to the aircraft. Canards experience up‐ wash and a nose‐down moment instead. These require less trimming to maintain level flight and consequently less drag. Canards are separated into two classes: lifting canards and control canards. Lifting canards share a portion of the lift with the main wing while also providing pitch control. This fact allows for main wing area to be reduced and provides for a small reduction in parasite drag. Lifting canards also provide beneficial stall characteristics. Most canard aircraft are designed such that the canard stalls before the main wing; when the aircraft does stall, the canards will lose lift first, and the aircraft will nose down out of the stall condition. In a control canard configuration, the main wing carries all the lifting load of the aircraft, while the canard is used for pitch control. These are typically employed on aft‐wing designs or in situations where an aft tail cannot be utilized. As with lifting canards, deflection of the control canards provides a significant pitching moment about the aircraft’s center of gravity, helping to counter pitch‐up and stall. The use of canards will provide some challenges in the design process because they are sensitive to placement; there are opportunities for flow interaction between the two surfaces because canards precede the main wing. This causes induced downwash on the main wing which adds to the induced drag of the aircraft. The use of canards will also impact the location of the center of gravity. By placing the wings further aft on the aircraft, this configuration reduces the flow interaction between the canard and the main wing. While this provides a benefit to cabin length, the long moment arm from the wings to the center of gravity will produce unwanted effects, such as requiring that the wing only stores a fraction of the fuel. Additionally, it is likely that complex trailing‐edge lift devices, such as Fowler flaps, will not be

feasible because of the increased nose‐down pitching moments during deployment. Despite these challenges, we believe that with careful design, the reduced trim drag can outweigh the potential drawbacks facing the canard design.

FORWARD‐SWEPT WINGS There are well‐documented advantages of forward‐swept wings (FSW). When compared to a back‐swept wing of the same area, forward swept wings provide the advantages of higher lift to drag ratio, higher range at subsonic speed, improved stall resistance and anti‐spin characteristics, improved stability at high angles of attack, a lower minimum flight speed, and a shorter take‐off and landing distance (Vatandas). Due to a lower leading‐edge sweep angle for the same quarter‐chord sweep, laminar flow transition is delayed on forward swept wings. This results in greater lift at slower speeds and lower angles of attack; this effect can be seen below in Figure 15. In addition, the forward‐swept configuration causes air to flow from the tips towards the fuselage. This means that the aircraft stalls first at the root, and thus retains aileron effectiveness even during stall conditions. As consequence, less air flows off at the tips and reduces the downwash and induced drag (Hepperle).

Figure 15: Laminar flow increase for a forward‐swept wing (Hepperle).

There are notable drawbacks to this design choice. Forward‐swept wings suffer from aero elastic divergence, a condition that can cause the wingtips to bend upward until structural failure occurs. Using composite materials and aero elastic tailoring, as seen in Grumman’s X‐29 experimental fighter design, addresses this problem. Therefore, by carefully orienting the composite laminate fiber directions, it is possible to design structural mechanisms to compensate for the bending tendency. By using this design technique, the X‐29 was able to achieve a 15% increase in lift‐to‐drag ratio (Vatandas). In order to reduce interference and improve control, FSW designs favor canards over aft‐mounted tails. As such, they have the same difficulties concerning center of gravity, fuel placement, and high‐lift devices as a back swept canard design. Small control canards are often used, as they can provide longitudinal control for the aircraft while minimizing flow interference over the wing.

BLENDED‐WING‐BODY In recent times, various companies have placed a greater emphasis on research into blended‐wing‐body (BWB) designs. Boeing and NASA studies have demonstrated that remarkable savings in fuel, weight, and drag are possible with existing technology. In one 2002 study, Boeing compared a BWB design to a conventionally‐configured aircraft. They sized both the conventional and BWB designs using the same payloads, engines, and technology factors; the only difference between the two sets of requirements was the shape of the aircraft. The BWB design exhibited 27% lower fuel burn, 15% lower takeoff weight, 12% lower operating empty weight, 27% lower total thrust, and a 20% increase in lift to drag ratio (Liebeck). Subsequent analyses performed by Boeing and NASA confirmed these initial findings. The success of the BWB design stems from several factors. First, the BWB design has substantially less interference drag compared to conventional designs. This is because the entire aircraft is one smooth shape and has very few sharp corners or connections to cause interference. Second, in a BWB, the fuselage generates lift. While the lifting efficiency of the body may not be as high as the wings, the large surface area still provides a significant amount of lift. This fuselage generated lift reduces the required surface area of the wings and consequently reduces the wetted surface of the aircraft. The reduction in surface area causes a

19 proportional reduction in viscous drag as compared to a conventional design. Third, the span loading of the aircraft allows for reduced structural weight. A conventional aircraft has the fuselage load concentrated in the center of the aircraft. As a result, heavier structure is needed to take the large bending loads present at the wing root. In a BWB aircraft, the load is distributed more evenly across the entire structure, as seen in Figure 16. The combination of these factors makes the BWB a very strong candidate to satisfy the design goals.

Figure 16: Loading comparison for BWB and conventional aircraft. (Liebeck)

While the BWB design excels aerodynamically, several issues need resolution in order to make a viable BWB business aircraft. First and foremost, the BWB design precludes window seats. Though it is possible to have windows towards the front of the cabin, the majority of the passenger area will not have windows adjacent to the seats. This is a major concern, as window location and size is often a major selling point for business aircraft. To help alleviate this problem, it may be possible in install skylights in the aircraft, allowing some natural lighting in the cabin. A slightly more radical solution is to place cameras outside the aircraft, with LCD screens on the back of the seats or on the side walls of the cabin. While this would effectively allow any seat to be a window seat, with the additional functionality of a ‘smart board’ to manage documents or music, it is unknown if this compromise would be acceptable to passengers. The second potential challenge with the BWB aircraft is the construction of the center body. Because the center body will be taking both pressurization loads from the cabin and bending loads as a wing, it will likely approach its design load during every flight. This subjects the center body to very large levels of fatigue. In this case, manufacturing the center body from aluminum would be prohibitively heavy, due to the large tolerances needed. Using composite

materials for the center body is a viable alternative due to their comparative immunity to fatigue (Liebeck). Much like the canard and FSW aircraft designs, the BWB will likely be unable to utilize high‐lift trailing‐edge devices. Because the BWB lacks a horizontal tail, there is no other surface to counteract the large moment generated from using such devices. As a result, the BWB will likely require a higher rotation angle for takeoff and landing. This could lead to passenger discomfort, and will require careful design of the wing sweep and aft configuration to prevent tail dragging. One of the most significant concerns facing the BWB design is in regards to scale. Nearly all of the research currently being conducted by Boeing and NASA is for a large jet transport BWB, for 200‐450 passengers. The design team has determined that a spacious cabin is possible for a business jet sized BWB; however, it is unknown whether this design would achieve the same measures of aerodynamic efficiency as the larger models. Investigation is ongoing, and further sensitivity studies will show if the BWB design is feasible.

