Summary

The following report is a detailed yet concise description of a design for the construction of a supersonic transport vehicle. The report begins with an introduction of the vehicle, motive for the design, and basic characteristics that make the design distinctive.

Following the introduction is a section outlining the Mission specifications and giving the account of a typical mission that the aircraft should undergo. Next calculations for the Take Off Weight are displayed, with assumptions and preliminary estimates of various values mentioned. Wing loading, area, and aspect ratio are also exhibited with details and explanations of how we arrived at the values.

The specific engine our team selected to power our aircraft is declared in the next section. Elucidation of the choice and some important information about the engines is written.

Next is a section detailing the aerodynamics and performance of the aircraft. Properties include speed for minimum drag, stalling speed, required thrust, and takeoff field length information. Graphs are plotted to show relation of various values and for use in calculations.

Fuel weight and volume is given with calculations as well as the range of the aircraft at full fuel load. A refined Estimate of Structural Weight available for building the aircraft is specified and charts are used to reveal weight distribution. Next wing shape, fuselage dimensions, and a detailed description of landing gear dimensions and placement are featured. The report is finished off with a three-view illustration of the aircraft.

The report largely endeavors to illustrate how the design meets the stated requirements and is able to achieve a typical mission successfully.

Team summary

Sridhar Easwar

Adam Plondke

Tim Lites

Adam made several initial assumptions and comparisons using the research of similar aircraft to sort of kick start the project. His initial work provided some framework to work with in getting together initial values. Siddharth expanded on these to come up with several calculations and values. All team members met to discuss and work on calculating various values for aerodynamic performance and properties of the aircraft many times. Eventually the team delegated certain responsibilities to each team member. Each person was responsible for making sure the calculations for his set of properties of the aircraft were correct, double checked, and put in final form. All team members contributed to all of these at some point, but individuals were responsible for these parts on the final report. It broke down more or less like this:

Tim –

Basic specifications such as;

Take Off Weight

Weight fractions

Payload

Wing loading, area, aspect ratio

Mission Specification and Typical Mission Profile

Siddharth –

Fuel weight and volume

Range at full fuel load

Structural weight

Wing geometry

Fuselage Dimensions

Landing gear dimensions and placement

3-view of the aircraft

Adam Plondke -

Speed for minimum drag

Stalling speed

Thrust

Takeoff field

Selection of engines and justification

Graphs and charts, empirical data for assumed values

Adam put together word files for the final report. Tim proof read, checked over, and further organized/compiled the report. Tim also wrote the summary and introduction for the final report. Although it was hard to get three people together sometimes due to different schedules, the team found many times it could meet or two could meet to work on the project. No one missed a meeting that they said they could come to. All team members feel very good about the contribution of the others to the project.

Contents

Summary-------------------------------------------page1

Team Summary-----------------------------------2

Contents--------------------------------------------3

I Introduction----------------------------------------4

II Mission Specifications---------------------------5

III Typical mission profile--------------------------5

IV Basic specification of our SST-----------------6

V Assumptions made-------------------------------7

VI Selections of engines----------------------------8

VII Speed for minimum drag------------------------9

VIII Stall speed-----------------------------------------11

IX Thrust required and thrust available-----------12

X Graph of thrust vs. speed------------------------13

XI Graph of thrust vs. altitude----------------------14

XII Takeoff field length------------------------------15

XIII Graph of runway length vs. altitude------------16

XIII Fuel Weight and Volume------------------------17

XIV Refined value for the structural weight--------18

XV Wing Shape----------------------------------------19

XVI Fuselage dimensions------------------------------20

XVII Landing gear and placement---------------------20

XVIII Multiviews of the aircraft------------------------21

XIX Addendum------------------------------------------21

XXI Calculations-----------------------------------------22

XXII Sources used----------------------------------------23

Introduction:

This is a conceptual design for a supersonic transport vehicle. Our team chose to develop this design after reasoning that the only Super Sonic Transport (SST) in operation, the Concorde, was manufactured over two decades ago and recent developments by the aviation industry in the fast paced era in which we live warrant something newer and better. Cutting edge technology allows us to ameliorate previous designs and build a modern SST that is faster and more efficient. We are able to design an aircraft that can carry up to 300 passengers, three times as much as the Concorde, and even more importantly, a supersonic airliner that is affordable for the average person. Being that the design is intended for commercial use, concerns such as airport accessibility and safety have also carried over into the final design.

