Systems Definition Review - Purdue EngineeringReview (SRR), while illustrating the steps that are...

Super Sonix Inc.

Systems Definition Review An analysis of performance, concepts, and sizing

Akshay Ashok, Nithin Kolencherry, Steve Skare, Michael McPeake, Muhammad Azmi, Richard Wang, Mintae Kim, Dodiet Wiraatmaja, Nixon Lange

3/12/2009

Super Sonix Inc. 12 March 2008 Team 2 Systems Definition Review

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Table of Contents

TABLE OF CONTENTS ........................................................................................................................... 2

TABLE OF FIGURES ............................................................................................................................... 4

LIST OF TABLES ...................................................................................................................................... 5

INTRODUCTION...................................................................................................................................... 6

Review of SRR ........................................................................................................................................................ 6

Opportunity Description ........................................................................................................................................... 6

Mission Statement .................................................................................................................................................... 7

Market Analysis and Concept of Operations ............................................................................................................ 7

Initial Sizing ............................................................................................................................................................... 7

CONSTRAINT ANALYSIS ...................................................................................................................... 8

1-g Supersonic Cruise ............................................................................................................................................. 8

2-g subsonic Maneuver ........................................................................................................................................ 10

Takeoff and Landing Ground roll .......................................................................................................................... 11

Second Segment Climb Gradient .......................................................................................................................... 13

CONCEPT SELECTION ......................................................................................................................... 16

Brainstorming ...................................................................................................................................................... 16

Pugh’s matrix ....................................................................................................................................................... 17

Hybrid Concepts .................................................................................................................................................. 18

Hybrid concept 1: .................................................................................................................................................... 19

Hybrid concept 2 ..................................................................................................................................................... 20

Hybrid concept 3 ..................................................................................................................................................... 21

Aft arrow wing concept .......................................................................................................................................... 22

ADVANCED TECHNOLOGIES ............................................................................................................. 25

Boom shaping technologies ................................................................................................................................. 25

Nose Design ......................................................................................................................................................... 25

Blunt Nose .............................................................................................................................................................. 25

Nose Keel ................................................................................................................................................................ 26

Dihedral Angle ..................................................................................................................................................... 27

Effective Area distribution ................................................................................................................................... 27

Aerodynamic Cowling ............................................................................................................................................. 27

Engine Nacelle Placement ...................................................................................................................................... 28

Canards ................................................................................................................................................................... 28

Efficient supersonic cruise ................................................................................................................................... 29

Active Flow Management .................................................................................................................................... 29

Engine Technology ............................................................................................................................................... 29

Modifications for Samara NK-321 Engine ............................................................................................................... 31

CABIN LAYOUT ..................................................................................................................................... 36

SIZING STUDIES .................................................................................................................................... 40

Sizing Approach ................................................................................................................................................... 40


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Component Weights ............................................................................................................................................ 41

Center of Gravity ................................................................................................................................................. 43

Shock Prediction .................................................................................................................................................. 44

N-wave .................................................................................................................................................................... 44

Plateau wave........................................................................................................................................................... 46

SUMMARY ............................................................................................................................................... 48

Compliance Matrix .............................................................................................................................................. 48

Walkaround and CAD Model ............................................................................................................................... 49

Next Steps ........................................................................................................................................................... 51

References ........................................................................................................................................................... 52

APPENDIX............................................................................................................................................... 54

Progress Update Chart ......................................................................................................................................... 54

Concept Selection Process ................................................................................................................................... 55

Initial Concept Drawings ...................................................................................................................................... 56

Detailed Design Criteria ....................................................................................................................................... 57

Design Mission Profile ......................................................................................................................................... 57

Historical Seating Data ......................................................................................................................................... 58

Combined Mass fraction Averages ....................................................................................................................... 58

Initial Sizing results .............................................................................................................................................. 59

Requirements Compliance Matrix ....................................................................................................................... 60

Next Step Flowchart ............................................................................................................................................ 61


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Table of Figures

Figure 1: Constraint Diagram ........................................................................................................ 15

Figure 2: Concepts used for the Pugh’s method analysis ............................................................. 17

