System Definition Review Report
Team 3
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Mission Statement To create an innovative and cost effective commercial aircraft capable of take-off and landing in extremely short distances, making it available to a larger number of runways, in order to open up more airports, primarily to relieve the continuous growing congestion of large hubs.
Mission Plans The goal of Team Arrival’s aircraft is to relieve congestion at major hubs. The three hubs that experience the most congestion (according to www.bts.gov) are Chicago O’Hare, New York LaGuardia, and Newark International. The first mission is to take off from Gary Chicago and land at Dallas Love Field. Gary Chicago airport is located 42 miles from Chicago O’Hare and Dallas Love Field is located 20 miles from Dallas International. This mission is intended to redirect some of the traffic from both Chicago O’Hare and Dallas International to secondary airports that experience less congestion. Team Arrival’s aircraft will have the capability of takeoff and landing on runways of 3000 feet or less making this mission possible. The second mission is to do a half-runway takeoff from New York LaGuardia and do a non-interfering spiral descent at Miami International. Both of the airports at LaGuardia and Miami have a significant amount of road traffic around them and it is Team Arrival’s feeling that most passengers will not want to fly into a secondary airport and then sit in street traffic for an extended period of time in order to get to their final destination. The third mission is to do a half-runway takeoff from Charlotte International, land at Essex County, then takeoff from Essex County without refueling, and return to Charlotte International and do a non-interfering spiral decent. This mission was planned as a round trip without refueling because Essex County is currently a small airport that may not have the equipment or personnel to refuel the aircraft and get it ready to takeoff again in a reasonably short amount of time. If this airport sees an increase in use in the future and increases it’s equipment and personnel this mission can be changed to refuel at Essex County. It would be more cost effective to refuel the aircraft because the aircraft would be lighter upon initial takeoff. This is hopefully something that can be done in the future, but Team Arrival is preparing for the current conditions.
Design Requirements Team Arrival’s current design requirements can be seen in Table (1) below. While much of this table is the same as what was seen in the previous review, there was a change in max takeoff weight. The max takeoff weight has now been set with a target of 100,000 pounds and a threshold of 150,000 pounds.
Mission Requirements Target Threshold Takeoff Runway Length ≤ 2500 ft 3000ft Landing Runway Length ≤ 2500 ft 3000ft Height to Passenger Door Sill at OWE ≤ 5 ft 9ft Height to Baggage Door Sill at OWE ≤ 4 ft 6ft Typical Cruise Mach Number ≥ 0.80 0.76 Range w/ Max Payload ≥ 2000 nmi 1500nmi Max Take-Off Weight ≤ 100,000 lb 150,000 lb
Max Passengers (single class) ≥ 170 pax 150pax $/seat- $/seat- Operating Cost ($US 2007) ≤ 0.08 mile 0.12mile Table (1) Current design requirements for threshold and target values
Aircraft Concept Selection
Aircraft concept selection was done using Pugh’s method. The first step of Pugh’s method was to develop the criteria that would be used to judge the concpets. The majority of the criteria that was used in Pugh’s method where developed during the Quality Function Deployment method. With the exception of rotation angle which was developed later to further evaluate and eliminate concepts. Concept design criteria where chosen roughly based on personal experience and the general requirements established by NASA. The ten criteria that were established are as follows: extremely short takeoff and landing, high lift over drag during cruise, high mach cruise number, high take off thrust, low door sill height, passenger comfort, low noise, low complexity, safety, and rotation angle. After the criteria were selected they were placed in a matrix which can be seen below in Table (2). After the matrix was formed each team member generated up to ten different concepts. From those seventy concepts; the team voted on them to get the top four designs which would move on for further clarification and evaluation. Those top four designs were the forward swept wing, high wing with engines mounted on bottom, high wing with engines mounted on top of the wing, and low mounted wing with engines mounted over the wing. The team took those four designs and compared them to the selected datum of a 737 NG+2. A comparison of two of those designs can be seen in Table (2). A simple +,-,S rating system was used to determined which concept would be selected as the team’s top choice. The top design was chosen based on the number of + and the number of – and S values. The concept with the highest number of (+) and the lowest number of (-) was determined to be the teams winning concept. The idea that was the best of the concepts was the low forward swept wing with blown flaps. Sketches of the major concepts that were considered are available in Appendix (A).