Mission Specifications

The Mission necessitates the following specifications:

Ø Aircraft must be able to carry 300 passengers, their luggage, and their supplies.

Ø Aircraft must have a range of 8000+ miles.

Ø Aircraft must be able to cruise at Mach 3 over an extended period of time.

In addition the following constraints have been imposed:

Ø Aircraft must be able to take off with three engines.

Ø Aircraft must be able to survive an engine failure anywhere

Ø Enough fuel reserve to be diverted to an airport 300 miles away from last stage of the landing approach, including a 1 hour loiter

A typical mission profile:

Our Aircraft will take off with a maximum payload and full fuel tank from Los Angeles airport and climb to 60,000 feet. The aircraft will then cruise at Mach 3 for a distance of 7,900 miles. When approaching the runway at Melbourne, Australia the aircraft will loiter for an hour in the air before landing. At this point the SST must have enough fuel to be diverted to an airport 300 miles away in the case of an emergency.

Basic specification of SST

To estimate take Off Weight we first had to find the payload and the payload fraction. Referring to the mission specifications it is evident that the payload carried by the aircraft will be the combined weight of 300 passengers plus their luggage plus 5 tons of cargo. In addition, it must carry the crew, which will consist of 12 members. We assumed that the average mass of each passenger would be 117.2 Kg. The average weight of luggage per passenger was assumed to be equal to 31.8 kg. The mass of supplies provided for per passenger (food, drink, etc.) was calculated to be 3.8 Kg. The total payload is 408029.5 N. The calculation is shown as follows:

(Weight of passengers + weight of cargo) * (300 + 12) = 408029.5 N

5 tons of cargo = 49,000 N

359,029.5 N

+ 49,000.0 N

408,029.5 N

The payload fraction was assumed to be .1 from comparisons to similar aircraft.

Now to find Take Off Weight we divided the payload by the payload fraction:

TOW = Payload = 3,550,213 N

Payload Fraction

For Wing Area we chose a value of 9,000 ft² (837.29 m²) based on comparison of similar aircraft and what is needed to provide the required lift. The Wing Loading is calculated by dividing the aircraft weight by the surface area of the wing.

W = 3,550,213 N = 4240.12 N/m²

S 837.29 m²

We assumed a suitable wingspan to be 51.83 m, based on comparison to similar aircraft. To calculate Aspect Ratio we divided the square of the wingspan by the surface area of the wing:

b² = (51.83 m)² = 3.208

S 837.29 m²

The engine we selected is the JSF-119, which had a thrust of 40000 lbs per engine.

Assumptions Made

The following assumptions were made while we designed our aircraft. They are as follows –

• The payload fraction was taken as .1

• The wing area was taken as 9000 square feet.

• The number of crew members was taken as 12

• The wingspan was taken 170 feet.

• The cruise altitude was taken as 60,000 feet.

• The value of Cl_max at take-off was taken as 1.6

• The value for co-efficient of parasitic drag was taken as 0.011

• The value for TSFC was taken as 0.52

• The span wise efficiency factor was taken as 0.95

• Net Thrust at takeoff was assumed as 0.3 TOW.

• The values for density and altitude were taken from tables

Note: Almost all the values were based on data acquired from similar aircraft and/or extrapolations of graphs.

Selection of engines

Our SST requires engines that are capable of flying for extended periods of time at Mach 3, but it also requires enough thrust for steady level flight. To get an estimate of the thrust required at takeoff, we assumed a value of 0.3 TOW. None of the engines we looked at initially seemed to be able to provide the thrust we required. The few engines that appeared suitable possibilities were the turbofan engines designed for the Boeing 777.