PROPFAN ENGINES Since the previous SSR report, further research regarding the use of propfans has revealed several issues in regards to noise levels. NASA Langley ran several test runs with propfan engines in the late 1980s and early 1990s. Reports of these tests show that noise levels for an aircraft at cruising altitude (30,000 feet) were often in excess of 60 dB at the ground (Garber and Willshire). In light of this research, progression in development of propfans has ceased in favor of alternative power plants. Although the estimated fuel savings would provide a significant contribution in reaching efficiency goals, the noise levels propfans generate will likely be unacceptable. Recent research programs have been restarted, thus we will continue observing the results of these studies will in order to determine whether the noise issues may be mitigated.

LAMINAR FLOW CONTROL Laminar flow control (LFC) technology has been in development since the early 1930s. The primary goal of the system is to extend the area of laminar flow over an aircraft surface.

Extending this flow and reducing turbulent zones results in an increase in lift and a decrease in drag. Since skin‐friction drag often accounts for 50% of the total drag of a subsonic aircraft, and laminar skin friction generates as much as 90% less drag than turbulent, laminar control can achieve significant savings. The provided benefits allow for better fuel efficiency, lower gross takeoff weight, and increased cruise lift to drag ratio. Current LFC technology uses a combination of passive and active systems to achieve greater laminar flow. Passive systems include designing surfaces for favorable pressure gradients and limiting wing sweep, while active systems often involve suction over the leading edge region of the surface. In general, rows of perforations are used. Flight tests done by Lockheed Martin in 1987 used a perforated LFC system along the leading edge and front spar of the wings of a Jetstar transport aircraft; these tests achieved 80‐90% laminar flow along the wing, resulting in an estimated 6% drag reduction for the entire aircraft (Joslin). Some considerations are necessary before the business jet design can implement LFC. First, it is likely that the business jet will be manufactured predominantly from composite materials. Currently, most of the LFC testing and research has been done using titanium or stainless steel. It remains to be determined if composite materials can be perforated without imposing larger maintenance or inspection requirements. Icing and insect contamination of the system is also a concern. Testing has demonstrated that effective use of a Krueger leading‐ edge device or liquid sprays are effective for preventing these problems. However, both of these systems add weight and complexity to the aircraft. Sensitivity studies will be necessary to determine if the potential benefits outweigh these disadvantages. As stated before, NASA’s N+2 goals are a driving factor in the business jet design. In order to meet these criteria, changes in aerodynamic design have the greatest potential for significant improvement. As a design project, these concepts are in a unique position; without constraints due to previous capital investment or threat of financial loss, the project team is free to pursue design features that would otherwise be considered too risky in the aerospace market. Thus, by investigating such technologies, there is greater potential for meeting the goals of marketability and environmental safety.

CONSTRAINT ANALYSIS AND DIAGRAMS

We derived the process of sizing the business aircraft from the constraint analysis. Since the System Requirements Review, three aircraft concepts were designed. The back swept wing aircraft, forward swept wing aircraft, and blended wing body aircraft each have different technical factors that influence the constraint analysis. Subsequently, the back swept wing design is a derivation of a conventional business aircraft; thus, we made newer and more accurate assumptions based off the previous designs. In Table 2, the updated assumed values are compared against the initial values from the SRR.

Table 2: Changes in Assumed Performance Values SRR Assumed Performance Values SDR Assumed Performance Values AR = 8.0 AR = 8.2 CD0 = 0.015 CD0 = 0.012 CL Take‐Off = 1.2 CL Take‐Off = 1.4 Take‐Off Ground Roll = 4000 ft Take‐Off Ground Roll = 4500 ft Landing Ground Roll = 2700 ft Landing Ground Roll = 3000 ft CL Landing = 1.5 CL Landing = 1.7

The updated value listed for is now based on rough calculations rather than a guess, hence its increased value. The calculations were very conservative and this value will likely be reduced in subsequent design iterations. Additionally, new constraint diagrams for the blended wing body aircraft and the forward swept aircraft were produced. The assumed values for all of the concepts were determined from engineering judgment and research. Each of these concepts has different performance constraints because of the diverse design features. The four primary constraints are Take‐Off (TO) Ground Roll, Top of Climb, Second Segment Climb, and Landing Ground Roll. The sensitivity of the various performance values on these constraints varies and depends heavily on the design itself. Take‐Off Ground Roll, Landing Ground Roll, and Second Segment Climb are very sensitive to the assumed values of the max coefficient of lift for take‐off and landing. When the coefficient of lift is increased, the performance constraints either move up or to the right. Aspect ratio and induced drag affects the Top of Climb as well as Second Segment Climb. As aspect ratio increases or induced drag decreases, the constraints move down or left. The distances of landing and take‐off ground roll affects Take‐Off and Landing Ground Roll.

When those distances are shortened, the aircraft is limited to a small area in the diagram. Now the performance values for each concept will be discussed individually.

BACK SWEPT WING AIRCRAFT CONCEPT The assumed values for the first concept, the back swept wing aircraft, were based on the design’s key features. The back swept wing aircraft has lifting canards, as well as an aft mounted wing. The canards give the aircraft lower trim drag but increase the induced drag. Thus, the induced drag is 0.02 and the max coefficient of lift is increased from a conventional business jet at 1.5 for TO and 1.8 for landing. Nikolai’s maximum lift over drag (L/D) equation found the value of 18.58 for the design (Nicolai). The lifting canards provide an increase in lift which is why this number is higher than a standard conventional jet’s L/D ratio. Since the design mission includes landing at secondary airports, the assumed TO and landing distances necessitate factoring in the requirements for those secondary airports. Furthermore, the assumed flight velocities were found from research done on aircraft of the same class. Table 3 shows the assumptions which were made for the Back Swept Wing Design.

Table 3: Assumptions for the Back Swept Wing Design Major Constraints Assumed Values CL Max 1.4 for Take‐Off; 1.7 for Land L/D (max) 18.58 Empty Weight Fraction ~0.5 to ~0.54 Engine Lapse Rate/SFC at Cruise 0.45 CD0 0.012 Oswald Efficiency 0.8 Flight Velocities Cruise: 0.85 M; Take‐Off: 149 ktas; Landing: 130 ktas; Stall: 100 ktas Aspect Ratio 8.2 Take‐Off Distance 4,500 ft Landing Distance 3,000 ft

The values that were input into the constraint analysis program produced a graph that demonstrates how the aircraft is constrained. Shown in Figure 17, this design is constrained by Second Segment Climb, Top of Climb, and Landing Ground Roll. Thus, a thrust to weight ratio at sea level of 0.33 and a wing loading of 80 lb/ft2 are optimal values for this aircraft.