Low Aft Low forward swept wing w/blown Criterion swept wing flaps Low straight wing w/plasma ESTOL + + High L/D (cruise) + S High M (cruise) S S High TO Thrust S S Passenger Comfort S S Low Door Sill Height + - Low Noise S - Low Complexity - S
Low Weight DATUM (737 NG+2) - S Safety + S Rotation Angle + +
+ 5 3 S No 3 6 - 3 2 Check 11 11 Pugh's Survivor? Yes Yes Table (2) Pugh’s method comparison
The initial winning concept was then evaluated to determine what areas required improvement. After the problem areas were identified, an attempt was made to combine the positive aspects of the selected concepts in order to improve on the weaknesses of other concepts. If those solutions did not work the team tried to think of new ideas with which to solve the problems. This resulted in a hybrid of the original winning concept. Once this process was completed it was once again ran through the matrix using the original winning concept as the datum and reexamining where the flaws lied in the new design. This lead the team to the winning concept of the low forward swept wing aircraft with blown flaps and plasma generators. The final design is shown in Figure (1).
Figure (1) Current aircraft configuration
Future work needs to be done on this design to determine the exact placement of the control surfaces and the need of the designed lifting surfaces. Also more work needs to be done in the aerodynamics and control of the aircraft. The canard will most likely be moved up and behind the main cabin door so that it will not interfere with ground operations. The overall need for the canard is still being assessed. If the desired rotation angle can be achieved by using a T-tail then the canard may be replaced.
Cabin/Fuselage Layout For the interior layout of the aircraft, Arrival plans on offering a two class and all economy class version of the plane. For the all economy version the planned layout includes 29 rows with 3 seats on each side of the aisle for a total of 174 passengers. The two class version will feature 4 rows of first class with 2 seats on each side of the row, and 24 rows of economy for a total of 160 passengers. For the first class section standard seat pitch will be 37 inches with a seat width of 20 inches. For the economy section, standard seat pitch will be 32 inches with variations around emergency exit locations, and a seat width of 18 inches. The preliminary interior layouts are pictured below; Figure (2) shows the economy only layout, and Figure (3) shows the two class arrangement.
Figure (2) Preliminary economy lay-out of cabin
Figure (3) Preliminary two class lay-out of cabin
Currently there are doors on both sides of the plane in the open space next to the fore and aft lavatories. There will be two more emergency exits on each side of the plane added in the cabin either over the wing, or in the middle of the fuselage depending on how far aft the wing is in the final design. Alterations will then be made in the seat pitch of the surrounding rows to accommodate these emergency exits. These two layouts were designed in CATIA to get an estimate of how long the cabin and fuselage would need to be. The current estimate for cabin size is 96 ft, and the fuselage is 128 ft long. The current cabin layouts will probably be altered in the final design to improve lavatory and galley placement, especially in the two class configuration. Another reason to alter the layout is to move the doors into a section of the fuselage that isn’t tapering, as recommended by Boeing. The final cabin configuration will also include detail on overhead bins, windows, lavatory interiors, and the cockpit configuration.
Constraint Analysis In order to perform the constraint analysis for the ESTOL aircraft, Arrival begun by using the mission requirements and generating other performance criteria with which to constrain the aircraft. The following quantities were considered when constraining the aircraft: take-off and landing ground roll, aspect ratio, Oswald’s efficiency factor, cruise Mach number, cruise altitude, service ceiling, parasite drag, lift-to-drag ratio, and maximum lift coefficient. Table (3) contains the values for which the constraint diagram was generated. These values were assumed from historical data, generated from technologies that are expected to exist by 2050, or generated by finding a trend over time and extrapolating that trend out to 2050. Values for Oswald’s Efficiency Factor (e) and parasite drag (CD0) were assumed to be typical for commercial transports. A reasonable aspect ratio of 10 was chosen for this simple analysis, but will be refined using carpet plots. The selection of CL,max, L/D, and We/W0 are discussed in the sizing studies section of this report. Other values used in constraint generation were determined from the system requirements. With the feedback from SDR, (L/D)max will likely increase to around 28, and this will lead to an increase of (L/D)cruise; however, this report does not reflect this change.
Parameter Value Description AR 10 Aspect Ratio e 0.8 Oswald's Efficiency
CD0 0.015 Parasite Drag
(L/D)max 23 Maximum Lift-to-Drag Ratio
(L/D)cruise 20 Lift-to-Drag in Cruise
(L/D)ssc 10 Lift-to-Drag during Second Segment Climb
STO1 1500 Take-Off Ground Roll for Constraint Set 1 (ft)
STO2 2500 Take-Off Ground Roll for Constraint Set 2 (ft)
SL1 500 Landing Ground Roll for Constraint Set 1 (ft)
SL2 800 Landing Ground Roll for Constraint Set 2 (ft) Table (3) Basis for constraint diagram
Two sets of constraints were maintained in this analysis, one for a 1500ft landing ground roll and one for a 2500ft landing ground roll. Each of these two sets of constraints used the same values for the other quantities listed in Table (3). In order to estimate the necessary thrust-to-weight and wing-loading for our aircraft, constraint diagrams were generated based on the system requirements and approximations of the typical physical limits by which aircraft are constrained. Constraints were computed for each of six criteria: 1. 2. Top-of-Climb 3. 4. 5. T 6.
