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SYSTEMS DEFINITION REVIEW

GREG FREEMAN

ANDY GRIMES

NICK GURTOWSKI

MOTOHIDE HO

VICKI HUFF

POORVI KALARIA

ROMAN MAIRE

TARA PALMER

SANJEEV RAMAIAH

JACK YANG

A A E 4 5 1

M A R C H 1 2 , 2 0 0 9

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TABLE OF CONTENTS

Executive Summary (Page 4)

I. MISSION OBJECTIVES (Page 5)

i. Introduction

ii. Mission Statement

iii. Systems Requirements Review

II. DESIGN REQUIREMENTS (Page 8)

III. AIRCRAFT CONCEPT SELECTION (Page 9)

IV. ADVANCED TECHNOLOGIES & CONCEPTS (Page 11)

i. Engine Selection

ii. Combustor Technology

iii. Inlet Design

iv. Nozzle Design

v. Compression Lift and Wing Tip Inclusion

vi. Material Selection

V. INITIAL CABIN LAYOUT (Page 16)

VI. CONSTRAINT ANALYSIS & DIAGRAM (Page 19)

i. Constraint Analysis

ii. Basic Assumptions Made For Each Constraint

iii. Constrain Diagram

VII. SIZING STUDIES (Page 24)

i. Current Sizing Approach

ii. Steps Toward Advanced Sizing

iii. Weight Breakdown

iv. Current and Future Steps

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VIII. CONCLUSION (Page 28)

i. Final Aircraft Concept

ii. Requirements Compliance Matrix

iii. Summary

iv. Next Steps

IX. REFERENCES (Page 31)

X. APPENDIX (Page 32)

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EXECUTIVE SUMMARY

As aviation advances, the desire for an economic, affordable Supersonic Transport (SST) has increased

rapidly. Since the Concorde, there have been no operable Supersonic Transports in the world. OceanAire

intends to design the world’s next SST, called Sky, which will solve the technical challenges that have

impeded the development of supersonic airliners for decades. As stated by NASA Aeronautics Research

Mission Directorate’s 2008-2009 University Competition, the technical challenges include supersonic

cruise efficiency, low sonic boom, and high-lift for take-off and landing. Other design specifications

decided upon were high cruise speed, long distance cruise, and reasonable passenger capacity, along

with providing a luxurious flight.

Sky will provide transportation for up to 49 passengers, mainly targeting first and business class

travelers. OceanAire has developed a concept to be the most effective of its kind, featuring

characteristics fit for an efficient supersonic aircraft and containing an optimized cabin layout. Sky will

incorporate various advanced technologies to improve its performance and surpass its competition. It

will offer these benefits while remaining both aerodynamically and structurally efficient and

environmentally friendly.

OceanAire proposes a Supersonic Transport to meet and exceed the requirements set forth by NASA,

Lockheed and the future market. Through research on past and present concepts, and ideas

recommended by Dan Raymer[10], tools have been used to generate a concept with features that will

allow Sky to meet objectives and perform necessary maneuvers. OceanAire has completed much

research to design an airplane that will be the best of its kind. These studies have led to an initial design

that will reasonably meet the aforementioned goals. This report will define what OceanAire has, thus

far, achieved for Sky.

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I. MISSION OBJECTIVES

Introduction

In the United States today, more passengers and cargo are moved by air than ever before. Over the

past 20 years, air travel has grown each year despite two major world recessions, terrorist acts, the

Asian financial crisis of 1997, the SARS outbreak in 2003, and two Gulf wars. According to NASA’s

Fundamental Aeronautics Program on October 2008, the total volume of air traffic is expected to double

or even triple over the next ten to fifteen years[4]. Boeing predicts an average growth in airline passenger

numbers between 2008 to 2027 of around 4.0 percent each year[5]. As a result, the demand for business

aircraft will increase dramatically by 2020. The need and desire for reduced travel time has and will

steadily increase as well. However, since the Concorde, there have been no operable Supersonic

Transports in the world to meet the growing market’s need for quicker travel times.

It has become clear that no new successful supersonic transport aircraft can be designed without first

identifying and solving the several technical challenges that have impeded the development and

production of one to date. Aeronautical advances will have to be made to solve many difficult

challenges. The main areas of progress are related to the environmental impact of the vehicle (NOx

emissions, generated sonic booms, exceptional amounts of noise) and to operational considerations

(manufacturing costs, operational costs, efficiency). These challenges will require progress in

aerodynamic improvement for sonic boom and drag reduction, combustion management for harmful

emission reduction, engine design to comply with noise regulations, and propulsion incorporation to

improve performance.

Meeting the challenges stated above is the objective of the NASA Fundamental Aeronautics University

Competition. OceanAire proposes a Supersonic Transport called Sky to meet the requirements set forth

by NASA and the growing airline passenger and traffic numbers. Through research on past (such as the

Concorde, Tu-144, and XB-70) and present concepts (Aerion Corporation’s SBJ, Lockheed Martin’s QSST,

Dassault Aviation’s HISAC, and Sukhoi’s S-21), and ideas recommended by Dan Raymer[10], goals and

benchmarks have been developed in order to facilitate the design of Sky.

Mission Statement

The two main mission objectives as set forth by OceanAire are to design an aircraft with supersonic

capabilities that is able to link major city pairs, as well as compete with other existing aircraft in the

market. Building on the first mission statement, it is imperative that Sky be designed as supersonic,

flying at speeds up to Mach 2.0. The aircraft will be designed for civilian purposes, enabling the

transport of passengers and cargo between city pairs across the world. The second mission objective

specifically is to design an efficient aircraft that can compete with existing ones in the market and still

remain profitable. With an assumed first flight in 2020, Sky’s competition will most likely be Aerion

Corporation’s SBJ, Lockheed Martin’s QSST, Dassault Aviation’s HISAC and Sukhoi’s S-21.

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The mission objectives as set forth by the NASA Fundamental Aeronautics University Competition are

to design a small supersonic airliner with a cruise speed of Mach 1.6 to 1.8, a design range of 4,000

nautical miles, a payload capacity of 35 to 70 passengers, a fuel efficiency of 3 passenger-miles per

pound of fuel, and a takeoff field length less than 10,000 feet. Furthermore, the NASA competition

states that the design should achieve one or more of the following: supersonic cruise efficiency, low

sonic boom (less than 70 PldB), and/or high-lift for take-off and landing. As stated in the System

Requirements Review, Sky will be designed with a cruise speed, fuel efficiency, and takeoff field length

as designated by NASA. The other mission requirements will specifically be a still air range of 5,450

nautical miles and a payload capacity of 49 passengers. Furthermore, Sky’s configurations and systems

will be chosen in order to emphasize supersonic cruise efficiency and high-lift for take-off and landing.

Even though the achievement of low sonic boom will not be stressed, Sky will be designed to be as eco-

friendly as possible.

Systems Requirements Review

The Systems Requirements Review was the initial design phase for Sky. The results of this first design

stage were the basis of the next design phase, the Systems Definition Review, as outlined in this paper.

The Systems Requirements Review focused mainly on obtaining customer requirements, an accurate

market study, and an initial sizing of the aircraft. After the market requirements and benchmarks were

known, concepts were generated based on these requirements.

After a preliminary market analysis, it was determined that Sky would be designed to serve business-

and first-class passengers who wish to significantly decrease their travel time. Up to 49 passengers will

be held in a luxurious and comfortable single-aisle configured cabin. The passengers of Sky will be able

to fly to a total of 17 global locations with a total of 19 global city pairs. The market analysis performed

for the year 2008 allowed for an estimated 2020 market forecast based on Boeing’s prediction of a four

percent increase in business travel per year[5]. From this forecast, it was predicted that 203 units of Sky

would be needed for the chosen city pairs. Using NASA’S Airframe Cost Model calculator[6], a crude cost

of each unit was estimated to be $123.4 million in the year 2009 (the year 2009 was chosen for

comparison between market competitors). With a sell price of $180 million, the profit for OceanAire is

predicted to be roughly $11.5 billion over a 10-15 year period. As a result, it is believed that 203 units

sold to a total of 19 city pairs is a sufficient number for profitable operations.

One major design constraint is that Sky will not be designed for supersonic flight overland. Per FAR36,

supersonic flight overland in the United States as well as 50 other countries is prohibited. Each of Sky’s

competition relies on this regulation being redefined; however, OceanAire believes that FAR36 will not

be changed in the near future, specifically by 2020. As a result, all of the 19 city pairs were chosen so as

to eliminate any supersonic flight overland. Despite not flying overland supersonically, Sky will be

designed to be as eco-friendly as possible.