Figure 17: Constraint Diagram for the Back Swept Wing Design

FORWARD SWEPT WING AIRCRAFT CONCEPT Assumed values for the second concept, the forward swept wing aircraft, were based on the use of canards and a forward swept wing. The canards on the forward swept aircraft are meant for control. However, unlike a conventional back swept aircraft, the forward swept design directs air flow to the root of the wing instead of the tips. The plane is therefore able to fly at higher angles of attack. Research done by NASA indicates that the L/D ratio for a forward swept aircraft is 15% greater than a typical conventional aircraft (Trippensee). So the maximum L/D for the second concept is approximately twenty. The aircraft can fly at higher angles of attack without stalling and can then fly at considerably lower speeds for take‐off and landing while still generating sufficient lift. This requires less distance to take off or land. The aspect ratio was assumed to be the same as the back swept wing aircraft because the planes are similar in design except for the sweep angle of the wings. Below in Table 4, are the assumed values for the constraint analysis.

Table 4: Assumptions for the Forward Swept Wing Design Major Constraints Assumed Values CL Max 1.6 for Take‐Off; 1.9 for Land L/D (max) 20 Empty Weight Fraction ~0.4 to ~0.54 Engine Lapse Rate/SFC at Cruise 0.45 CD0 0.012 Oswald Efficiency 0.8 Flight Velocities Cruise: 0.85 M; Take‐Off: 130 ktas; Landing: 115 ktas; Stall: 80 ktas Aspect Ratio 8.2 Take‐Off Distance 4,300 ft Landing Distance 2,800 ft

Figure 18 shows that this design is constrained by Second Segment Climb, Top of Climb, and Landing Ground Roll. Thus, a thrust to weight ratio at sea level of 0.32 and a wing loading of 85 lb/ft^2 are optimal values for this aircraft.

Figure 18: Constraint Diagram for the Forward Swept Wing Design

BLENDED WING BODY AIRCRAFT CONCEPT Assumed values for the last concept, the blended wing body aircraft, were derived from engineering judgment and research on this new design. The blended wing body is a lifting body

across its entire surface. This type of aerodynamic loading should significantly lower the wing loading. Research done by NASA indicates that the L/D ratio for a blended wing body is 15% greater than a typical conventional aircraft (NASA). So the maximum L/D for this concept is roughly around twenty. Since the entire surface is lifting, the aspect ratio is more important in the sizing of the aircraft. The aspect ratio for this design is at 7.8 because a shorter span is necessary to compensate for the runway width at most smaller airports. The induced drag on this aircraft is different from a conventional aircraft at 0.010 rather than 0.012. Like the forward swept aircraft design, the take‐off and landing speed as well as the take‐off and landing distance are shorter because of the increase in the L/D. However, the coefficient of lift values are lower than the forward swept aircraft because this design is not able to fly at a high angle of attack, since it has a back swept wing. This is due to the aircraft being more streamlined than a conventional aircraft. Below in Table 5, are the assumed values for the constraint analysis.

Table 5: Assumptions for the Blended Wing Body Design Major Constraints Assumed Values CL Max 1.5 for Take‐Off; 1.8 for Land L/D (max) 20 Empty Weight Fraction ~0.4 to ~0.54 Engine Lapse Rate/SFC at Cruise 0.45 CD0 0.010 Oswald Efficiency 0.8 Flight Velocities Cruise: 0.85 M; Take‐Off: 130 ktas; Landing: 115 ktas; Stall: 100 ktas Aspect Ratio 7.8 Take‐Off Distance 4,800 ft Landing Distance 3,000 ft

Shown in Figure 19, this design is constrained by Second Segment Climb, Top of Climb, and Landing Ground Roll. Thus, a thrust to weight ratio at sea level of 0.3 and a wing loading of 60 lb/ft^2 are optimal values for this aircraft.

Figure 19: Constraint Diagram for the Blended Wing Body Design

As shown from the constraint diagrams above, the blended wing body can fly at a lower thrust to weight and wing loading than the other two designs. The back swept wing concept has lower drag and more lift than a conventionally designed aircraft; however, it is also less stable. The forward swept with canard design has an increased L/D as well as a shortened take‐ off and landing distance. The blended wing body has lower drag because of the fuselage being a lifting body surface as well as an increased Lift‐to‐Drag ratio.

TRADE STUDIES Trade studies are continually being produced in order to maximize performance values for each aircraft design. Through the constraint diagram, the main variables to be considered are the thrust to weight at sea level and the wing loading. By increasing or decreasing the thrust to weight at sea level by 0.1, the weight values changed by roughly 500 pounds and when the wing loading was increased or decreased by 10, the outputted weights changed by roughly 200 pounds. These values were derived from the sizing code. We concluded that wing loading is not as important as the thrust to weight ratio because the change of thrust to weight ratio by 0.1 is

28 more likely to occur than the change of wing loading by 10. When weight increases, cost of the aircraft will increase as well. As far as performance constraints, when Mach number increases, the top of climb constraint line rises. When the top of climb increases, the constraint area is smaller and increases the thrust to weight ratio. This in turn increases weight; however, the travel time for the customer is lowered. An example of this is shown in Figure 20.

Figure 20: Trade Study for increase of Mach

Additionally, since the concepts are mainly constrained by the landing ground roll, second segment climb and top of climb, increasing the take off distance did not play a role in the sizing of the concepts. When each variable was analyzed, there did not appear to be a feature that impacted the performance requirements drastically. However, by keeping the thrust to weight ratio low and the wing loading higher, all concepts will maintain the lowest weight possible.

The approach to this trade study used The Bayesian Team Support approach from Trade Studies with Uncertain Information (Ullman). This method is a way of dealing with uncertain and evolving information revolved around quantitative and qualitative information.

SIZING STUDIES

The aircraft mission is composed of a two parts: a design segment and a detour segment. The mission at which the aircraft is sized includes: 10 passengers, pilots and crew to have a range of 7,100 nmi while cruising at Mach 0.85. For the detour, the aircraft must be able to travel an additional 200 nmi to an alternate destination. The design range chosen classifies the aircraft as a long range business jet. Furthermore, another design choice is for a maximum capacity of 19 passengers. This requirement, however, does not play a direct role in the current sizing process because the sizing is performed for a typical operating mission. Currently the main sizing code is only capable of sizing the back swept canard concept. The constraint diagram for the Back Swept Wing states that the minimum thrust to weight at sea level required is constrained by the top of climb and landing ground roll on high and hot days. The diagram for the back swept wing shows that at sea level a minimum thrust to weight of 0.33 is required and that the aircraft has a wing loading is 68 lb/ft2.The following table, Table 6, shows the conditions which the aircraft is sized.

Table 6: Constraints Constraining variables Maximum Range 7,100 nmi Payload 10 passengers @ 225 lbs/person Wing Loading 68 lb/ft2

Minimum (T/W0)SL 0.33 Cruise Mach 0.85 Cruise Altitude 42,000 ft

CURRENT APPROACH The approach taken to estimate the gross weight of the aircraft comes from the course text “Aircraft Design: A Conceptual Approach Fourth Edition” by Daniel P. Raymer. A rubber engine assumption was used, thus allowing the size and performance of the propulsion system to change with the weight of the aircraft.