Figure (4) Constraint diagram for 1500 ft take-off ground roll case
Figure (4) is the constraint diagram for the 1500 ft take-off ground roll case. In this diagram, a wing-loading of approximately 74 psf and a thrust-to-weight ratio of 0.283 is necessary to meet the desired performance criteria – this point is marked on the graph.
Figure (5) Constraint diagram for the 2500 ft take-off ground roll case
Figure (5) is the constraint diagram for the 2500 ft take-off ground roll case. For this set of constraints, a wing loading of approximately 131 psf and a thrust-to-weight ratio of 0.30 are necessary to meet these performance criteria. As is evident in both constraint diagrams, the major constraining factors for this aircraft are the extremely short take-off and the ability to perform a 1.5g maneuver in cruise at cruise altitude. It was expected that the second segment climb would have constrained the aircraft; this was not the case, however, the second segment climb constraint’s importance is second only to the requirement of performing a 1.5g maneuver at cruise altitude. The approximations used to determine the constraints for the aircraft are given as equations below. Equation (1) was used to determine the constraint lines for Constraints 1-3. Equations (2), (3) and (4) were used to determine the constraint lines for Constraints 4, 5, and 6 respectively.
⎧ ⎡ 2 ⎤ ⎫ T β ⎪ q CD 1 ⎛ηβ ⎞ ⎛W ⎞ 1 dh 1 dV ⎪ SL = ⎢ 0 + ⎜ ⎟ ⎜ 0 ⎟⎥ + + (1) ⎨ ⎢W ⎜ ⎟ ⎥ ⎬ W0 α ⎪β 0 πARe ⎝ q ⎠ ⎝ S ⎠ V dt g dt ⎪ ⎩ ⎣⎢ S ⎦⎥ ⎭
W0 S L ρCL,max gμ = 2 (2) S ()Vland Vstall β W V V 2 β 0 T ()takeoff stall SL = S (3) W0 ρCL,maxαSTO T ⎛ number.of .engines ⎞⎛ 1 ⎞ SL ⎜ ⎟ = ⎜ ⎟⎜0.024 + ⎟ (4) W0 ⎝ number.of .engines −1⎠⎝ ()L D ssc ⎠
Sizing Studies
Team Arrival’s sizing approach used techniques outlined in Daniel Raymer’s text, “Aircraft Design: A Conceptual Approach”. These techniques were applied in MATLAB scripts and used to arrive at initial sizing figures for Team Arrival’s aircraft. Using values for wing loading and thrust-to-weight ratio for 1500 ft and 2500 ft ground rolls obtained from our constraint diagrams, ranges from our target and threshold maximum range values, as well as ranges for all of our missions, it was possible to obtain initial sizing output from MATLAB and then sized our aircraft to the “worst case” mission output. The fixed range and payload inputs allowed for sizing based off of weight, which is easier to work with in order to obtain necessary wing-span and control surface sizing given required the thrust-to-weight and wing loading ratios. MATLAB scripts allowed Team Arrival the flexibility to apply various technology and weight saving factors where appropriate to obtain as representative an approximation for weight and size values as possible. For instance, advanced technology engines such as the Geared Turbofan (GTF) are projected to have a specific fuel consumption (SFC) as low as 0.36 according to optimistic fuel burn values from Bombardier. SFC values factor in greatly to calculating the needed fuel capacity for our aircraft. Fuel capacity will also have a great impact on Take-Off Gross Weight (TOGW). Advanced materials, such as Carbon Fiber Reinforced Plastics (CFRP) were also factored in. Using the data of empty weight fraction for the CFRP 787 Dreamliner versus other contemporary wide-bodies such as the 777 family and A330 series, and applying it to today’s 737NG and A320 family, Team Arrival predicted a weight savings of 15% over today’s narrow-body jets. This allowed us to reduce the original calculated TOGW using Raymer’s methods by 15% to reflect the weight savings from advanced materials. Further weight savings from CFRP may be factored in to reflect a higher percentage use of composite materials that Team Arrival feels will very likely be possible in the next 50 years. The sizing approach started by taking a look at the three design mission scenarios that Team Arrival specified as typical uses for the concept aircraft. Table (4) shows the suggested city pairs and the ranges between them.