Figure 3: Hybrid concept 1 ............................................................................................................ 19



Figure 6: Aft Arrow wing concept [12] ............................................................................................ 22

Figure 7: 3-view and isometric drawing of final design ................................................................ 24

Figure 8: Blunt nose effect on boom overpressure [23] ................................................................. 26

Figure 9: Dihedral angle design..................................................................................................... 27

Figure 10: Non-Axisymmetrical fuselage shape [22] ...................................................................... 28

Figure 11: Engine Database .......................................................................................................... 30

Figure 12: Thrust-to-weight ratios for viable engines .................................................................. 31

Figure 13: Proposed engine exhaust configurations [14] ............................................................... 33

Figure 14: Directivity of Overall sound pressure level [14] ............................................................ 34

Figure 15: Perceived Noise Level [14] ............................................................................................. 35

Figure 16: Passenger Seating Options (pitch) ............................................................................... 36

Figure 17: Passenger Seating Options (width) .............................................................................. 37

Figure 18: Passenger Seating vs. Overall a/c Length .................................................................... 37

Figure 19: Seating Layout with Galleys and Restrooms ................................................................ 38

Figure 20: Seating Dimensions, Coach (L) First Class (R) .............................................................. 39

Figure 21: Sizing Approach flowchart ........................................................................................... 40

Figure 22: Component weight mass fraction ................................................................................ 42

Figure 23: Center of Gravity Calculator ........................................................................................ 44

Figure 24: Shape factor and pressure amplification factor charts ............................................... 45

Figure 25: Walkaround chart ........................................................................................................ 49

Figure 26: 3-View of Super Sonix Aircraft ..................................................................................... 50


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List of Tables

Table 1: Flight Conditions and Aircraft Parameters for 1-g steady level flight............................. 10

Table 2: Parameters for subsonic 2-g maneuver constraint ........................................................ 11

Table 3: Landing and Takeoff parameters for DXB airport ........................................................... 12

Table 4: Takeoff and Landing parameters for JFK ........................................................................ 13

Table 5: Second segment climb constraint parameters ............................................................... 14

Table 6: Assumptions Sensitivity Table ......................................................................................... 16

Table 7: First run of Pugh’s matrix ................................................................................................ 18

Table 8: Second run of Pugh’s matrix ........................................................................................... 23

Table 9: Viable supersonic engines [5] ........................................................................................... 30

Table 10: Current Sizing Attributes ............................................................................................... 41

Table 11: Definition of variables for overpressure calculation .................................................... 45

Table 12: Input Parameters for N-wave overpressure calculation ............................................... 46

Table 13: Input Parameters for Plateau wave overpressure calculation ..................................... 47


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Introduction

Over the last few years, air traffic has seen a high increase in capacity, range and efficiency. In

addition to these, the need for a faster aircraft is imperative at the present time. Transporting

people over longer distances within a short span of time will be the objective of the airline

industry in the near future. This report addresses the systems definitions for a small supersonic

airliner with Initial Operating Capability (IOC) in 2020. A progress update chart is presented in

the appendix, and summarized tasks that were performed earlier in the Systems Requirements

Review (SRR), while illustrating the steps that are taken as part of the Systems Definition

Review (SDR) process. The SRR included an initial sizing exercise based on historical data, to

estimate the aircraft’s takeoff gross weight. Market analysis was performed to determine the

focus markets and passenger loads, operating costs and aircraft utilization. Customer

requirements and engineering design characteristics were detailed in the House of Quality as

part of the Quality Function Deployment (QFD) phase.

To better define the aircraft system, as part of the SDR, additional work is done to identify

performance constraints using a constraint diagram. The process drives the need for a

comprehensive engine database that is used to select an engine, or a concept, that fits the

initial performance requirements of the Super Sonix aircraft. Concept generation is an

important phase of the SDR which allows the aircraft to be visualized by taking on a geometrical

form. The brainstorming of ideas and the Pugh’s selection matrix exercises leads to further

research in advanced technological concepts. These advanced considerations allow a validation

of shortlisted concepts, and enables a preliminary calculation of the sonic boom overpressure.