The sizing of the aircraft involves beginning with desired capabilities of the aircraft, estimating remaining values and iterating until the specified constraints are obeyed. The gross weight is comprised of the crew weight, payload weight, fuel weight, and aircraft empty weight. The calculations for fuel weight begin various fuel fraction estimates for different segments of the mission. The following table shows where this approach was taken and the corresponding values.

Table 7: Weight Fractions for the Sizing Code (Raymer)

Taxi & Takeoff W1/W0 0.97 Climb W2/W1 1.0065‐(0.0325×M) Cruise W3/W2 Discussed later Landing W4/W3 0.995 Missed approach W5/W4 0.97 Climb to Detour W6/W5 1.0065‐(0.0325×M) Cruise To Alt W7/W6 Discussed later Detour Loiter W8/W7 Discussed later Detour Landing W9/W8 0.995

The segment fractions and climb estimates were all taken from the course textbook. The Brequet range equation was used for determining the loiter fuel fraction, here it is in original form: ln Equation 1: Breguet range equation (Raymer) It can be rearranged slightly in order to form the “endurance equation”, as seen here:

ln Equation 2: Endurance equation (Raymer)

This can further be rearranged to form the loiter weight fraction, Equation 3.

Equation 3: Loiter weight fraction (Raymer) Both cruise segment fuel weights were directly approximated with drag and specific fuel consumption predictions, both of which will be outlined later. The analysis involves dividing the cruise segments into smaller intervals where weight, SFC, and drag were held constant. With 31

the flight conditions and geometry, drag was determined from a subroutine. The aircraft was assumed to be at steady state over these intervals, therefore the thrust required was simply that of the drag produced. Equation 4: Thrust required Using flight conditions and the thrust requirement, specific fuel consumption was estimated. This produced enough information to determine fuel flow rate, as seen in the following equation:

Equation 5: Fuel weight flow rate With a specified cruise speed and range interval, a flight time for the interval can be calculated. The previous equation states the fuel flow per unit time, with a flight time, the actual weight of the fuel was found: Equation 6: Fuel weight consumed The process was repeated until the cruise segment was completed, the initial form of the code breaks the each cruise segment into 20 intervals. The structure of the code is designed such that it can scale in sophistication to allow for changes in altitude and Mach over the cruise segment. With the fuel weight estimated, the empty weight of the aircraft needed to be determined next. Without a complete component weight build, the following equation was used for approximating the aircraft empty weight.

Equation 7: Empty weight fraction from statistical equation (Raymer) In the previous equation, W0 is the gross weight, is the aspect ratio, is the

thrust to weight ratio, is the wing loading and Mmax is the maximum Mach number. A database of aircraft including the Gulfstream G650, G550 and the Bombardier Global Express

was created and the coefficients a, b, and C1 through C5 are found by solving a matrix equation. In performing the analysis, the coefficients found to give the least error in the correlation of the five aircraft properties did not follow logical correlation. For example, it is obvious from the previous equation that if the gross weight were to increase the empty weight

fraction should decrease, thus C1 should be a negative value. In an effort to arrive at a realistic correlation, the database of thirteen business jets was reduced to only four which still failed to give a logical fit. It was concluded that the aircrafts in the database were too diverse to allow for this method of determining an empty weight fraction. In order to determine empty weight fraction, a correlation between jet transport aircraft found in “Aircraft Design: A Conceptual Approach Fourth Edition” was used. The resulting coefficients are shown in Table 8.

Table 8: Empty weight fraction coefficients (Raymer) Coefficients for Jet Transport 0.32 0.66

C1 ‐0.13 C2 0.30 C3 0.06 C4 ‐0.05 C5 0.05

A correction factor of 1.08 was determined by averaging the difference in empty weight ratios calculated from Equation 7 and actual empty weights of some business aircrafts in the database. Thus, Equation 8 is the final empty weight fraction equation used in the sizing.

. . . . . 1.080.32 0.66

Equation 8: Empty weight fraction from statistical equation (Raymer)

Upon reaching an estimate of the gross weight, a logic statement is posed to determine if the difference between the guessed gross weight and calculated gross weights is less than 1000 lb. If the gross weight does not meet the tolerance, the process is iterated until the constraints are met. Figure 21 is a flow chart of how the main sizing predicts the gross weight of the conventional back swept concept.

Figure 21: Main Sizing Code Flow chart In order to determine the weight estimate for the blending wing body design, equations and estimates were used from Daniel Raymer’s text. Estimations for the weights of the wings, vertical tail, landing gear, avionics, crew, passengers, and lavatories were obtained using the rough estimates. For the fuselage and aft body weight estimates, equations from the NASA report “A Sizing Methodology for the Conceptual Design of Blended‐Wing‐Body Transports” by Kevin Bradley was used. This report detailed the process of adapting the Flight Optimization System (FLOPS) to provide analysis for BWB aircraft. In this report, equations were provided that related the weight of the fuselage and aft body to the surface area of the cabin, the area of the aft body, and the number of engines. These equations are listed below.

Equation 9: Weight of Fuselage using W0 of conventional A/C (Bradley)

Equation 10: Aft‐Weight of Fuselage using W0 of conventional A/C (Bradley)

In estimating the gross weight of the forward swept wing design, the gross weight estimate of the conventional concept was again used. It was assumed that additional weight for structural support and avionics for extra control would be required for this design.

DRAG PREDICTION The drag prediction of the sizing code approximates the drag on the aircraft by summing three drag components together: induced drag, parasite drag, and wave drag. The wave drag is difficult to calculate without a final geometry so a conservative estimate of 20 counts of drag was added to the calculated coefficient value to account for wave drag. Parasite drag is determined in a manner similar to a component weight build up. The parasite drag is calculated by summing the component drag values together. These component values are based on form factors, wetted areas and other factors as discussed in Chapter 12 of the Raymer textbook. The induced drag was calculated using a two horse shoe vortex model. This simple model is comprised of two horse shoe vortices located at each of the aerodynamic centers; one is located at the main wing's aerodynamic center and one at the canards. The downwash was then calculated at the respective aerodynamic centers using the Biot‐Savart Law. The induced angle of attack was then calculated using a small angle approximation between the values of downwash and the forward velocity. Induced drag was calculated directly using the same small angle approximation and multiplying the induced angle of attack by the lift of the surface. The parasite drag and the induced drag were then summed together and returned to the sizing code's main functions.