Table (4) 3 Mission Scenarios for Team Arrival
These ranges were used as inputs parameters for the initial sizing. The initial sizing methods called for the concept aircraft to be capable of carrying 170 passengers plus two crew members for these distances. The most difficult mission is the Charlotte to Newark mission which calls for the aircraft to not be refueled, rather it will drop off its passengers and pick up new ones and return to Charlotte, giving a 920 nm round trip. This had the effect on sizing of essentially adding another set of weight fractions to the sizing code. In addition to the three mission scenarios listed in Table (4), the target maximum range of 2000 nm and the threshold maximum range of 1500 nm were used as inputs for a generic mission scenario. These range values along with an SFC of 0.36, a cruise L/D of 23 based off of the 50 year trend, a cruise Mach of 0.78 at an altitude of 35,000 ft Mean Sea Level (MSL), and a 170 passenger payload capacity formed the basis for Team Arrival’s sizing. Additional factors such as a 100 nm alternate airport range and a 45 minute fuel reserve mandated by the FAA were worked into the sizing calculations.
L/D by Year
30
25
20
15 L/D
10
5
0 1920 1940 1960 1980 2000 2020 2040 2060 2080 Year
Figure (6) Plot of projected L/D from historical trends
The initial sizing split each of the missions into phases of flight which had their own specified or calculated weight fractions. For the maximum range, Chicago to Dallas, and New York to Miami missions, the flights were divided into takeoff phase, climb phase, cruise phase, holding phase, and landing. For Charlotte to Newark, these phases are repeated. The takeoff, climb, and landing weight fractions were obtained from historical values laid out in the Raymer text. The cruise weight fraction factored in mission range, aircraft SFC, cruise velocity, and cruise L/D. The holding phase also factored in mission range and aircraft SFC, but used maximum L/D instead of cruise velocity and cruise L/D. Once all of the weight fractions for each phase were calculated, they were all multiplied together into a total weight fraction value. This value was subtracted from one and multiplied by a 1% trapped fuel factor to find the fuel weight fraction using Equation B1. Using an input TOGW estimate and applying the 15% weight savings of CFRP materials, an empty weight fraction (Equation 5) was found which with the payload weight and fuel weight fraction were used to find the design TOGW using Equation B2. Design TOGW, the design empty weight fraction found using the obtained design TOGW, and the fuel weight fraction were then used to find the empty weight (Equation B3) of the aircraft and the amount of fuel necessary (Equation B4) for the aircraft to complete the mission scenarios. This process was iterated until the design TOGW matched the estimated TOGW.
(−0.0434) WeWo = 0.85 * (0.9457 *Wo ) (5)
The initial sizing results for the aircraft obtained using Raymer’s methods and the weight-sizing MATLAB code are tabulated weight according to range, airframe size based off of weight and dictated by ground roll, and engine thrust requirements based off of weight and dictated by ground roll. Table (5) shows the initial weight sizing results for the 2000 nm target range while Table (6) shows the initial weight sizing results for the 1500 nm threshold range.
Preliminary Weight Sizing Results (2000 nm Range) Parameter Sizing Value TOGW [lbs] 100,000 We [lbs] 50,000 We/Wo 0.49 Fuel Weight [lbs] 13,000 Table (5) Weight Sizing for 2000 nm Range
Preliminary Weight Sizing Results (1500 nm Range) Parameter Sizing Value TOGW [lbs] 95,000 We [lbs] 50,000 We/Wo 0.49 Fuel Weight [lbs] 11,000 Table (6) Weight Sizing for 2000 nm Range
The planned use of lightweight materials such as CFRP has thus far been able to reduce the weight of Arrival’s aircraft to 50,000 lbs empty. Such reductions in weight will be critical in developing an ESTOL aircraft capable of carrying the required number of passengers and payload. Currently, Arrival is looking at the weights for the target 2000 nm range for the aircraft as these values do a better job of encompassing the “worst case” values calculated for the Charlotte to Newark mission which calls for a round trip without refueling. The engines are currently “rubber engines” sized to the requirements and constraints of the aircraft as well as TOGW. Table (7) shows the initial engine sizing results based off of the weight of the aircraft and the desired thrust-to-weight ratios for the 1500 ft and 2500 ft ground rolls. The thrust was then sized up to ensure that the values could meet the requirements of all mission scenarios, including the worst case mission using Equation B5.