A component weight database is built to aid in determining component weight fractions, and

leads to a better resolved takeoff gross weight prediction. Along with the cabin layout, these

weights are used to calculate the center of mass of the aircraft system. The next section serves

as a brief review of the detail of the analysis done in the SRR.

Review of SRR

Opportunity Description

NASA ARMD 2008-09 University Competition provides a framework that is used to guide the

conceptual design of a supersonic aircraft. The aircraft is expected to meet a set of goals

specified by the competition. These include, but are not limited to:

Mach cruise speed of 1.6 -1.8

Design Range of 4000 nautical miles

Accommodate 35-70 passengers (preferably in a mixed class configuration)

Fuel efficiency of 3 passenger-miles per pound of fuel or better

Takeoff field length < 10,000 ft


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Additionally, the aircraft is expected to achieve supersonic cruise efficiency, have a low sonic

boom (<70 PldB) and high lift for takeoff or landing – all while making a reasonable profit for

the company.

Mission Statement

The Super Sonix mission statement is as follows:

A cost-effective, advanced, high-speed commercial air transport that connects major worldwide

hubs

Super Sonix has set out several key design goals that is expected to be met and exceeded

through the course of the conceptual development:

Supersonic flights over land (boom overpressure less than 0.3 lb/ ft2)

Capture significant market of Supersonic Business Jets and supersonic transports

Initial Operational Capability (IOC)in 2020

Manufacturing capabilities exist for the aircraft

Payload of 60 passengers in a twin class configuration

Still-air ground range of 4000nm

Market Analysis and Concept of Operations

Three focus markets emerge from the market analyses: Trans-Continental, Trans-Atlantic and

Inter-Asia. Through detailed studies on market viability and cost models, potential worldwide

hubs are selected to set forth a viable business proposition for customer airlines. These are: Los

Angeles International Airport (LAX), John F. Kennedy International Airport (JFK), London

Heathrow International Airport (LHR), Dubai International Airport (DXB) and Beijing

International Airport (PEK). A hub and spoke structure is to be employed, with the reliance on

the growing number of Low Cost Carriers to provide regional connections.

Initial Sizing

The initial sizing process is performed using an iterative approach. First, a database of similar

aircraft is compiled. Next, a least squares regression is used on the database to find the

coefficients of the basis function. The design mission is then used to determine Wf/W0. The

sizing equation was then iterated to find the final weight of the aircraft. The results of this initial

sizing are plotted against the database aircraft to validate the results. These results are

summarized in the appendix.

In the following section, constraint studies are performed to determine the limits that are

placed upon the Super Sonix Aircraft during the operations to and from hub airports.


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Constraint Analysis

During the operation of the Super Sonix aircraft, there are several performance constraints that

the aircraft needs to meet in order to successfully carry out its mission, perhaps even to be

airworthy in the first place. Several operational conditions are considered in this analysis: 1-g

supersonic steady level flight, 2-g subsonic maneuver, takeoff and landing ground roll during a

hot day and short runway, and a 3% second-segment climb gradient. Several values are based

on the design mission (which is included in the appendix). The constraint diagrams aid in

identifying a viable range of operating wing loading and thrust-to-weight ratios.

1-g Supersonic Cruise

The ability to maintain supersonic cruise is vital to the success of the Super Sonix aircraft. The

derivation of the equation that relates thrust to weight TSL/W0 to wing loading W0/S begins with

the specific excess power relation:

Equation 1

V is the aircraft velocity, T is the thrust at altitude, D is the drag force, and W is the current

aircraft weight. We note that the thrust of a turbofan engine decreases with altitude by the

following definition:

Equation 2

Where the thrust lapse rate 𝛼, is approximated as:

Equation 3

For subsonic conditions, the density at altitude 𝜌𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒 can be found from standard

atmospheric tables. However, for supersonic flight, the lapse rate is approximated using an

effective density 𝜌𝑒𝑓𝑓𝑒𝑐𝑡𝑖𝑣𝑒 . This effective density is obtained using normal shock jump

relations, assuming a normal shock in front of the body. This approximation will be justified in a

later part of this report, where it is in fact seen that a bow shock is preferred for boom

mitigation. It is seen that this approximation yields the effective density as seen by the engine

inlet, and as such will be compatible with the estimate obtained for subsonic conditions.