Despite the inability for supersonic flights across land, Sky has many market strengths compared to

Aerion’s SBJ, Lockheed Martin’s QSST, Dassault Aviation HISAC, and Sukhoi’s S-21. Sky is able to hold

more passengers providing for better per-seat efficiency, has a longer range capability, and has one of

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the highest Mach cruise speeds. Specifically, Sky will be designed to have a maximum cruise speed of

1.8, a total still air range of 5,410 nautical miles, and a maximum cruise altitude of 50,000 feet. From the

quality function deployment and house of quality studies, it was shown that cruise Mach cruise number

and cruise efficiency would be the prime design focuses.

Furthermore, the System Requirements review showed that the design mission profile is based on the

San Francisco to Seoul route with head winds of 100 knots and added range capability to reach an

alternative airport 200 nmi away as well as loitering there for 30 minutes. The design range of the design

mission profile cruise segment was calculated to be 5,155 nmi. The economic mission profile was based

on the most route flown most frequently, New York (JFK) to London (HEA).

The initial sizing of Sky entailed the use of an aircraft database of numerous types of supersonic

aircraft and the gross takeoff weight, empty weight, empty weight fraction, aspect ratio, thrust-to-

weight ratio, wing loading, and maximum Mach number for each aircraft. From the database, estimates

for Sky’s aspect ratio and wing loading were predicted to be 2.2 and 100 lbs/ft2, respectively. The thrust-

to-weight ratio was estimated as 0.3. The SFC was predicted to be 0.78 1/hr on the presumption that a

low to medium bypass turbofan will be used. MATLAB was used to determine the empty weight fraction

equation. The empty weight fraction was imported into another section of the aircraft database

together with mission and design parameters. From an iterative process, the estimate for the gross

takeoff weight was predicted as 245,283 lbs with an empty weight fraction of 0.41355.

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II. DESIGN REQUIREMENTS

As discussed in the Systems Requirements Review, OceanAire developed the major design

requirements for Sky using a house of quality. After identifying Sky’s four primary customers (airlines,

passengers, the public, and NASA/Lockheed Martin), the needs and wants of these customers were

evaluated. The following quantifiable design requirements were chosen in order to meet every

customer’s needs:

• Takeoff field length • Operating cost

• Landing field length • Cruise altitude

• Door height above ground • Cruise efficiency

• Airframe life • Cumulative certification noise

• Range • Stall speed

• Number of passengers • Wing span

• Cruise Mach number • NOx emissions

• Cabin volume per passenger

Takeoff and landing field lengths are important design requirements that OceanAire must meet

because if Sky cannot operate at the airports of chosen city pairs, it will not be successful. The current

shortest runway that Sky will operate on is 10,081 ft in Boston. Sky’s door height above the ground must

be taken into consideration to meet airlines’ need for airport compatibility. In order for Sky to be

profitable it must have an airframe life of at least 20 years. The aircraft must have a range of 5,440

nautical miles in order to avoid flying supersonic overland and to reach its furthest city pair. Sky will hold

a total of 49 passengers: 10 in first class and 39 in business class. This number was chosen so that an

optimum seating arrangement could be utilized. Sky will cruise at a Mach number of 1.8 in order to

reach NASA’s design requirements. To provide adequate room, a cabin volume per passenger of 8.5

ft3/pax will be proposed. Operating cost for the development of this supersonic aircraft must be

reasonable in order for Sky to be profitable. The Concorde failed as a supersonic transport aircraft

because it made less money than its operating cost. A cruise altitude of 50,000 ft must be obtained so

that supersonic flight can take place. The current estimated cruise efficiency is 0.36 lb fuel/pax-nm. This

value still needs to improve so that fuel can be saved and range can be improved. Since Sky will not be

flying supersonic overland, sonic boom is not a design issue. However, OceanAire still wants to choose

or develop engines that mitigate the sonic boom. Stall speed will be an important design requirement so

that stall does not occur during takeoff and landings. This is a concern since one of NASA’s technical

challenges is to have high-lift for takeoff and landing. Sky’s wing span needs to meet airline regulation

for gate clearance and stay within the Mach cone. Finally, the emission of NOx into the atmosphere is an

important design requirement. This is due to the fact that the world is becoming more concerned about

keeping the earth clean, and the public will not tolerate an airplane that emits pollution through its

engines.

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III. AIRCRAFT CONCEPT SELECTION

Pugh’s Method was chosen to be used as the evaluation tool for the aircraft concept selection. It was

chosen because it is effective in comparing different designs, it allows for scoring concepts relative to

one another, and it is iterative and efficient. Prior to the concept brainstorming, it was necessary to

determine what criteria the designs would be compared through as well as the descriptors by which the

designs would be characterized. The lists of these categories are listed below. The criteria chosen were

based on market research, the Quality Function Deployment, and the NASA requirements given.

Once these categories had been discussed and evaluated, the next step was for each group member

to design his or her own custom supersonic transport concept based on the descriptors listed above. For

the initial evaluation of the generated concepts, the Concorde was chosen as the datum. It was chosen

due to the fact that it is one of the only successful commercial supersonic transports to be produced.

The initial concepts can be viewed in the appendix as well as the two completed iterations of Pugh’s

Method. Initially, nine concepts were evaluated and compared to the Concorde based on the design

criteria. Positives and negatives were used to determine what characteristics of each concept were

better or worse than those of the Concorde. This allowed for the best components from each concept to

be chosen and incorporated into three hybrid designs for the second iteration of Pugh’s Method. In the

second iteration, the “best” design from the first comparison was chosen as the new datum, and the

three new hybrid concepts were compared to it, again via the design criteria. Following this iteration, a

“winning” concept was established. A detailed sketch of this design is shown below in Figure 1.

Design Criteria Concept Descriptors

Airport Compatibility Nose Type

High Supersonic Cruise Efficiency Canards (Yes or No)

Low Certification Noise Fuselage Design

High Lift for Takeoff and Landing Wing Type

High Cabin Volume per Pax Engine Placement

Low Wave Drag Engine Inlet and Nozzle

Aerodynamic Supersonic Regime Tail Configuration

Stable Flight Gear Type

Low MX Cost Door Placement

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Figure 1: Current Sky Concept

The current concept for Sky is based on the best characteristics of the two iterations of Pugh’s

Method. The nose will not incorporate a quite spike, but will be long and slender with a gradual taper. A

quite spike was not chosen due to the fact that Sky will primarily operate over the ocean and, therefore,

will not need to reduce boom overpressure. Weaker shocks will be created via the long and slender

nose, which will allow for lower pressure losses. Sky will also implement canards aft of the cockpit to

add stability in supersonic flight, as well as additional lift. A Sears-Haack fuselage design will be used to

allow for a more gradual increase in cross-sectional area as the wings begin to develop. This will create a

decrease in wave drag. The wings will incorporate a dual sweep to not only allow them to stay inside the

Mach cone, but to also allow for a relatively larger surface area in the second segment to create more

stability in subsonic flight. Sky may also use wing tips that will fold down ninety degrees in supersonic

flight. If this design concept is chosen, Sky will also take advantage of compression lift to add lift in

supersonic flight. Compression lift will be discussed in more detail in the advanced concepts section.

Based on trade studies and constraint diagrams, the use of three engines is currently being considered

to create enough lift for Sky. Two of the engines will most likely be placed under the wings close to the

fuselage similar to the Concorde. The third engine will be mounted below the fuselage along the

centerline with its exhaust exited via the tail. The placement of the third engine, however, may create

unacceptable drag and affect the configuration of the tail. Thus, using three engines is only an initial

design consideration. The inclusion of a forth engine may be deemed necessary further into the design

of Sky. Sky will also integrate variable inlet and nozzle geometry to allow for a better match between

supersonic and subsonic regimes. These variable geometries will also be described in more detail in the

advanced concepts section.

Sky will use a tricycle gear configuration, and placement will be decided upon later. Due to the

inclusion of canards, a horizontal tail has been considered unnecessary at this stage in the design.

However, Sky will use a large vertical tail. Currently, door placement will be aft of first class and forward

of business class to allow for better airport compatibility due to the use of canards. At this phase in the

design, the current concept for Sky will allow for the most efficient supersonic transport. However, as

the design develops, some of these components may change.

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IV. ADVANCED TECHNOLOGIES AND CONCEPTS

Based on research done on supersonic transport flight, there are six major technologies to consider

for the design of Sky. These are engine selection, combustor technology, inlet design, nozzle design and

selection, wing tip inclusion, and material selection. Other technologies to research are the use of

compression lift as well as possible skin and structural materials to be used.

Engine Selection

When dealing with engine selection, there are a few considerations that should be taken into account.