Preliminary Engine Sizing Results (Both Ranges) Parameter Sizing Value Thrust/Engine 1500 ft Ground Roll [lbf] 14,000 Total Thrust 1500 ft Ground Roll [lbf] 28,000 Thrust/Engine 2500 ft Ground Roll [lbf] 15,100 Total Thrust 2500 ft Ground Roll [lbf] 30,200 Table (7) Weight Sizing for 2000 nm Range
The sizing for the wings will be based off of the weight of the aircraft over the required wing loading values for the 1500 ft and 2500 ft ground rolls. The initial wing dimensions were obtained using basic Equations B6, B7, B8. The sizing of the aircraft vertical stabilizer and canards will then be based off of the dimensions of the wing, volume coefficients obtained from Raymer Table 6.4, and the preliminary lengths of the moment arms relative to the wing quarter chord using Equations B9 and B10. Table (8) shows the initial airframe sizing results based off of weight.
Preliminary Airframe Sizing Results (Weight Based) Parameter Sizing Value Est. Wing Span 1500 ft Ground Roll [ft] 113 Est. Wing Span 2500 ft Ground Roll [ft] 84 Est. Wing Area 1500 ft Ground Roll [sq ft] 1,270 Est. Wing Area 2500 ft Ground Roll [sq ft] 703 Fuselage Length (Weight Based) [ft] 106 V. Stab. Area 1500 ft Ground Roll [sq ft] 251 V. Stab. Area 2500 ft Ground Roll [sq ft] 103 Canard Area 1500 ft Ground Roll [sq ft] 202 Canard Area 2500 ft Ground Roll [sq ft] 83 Table (8) Weight Sizing for 2000 nm Range
It should be noted that while a fuselage length of 106 ft is specified in Table (8), this value is based completely off of design TOGW (Equation B11). The fuselage will likely be a larger size for optimum passenger comfort and optimum efficiency while keeping with required Federal Aviation Regulations for number of exits, cabin lighting, and aisle width. Initially, fuselage sizing coefficients found on Raymer Table 6.3 were used to find the weight based fuselage length.
Advanced Technologies
• Forward-Swept Wing There are various advantages we can obtain from using forward-swept wings. The first advantage will be improved leading-edge sweep/shock sweep relationship. A typical transport aircraft, which cruises at a high subsonic Mach number, encounters extensive regions of supercritical flow on the wing upper surface. Recent design experience on supercritical wings demonstrates that the effective aerodynamic wing sweep is the shock sweep. The current design practice for transport wings is to maintain a shock well aft on the wing, swept at a constant percent chord in order to minimize compressibility drag.
Figure (7) Leading-Edge/Shock-Sweep Relationship
As shown in Figure (7), for equal leading-edge sweeps, the shock sweep on a tapered forward swept wing will be appreciably higher than on an aft swept wing (ASW). The shock is assumed to be fixed at the 70 percent chord location by proper wing twist and camber. For a transport aircraft, the leading edge sweep/shock sweep advantage might be exploited to achieve higher drag divergence Mach numbers, to reduce the wing sweep, or to increase the wing thickness-to-chord ratio. The higher drag divergence Mach numbers would lower operating costs by increasing fuel efficiency. An increased wing thickness-to-chord ratio would increase fuel volume and result in a wing weight reduction which would in turn lower acquisition costs.