ENGINE MODEL As stated earlier, a rubber engine approach was used in the sizing study, which requires the use of a scaling factor in conjunction with a baseline engine. The engine data used as a basis for the calculations was provided by the course instructor. The data is for a high bypass turbofan rated for approximately 13,600 lbs of thrust at sea level static conditions. With a takeoff gross weight in the range of 90,000 lbs, the thrust requirement from each engine was expected to be around 15,000 lbs at SLS for a thrust to weight of about 0.33 and two engines. This produced a scale factor of:

15000 1.1029 13600 Equation 11: Scale factor This value was deemed close enough to one and the data was used for the rest of the calculations. To begin, all of the data for this baseline engine was curve fitted for different altitudes and Mach numbers. Fourth order polynomials were used, on the suggestion from the instructor, for each altitude and Mach number. Equation 12: Curve fits Where T is the thrust required for that flight condition, the coefficients a, b, c, and d are not related to those used for the empty weight fraction estimates. An optimal algorithm has yet to be determined for how the TSFC should be approximated for all Machs and altitudes; however what is in place is a method for approximating the TSFC for cruise. First the thrust requirement for the engine is found: Equation 13: Curve fits From there, TSFC’s for different Mach numbers at altitudes of 35,000 ft, 37,000 ft, and 43,000 ft were found from the curve fits. MATLAB’s interpolation function with Hermite interpolation was then used to estimate the TSFC for the given Mach and altitude. With the rubber engine approach, the scale factor changes as a new estimate of gross takeoff weight is calculated. Thrust to weight is a design variable which stays constant, as is the number of engines, thus the SLS thrust can be recalculated on each iteration, which results in a new scale factor. The current design takeoff gross weight of the back swept design with a canard is 96,000 lbs. Table 9 includes more estimates determined by the sizing code.

Table 9: Weight Estimates for the Back Swept Wing Design

W0 95,000 lb Wf/W0 0.41 We/W0 0.53 Wf 39,000 lb We 56,000 lb

Using the gross weight estimate of the forward swept canard concept is estimated to be 99,000 lbs. Table 10 includes further information about the estimated weight.

Table 10: Weight Estimates for the Forward Swept Wing Design

W0 99,000 lb Wf 39,000 lb We 60,000 lb

Using the gross weight estimate of the back swept wing design, the blended wing design concept was estimated to have a gross weight estimate of 88,000 lbs. Table 11 includes further information about the estimated weight.

Table 11: Weight Estimates for the Blended Wing Body Design

W0 88,000 lb Wf/W0 0.38 We 53,000 lb

ENGINE AND PROPULSION SELECTION

We considered several propulsion methods for our business jet. From these propulsion methods, several base categories were derived in order to evaluate the effectiveness of the engine for our business jet. These categories include: 1. Thrust required by the business jet (determined by drag) 2. Power required by the business jet (determined by drag and velocity) 3. Weight of engine 4. Type of propulsion and hence limitations of the type (propeller, turbine, other) 5. Overall Efficiency 6. Fuel consumption (SFC or TSFC) 7. Power loading (hp/W) Through deliberation of these topics, we appropriately selected the different types of engines that would be effective on a business jet. These include turbofans and unducted fans. Even though the end decision was to evaluate these two types of engine, we discussed many other types of engines. These engines included turboprops, high and low bypass turbofans, ramjets, scramjets, turbojets, turbine based combined cycle, rocket based combined cycle, 37

counter‐rotating fans, geared turbofans, and unducted fans. For our conceptual long range business jet, we have chosen to implement two turbofan engines that each provides 17,000 lbs of thrust. This decision was derived from discussion and our engineering judgment, where those two satisfied the desired efficiency and feasibility for a long range business jet. Since we have chosen a turbofan engine to be used on our jet, it is appropriate to discuss the specific types of turbofan engines. First, is the low and high bypass turbofan. This is the “conventional” type of turbofan used on current aircraft. A low bypass turbofan usually has a bypass ratio lower than 3 while a high bypass turbofan will have a bypass of more than 5. A turbofan engine is the most modern variation of the basic gas turbine engine. As with other gas turbines, there is a core engine, which is surrounded by a fan in the front and an additional turbine at the rear. The fan and fan turbine are composed of many blades, like the core compressor and core turbine, and are connected to an additional shaft. As with the core compressor and turbine, some of the fan blades turn with the shaft and some blades remain stationary. The fan shaft passes through the core shaft for mechanical reasons. This type of arrangement is called a two spool engine (one "spool" for the fan, one "spool" for the core). Some advanced engines have additional spools for even higher efficiency as discussed in some of the other turbofans below. Second is the Direct Drive Turbofan (DDTF), a concept developed by Rolls‐Royce, which is the standard design of today, but could be improved resulting from increasing the bypass ratios, increasing the turbine inlet temperatures, etc. Material developments would help contain the weight growth as well as increase temperatures. The geared turbofan reduces the number of turbine stages by allowing use of higher rotational speed in conjunction with a gearbox for the slower rotating fan. The gearbox weighs more, but is somewhat offset by weight savings from fewer turbine stages. They are also more complex, but would help decrease fuel burn. (Saravanamuttoo, Rogers and Cohen) The contra‐rotating fan incorporates two fans rotating in opposite directions to reduce noise (Snecma). As a result of the higher bypass ratio the fan speed and pressure ratio will be reduced. This engine type would, therefore, trade off reduced noise and emissions for increased weight, which is the main goal of our business jet.

Third, a geared turbofan being developed by Pratt and Whitney and MTU Aero has shown gains in efficiency over the direct drive turbofans of today, a photograph can be seen in Figure 22. Additionally, P&W flew a concept demonstrator in 2008 and has accumulated many flight hours and engine cycles since then. There does not appear to be information of a contra‐ rotating fan in development, much less close to readiness by 2020.

Figure 22: Geared Turbofan (Flight International)

Pratt and Whitney’s geared turbofan, the PurePower series, aims at improving propulsive efficiency, , by gearing the main fan to rotate slower. Current turbofans link their fans and low spool compressors to the same low pressure turbine, which forces each component to operate outside of their optimal regions. The gearing of the fan reduces the fan pressure ratio, improving propulsive efficiency, and allows the compressor to operate at higher rotor speeds, improving their efficiency. Maximizing allows for reductions in fuel burn, noise production, and emissions, all of which are emphasized in NASA’s N+2 goals. This efficiency is shown in Equation 14. 2 ⋯ 1 Equation 14: Propulsion Efficiency

This is maximized when the magnitude of exit flow velocity is close to the magnitude of freestream flow velocity. However, since jet thrust is defined by how much a given mass flow is accelerated, the lack of acceleration would also produce a lack of thrust. Therefore, the mass flow rate of air is needs to be increased in order for the propulsion system to be useful. This results in bypass ratios of 8:1 and 10‐12:1 for these engines (Norris), this results in larger engines, which presents a problem with engine placement for smaller aircraft. Underwing placement would be restricted for many of them and larger pylons would be required for standard aft fuselage installation. Pratt and Whitney intends on developing a smaller version of their geared turbofan for Mitsubishi’s regional jet. The PW1215G and PW1217G are rated for between 14k to 17k lbs at