This first is operational power. In most subsonic flight, engines operate at partial power. However, in

supersonic flight, engines must usually operate at near full power for hours at a time. This introduces

very high temperatures in the engine components which cannot be ignored. Another consideration is

the limit on pressure ratios created in the compressor section. The increase in air flow increases the ram

temperature. This in turn increases the compressor temperature which decreases the pressure ratio

that is achievable. Engine noise is also a problem with supersonic flight. High exhaust velocities from the

nozzle create high noise. Nozzle geometry selection needs to take this into account. The last

consideration is emissions. The emissions of a supersonic engine are quite high and need to be

addressed.

Figure 2: Variable Cycle Engine

The best type of engine to consider for Sky will be a medium bypass turbofan incorporating variable

cycle technology, as shown in Figure 2. A medium bypass turbofan is a better choice than a high or low

bypass due to a few factors explained in Koff’s paper entitled “Engine Design and Challenges for the

High Mach Transport.”[8] In it, he describes that a high bypass turbofan will not create the required

thrust while maintaining the necessary reduction in velocity from supersonic to subsonic. The fan for a

high bypass turbofan is too large and will be too prone to damage from the shocks. A low bypass

turbofan will address the thrust issue but will create too much noise to operate commercially. Variable

cycle technology will be used in order to vary the inlets and exits of various engine components. Analysis

done by Sipple in “Research on TBCC Propulsion for a Mach 4.5 Supersonic Cruise Airliner” predicts that

this technology offers superior thermal and propulsive efficiencies at various mach numbers and

altitudes compared with non variable cycle engines[11]. It will also allow for a vast range of bypass ratios

to be achieved, which in turn will give lower TSFC and will help reduce turbine temperatures.

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Combustor Technology

Combustor technology is also worthwhile to consider in engine selection and component design.

Temperature is the main concern when dealing with a jet engine combustor. The temperature of the

combustor and of the turbine inlet is a direct function of the gas temperature at which the fuel-air

mixture is burnt[3]. Current technology stipulates that combustor temperatures cannot remain above

3,300 degrees Fahrenheit for too long due to the production of NOX emissions. The main tradeoff in

combustors is that turbines operate more efficiently at high temperatures, but higher temperatures in

the combustor create high levels of NOX. One way to improve this would be more effective mixing of the

fuel and air. Another way is to vaporize the fuel better before its injection into the air flow. In Figure 3

shown below from Koff’s paper on combustor technology, having a more air-rich or even fuel-rich

mixture creates less emissions than a stoichiometric mixture. With the utilization of variable cycle

technolgy as well, better fuel/air mixtures could be achieved through bypass air entering the core flow

after combustion.

Figure 3: NOx Emissions

Inlet Design

Due to the differences between air mass flow required for subsonic and supersonic flight, inlet design

is also quite important. The inlet must capture core and bypass air to be used for combustion as well as

cooling for the various engine components. It is also important because the inlet must slow the

incoming air from supersonic to subsonic in order to keep the components of the engine (primarily the

fan) from being damaged. Based on these qualities, the inclusion of an inlet ramp with variable

geometry is being considered. Shown in Figure 4 below from Raymer’s book “Aircraft Design: A

Conceptual Approach,” this type of inlet would allow for the necessary changes to be from a subsonic to

a supersonic regime[10].

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Figure 4: Variable Inlet Geometry

The ramp will open in supersonic flight to create the necessary oblique shock waves that will slow the

flow without agitating it and decreasing the pressure too much. At subsonic speeds, the ramps will close

to allow for the maximum amount of mass flow to enter the engine. The gradual taper to the inlet

throat will create oblique shocks rather than an immense normal shock, which would destroy the

pressure of the incoming air flow. In a study done by Evelyn of the Boeing Corporation, it was found that

an axisymmetric inlet with variable geometry will create the least amount of drag for a supersonic

transport compared to a simpler spike or conical inlet[7].

Nozzle Design

In subsonic and low supersonic flight regimes, nozzle design also becomes important. Because Sky

needs the exit velocity from the nozzle to be supersonic, a converging-diverging nozzle must be used

with a chocked throat. Variable nozzle geometry with an ejector will also be used to allow for flow

control. In the subsonic regime, the maximum exit area will be achieved to allow for efficient flight. In

supersonic flight, the ejector will inject bypass air into the core flow, and the variable nozzle will

accelerate the chocked flow to a supersonic Mach number of 1.8. The utilization of flow vectoring may

also be considered if additional control is needed later in design.

Compression Lift and Wing Tip Inclusion

As mentioned earlier, folding wing tips may also be used of Sky. This will create inherent advantages

and disadvantages to the design. One advantage is that the aspect ratio will be reduced during cruise

when the wing tips are folded down. They will also add more stability surfaces to Sky. The main focus in

using wing tips is the incorporation of compression lift in the design. As the shock waves are created on

the underside of the aircraft body, they will be reflected back under the wing via the wing tips. This will

create a greater pressure reduction under the wing and will create more lift. The tips will, however,

create more weight and complexity and may also interfere with landing constraints given a malfunction.

Although the disadvantages seem high, the inclusion of the wing tips and the use of compression lift will

be advantageous to use.

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Material Selection

Various materials that could be used on Sky were considered, and it was realized that there would be

a number of factors that could affect material selection. Skin temperature increases more rapidly at

higher speeds, and since Sky will be flying at high speeds, material performance at high temperatures

will be important. According to Raymer, flying between Mach 1.6 and 1.8 would produce an average

skin temperature of about 350 degrees Fahrenheit, which is not particularly challenging but could be

limit material selection[10]. Other factors will be affordability, as this was one of the main reasons

Concorde fell short; efficiency, which includes corrosion and service life of the material and aircraft; and

availability of the material, determining if a certain material or manufacturing process has been or has

to be developed. However, at this point, only standard materials used in the aerospace industry were

researched.

One of the most widely used materials on aircraft is obviously aluminum and its alloys, but aluminum,

like other materials, has its pros and cons. Aluminum is abundant, affordable, and has a great strength-

to-weight ratio, especially its 7075 alloy. Some aluminum alloys, such as aluminum lithium, offer the

same weight savings as composites and can be formed using standard techniques. However, aluminum

has a maximum operating temperature of about 250 degrees Fahrenheit, which limits its use in

supersonic cruise, and is weak in fracture toughness and creep resistance[10].

Titanium alloys are another widely used material in the aerospace field. They are very stiff and

resistant to high temperatures and corrosion. They also have a high strength to weight ratio. However,

titanium alloys are difficult to form and are excessive in weight, which is extremely disadvantageous for

Sky. Also, titanium costs about 5 times as much as aluminum. For this reason, titanium is mainly used on

the leading edges of wings and tails and on engine components and landing gear[10].

One of the most remarkable materials to enter the world of aviation is composites. They are light in

weight, and composite material of the filament reinforced form offers a great strength to weight ratio.

The most commonly used composite in aircraft structure is graphite epoxy which also has a high

strength to weight ratio, but could be very expensive; up to 20 times more expensive than aluminum.

Composites that use epoxy as its matrix also have a maximum temperature about 350 degrees

Fahrenheit, which could limit its performance in supersonic transport[10].

As mentioned in the Systems Requirements Review, composites are good but are not the answer to

everything. It’s been noted that composites cannot accept concentrated loads, and that the strength of

composites can be affected by many factors such as moisture, cure cycle, and temperature exposure, to

name a few. Composites are also susceptible to damage, and internal damage can be particularly

difficult to find. When dealing with an aircraft that may be the first of its kind, internal damage is not

something that should be looked over. Composite materials are also hard to repair, which can incur

maintenance costs. Lastly, composites have very complex material properties that may be difficult to

fully comprehend[10].

The material selections of some of OceanAire’s competitors and benchmarks were also researched.

Aerion’s business jet incorporates composite materials, namely carbon epoxy, in its wings, but uses a

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metal coating on the leading edge for erosion resistance. The fuselage is made out of aluminum and

composites[1]. Lockheed mentions, on its website, that no new breakthrough materials were used or

needed on QSST, so the structural technology exists today to built such a supersonic transport[9]. Lastly,

the XB-70 has stainless steel and titanium components and utilized a sandwich honeycomb, which is a

certain type of composite made by attaching two thin, stiff skins to a thick, lightweight core[14].This

allows for high bending stiffness but low density. The sandwich honeycomb is a commonplace concept

today.

Advances in composites are constantly being made, so there is no telling what the future in materials

holds, even in 2020. This is especially true for composite materials for which advances are being made

constantly. It is also very possible for the cost of composites to decrease by 2020. However, it’s been

realized that, as far as material selection goes, the two main focuses for OceanAire will be weight and

temperature. Weight will be important because assuring that Sky remains as lightweight as possible is

crucial and temperature will be important because travelling at supersonic speeds requires temperature

to be a major concern. Also, since different materials prove advantageous in certain locations on the

aircraft than others, different materials will most likely be used in different locations based on the

specific needs on that location. Working with sizing with definitely be a must since maximum loads will

have to determined and met, and since the chosen materials will have an impact on empty weight

prediction. Material selection will also affect the cost of Sky, and this will be taken into consideration as

well. Next steps will be to begin looking at joint methods and sealants, as it becomes necessary.