The problem was that the engines designed for the 777 were too big and heavy to be used for our aircraft. Such big engines would have created an enormous amount of drag, which would greatly affect the efficiency. Moreover, these engines were designed for airplanes that had a cruising speed of Mach 0.85 while our SST had to fly at a much higher supersonic speed (Mach 3).

The engine we finally decided on was the Pratt and Whitney JSF – 119 engine. As the name implies, the engine was designed for the Joint Strike Fighter. The thrust that these engines provide is 40,000 pounds each. As we have a total of four engines, the total thrust provided by these engines is 160,000 pounds, which satisfies our requirements.

The value for TSFC was taken as 0.52. This was based on the few details provided by Pratt & Whitney and also on educated estimates. The reason for this is that usually TSFC is calculated from thrust and fuel flow rate, a quantity we did not have. However, we found the fuel flow rate from the thrust and the TSFC.

The bypass ratio (the ratio of the mass of cold air leaving the engine to the mass of hot air) wasn’t available for this engine on the Pratt & Whitney page but we assume that it is very close to one.

The placement of the engines is approximately half the way along the wing. A more detailed explanation for placement of the engines is provided later in this paper.

Speed for minimum drag

The speed for minimum drag is when L = W

Therefore the equation for minimum drag speed is:

Uinf^2 = 2W/(rho*S*Cl)

Substituting:

Uinf^2 = 2*4080301N /rho * 837.29 * .107174

Uinf = sqrt [8160602/rho *89.73571846]

As these graphs show, the speed for minimum drag increases as the altitude increases.

Stall Speed

Vs^2 = 2*W/rho S Clmax

Substituting:

Vs = sqrt (8160602/rho 837.29 Clmax)

There are two different plots because a range of 1.6 to 1.8 was given as CL max.

Thrust required and Thrust available

Drag = Cdo + Cdi

Cdi = .0009077

Cdo = .011/(Minf^2 –1)

CD = (.011/(Minf^2 –1))+.0009077

D = .5*rho*Uinf^2*S*CD

D = .5*rho*Uinf^2*837.29*((.011/(Minf^2 –1))+.0009077)

D = rho*Uinf^2*837.29*((.011/(Minf^2 –1))+. 0009077)

at low speed D = rho*Uinf^2*837.29*.0119077

At min drag speed

These graphs show that the required thrust will equal to the available thrust at about 1000 m/s and altitude 2000m. Steady level flight beyond this point is impossible because the required thrust exceeds the available thrust

Takeoff field length

Runaway Length = (1.2*Stall speed)^2/ net acceleration

Net acceleration at takeoff = net Thrust/ mass of aircraft
= / 4080301= 2.646 m/s

Fuel Weight and Volume

The calculation for the fuel weight was based on the mission specification that the aircraft must be able to travel about 8,000 miles. In addition to this, the SST should be able to fly an additional 300 miles in the unforeseen circumstance that it should be diverted to another airport. We also had to take into consideration the fuel that needed to be burned to reach the desired cruise altitude of 60,000 feet. We assumed this to be an extra ten percent of fuel weight.

Distance flown per pound of fuel = V/(c * D)

Where

V = velocity of the aircraft = 1980.58 mi/hr

c = specific fuel consumption at cruise altitude = .52 lbs of fuel per pound of thrust

Per hour

D = Total Drag at cruise altitude = 33227.63 pounds

• Hence the distance flown per pound of fuel = .0257 mi/lb

But as our cruise range is 8300 miles we get fuel required as –

• Distance traveled/distance traveled per pound of fuel = 322668.42 pounds.