SLS and are claimed to reduce fuel burn by 12‐15% versus current engines. NOx and CO2

emissions are reduced due to an advanced combustor, NOx emissions being 50% under CAEP (Committee on Aviation Environmental Protection) 6 standards. The smaller geared turbofans are also expected to reduce noise levels 15 decibels below stage four regulations (MTU Aero Engines GmbH). The PW1215G and PW1217G are to have fan diameters of 56 inches, which is translates to the engine being slightly larger than other high bypass turbofans in this thrust class. Rolls Royce’s BR710 is rated for 15500 lbs at SLS and has a fan diameter of 48 inches. Their BR725 at 17000 lbs SLS is about 50 inches in diameter (Rolls Royce). We are conducting studies examining the effects of added weight and drag due to the pylons and engine. Pratt and Whitney claims that the PW1000G series are lighter than current high bypass turbofans, but does not quantify this amount. Nonetheless attempts will be made to study the impact of a lighter engine. NASA’s N+2 goals call for a noise reduction of 42 dB from stage four, NOx emissions reduced 75% below CAEP 6, and a 50% reduction in fuel burn. The power plant choice alone will not satisfy targets, but should make significant progress towards being able to fly an environmentally sensitive business jet. Another considered propulsion choice, as discussed previously in the Advanced and Unconventional Technologies Under Consideration section, is the Propfan engine. This engine design has difficulties concerning noise but is expected to be more fuel efficient. The engine specifics were mentioned previously in the paper in Advanced and Unconventional

Technologies under Consideration. The propfan is also known as the Counter‐rotating turbofan. This layout simply means that there are two independent shafts, rotating in opposite directions. At the other end of the low‐pressure section, they are joined to a low‐pressure turbine with several stages of counter‐rotating blades. For a given aerodynamic load, this configuration will reduce the fan rotation speed by 30% or more. This fan concept offers the same performance as a conventional fan, but with slower tip speeds: which in effect reduces fan noise. Therefore, engines development would be able to follow a new trend enabling a higher bypass ratio with lower fuel burn. This would provide a good start to meeting NASA’s N+2 goals by reducing noise and fuel emissions. As shown, there are many benefits to using a propfan engine; however, there are important drawbacks which the benefits cannot overcome. Open rotor systems would maximize propulsive efficiency, but has issues with noise, installation, and certification dealing with blade containment (Saravanamuttoo, Rogers and Cohen). Currently there is a lack of detail about the ability to control the noise and vibration of the propfan engine. The noise heard inside the cabin of the aircraft must be reduced to avoid passenger discomfort during flight, and there must be a way to delay acoustic fatigue in the structure of the fuselage. In 1983, the acoustic noise measured from the fuselage was 150 dB. This high level of noise is caused from the high speeds of the tip blades; however, even when the tips are operating at subsonic speeds, the noise imparted to the fuselage is still significant. (Bauer) There have been advanced tests to determine if acoustical resonators could be placed in the walls of the fuselage to decrease the noise and vibrations caused from the propfan on the passengers. The tones produced by the engine require a large amount of transmission loss through the fuselage which resulted in a significant weight addition from the resonators used. One sidewall treatment produced a cabin noise between 80 and 85 dB (H.L.Kuntz). When adding such resonators to reduce the cabin noise, the takeoff gross weight (TOGW) of the aircraft is increasing. Additionally, the SFC benefits of the propfan are limited since the TOGW is much higher than an aircraft without resonators in the panel.

Even with current technologies being developed to control cabin noise, there is not enough information to prove that propfan noise and vibrations can be controlled. Since one of the main features of this aircraft is customer comfort, the ability to control the noise is extremely important. Installation of open rotor systems is also an issue with these technologies. The engine must be moved from under the wing to above the tail, or there must be a high wing to allow clearance for the fan blades. If the engines are tail mounted, there must be new ways to install the engines since this is not a common engine location. There have been some aircraft which have tail mounted engines, but this will negatively affect the stability of the aircraft, and may cause additional unforeseen problems once implemented (Thomas). If the engine is wing mounted, open rotor systems must deal with the consequences of having unducted fans, and the associated concerns associated with engine failure. Currently there are no certification standards with blade safety for open rotor systems, but they are expected to be more severe than those in place for turbine or propeller standards (Thomas). If an engine malfunctions, there must be proof that blade impact will not hinder the safety of the passengers. There has been testing for fan‐blade containment cases which would help eliminate the risk associated with engine malfunctions; however, there is not enough information to assume this will be feasible by 2020 (Documents for Small Businesses and Professionals). Another engine that potentially viable is the turbojet. If development in the near future to allow for the engines to provide more than 15,000 lbs of thrust, a turbojet will be viable. Until we can obtain data that proves that this will be possible by 2015, we will not be able to viably consider a turbojet. A significant problem is engine placement. Below are the advantages and disadvantages of placing an engine in particular arrangements (LeBeau). Figure 23 shows an overview of an Engine Location Summary Table.

Figure 23: Engine Placement (LeBeau)

Rear‐Mounted Engines: ‐Advantages 1. Aerodynamically clean, particularly wing interference 2. Small asymmetric moments 3. Reduced cabin noise levels ‐ Disadvantages 1. Heavier wing structure for lost bending relief 2. c.g. further aft, resulting in larger tail to balance moments 3. Long slender body causes a tendency for deep stall 4. Engines closer to passengers in crashes 5. More difficult to reach for maintenance Wing Mounted Engines ‐Advantages 1. Can be placed for optimal bending relief on wing, area ruling and flutter 2. Easier to maintain and replace

3. Closer to c.g., thus easier to balance and install different engine sizes ‐Disadvantages 1. Higher drag, both nacelle and interference 2. Higher noise, both for passengers and surroundings 3. Large range of intake angle of attack results in inlet inefficiencies 4. Engines closer to ground (FOD, clearance for smaller aircraft) Additionally the wing mounted engines can be mounted above or below the wing. Mounting the engine in this position causes several effects. Assuming that we can design a structurally stable design with the engines above the wings, the above wing mount would cause a slight reduction in external noise as seen in Figure 24.

Figure 24: Sound effects Due to Engine Placement (LeBeau)

Considering all of these advantages and disadvantages allows our team to choose whether or not an engine can be located in the ideal place on the aircraft. At the current stage the engines are to be mounted in the optimal location according the design. For example, the BWB will have the engines mounted on pylons at the rear of the plane.