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V. INITIAL CABIN LAYOUT

The initial cabin layout has been modified since the Systems Requirements Review. The new

configuration can be seen in Figure 5. The aircraft will still be divided into two different classes that can

hold 49 passengers. Those two classes will be first class and business class. First class will be able to

accommodate up to 10 passengers while business class will accommodate the remaining 39 passengers.

First class will be divided into 5 rows with 2 seats in each row. Business class will be divided into 13 rows

with 3 seats in each row in a 2-1 configuration. There will be 4 crew members, 2 pilots, and 2 flight

attendants. Even though our passenger capacity is under 50, two flight attendants were chose because

of the fact our aircraft will be a luxury aircraft, and having prompt service will be a desirable benefit. The

flight attendants will have their jumps seats located in the back facing forward. The cabin will have 2

lavatories, one located in front of first class and one located aft of business class. The galley will be

located behind the lavatory behind business class so that flight attendants can have easy access to the

galley.

Figure 5: Cabin Layout

There will be 4 doors on the airplane, as seen in Figure 6. Most other airplanes with a passenger

capacity of about 40 - 60 have 4 emergency doors as well[10]. One of those doors will be a boarding door

on the right side of the aircraft while the remaining three doors will be emergency doors. One of these

doors on the left side can also be used as

boarding door will be located in front of business class

fact that there will be canards towards the front of the plane

somewhere aft of the canards.

Figure 6

The dimensions inside the fuselage have changed since the first systems review.

dimensions can be depicted from Figure

business class seat pitch is 42 inches.

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be used as a maintenance door or as a loading door for the galley.

boarding door will be located in front of business class, but directly behind first class.

here will be canards towards the front of the plane. Therefore, the plane m

Figure 6: Boarding and Emergency Door Locations

The dimensions inside the fuselage have changed since the first systems review.

dimensions can be depicted from Figures 7 and 8. The first class seat pitch is now 46 inches, while the

business class seat pitch is 42 inches.

Figure 7: Seat Pitch

loading door for the galley. The

but directly behind first class. This is due to the

the plane must be boarded

The dimensions inside the fuselage have changed since the first systems review. These new

seat pitch is now 46 inches, while the

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The aisle width in first class is 26 inches and the aisle width in business class is 20 inches. The inside

cabin length will be 88 feet 8 inches, and the internal cabin width is 8 feet 1 inch. These dimensions are

subject to change if it is decided that the delta wing will be placed in a different location on the aircraft,

due to the area rule. The overall dimensions of the external fuselage have been adjusted to be able to

hold all the passengers. The fuselage length is now 196 feet, and the external width of the fuselage is 9

feet at the maximum diameter while the width at the tail end of the plane is 4 feet.

Figure 8: Cabin Dimensions

19

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VI. CONSTRAINT ANALYSIS AND DIAGRAM

Constraint Analysis

The optimal performance requirements of an aircraft necessitate a functional relation between thrust

to weight ratio (��) and wing loading (

�� �. It is crucial for OceanAire to perform at optimal conditions for

cruise, subsonic maneuver, takeoff ground roll, landing ground roll, and second segment climb.

OceanAire wants to be able to have the opportunity to design their aircraft in the best region of thrust-

to-weight ratio and wing loading region. There are 6 major performance constraints being considered.

These are below:

• Cruise

� 1 g Steady Level Flight, M = 1.8 at h = 50,000 ft

� Assuming Standard Atmosphere Conditions

This is the cruise mode of OceanAire. The altitude of cruise is not yet fully determined and is still

subject to change. A trade study of the different efficiencies at different altitudes will be carried in later

steps.

• Subsonic Maneuver

� 2 g turn at 250 knots at h= 10,000 ft

� Assuming 92% of the takeoff weight

OceanAire considered that it would be important to be able to perform such a maneuver in order to

blend in the subsonic traffic around airports in a safe manner. The 2g turn is still not fixed, and the team

will research federal regulations in order to eventually use a lower rate of turn.

• Takeoff Ground Roll

� 6000 ft at h = 0 ft

� +15° Hot Day

Takeoff ground roll is the distance required before the wheels leave the runway. Takeoff distance is

normally the total distance required for the aircraft to depart and clear an obstacle of a known height

past the end of the runway. OceanAire wants to be able to comply with the NASA 2009 design

competition requirements of 10,000 ft field length. The field length is the ground roll plus 66%, which

gives a 6,000 ft ground roll. OceanAire will be flying to airports in a hot environment. Therefore, an extra

15 degrees were added to the standard atmosphere.

• Landing Ground Roll

� 6000 ft at h = 0 ft

� +15° Hot Day

The landing ground roll is the same as the take ground roll. OceanAire used a ground roll of 6,000 ft to

be able to land at an airport with a 10,000 ft field length. The use of thrust reversers was not useful

20

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here, as the landing constraint on a 6,000 ft ground yielded a wing loading of about 130 psf, which is far

from the design region that OceanAire will use.

• Second Segment Climb Gradient

� Above h = 0 ft

� +15° Hot Day

Second segment climb gradient requirements from FAR regulations were used in this constraint. The

only configuration showing in the diagram is the one with 3 engines, which is the one that is optimum

for the set up. The use of a 2 engine configuration yields a thrust-to-weight ratio close to 0.7, which is

unfeasible; and 4 engines yields a thrust to weight ratio of about 0.3, which is much lower than the

design region and therefore won’t be necessary.

• Second Segment Climb Gradient

� Above h = 30,000 ft

� Level flight

This constraint was established to see how the aircraft would be able to accelerate through the

transonic regime. Following the results found in the report of the conceptual design of a Mach 1.6

airliner, the optimum transonic acceleration must happen at an altitude of 30,000 to decrease the wave

drag[12].

Basic Assumptions Made for Each Constraint

The 5 constraints assumptions which include cruise, subsonic maneuver, take off ground roll, landing

ground roll, and 2nd segment climb gradient are tabulated in the Table 1 below with respect to major

parameters including engine lapse rate, weight fraction, aspect ratio, leading edge angle, CLmax, Cd0, CDW,

engine quantities, climb gradient, and distance constraint. This table offers a summary of the main

assumptions made during the constraint analysis.

Cruise Subsonic

Maneuver

Take Off

Ground

Roll

Landing

Ground

Roll

2nd

Segment

Climb Gradient

Engine Lapse Rate (�) (%) 42% 82% 99% 99% 99% (25%

Reverse T)

Weight Fraction (Wi/Wo)

(%) 91% 92% 100% 100% 100%

AR

1.9 2.6 2.6 2.6 2.6

Oswald Efficiency (%)

82%

LE Angle (deg)

60°

CLmax

1.2 1.2 1.2

21

.

Cd0

0.0018 0.018

CDW

0.022

Number of Engines

3

Climb Gradient

(%) 2.7%

Distance Constraint

(ft) 6000ft 4000ft

Table 1: Major Parameters for Constraints

The engine lapse rate (�) was initially calculated using thrust, density at a specified altitude to thrust,

and density at sea level. This calculation is given by Equation 1.

� � ���

� ���

(Equation 1)

However, this relation is simply an approximation for obtaining the engine lapse rate, and the value

yielded through this method was clearly unrealistic (�� ���� = 15.23%). To verify the invalidity of this

relation for this application, the Concorde’s parameters were used to compute the lapse rate for

Concorde itself. This yielded a rather invalid cruise lapse rate of �� ���� = 9.15% (Concorde’s actual lapse

rate is �� ���� = 37%). By using the Concorde’s parameters to come up with its constraint diagram, it was

evident that the thrust-to-weight ratio was too high. The lapse rate approximation is more applicable for

high bypass turbofan engines, such as that used on the Boeing 747. Thus, its incompatibility for other jet

engine such as the Concorde’s twin spool Rolls-Royce/Snecma Olympus 593 turbojet engine is

understandable. It was concluded that an alternative approach for estimating the engine lapse rate for

OceanAire was necessary.

The lapse rate for Sky was instead estimated through Dassault, Aerion, Lockheed Martin, and

Concorde’s design lapse rate. OceanAire does plan on using new technologies when implementing its

propulsion system, which will have better performance and be more adaptable than the current

technology. For more information refer to the Advanced Concepts section.

Most of the values used in the constraint analysis were estimated roughly for a supersonic transport

type airplane or were found from historical data. The limitation of these assumptions will be clearly

decreased as more accurate values will be found from the different analyses that will be performed in

the upcoming stage of the project. For instance, the maximum lift coefficient was assumed to be 1.2

because it is a feasible value for a delta type of wing. Sky’s wing does not exactly meet the delta wing

specifications but can be approximated as being close for now.