In addition we will add an extra ten percent of fuel weight that will be required to attain the cruise altitude of 60,000 feet

• Total fuel weight = 1.1 * 322668.42 = 354936.1 lbs of fuel

Hence we arrive at the fuel fraction = weight of fuel/takeoff weight = 38.7 %

The density of aviation fuel is around 790-kg/cubic meter

The weight of fuel in kg is 161334.59 kg

• Hence the volume of the fuel will be the weight / density = 204.221 cubic meters

Refined value for the structural weight

As seen in the previous page, the fuel fraction was 38.5 percent. The payload fraction was approximated as 10 percent based on airplanes of similar designs. The average weight of the four engines can be approximated as 15 percent of TOW. Hence the remaining weight of the aircraft can be attributed to the structural weight. Hence a refined value for the structural weight fraction will be 36.5 percent.

Aircraft component	Percentage of Aircraft Weight (TOW)
Fuel	38.7
Engines	15
Payload	10
Structure	36.3

The distribution of the weight of the aircraft can be visualized best from the following pie chart –

One might notice that the structural weight is a little higher than usual. This value for structural weight allows us to experiment with more stronger and heavier alloys for the frame of the airplane. We must also take into consideration that as our aircraft will be cruising at Mach 3, a sufficient amount of material with high thermal resistance and coefficient of thermal expansion should be incorporated to protect the airplane from the adverse heating affects of high-speed flight. Such materials will only add to the overall structural weight of the aircraft. Hence a higher percentage of the weight of the aircraft must be devoted to the structure.

Wing Shape

The fact that our aircraft needs to cruise at an extremely fast velocity is reason enough for a high degree of sweep on our airplane. The reason for this is to reduce the resultant velocity with which the air hits the wing. The higher the sweep angle, the lower is the effective velocity and the heating effect of air friction, a very important aspect.

After it was decided that a high angle of sweep was an absolute requirement, we had to decide on the whether we should sweep the wings forward or backward. Using rough dimensions, we calculated the center of mass for the aircraft for a forward sweep and a backward sweep using the mass properties of the Mechanical Desktop software. We found that with a forward sweep, the center of mass of the aircraft was too close to the front and this would make the aircraft highly unstable. Consequently, it was decided that a backward sweep was the best solution.

Another problem that needed to be tackled along the way was the extremely high value for the platform area (837.29 square meters). Using conventional swept-back wings would have meant that our wingspan would have been extremely large. Such a large value for wingspan would have meant that we would be unable to use a vast majority of the airports in the world, as they can’t accommodate planes with very large wingspans.

The necessity for a swept wing and the limitation for an acceptable wingspan led us to the conclusion that the best alternative was a Delta Wing. The Delta wing could provide us with a large platform area and high sweep but at the same time provide us with smaller wingspans as compared to normal swept back wings.

In fact, when we did a comparison with similar planes such as the Concorde, the Boeing 2707 SST and the Tupolev Tu-144, we found out that all employed Delta Wings.

Using estimations based on extrapolations of the attributes of other aircraft, we found out that an ideal wingspan would be in the vicinity of 170 ft (51.83 square meters). Using this value, the altitude of the wing would be 32.30 meters.

Fuselage Dimensions

The aircraft was designed to carry 300 passengers, their luggage and limited cargo (5 tons). Taking all these parameters into consideration, we decided the approximate length of the aircraft to be 65 meters. The average cross-sectional diameter of the aircraft will be 5 meters. As mentioned earlier, the wingspan will be about 51 meters.

The angle of sweep for the wings will be around 38.55 degrees. The distance from the base of the wing to the tip will be about 32 meters. The average wing thickness will be about .25 meters.

Another point to note is that the front wheels will have a larger diameter and will be thicker than the rear wheels. This is to sustain the greater load placed on the front wheels. The rear wheels have a diameter of 0.75 meters while the front wheels have a diameter of 0.90 meters.

The average thickness of the tail wings will be 0.125 meters. The height of the tail will be 5 meters and its average thickness will be 0.2 meters.

Landing Gear And Placement

The landing gear will extend 4 meters to the ground. In addition the exit hatches will be 1.5 meters above this level. Hence the gate will be 5.5 meters above the ground, which in turn corresponds to the second level of most buildings.