The 2nd advantage will be provided by using a canard trim surface on a forward swept wing transport. Preliminary aerodynamic investigations conducted at Lockheed indicate the attractiveness of combining a canard with a forward swept wing. On a forward swept wing, the wing root tends to be highly loaded relative to the tip, whereas the opposite is true for aft sweep. For an aft swept wing, the canard downwash on the wing root region increases the washout twist requirement. However, for the forward swept wing, the canard downwash impinges the highly loaded root region. This should permit a reduction in wing twist and/or an increase in thickness at the root section. A canard arrangement also produces more lift than a conventional set-up when total lift produced is considered. Because the canard generates upward lift, unlike with a tail plane which produces downward or negative lift, there is a reduction in the lift required from the main wing. This reduction in the required lift generation by the wing to over come the weight of the aircraft a reduction in lift-induced drag by the wing. As well as removal of the negative lift generated by the tail plane and the associated lift-induced drag. Overall drag and lift requirements of the aircraft is reduced. Another advantage of FSW will be the improved low speed characteristics. Improvements in low speed handling qualities and in low speed maximum lift coefficient are possible for a forward swept wing/canard configuration. On a forward swept wing, flow separation and stall tend to occur initially near the wing root, in contrast to an aft swept wing where the tip tends to stall initially. Thus, tip devices on a forward swept wing provide effective roll control to higher lift coefficients. An improvement in the low speed trimmed maximum lift coefficient will further enhance the low speed performance. However, there is a salient problem when using FSW which is the aeroelastic divergence occurring at lower mach number than aft swept wing. A reduction in divergence speed on the order of 90%-can be expected when a wing is-swept-forward from 0 to 28 degrees. Although it will remain a fact that the divergence dynamic pressure of swept-forward designs will always be lower than their unswept counterparts, the important consideration that designers face is associated with the amount of additional structural weight needed for the increased stiffness required t o insure the absence of divergence within the operating performance envelope of the aircraft . This weight penalty is known to be very severe for conventional metal wings. The unique properties of advanced composites, the basic material properties can be tailored to suit a particular loading condition. This additional design parameter is the key to controlling and effectively adjusting the wing stiffness characteristics to combat the weight problem caused by the lower divergence speeds associated with swept- forward wings. FSW has been studied and developed since the 40’s and has been exhibited in a number of aircrafts both in military uses and civilian uses, such as Grumman X-29, Sukhoi Su-47, and HFB-320 Hansa Jet. By the time of 2058, we are expecting the FSW concept to have a NASA Technical Readiness Level (TRL) of 9.
• Composites Weight Savings The Boeing 787's all-composite fuselage makes it the first composite airliner in production. While the Boeing 777 contains 50% aluminum and 12% composites, the numbers for the new airplane are 15% aluminum, 50% composite (mostly carbon fiber reinforced plastic) and 12% titanium. Each fuselage barrel will be manufactured in one piece, and the barrel sections joined end to end to form the fuselage. This will eliminate the need for about 50,000 fasteners used in conventional airplane building. The superior strength of the composite fuselage will allow higher pressurization in the passenger cabin, making it easier to control temperature, humidity and ventilation. Composite materials are also more durable than aluminum, because of corrosion and fatigue benefits, as well as a dramatic reduction in fasteners. There are various parts that can be made of composites such as airframes (flight control surfaces, panels, doors, fairings, etc.), nacelles (pylons, cowlings, acoustic panels, cowl doors, etc.), engine components (bonded honeycomb acoustical, ducts, fairings, liners, etc.), thrust reversers (thrust reverser, translating cowls/sleeves, support assemblies, blocker doors), fuselage/wing panels (skin panels, access panels, internal panels, fuel panels, etc).
Figure (8). Capabilities of composite material.
By comparing the empty weight fraction of Boeing’s 787 to similar size aircrafts such as Boeing 777 and Airbus A330, we have a trend of weight savings shown in Figures (9) and (10). Following this trend we have estimated that our aircraft could have about 15% weight savings factor on the empty weight compared to current similar size aircrafts. This weight saving can also be the solution to forward-swept wing’s divergence phenomenon.
Empty Weight Fraction Comparison
0.7
0.6 -0.5221 y = 3853.1x-0.6833 y = 549.16x 0.5 Airbus A330 0.4 Boeing 777 Boeing 787 0.3 Power (Boeing 787) Power (Boeing 777) 0.2 Empty Weight Fractio
0.1
0 0 200000 400000 600000 800000 1000000 TOGW [lb]
Figure (9). Empty Weight Fraction Comparison
Empty Weight Fraction Material Comparison
0.70 y = 233.08x-0.5017 0.60 y = 851.58x-0.6321 Boeing/Airbus Current 0.50 Boeing/Airbus CFRP 0.40
0.30 Power (Boeing/Airbus Current) 0.20 Power (Boeing/Airbus CFRP) Empty Weight Fraction Empty Weight 0.10
0.00 0 50000 10000 15000 20000 25000 0 0 0 0 TOGW [lb]
Figure (10). Empty Weight Fraction Material Comparison
Due to the benefits of composite materials, composites are worth the effort and are clearly on a fast track in the aircraft industry. We expect the use of composites will have NASA TRL of 9 as well as FSW by 2058.
• Plasma Actuators Plasma actuators consist of two electrodes that are separated by a dielectric material. One of the electrodes is typically exposed to the air. The other electrode is fully covered by the dielectric material. A schematic illustration is shown in Figure (11). A high voltage alternating current potential is supplied to the electrodes. When the alternating current amplitude is large enough, the air ionizes in the region of the largest electric potential. This generally begins at the edge of the electrode that is exposed to the air, and spreads out over the area projected by the covered electrode.