Also, the weight fractions for each of the constraints were roughly approximated from the mission of

the aircraft, but they will be updated with better accuracy once the sizing code is completed, tested, and

ran for this project.

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.

It should be noted that the aspect ratio varies in different regimes of flight of the aircraft. This is

because OceanAire considers using compression lift with folding wing tips, which inevitably will change

the aspect ratio. Of course, very little information is known about the aerodynamics performances of

this concept, and more accurate aspect ratios might be used in different phases of the mission of the

aircraft. The current aspect ratios were found using hand sketches of the airplane respecting scale ratios

(length of the airplane and wing surface area).

OceanAire initially made estimations for various parameters utilizing Concorde’s performance

attributes as the basis for estimating OceanAire’s unknown parameters. Accordingly, some parameters,

such as the zero lift drag coefficient Cd0, were challenging to be predicted in this stage of the project due

to lack of information regarding airfoil shape. Airfoil geometry for Sky has not yet been decided.

The drag coefficients were roughly estimated as well. The wave drag coefficient Cdw, was computed

with the following formulas (Equation 2 and 3) provided by Raymer[10].

(������� � ��� �1 � 0.386 ! � 1.2�#.$% &1 � 'Λ()*+,-...

/## 01 ������ �23���4 (Equation 2)

������� � 5'

6 789:; �6 (Equation 3)

Equations 2 and 3 are linked with the Sears-Haack body wave drag for M < 1.2. The major assumption

made in this correlation is the utilization of the wing area from initial sizing.

The Oswald Efficiency Factor (e) is another factor that had to be roughly estimated in this step. A fair

estimation generally falls within the following range for transport aircrafts:

0.75 < e < 0.85

It was assumed that the Oswald Efficiency factor would tend to be high for Sky due to aerodynamic

efficiency concepts that could be used later on, and also due to the variable geometry of the wingtips

which allows for a greater aspect ratio at subsonic speeds.

The leading edge angle was found using the supersonic flow behaviors. Since OceanAire is planning on

using compression lift, a device enabling the formation of an oblique shock parallel to the leading edge

of the wing will have to be implemented, such as a central engine air inlet. To maximize the surface of

overpressure from the shock compression, it is more beneficial if the oblique shock is parallel to the

leading edge. After calculation of the oblique shock angle from a 5o deviation, it was found that the

leading edge angle of the wing should be 60o.

The transonic acceleration was calculated using the Equation 4.

�� � =

> ? �� @ /

�ABAC @ /

DA�ACE (Equation 4)

The coefficients alpha and beta were assumed to be 60% and 92%, respectively. The drag to weight

coefficient was taken from Torenbeek and Laban’s Figure 10[12]. There was no climbing, therefore the

second term in the brackets was equal to zero and the acceleration dv/dt was found by adjusting the

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constraints to the rest of the maneuvers of the aircraft such that the acceleration would not be a

limiting factor. This process yielded an acceleration of 6 ft/s2 when the drag is at its highest.

Constraint Diagram

Figure 9 shows the constraint diagram including all 6 of the main limiting maneuvers that will need to

be performed by Sky. The three maneuvers determining where the design region will lie are the cruise

segment, the subsonic 2g turn, and the landing. The shaded region shows the design region of the

aircraft, which should remain below a thrust-to-weight ratio of 0.5 for efficiency issues and between

wing loading values of 85 and 110 psf. The design value chosen right now is 104 psf wing loading and

0.45 thrust to weight ratio, but this can be subject to change depending on later trade studies

conducted by the design team.

Figure 9: Constraint Diagram

70 80 90 100 110 120 130 140 1500

0.1

0.2

0.3

0.4

0.5

0.6

Wing Loading [psf]

Thr

ust

to W

eigh

t ra

tio T

/ W

0

Constraint Diagram

CruiseSubsonic 2g Maneuver

Take Off

Landing

2nd Segment ClimbAcceleration

24

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VII. SIZING STUDIES

Current Sizing Approach

Between the Systems Requirements Review and Systems Definition Review, much advancement has

been made in predicting aircraft size. It was determined that the Excel spreadsheet originally being used

was not detailed enough to allow for accurate weight estimates of a supersonic transport. Although

there are numerous sizing codes available, it seemed beneficial to write an original code using MATLAB,

as the design of a supersonic transport is fairly unique.

In the previous design phase, empty weight estimations were based on the empty weight fraction,

We/W0. In order to include more detail and accuracy in the sizing estimates, one important addition to

sizing prediction is the estimation of component weights. Other important considerations are the

variables that need to be taken into account for supersonic flight due to the continuous updating of

flight conditions throughout the design mission, along with more detailed equations for different

segments of the design mission. These will be discussed in further detail below.

However, as of now, the Excel spreadsheet is still being used to predict aircraft weights. Since the

Systems Requirements Review, weight has been modified due to the constraint diagrams. From the

constraint diagram, values for W/S and T/W0 ratios that allow for flight of the aircraft were determined

and used in the spreadsheet rather than the previous values predicted from past aircraft. This allowed

for a more accurate weight prediction of W0 = 300,000 lbs, Wf = 170,000 lbs, and We = 118,000 lbs.

Steps Toward Advanced Sizing

As the design of the aircraft advanced, new and more detailed equations were chosen for advanced

sizing, which required many numbers that were input as constants in the initial sizing to be made into

functions of other parameters. This was necessary because as the conditions of flight change, so do

many variables. Some examples of the variables that need to be accounted for are the lift and drag of

the aircraft, along with SFC, which changes based on flight conditions. Along with these parameters

being made into functions, so were the design mission segments. This is due to the fact that many of

these parameters will change during the mission due to the burning on fuel (the change in fuel weight

effects the current weight of the aircraft). This is based on the fact that current aircraft altitude is

constantly changing, which means the density, speed of sound, etc. are also changing. As a result, these

variables will affect the instantaneous weight of the aircraft. The breakup of the individual design

mission segments into smaller segments accounts for continuous atmospheric and aerodynamic changes

based on altitude, along with calculating relatively continuous changes in weight based on fuel

consumption.

Weight Breakdown

When the component weights are estimated, they will be tabulated and summed to determine the

empty weight. The fuel weight can then be adjusted to yield the takeoff weight, which is the sum of the

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payload, crew, empty, and fuel weights. If the empty weight is higher than expected, there may not be

enough fuel to complete the design mission. As a result, the aircraft must be resized and optimized.

No component equations exist specifically for supersonic transport. As a result, the specific

component weights were computed using Dan Raymer’s fighter/attack weights (Section 15.3)[10]. These

equations were chosen due to the supersonic capabilities of Sky. For consistency, only the fighter

equations were used. Since Sky is not a fighter aircraft, a “calibration factor” would be needed in order

to accommodate the difference in component weights. The “calibration factor” was determined by first

applying Raymer’s equations to the Concorde, the only supersonic transport to date. Since no individual

component weights were found for the Concorde, all component weight equations were calculated,

summed, and then compared with Concorde’s published empty weight. The weights of the following

components were calculated: wing, vertical tail, fuselage, main landing gear, nose landing gear, engine

mounts, engine section, air induction system, engine cooling, oil cooling, engine controls, starter, fuel

systems and tanks, flight controls, instruments, hydraulics, electrical, and avionics.

The component weight equations that were used included 57 constants specific to the aircraft. Many

of the constants were not published values for the Concorde and had to be estimated. Conservative

estimates were taken in order that the “calibration factor” would be more accurate. As a result, when

the “calibration factor” is applied to Raymer’s equations to determine the component weights for Sky,

the values will be more accurate. Some constants that were estimated for the Concorde include the

taper ratio for both the wing and vertical tail, tail length, air induction duct length, nose and main

landing gear lengths, electrical routing distance, ultimate load factor, ultimate landing load factor,

number of hydraulic utility functions, system electrical rating, sweep of vertical tail, and uninstalled

avionic weight.

Once all the fighter equations were calculated using Concorde’s specifications in MATLAB, the

predicted empty weight was found to 84,923 lbs. The published empty weight of the Concorde is

173,500 lbs[2]. This gives a “calibration factor” of 2.043. This means that in order to effectively predict

the component weights for Sky using the fighter equations, the predicted component weight needs to

be multiplied by 2.043 in order to get a more accurate estimate. The new component weight estimates

will then be added up to determine Sky’s new predicted empty weight.