The rear landing gear will be approximately 14 meters from the axis of the plane on each side and about 5 meters in front of the rear of the wings.

The front landing gear will be on the axis of the plane and will be about 5 meters ahead of the front of the wings.

The engines will be placed exactly in between the rear landing gear and the axis on either side.

The wingspan will be 51 meters, which allows the plane to land in virtually any airport.

Multiviews of the Aircraft

Note: The following views were created with AUTOCAD (Mechanical Desktop 3).

The engines have been hidden to show a better view of the landing gear. The engines are positioned as mentioned in the previous page.

The multiviews were done according the American Standard of Third Angle Projection.

The isometric drawing was included to give a better overall idea of our SST.

Addendum –

The following material was used to estimate various quantities.

Airplane	# of Passengers	Max TOW lbs	TOW Kg	Max TOW N	Range mi	Wing span m	Wing Area ft^2	Wing area m^2
NASA TCA	310	738,550	334,932	3,282,338				0
Tu-244	300	771,625	349,932	3,429,333	5,716		12,916.70	3937.01016
S-21	10	114,200	51,790	507,539			1,506.90	459.30312
Concorde	100	408,000	185,028	1,813,274	3,180	26	3,856.00	1175.3088
S-51	68	195,105	88,480	867,105	5,715		3,230.00	984.504
Boeing SST	300	600,000	272,100	2,666,580	3,900	37	9,000.00	2743.2
Tu-144	140	414,000	187,749	1,839,940	4,000	29

Calculations

Payload Fraction	0.1
Average Passenger mass (kg)	81.8181
Average luggage mass per pass.(kg)	31.8181
Supplies mass per passenger (kg)	3.786
Total Mass per passenger (kg)	117.4222
Max Passengers + crew	312	Passengers	300
Max Payload (kg)	41635.73	Crew	12
Acceleration Due to gravity (m/s^2)	9.8	Cargo (Kg)	5000
Weight of Max Payload (newtons)	408030.1
Takeoff weight (newtons)	4080301
Wing area (square meters)	837.29
Wing Loading (n/m^2)	4873.223
Wing span (meters)	51.83
Aspect ratio	3.208385
Gamma	1.4
Cruise Mach Number	3
Cruise Altitude (Meters)	18292
Air pressure at cruise altitude (N/m^2)	7217.5
Dynamic pressure at cruise altitude	45470.25
Lift Coefficient	0.107174
Angle of attack (radians)	0.303134
Max. Lift Coefficient at takeoff	1.6
Density at zero altitude (kg/m^3)	1.225
Stall Velocity (m/s)	70.51721	Stall Velocity (km/h)	253.862
Takeoff Velocity (m/s)	84.62065	Takeoff Velocity (km/h)	304.6343
Thrust at Takeoff (Newtons)	1224090
Spanwise efficiency factor	0.95
Induced drag coefficient	0.0012	Coefficient of Parasitic Drag	0.011
Speed of sound at cruise altitude (m/s)	295.07	Cdo at cruise speed	0.003889
Air density at cruise altitude	0.11606	Parasitic Drag	148070.6
Induced drag (newtons)	45670.85
Number of engines	4
Average thrust per engine at takeoff (N)	306022.6
Net Thrust at takeoff (newtons)	1101681
Mass of airplane in kg	416357.3
Specific Range (mi/lb)	0.025723
Fuel required (lbs)	322669.2
Range (miles)	8300
Extra fuel required to attain cruise altitude (lbs)	354936.1
Extra fuel required to attain cruise altitude (newtons)	1581079
Drag at cruise (Newtons)	148070.6
TSFC (lb/lb.hr)	0.52
Fuel fraction	0.387491

Sources Used

1. Jane’s all the World’s Aircraft

2. Fundamentals of Flight – Richard S Shevell

3. Http://www.adl.gatech.edu

4. Http://www.pratt-whitney.com

5. Http://www.Boeing.com