Figure (11) Schematic drawing of asymmetric electrode arrangement for plasma actuators used in experiments.
The plasma actuators have been successfully used in numerous flow control applications. These included 3-D boundary layer instabilities on a sharp cone at Mach 3.5, lift augmentation on wings, separation control for low-pressure turbine blades, leading-edge separation control on wing sections, and control of the dynamic stall vortex on oscillating airfoils. An example of leading-edge separation control on a NACA 0015 airfoil is shown in Figure (12), this shows the lift coefficient versus angle of attack for the airfoil with the leading-edge plasma actuator off and on in” steady” operation. The measurements are shown as the symbols. The lift forces were measured by a force balance. The curves correspond to numerical simulations using a modified version of CFL3D .16
Figure (12) Lift coefficient versus angle of attack for NACA 0015 airfoil with leading-edge actuator off and on (steady
With the actuator off (square symbols), the lift increases linearly up to the stall angle, which is approximately = 14 degrees. The solid curve is the numerical prediction without the actuator. Stall corresponds to full leading-edge separation, with a separation bubble that covers the total suction surface of the airfoil. When the actuator is on in “steady” operation, the stall angle increases to 18 degrees. With plasma actuators’ ability to delay separation over the airfoil and in turn ability increase stall angle of attack, the primary advantage this technology offers is decreased takeoff distance. Because the stall angle is increased, the takeoff angle of attack can be increased further which will get the plane off the runway quicker. It should be noted however, that the body of the plane will still be the limiting factor in takeoff angle of attack. This is because no matter how steep the stall angle, it is unacceptable to strike the tail on the runway during takeoff. Therefore, the second advantage of the plasma actuators is that the increased stall angle will allow for optimum placement of the wings and thus landing gear can also be placed to maximize takeoff angle of attack. While the advantages of this technology are obvious, it is important to recognize that that the TRL of plasma actuators is currently 6. This is because the plasma actuators have are currently being used on sailplanes; this qualifies them for relevant environment use. Based on the fact that this technology is currently in successful use on functioning aircraft, it is the expectation of Team Arrival that 30 years from now the TRL of plasma actuators will be 9 and ready for use on planes entering service. However, it should be noted that at this time, max stall angle and Cl calculations do not take into account the effects of plasma actuators.
• Upper Surface Blowing Upper Surface Blowing (USB) is another technology that is currently being examined for use on the aircraft developed by Team Arrival. As can be seen in Figure (13) below, the primary difference in USB is that the engines are mounted above the wings as opposed to the traditional mounting of wings below the wings.
Figure (13) Depiction of Upper Surface Blowing
In the early 1970’s, Langley research center conducted an experiment on upper wing surface mounted engines and the results of this experiment showed that the incremental lift provided by thrust vectoring of lower-surface engines was limited to the vector component of thrust with no appreciable induced circulation for the particular configuration tested. However, significant additional circulation lift was produced by upper-surface blowing obtained by deflecting the exhaust of upper surface mounted engines down on to the wing surface. The additional lift created from these upper surface mounted engines generated a large buzz in the aircraft industry because of the possibility of allowing short take and landing. In the 1970’s this technology was put into place on the first models both the Boeing YC-14 and the Antonov An-72. Since that time, upper surface blowing has been in limited use on several different aircraft because of its ability to significantly decrease takeoff and landing distances and to increase cruise and max Cl. Currently, the concept developed by Team Arrival utilizes USB and its affects are taken into consideration in the calculation of Cl max and also the approximated takeoff and landing distances. Because this technology has been in use since the mid 1970’s the TRL is considered to be 9.
• Specific Fuel Consumption With the consistent increase in fuel prices and the expectation that this trend will continue, one vital portion of success in the aircraft industry is specific fuel consumption. As it is the current projection that geared turbofan engines will be used an unducted turbofan engines will possibly be used, Team Arrival has investigated the research done by Bombardier on the SFC of such engines which can be seen in Table (9) below.
Power Series Projection UDF SFC UDF SFC GTF SFC GTF SFC Certification (15% (25% (15% (20% Year SFC savings) Savings) Savings) Savings) 2010 0.59 0.51 0.45 0.51 0.48 2015 0.57 0.49 0.43 0.49 0.46 2020 0.56 0.47 0.42 0.47 0.44 2025 0.54 0.46 0.40 0.46 0.43 2030 0.52 0.44 0.39 0.44 0.42 2035 0.50 0.43 0.38 0.43 0.40 2040 0.49 0.41 0.36 0.41 0.39 2045 0.47 0.40 0.35 0.40 0.38 2050 0.46 0.39 0.34 0.39 0.36 Table (9) Power series projection for SFC value
In addition to the research done by Bombardier, in order to develop an accurate assumption as to the aircrafts SFC, Team arrival has also compiled data on SFC of current aircraft fleets as well as SFC data from aircraft that are exiting service to determine the trend in SFC over the next 50 years. The results of this process can be seen in Figure (14) below.