The “calibration factor” of 2.043 means that the fighter equations given by Raymer predict an empty

weight that is a little more than half the actual value for a supersonic transport. It may be assumed that

the equations are poorly correlated. However, there is no direct correlation between any known

component weight equations and supersonic airliners. The only way to improve this “calibration factor”

is to combine both the fighter and cargo/transport equations given by Raymer. The same combination

of equations applied to the Concorde to determine the new “calibration factor” will have to be applied

to Sky. After combining the cargo and fighter equations, it was found that the cargo/transport equations

given by Raymer predict the empty weight of supersonic transports (as shown by calculating the

equations in relation to the Concorde) more accurately. The specific component weights calculated with

the cargo/transport equations were the weights of the following: wings, fuselage, vertical tail, horizontal

tail, main and nose landing gears, nacelles and contents, starter (pneumatic), hydraulic systems,

26

.

electrical systems, installed APU, furnishing, air condition system, avionics, instruments, anti-ice system,

and handling gear. The component weights computed from the fighter equations that were kept were of

the engine mounts, engine cooling, oil cooling, fuel systems, flight controls, and engine controls. This

combination of weight equations used for the Concorde decreased the “calibration factor” to a value of

1.08. Since this new value is nearly one, it shows that the current combination used to calculate

component weights (and subsequently the empty weight) is more accurate.

Once the “calibration factor” had been calculated, a new component sizing code in MATLAB was

written for Sky. However, the component weights could not be determined in this design phase due to

the limited knowledge of the aircraft’s specifications and design characteristics. There are too many

estimations at this point in the design phase, and an accurate prediction of the weight components

could not be reached. The next step in the design phase is to calculate the needed constants used in the

component weight equations. Some constants for Sky may never be accurately predicted, such as

electrical routing distance. One solution to this problem may be determining the component weights

that cannot be calculated with Raymer’s equations as a percentage of other component weights, such as

the wings or fuselage. This percentage can be found from historical values. Once the component weights

are calculated, the new empty weight prediction can be incorporated into the sizing code to more

accurately predict the fuel weight and takeoff gross weight. Furthermore, the center of mass can then

be estimated for the aircraft. Once the center of mass is found, an analysis of the dynamics of the

aircraft can then ensue. The static margin can be determined in order to analyze the stability of Sky. The

vertical tail and canards can then be sized more accurately as well.

Current and Future Steps

Although many forward steps have been made in developing an accurate weight prediction, there is

still much left to be completed. The first is to include different parameters based on engine

specifications. As of now SFC is constant, which will not be true throughout flight. As altitude increases

and density changes, SFC will change. Lift will also change as weight changes during different segments

of flight. This will need to be updated. Perhaps most importantly is the need for a function to predict

drag. There are many components of drag, which include induced drag, parasite drag, and most notably

wave drag. Furthermore, these need to be calculated for the many different components of an aircraft

such as wings, fuselage, nacelles, canards, etc. Accurate drag estimations can allow for an accurate

weight prediction, since the amount of fuel needed to achieve a certain range greatly depends on the

amount of drag an aircraft produces, and since fuel accounts for a significant portion of gross takeoff

weight. As of now, accurate methods for prediction of parasite and induced drag have been determined,

but overall drag cannot be calculated since, at this point, there is not a valid wave drag prediction.

Finally, the advanced sizing code needs to be completed. This incorporates much of what has been

stated above, along with the completion of any functions already created that rely on the above

parameters. Once completed, validation should be done to show that the code predicts the weights

accurately. This can only be done using numbers for existing aircraft, such as the Concorde, to see how

closely the code predicts its weight compared to the published value.

27

.

Overall, the sizing code is advancing, and although there are still variables to account for in order to

obtain accurate predictions for empty weight and gross takeoff weight, this greater accuracy will allow

for a more reliable prediction of how the aircraft will truly perform. Current sizing code can be found in

the Appendix.

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VIII. CONCLUSION

Final Aircraft Concept

Figure 10 displays an isometric view of OceanAire’s Sky, to scale. Layout and arrangement were

discussed in the Initial Cabin Layout section above.

Figure 10: Isometric of Sky

Requirements Compliance Matrix

The requirement compliance matrix is a tool used to keep track of the achieved design objectives. The

matrix, shown in Figure 11, contains various engineering parameters with target, threshold, and current

values for those parameters.

Requirement Unit Condition Target Threshold Design (to Date)

Takeoff Field Length [ft] < 10,000 11,800 11000

Range [nmi] > 5410 4000 5410

Payload [pax] > 49 35 49

Cruise Mach # [N/A] > 1.8 1.6 1.8

Cruise Efficiency [lb fuel/pax-nmi] < 0.25 0.33 0.36

Certification Noise [PldB] < 50 70 69

Cabin Volume per Pax [ft^3/pax] > 10 8 8.55

Cruise Altitude [ft]

50000 60000 50000

Aircraft Life [years] > 30 20 28

Aspect Ratio [N/A] < 2.6 1.9 1.9

Thrust to Weight Ratio [N/A] > 0.37 0.3 0.45

Wing Loading [N/A] > 125 95 104

Crew [crew] < 3 5 4

Figure 11: Requirements Compliance Matrix

It can be seen from the above matrix that all the design parameters achieved up to now are in the

range between their respective target and threshold values. The only exceptions are the values obtained

29

.

for cruise efficiency and the thrust to weight ratio. Further work will be done in the next phase of the

design process to improve cruise efficiency and bring it into the acceptable range. As far as the thrust-

to-weight ratio is concerned, the current value was obtained from the constraint analysis and is fixed for

a particular mission.

Summary

The second phase of this project was a very crucial phase in the design process. It dealt with the

process of concept generation and a concept selection; research was conducted into advanced

technologies pertaining to different components and the performance of the aircraft; an initial cabin

layout was designed; a constraint analysis was performed on the selected mission; and further work was

done on the sizing of the aircraft.

The initial stages of the second phase of this degisn focused on concept generation and selection. As

for the concept generation process, each member of the team came up with a design for the aircraft to

meet the mission requirement. A total of thirteen concepts were generated. All the generated concepts

were then evaluated using the Pugh’s matrix to narrow down on a final concept. During this evaluation

process, all the generated concepts were evaluated against the Concorde, which was chosen to be the

datum for the required aircraft. The winning concept in this evaluation concept would be the most

efficient and most suited to the mission requirements among all other concept generated.

Once the concept evaluation process was finished and a concept was narrowed down on, studies

were conducted to look into advanced technologies for different components of the aircraft. This was

done to identify technologies that could be incorporated into the aircraft to improve its performance

and efficiency. The study was conducted on two major aspects of the aircraft, viz.: the engine selection

and the materials. The study was conducted by researching numerous papers on the above aspects of

the aircraft. For the engine selection, it was found that a medium bypass turbofan engine would suit the

thrust requirements. It was also determined that the supersonic efficiency of the engine can be

maximized by employing the variable cycle technology in the engine. Further research was conducted

into inlet geometry for the engine, as shock waves at inlet during supersonic flight play an important in

deciding the engine efficiency. It was found that a variable geometry inlet would be very beneficial in

terms of meeting the mass flow requirements and control over shock waves.

Along with engine selection, studies were conducted on materials to be used in the construction of

the aircraft. The main factors that dictated material selection were their performance at high

temperatures, affordability, efficiency, and availability of the material. The materials studied were

aluminum, titanium and composites. The materials present on some of Sky’s competitors’ planes were

also looked into. It was determined that the material focuses are the weight of the material and the

temperature resistance, as the aircraft is in supersonic flight for most of its mission.

An initial constraint analysis was conducted on the mission profile to obtain constraints for the thrust-

to-weight ratio and wing loading. It was found from the constraint analysis that the main limiting factors

during aircraft design would be the subsonic 2g maneuver, the landing of the aircraft and cruise. Once

the constraint analysis was completed, initial values for the thrust-to-weight ratio and for wing loading

30

.

were obtained. These values are not final and will be subject to change as more detailed analysis will be

conducted on the aircraft and better estimates will be available.

The sizing process used to predict the weights of the aircraft and its components saw progress in this

phase. The Excel sizing spreadsheet used in phase one was updated, and OceanAire also developed its

own sizing code in MATLAB.

Next Steps

In phase three of the design process, OceanAire wants to improve its sizing algorithm to more

accurately predict the required weights, and to include aircraft dimensions in the sizing process. More

advanced technologies will be looked into, and it will be determined how to best employ the newest and

best technologies into Sky in order to acquire the best performance. Further work is needed in the solid

modeling of the aircraft, and OceanAire intends to improve its CATIA model during the third phase. It

was also found during phase two that more work and analysis will be needed on the vertical and

horizontal tails. Hence, structural and dynamical analysis will be performed on the vertical and

horizontal tails, which may also affect their sizing. Finally, efficient use of carpet plots will be made in

phase three, which will help OceanAire to better visualize the trends in thrust-to-weight ratio and wing

loading on Sky.