APPROACH TO MODEL ENGINE In order to size an engine that would fit into the optimal configuration of the business jet, an engineering model will need to be constructed. This engineering model consists of an electronic program such as ONX‐OFFX or another from the trade. In order to construct this model, a baseline engine is chosen. For our business jet, a high‐bypass (bypass ratio of 5+)

turbofan engine was chosen as the baseline due to the fact that it most closely represents the current advances in propulsion technology. The smaller turbofans, while being more of the size that is desirable, are slightly outdated since they were created multiple decades ago. Our ideal choice for a baseline engine, therefore, would be the Trent 1000. This engine has very little data due to its recent development, so it would be hard to obtain a model from it. A slightly older engine was therefore chosen to act as our baseline engine. The current baseline that has been chosen is the GE90, the engine that was used on the Boeing 777 and began production in 1990. It still is a much older engine than we would ideally have for a baseline, but this is the best option available until we can obtain data for one of the newer engines. With the baseline engine, the resizing process will be conducted by iterating through a process where the program will alter values of the engine proportionally until the projected thrust values are met. A sample of this process of inputs is shown in Figure 25.

Figure 25: Process of Sizing (NATO)

STABILITY

Initial stability was calculated for each design by using a simple component build up process. In order to accomplish this, we drew sketches for each design in their optimal configurations chosen from our earlier analysis of each aircraft concept. From these sketches, we were able to closely estimate the position of each component for every concept. We then 45

calculated the center of gravity for each design from Equation 12 where CG is the design’s center of gravity, W is the component’s weight, and L is the location of the component’s weight measured from the nose of the aircraft. Equation 15: Center of Gravity

The aerodynamic center for the FSW and BWB aircraft was found by estimating it to be located at the quarter chord of the mean chord line. An additional factor had to be incorporated with our back swept design because of our use of a lifting canard that will bring the aerodynamic center forward. For this, a MATLAB code was written to estimate the effect on the canard on the aerodynamic center. Once we had obtained both the center of gravity and aerodynamic center for each of our three designs we could calculate the static margin. The location of both the center of gravity and the aerodynamic center were non‐dimensionalized with respect to the aircraft’s reference length. The static margin was then found simply from the difference between the aerodynamic center and the center of gravity. Table 12 shows the weights and locations of all components for our back swept design. Almost the entirety of our weight lies in the back of the design due to aft fuselage mounted engines, the aft vertical tail, and the wings being placed as far back from the canard as possible. This, unfortunately, moves the center of gravity for the design back to 63.6% of the design characteristic length. With our estimated aerodynamic center at 68.7% of the design characteristic length this produces a static margin of a mere 5.1%. Although this is ultimately stable, it does not have a desirable margin of error. Therefore, we are currently exploring options to move the center of gravity forward to increase our static margin, such as pushing fuel into the fuselage ahead of the wing. Further analysis will also attempt to determine a more optimal locations for the both the canard and main wing.

Table 12: Component Weights and Locations for the Back Swept Design Wing Tail Canard Engines Gear Fuel* Avionics Crew Passenger Fuselage Weight (lbs) 5200 500 500 7200 4100 39100 989 900 2250 34761 Location (ft) 80.21 110 16.5 104.5 62.33 80.21 5.5 5.5 60.5 55 *Assumed a Wet Wing 46

Table 13 shows the weights and locations of all components for our forward swept design. Similar to our back swept design, most of the weight is in the rear of the aircraft. Here, however, we were able to take advantage of our mid‐wing design and use the extra space underneath the wing to house the nacelles. This moves the engine weight farther forward and, therefore, also moves the center of gravity for this design forward. The estimated center of gravity for our forward swept design is currently at 61.6% of the design characteristic length. Also, by design, forward swept wings allow the wings to be placed in the far aft portion of the aircraft pushing our aerodynamic center back to 75.0% of the design characteristic length. This gives our forward swept design a static margin of 13.4%, much better than our back swept design.

Table 13: Component Weights and Locations for the Forward Swept Design Wing* Tail Canard Engines Gear Fuel** Avionics Crew Passenger Fuselage Weight (lbs) 6000 500 500 7200 4100 39100 989 900 2250 38061 Location (ft) 82.5 110 16.5 77 55 82.5 5.5 5.5 38.5 55 *Added fudge factor for aeroelastic tailoring **Assumed Wet Wing

Table 14 shows the weights and locations of all components for our blended wing design. The component build up for this design was much more complicated to compile. Most of the weights and locations came from NASA’s currently flying blended wing body scaled down to business jet size. The locations of these weights were estimated to the best of our ability. We were able to capitalize on the enlarged fuselage volume inherent in blended wing designs and move the fuel forward out of the wings. We were therefore able to obtain a center of gravity of 58.4% of the design characteristic length for this design. The aerodynamic center is at 68.6% of the design characteristic length giving this design a static margin of 10.2% which is acceptably stable.

Table 14: Component Weights and Locations for the Blended Wing Body Design Wing Tail Engines Gear Fuel Avionics Crew Passenger Fuselage Weight (lbs) 5100 700 7200 4100 35000 989 900 2250 31500 Location (ft) 72.5 83.125 83.125 39.375 52.5 4.375 4.375 30.625 43.75

INITIAL VERTICAL TAIL SIZING Initial tail sizing is a key step in any aircraft design process. Any aircraft that flies must be stable, or at least controlled to be stable, in yaw. Without a large enough vertical tail, aircraft could become unstable and not able to fly. In order to prevent this, the vertical tail on our designs must be large enough to be able to overcome any unwanted yawing moments with a maximum of a 20 degree rudder as a safety factor. In order to ensure our aircraft could fly under any circumstance, tail sizing was conducted for two cases. The first case was to make sure our design could fly with one engine out. This being the case, one engine would provide a positive yawing moment around the design’s center of gravity from its thrust, while the other engine would also provide a positive moment due to the extra drag. This yawing moment must be overcome by the vertical tail in order to ensure stable flight. The vertical tail can be properly sized by setting the sum of the moments from both engines and the vertical tail to be zero, as shown in Equation 16 where Meo is the moment provided by the operable engine, Meu is the moment provided by the inoperable engine, and Mvt is the moment provided by the vertical tail.

0 Equation 16 : Sum of the Moments when sizing the Vertical Tail (Raymer)

The moment on the operable engine can be found easily from the thrust given by the engine and the engine’s distance from the centerline of the aircraft. The drag on the inoperable engine can be estimated using an equation from Raymer’s book shown here in

Equation 17, where D is the drag on the inoperable engine, Aeff is the effective area of the front of the nacelle, and q is the dynamic pressure.

0.3 Equation 17 : Drag of the Inoperable Engine(Raymer)

From here, one must simply solve Equation 16 for the moment of the vertical tail and then divide this by the distance from the aircraft’s center of gravity to the vertical tail to get the force the vertical tail must produce to overcome these moments. Once the necessary force for the vertical tail was obtained it became possible to calculate the size of the vertical tail. This was done by taking the definition of a lifting force and

solving for the vertical tail area as shown in Equation 18 and Equation 19where N is the total yawing force, q is the dynamic pressure, S is the vertical tail area, cnβ is the coefficient of lift due

to sideslip, β is the sideslip angle, cnδr is the coefficient of lift due to rudder deflection, and δr is the rudder deflection angle. Equation 18: Total Yawing Force

Equation 19 : Vertical Tail Area

Table 15 shows a table of the calculated sizes for the vertical tail assuming one engine out for each of our three designs. As we expected, the swept forward has a much larger vertical tail than either of our other two designs. This is due to the fact that the engines on this design are mounted on the wings. This puts them farther from the centerline of the aircraft than on the other two designs and thus they provide much greater moments around the center of gravity that the vertical tail must overcome.