SFC vs. Certification Date
0.7
0.6
0.5 y = 2E+44x‐13.476
0.4
SFC SFC 0.3 Power (SFC) 0.2
0.1
0 1960 1980 2000 2020 2040 2060 2080 Certification Date
Figure (14) Plot used to determine SFC value
Based on the precision between both of these SFC projections, Team Arrival expects that a geared turbofan in service in the year 2058 will have an SFC of 0.36. This will not only significantly decrease operating costs, but will also increase possible mission range for the aircraft.
Summary
Team Arrival has been able to specify the specific design requirments needed to meet our mission cases and overall objectives of being able to service secondary airports. This has allowed the team to consider several concepts and ultimately determine a preliminary design for our aircraft. Two different layouts for the cabin were generated and are currently being updated in order to properly fit with the exterior design. A full constraint analysis was completed determining specific restrictions of the aircraft’s design and performance. Baseline dimensions and weights were determined through continuing sizing studies. More research was done on possible advanced technologies giving projected values for the year 2045 for such parameters as L/D and SFC. The advantages and disadvantages of both forward swept wings and plasma generators were also investigated. There is still a lot of work that needs to be done on the aircraft, but the current values being achieved fit within the compliance matrix which can be seen in Table (10) below. This gives Team Arrival a good foundation to build on.
Mission Requirements Target Threshold Current Takeoff Runway Length ≤ 2500 ft 3000ft 2500 ft Landing Runway Length ≤ 2500 ft 3000ft 900ft Height to Passenger Door Sill at OWE ≤ 5ft 9ft 8ft Height to Baggage Door Sill at OWE ≤ 4ft 6ft 6ft Typical Cruise Mach Number ≥ 0.8 0.76 0.78 Range w/ Max Payload ≥ 2000nmi 1500nmi 2000nmi Max Take-Off Weight ≤ 100,000lb 150,000 lb 100,000lb Max Passengers (single class) ≥ 170pax 150pax 170pax Operating Cost ($US 2007) ≤ 0.08$/ASM 0.12$/ASM 0.05$/ASM Table (10) Compliance matrix
Next Steps
Team Arrival next objectives include finishing the quantifying of advanced concepts in order to factor in the entire benefit that advanced technologies will be giving. We want to complete the aircraft sizing in order to finalize the weight and dimensions of the aircraft. We would like to develop the final design details and finalize the performance characteristics. We would like to get a final estimation of the total cost of the aircraft including cost per seat mile. We would also like to determine the environmental impact the aircraft will have as far as it’s level of pollution compared to other existing and future aircrafts. We would also like to determine the component weight breakdown in order to see where the weight is distributed on the plane and to make sure that our center of gravity and aerodynamic center are located in appropriate places.
Appendix A
Sketches of major concepts considered for Pugh’s method
Appendix B
Sizing study equations
FuelWeightFraction =1.01(1 − TotalWeightFraction) (B1)
(Wcrew + W payload ) Wo = (B2) ()1 − ()W f Wo − ()WeWo
We = WeWo *Wo (B3)
W f = W f Wo *Wo (B4)
W *T _W T = o (B5) # Engines
W S = o (B6) W _ S
b = S * AR (B7)
S c = (B8) b
cvt * b * S wimg S vt = (B9) Lvt
ccan * b * S wing S can = (B10) Lcan
c L fuse = a *Wo (B11)
Symbol Discription Units Wo Max takeoff weight lbs Wcrew Weight of crew lbs Wpayload Weight of payload lbs Wf Weight of fuel lbs We Empty weight lbs T Thrust lbf T_W Thrust-to-weight ratio N/A S Wing area ft2 W_S Wing loading psf b Wing span ft AR Aspect ratio N/A c Wing chord ft cw Vertical tail volume coeff N/A Lvt Vertical tail moment arm ft 2 Swing Wing area ft 2 Scan Canard area ft Lcan Canard moment arm ft ccan Canard volume coeff N/A 2 Svt Vertical tail area ft a Fuselage sizing coeff N/A Lfuse Length of fuselage ft C Fuselage sizing coeff N/A