31

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IX. REFERENCES

1Aerion Supersonic Business Jet. Aerion Corporation http://www.aerioncorp.com/home

2“Aerospatiale_BAC Concorde”. Aircraft-Info. http://portal.aircraft-info.net/article11.html

3Bernard Koff, TurboVIsion, Inc., Miami, FL; Steven Koff, TurboVIsion, Inc., Miami, FL

AIAA-2007-5344 . 43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,

Cincinnati, OH, July 8-11, 2007

4Brown, Diane. “Fundamental Aeronautics”. NASA. October 17, 2008.

http://www.aeronautics.nasa.gov/fap/

5“Current Market Outlook 2008-2027”. Boeing. http://www.boeing.com/commercial/cmo/index.html

6Cyr, Kelly. “Airframe Cost Model”. NASA. May 2007. http://cost.jsc.nasa.gov/airframe.html

7EVELYN, G. B., Boeing Commercial Airplane Co., Seattle, Wash.; JOHNSON, P. E., Boeing Commercial

Airplane Co., Seattle, Wash.; SIGALLA, A., Boeing Commercial Airplane Co., Seattle, Wash. AIAA-

1978-1051. American Institute of Aeronautics and Astronautics and Society of Automotive

Engineers, Joint Propulsion Conference, 14th, Las Vegas, Nev., July 25-27, 1978, AIAA 14

8Kauser, Fazal B., California State Polytechnic Univ., Pomona

AIAA-1994-2828 . ASME, SAE, and ASEE, Joint Propulsion Conference and Exhibit, 30th,

Indianapolis, IN, June 27-29, 1994

9QSST Quiet Supersonic Transport. Lockheed Martin http://www.saiqsst.com/

10Raymer, Daniel P. “Aircraft Design: A Conceptual Approach.” Fourth Edition. 2006. AIAA Education

Series. Conceptual Research Coorporation, Playa del Rey, California

11Sippel,Martin. DLR, German Aerospace Research Center, Cologne; AIAA-2006-7976. 14th AIAA/AHI

Space Planes and Hypersonic Systems and Technologies Conference, Canberra, Australia, Nov. 6-

9, 2006

12Torenbeek, Jesse, and Laban, Conceptual Design and Analysis of a Mach 1.6 Airliner, AIAA2004-4541,

10th Multidisciplinary Analysis and Optimization Conference 2004.

13United Airlines Embraer ERJ-145 (ER4) February 17, 2009 www.seatguru.com

14“XB-70 Valkyrie.” Wikipedia.com. March 9 2009. http://en.wikipedia.org/wiki/XB-70_Valkyrie#Design

32

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X. APPENDIX

Pugh’s Method: Iteration 1 & 2

33

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Initial Sky Concepts

34

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MATLAB Sizing Codes

EmptyWt.m

function We = EmptyWt(W0guess,AR,T_W0,W0_S,Mmax) % Empty Weight Fraction We_W0 = 2.808524*(W0guess^-0.08453959)*(AR^0.1377132)*(T_W0^0.1351319)*(W0_S^-0.1789255)*(Mmax^0.01676361); % We/W0 using composites composites = 0.95*We_W0; % Empty Weight We = composites * W0guess;

Mission.m

function Wfuel = Mission(Mcruise,dmr,sfc_cr,velCr,suprsonic_L_Dcr,al tAP,suprsonic_L_Dmax,resltr,sfc_ltr,subsonic_L_Dmax,W0guess,velMax) % ***** CHECK UNITS & DIMENSIONS ***** % Inputs numpax = 49; %number of passengers numcrew = 4; %number of crew AR = 2.55; %aspect ratio dmr = 5100; % Design mission range altAP = 180; %alternate airport range resltr = 0.75; %reserve loiter time T_W0 = .45; %thrust to weight ratio guess W0_S = 106; %wing loading guess Mmax = 2.0; %mach number sfc_cr = 0.78; %SFC for cruise $$$$$$$$$$$$$$$$$ Cl_cr = ; %$$$$$$$$$$$$$ sfc_ltr = 1; %sfc for loiter $$$$$$$$$$$$ Mcruise = 1.8; %cruise mach number % Constants g = 32.2; % Gravity Constant, ft/sec^2 % Calculations Based on Inputs W_pyld = 220*numpax; W_crew = 200*numcrew; % Supersonic Calculations suprsonic_L_Dmax = 11*Mcruise^-0.5; % $$$$$$$$$$$$$$$$$4 suprsonic_L_Dcr = 0.86*suprsonic_L_Dmax; % $$$$$$$$$$$$ velCr = Mcruise*968.1/1.689; velMax = Mmax*968.1/1.689; % Subsonic Calculations subsonic_L_Dmax = (1.4*AR)+7.1; %$$$$$$$$$$$$$$$$ subsonic_L_Dltr = subsonic_L_Dmax; % $$$$$$$$$$$$$$$$$

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% Design Mission % Wb_W0: Engine Start, Warmup, Taxi Wb_W0 = 0.97; Wb = Wb_W0 * W0guess; % Wc_b: Takeoff/1st segment climb : pages 581,549 T_tos = ; % Takeoff static thrust BPR = ; % Bypass Ratio G = ; % Gamma_climb - Gamma_min CALCULATE $$$$$$$$$$$$$ $$$$$$$ Cl_climb = ; % Cl at climb speed (1.2Vstall) % $$$$$$$$$$$$$$$ Cl_max = ; % Cl max ????? h_obs = 35; % Height of obstacle, ft rho_sl = ; % Density at Sea Level Wc = Takeoff(Wb,T_tos,BPR,G,Cl_climb,Cl_max,h_obs,r ho_sl,W0_S); % Wd_c: 2nd segment climb: pages 536-537, 582 *Subs onic Vel2 = 1.25*Vstall; h1 = 35; h2 = 1500; Wd = Climb(Vel2,h1,h2,Wc); % We_d : 3rd segment climb: climb to 10,000 ft at 250 KCAS *Subsonic h3 = 10000; Vel3 = ; % 250 knots We = Climb(Vel3,h2,h3,Wd); % Wf_e: Accelerate to best climb speed % DO WE NEED THIS??????????????????????????? % Wg_f: 4th segment climb *Supersonic h4 = 50000; Vel4 = ; Wg = Climb(Vel4,h3,h4,We); % $$$$$$$$$$$$$$$Possibly Wf % Wh_g: Best cruise speed and altitude *Supersonic initial_alt_cr = 50000 ; VelCr = M_spr_cr*a; range = ; Wh = Cruise(VelCr,Wg,range,initial_alt_cr,S); % Wi_h: Descend to 10,000 ft : No range descent % Wj_i: Decelerate to 250 knots % DO WE NEED THIS?????????????????? % Wk_j: Descend to 1500 ft at 250 knots: No range d escent % Wk: Loiter for 30 min E = 30; % min Wk = Loiter(E,Wh); % $$$$$$$$$$$$$$$$$$$$ Possibly Wj % Wl_k: Approach

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% ASK ABOUT THIS % Wm_l: Missed approach, climb 10,000 ft $$$$ VERI FY $$$$ Vel5 = 1.5*Vstall ; % h5 = ; h6 = 10000; Wm = Climb(Vel5,h5,h6,Wk); % $$$$$$$$$$$$$$$$$ Possibly Wl % Wn_m: Cruise for best alternate airport (200 nmi) VelCr2 = M_sub_cr*a; range2 = ; Wn = Cruise(VelCr2,Wm,range2,Cl_cr); % Wo_n: Decelerate to 250 knots and descend to 1,5 00 ft: No range descent % Wo: Loiter for 30 min Wo = Loiter(E,Wn); % Wp_o: Approach % ASK ABOUT THIS % Wq_p: Land over 50 ft obstacle Wq = 0.995*Wo; % Possibly Wp % Fuel Fraction wfuel_w0 = 1.01*(1-Wq/W0guess); % Fuel Weight Wfuel = wfuel_w0*W0guess;

Climb.m

function Weight = Climb(Vel,h1,h2,W) Weight = W; Delta_h = (h2-h1)/50; for i = 1:50; % sfc = will be a function % T = will be a function % D = will be a function WeightRatio = exp(-sfc*Delta_h/(Vel*(1-D/T))); Weight = WeightRatio(i)*Weight(i); end Weight;

Cruise.m

function Weight = Cruise(VelCr,W,range,initial_alt,S) segrng = range/300; Weight = W; for i = 1:300 rho = atmosphere4(atmosphere4(initial_alt,0)); q = 0.5*rho*VelCr^2 Cl = Weight/q*S

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L = Weight %D L_D = Lift/Drag; % sfc WeightRatio = exp(-(segrng)*sfc/(VelCr*L_D)); Weight = WeightRatio*Weight; end Weight;

Loiter.m

function Weight = Loiter(E,W) segE = E/10; Weight = W; for i = 1:10 % D L = Weight L_D = Lift/Drag; % sfc WeightRatio = exp(-(segE)*sfc/L_D); Weight = WeightRatio*Weight; end Weight;