Table 15: Initial Vertical Tail Sizing Calculations Design Vertical Tail Area Conventional Canard 99 ft2 Swept Forward Wing 176 ft2 Blended Wing 84 ft2

The second case that was needed to properly size the vertical tail was a twenty degree landing sideslip angle. For this case, the vertical tail must overcome the moments provided by the fuselage and the wings about the center of gravity due to this sideslip angle. In a similar manner to the one engine out case, we calculated the necessary vertical tail size by summing the moments around the center of gravity. Before this can be done, however, the moments needed for this case must be calculated. Raymer’s book provides equations that allow us to calculate both of the unknown moments needed for this case. Equation 20 provides an estimation of the coefficient of yaw for a fuselage surface due to sideslip. In this equation cnβf is the coefficient of yaw for a fuselage

surface due to sideslip, V is the fuselage volume, Sb is the fuselage wetted area, L is the fuselage length, Df is the fuselage parasite drag, and Wf is the fuselage weight.

1.3 Equation 20: Yaw Coefficient for a Fuselage Surface

Equation 21 provides an estimation of the coefficient of yaw for the wings due to sideslip. In this equation cnβw is the coefficient of yaw for the wings due to sideslip, cl is the

wing coefficient of lift, A is the wing aspect ratio, Λ is the wing sweep angle, xac is the position of

the wing aerodynamic center, and xcg is the position of the center of gravity of the aircraft.

1 tan sin cos 6 4 4cos 2 8cos

Equation 21: Yaw Coefficient for the Wings

Once calculated, these coefficients can be used to determine the yawing force, using the definition of these coefficients, and then sequentially the yawing moment, by multiplying the yawing force by the distance from the force location to the center of gravity. These moments can then be summed to find the necessary moment the vertical tail needs to be. From here, using the exact same steps as in the engine out case, the size of the vertical tail can be found using equations Equation 18 and Equation 19. Unfortunately, at this time, our group is having a lot of trouble trying to code this. This case is sizing our tail on the magnitude of 10‐3 ft3. This is way too small for a vertical tail to be effective, so we are assuming an error in coding. We are in the process of troubleshooting this error; however, we believe that this may not be possible until we obtain some better drag values.

SUMMARY AND NEXT STEPS

Currently, we have three viable concepts for our business aircraft proposal. Our next step would be to determine exactly which designs we wish to move forward with and begin a detailed analysis on those aircraft. This analysis will consist of refining the sizing code used to determine the empty and fuel weight of the aircraft, more accurately determining the drag over

the aircraft and pinpointing the exact location of the center of gravity and aerodynamic center using finalized component weights and locations. Once the aircraft is accurately sized, a detailed CAD model must be created to be able to visualize and verify the dimensions of the aircraft.

WORKS CITED 1. Bauer, A. B. "Simulation of Propfan Noise Impact on a Fuselage." AIAA'83. Atlanta, 1983. 6. 2. Bradley, Kevin R. A Sizing Methodology for the Conceptual Design of Blended‐Wing‐ Body Transports. Hampton: NASA, n.d. 3. "Documents for Small Businesses and Professionals." Ducted‐Prop/Propfan Technologies. March 2010 . 4. Flight International. "Civil Engines, Pratt and Whitney Gears up for the future with GTF." 2007. 2010 . 5. Garber, Donald P. and Jr., William L. Willshire. En Route Noise Levels From Propfan Test Assessment Airplane. Langley Research Center: National Aeronautics and Space Administration, 1994. 6. H.L.Kuntz, R.A.Prydz, and F.J.Balena. "Development and Testing of Cabin Sidewall Acoustic Resonators for the Reduction of Cabin Tone Levels in Propfan‐Powered Aircraft." 1991. 7. Hepperle, Martin. "MDO of Forward Swept Wings." 15 April 2008. KATnet. 26 February 2010 . 8. Joslin, Ronald D. "Aircraft Laminar Flow Control." Annual Review of Fluid Mechanics 30 (1998). 9. LeBeau, Raymond P. "Engine Selection as a Part of the Aircraft Design Proccess." 10. Liebeck, R. "Design of the Blended‐Wing‐Body Subsonic Transport." 40th AIAA Aerospace Sciences Meeting and Exhibit. Reno: American Institute ofAeronautics and Astronautics, 2002. 11. MTU Aero Engines GmbH. "PurePower PW1000G Engine Datasheet."

12. NASA. The Blended Wing Body: A Revolutionary Concept in Aircraft Design. 2009. 03 2010 . 13. NATO. "ONX Program User Guide." . 14. Neely, J. L. Taylor and M.A. "Integrated Technology Assessment Center Update." 7‐10 July 2002. February 2010 . 15. Nicolai, L. Fundamentals of Aircraft Design. 1975. 16. Norris, Guy. "Aeronautics/Propulsion Laureate; Pratt & Whitney's Geared Turbofan Development Team." AviationWeek 16 March 2009. 17. Ramjet Propulsion. 11 July 2008. February 2010 . 18. Raymer, Daniel P. Aircraft Design: A Conceptual Approach Fourth Edition. Blacksburg: American Institude of Aeronautics and Astronautics, 2006. 19. Rolls Royce. "Rolls Royce Engine Datasheets." 20. SAE International. "Aerospace Engineering Online." 2010. 2010 . 21. Saravanamuttoo, HIH, et al. Gas Turbine Theory, 6th Edition. Harlow: Pearson Education Limited, 2009. 22. Snecma. "Research and Technology at Scecma to ensure sustained development of air transportation." 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Tuscon, Arizona, 2005. AIAA 2005‐4202. 23. "The Turbojet Engine." 2008. February 2010 . 24. Thomas, Geoffrey. "Rolls‐Royce Pursues Open Rotor." November 2009. March 2010.

25. Trippensee, Gary and David Lux. "X‐29A Forward Swept Wing Flight Research Program Status." 1987. 26. Turboprop Engines. 23 December 2008. February 2010 . 27. Ullman, D.G. and B.P. Spiegel. "Trade Studies with Uncertain Information." Sixteenth Annual International Symposium of the International Council On Systems Engineering. 2006. 28. Vatandas, Erguven. "Geometrical and positional optimization of the foward swept lift producing surfaces in 3D flow domains." Aircraft Engineering and Aerospace Technology: An International Journal 2007. 29. What's a Scramjet? 22 November 2007. February 2010 .

APPENDIX A: HOUSE OF QUALITY