Takeoff.m

function Weight = Takeoff(Wb,T_tos,BPR,G,Cl_climb,Cl_max,h_ obs,rho_sl,W0_S) % CHECK on Tavg, Tavg_W0 % Sfc = function rho = atmosphere4(atmosphere4(initial_alt,0)); %%%%%%%%%RHO for what altitude?! U = 0.01*Cl_max + 0.02; % For Flaps in takeoff position Tavg = 0.75*T_tos*((5 + BPR)/(4 + BPR)); Tavg_W0 = Tavg/Wb ; % $$$$$$$$$$$$$$$$$$$$ d = (0.863/(1+2.3*G))*(W0_S/(rho*g*Cl_climb) + h_ob s)*(1/(Tavg_W0 - U)+2.7) + (655/sqrt(rho/rho_sl)); WeightRatio = 1-sfc*d*(T/W); % (T/W)i Weight = WeightRatio * Wb;

Main.m

% Main Sizing Code clear all clc counter = 1; W0guess(counter) = 300000; %-------------------------------------------------- ---------- % ITERATION FOR WEIGHTS % Fuel Weight Wf(counter) = Mission(Mcruise,dmr,sfc_cr,velCr,suprsonic_L_Dcr,al tAP,suprsonic_L_Dmax,resltr,sfc_ltr,subsonic_L_Dmax,W0guess(counter),velMax); % Empty Weight

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We(counter) = EmptyWt(W0guess(counter),AR,T_W0,W0_S ,Mmax); % Gross Takeoff Weight W0(counter) = We(counter) + Wf(counter) + W_pyld + W_crew; % end while abs(W0(counter)-W0guess(counter)) > 0.001 counter = counter+1; W0guess(counter) = W0(counter-1); % Fuel Weight Wf(counter) = Mission(Mcruise,dmr,sfc_cr,velCr,suprsonic_L_Dcr,al tAP,suprsonic_L_Dmax,resltr,sfc_ltr,subsonic_L_Dmax,W0guess(counter),velMax); % Empty Weight We(counter) = EmptyWt(W0guess(counter),AR,T _W0,W0_S,Mmax); % Gross Takeoff Weight W0(counter) = We(counter) + Wf(counter) + W _pyld + W_crew; % end end fprintf( 'The gross takeoff weight Wo = %.2f\n' , W0(counter)) fprintf( 'The empty weight We = %.2f\n' , We(counter))

atmosphere4.m

function [temp,press,rho,Hgeopvector]=atmosphere4(Hvector,G eometricFlag) %function [temp,press,rho,Hgeopvector]=atmosphere4( Hvector,GeometricFlag) % Standard Atmospheric data based on the 1976 NASA Standard Atmoshere. % Hvector is a vector of altitudes. % If Hvector is Geometric altitude set GeometricFla g=1. % If Hvector is Geopotential altitude set Geometric Flag=0. % Temp, press, and rho are temperature, pressure an d density % output vectors the same size as Hgeomvector. % Output vector Hgeopvector is a vector of correspo nding geopotential altitudes (ft). % This atmospheric model is good for altitudes up t o 295,000 geopotential ft. % Ref: Intoduction to Flight Test Engineering by Do nald T. Ward and Thomas W. Strganac % index Lapse rate Base Temp Base Geopo Alt Base Pressure Base Density % i Ki(degR/ft) Ti(degR) Hi(ft) P, lbf/ft^2 RHO, slug/ft^3 format long g D= [1 -.00356616 518.67 0 2116.22 0.00237691267925741 2 0 389.97 36089.239 472.675801650081 0.000706115448911997 3 .00054864 389.97 65616.798 114.343050672041 0.000170813471460564 4 .00153619 411.57 104986.878 18.1283133205764 2.56600341257735e-05

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5 0 487.17 154199.475 2.31620845720195 2.76975106424479e-06 6 -.00109728 487.17 170603.675 1.23219156244977 1.47347009326248e-06 7 -.00219456 454.17 200131.234 0.38030066501701 4.87168173794687e-07 8 0 325.17 259186.352 0.0215739175227548 3.86714900013768e-08]; R=1716.55; %ft^2/(sec^2degR) gamma=1.4; g0=32.17405; %ft/sec^2 RE=20926476; % Radius of the Earth, ft K=D(:,2); %degR/ft T=D(:,3); %degR H=D(:,4); %ft P=D(:,5); %lbf/ft^2 RHO=D(:,6); %slug/ft^3 temp=zeros(size(Hvector)); press=zeros(size(Hvector)); rho=zeros(size(Hvector)); Hgeopvector=zeros(size(Hvector)); % Convert from geometric altitude to geopotental al titude, if necessary. if GeometricFlag Hgeopvector=(RE*Hvector)./(RE+Hvector); disp( 'Convert from geometric altitude to geopotential al titude in feet' ) else Hgeopvector=Hvector; %disp('Input data is geopotential altitude in feet' ) end ih=length(Hgeopvector); n1=find(Hgeopvector<=H(2)); n2=find(Hgeopvector<=H(3) & Hgeopvector>H(2)); n3=find(Hgeopvector<=H(4) & Hgeopvector>H(3)); n4=find(Hgeopvector<=H(5) & Hgeopvector>H(4)); n5=find(Hgeopvector<=H(6) & Hgeopvector>H(5)); n6=find(Hgeopvector<=H(7) & Hgeopvector>H(6)); n7=find(Hgeopvector<=H(8) & Hgeopvector>H(7)); n8=find(Hgeopvector<=295000 & Hgeopvector>H(8)); icorrect=length(n1)+length(n2)+length(n3)+length(n4 )+length(n5)+length(n6)+length(n7)+length(n8); if icorrect<ih disp( 'One or more altitutes is above the maximum for thi s atmospheric model' ) icorrect ih end % Index 1, Troposphere, K1= -.00356616 if length(n1)>0 i=1; h=Hgeopvector(n1); TonTi=1+K(i)*(h-H(i))/T(i); temp(n1)=TonTi*T(i);

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PonPi=TonTi.^(-g0/(K(i)*R)); press(n1)=P(i)*PonPi; RonRi=TonTi.^(-g0/(K(i)*R)-1); rho(n1)=RHO(i)*RonRi; end % Index 2, K2= 0 if length(n2)>0 i=2; h=Hgeopvector(n2); temp(n2)=T(i); PonPi=exp(-g0*(h-H(i))/(T(i)*R)); press(n2)=P(i)*PonPi; RonRi=PonPi; rho(n2)=RHO(i)*RonRi; end % Index 3, K3= .00054864 if length(n3)>0 i=3; h=Hgeopvector(n3); TonTi=1+K(i)*(h-H(i))/T(i); temp(n3)=TonTi*T(i); PonPi=TonTi.^(-g0/(K(i)*R)); press(n3)=P(i)*PonPi; RonRi=TonTi.^(-g0/(K(i)*R)-1); rho(n3)=RHO(i)*RonRi; end % Index 4, K4= .00153619 if length(n4)>0 i=4; h=Hgeopvector(n4); TonTi=1+K(i)*(h-H(i))/T(i); temp(n4)=TonTi*T(i); PonPi=TonTi.^(-g0/(K(i)*R)); press(n4)=P(i)*PonPi; RonRi=TonTi.^(-g0/(K(i)*R)-1); rho(n4)=RHO(i)*RonRi; end % Index 5, K5= 0 if length(n5)>0 i=5; h=Hgeopvector(n5); temp(n5)=T(i); PonPi=exp(-g0*(h-H(i))/(T(i)*R)); press(n5)=P(i)*PonPi; RonRi=PonPi; rho(n5)=RHO(i)*RonRi; end % Index 6, K6= -.00109728 if length(n6)>0 i=6; h=Hgeopvector(n6);

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TonTi=1+K(i)*(h-H(i))/T(i); temp(n6)=TonTi*T(i); PonPi=TonTi.^(-g0/(K(i)*R)); press(n6)=P(i)*PonPi; RonRi=TonTi.^(-g0/(K(i)*R)-1); rho(n6)=RHO(i)*RonRi; end % Index 7, K7= -.00219456 if length(n7)>0 i=7; h=Hgeopvector(n7); TonTi=1+K(i)*(h-H(i))/T(i); temp(n7)=TonTi*T(i); PonPi=TonTi.^(-g0/(K(i)*R)); press(n7)=P(i)*PonPi; RonRi=TonTi.^(-g0/(K(i)*R)-1); rho(n7)=RHO(i)*RonRi; end % Index 8, K8= 0 if length(n8)>0 i=8; h=Hgeopvector(n8); temp(n8)=T(i); PonPi=exp(-g0*(h-H(i))/(T(i)*R)); press(n8)=P(i)*PonPi; RonRi=PonPi; rho(n8)=RHO(i)*RonRi; end