𝑃𝑠 =𝑉(𝑇 − 𝐷)

𝑊=

𝑑𝑕

𝑑𝑡+𝑉

𝑔

𝑑𝑉

𝑑𝑡

𝑇 = 𝛼𝑇𝑆𝐿

𝛼 =𝜌𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒

𝜌𝑆𝐿


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The normal shock jump relation for density yields:

Equation 4

For subsonic speed, drag at any point in flight is given by

Equation 5

where e is Oswald Efficiency Factor typically between 0.7 and 0.85. For supersonic flight, drag

value is estimated using the equation

Equation 6

where

Equation 7

Equation 8

Equation 9

𝐶𝐷0 is approximated to be 0.015 for a “clean jet” aircraft as given by Raymer[1]. The current

weight of the aircraft is obtained by multiplying the takeoff gross weight W0 by the fuel fraction

β. The load factor n is set to 1 for this flight condition. Combining the equations and setting the

specific excess power to zero, we obtain the following relationship between 𝑇𝑆𝐿

𝑊0 and

𝑊0

𝑆:

Equation 10

The variable q is the dynamic pressure set by the cruise velocity and density: 𝑞 =1

2𝜌𝑉2. Table 1

shows flight conditions and aircraft parameters that are used for the 1-g steady level flight.

𝜌𝑒𝑓𝑓𝑒𝑐𝑡𝑖𝑣𝑒

𝜌𝑎𝑙𝑡𝑖𝑡𝑢𝑑𝑒=

𝛾 + 1 𝑀2

𝛾 − 1 𝑀2 + 2

𝐷 = 𝐶𝐷0 +𝐶𝐿

2

𝜋𝐴𝑅𝑒+ 𝐶𝐷𝑤

𝑞𝑆

𝐷 = 𝐶𝐷0 + 𝐾𝐶𝐿2 + 𝐶𝐷𝑤

𝑞𝑆

𝐾 =1

𝜋𝐴𝑅 1 + (𝑀2 − 1)

𝜋𝐴𝑅

4

2

𝐶𝐷𝑤= 𝐸𝑊𝐷

1 − 0.386 𝑀 − 1.2 0.57 1 −𝜋ΛLE

0.77

100

9𝜋

2

𝐴𝑚𝑎𝑥

𝜄

1

𝑆

𝐶𝐿 =𝑛𝑊

𝑞𝑆

𝑇𝑆𝐿𝑊0

=𝛽

𝛼 𝑞

𝛽 𝐶𝐷0

+ 𝐶𝐷𝑤

𝑊0 𝑆 +

1

𝜋𝐴𝑅𝑒 𝑛𝛽

𝑞

2

𝑊0

𝑆 +

1

𝑉

𝑑𝑕

𝑑𝑡+

1

𝑔

𝑑𝑉

𝑑𝑡


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1g steady, level flight, M =1.8 @ h=50000ft

Altitude 50000 ft

0.000362 slug/ft3

0.44

0.91762

M 1.8

V 1742.13 ft/s

CD,0 0.016

AR 2.1

K 0.4037

EWD 1.9

LLE 60 deg

dmax 13 ft

Length 180 ft

Wing Area 3335 ft2

CD,W 0.00404

q 549.34 lb/ft2

n 1

dh/dt 1.67 ft/s

g 32.17 ft/s2

dV/dt 0 ft/s2

Table 1: Flight Conditions and Aircraft Parameters for 1-g steady level flight

Flight condition such as Mach number, altitude and aircraft parameters such as aspect ratio

(AR), leading edge sweep (LLE), maximum diameter (Dmax), length and wing area are

determined from sizing exercises and the mission profile. The value of 𝐶𝐷0is increased slightly to

0.016 to provide a conservative estimate. The value of is chosen based on the design mission

of the Super Sonix aircraft, specifically the fuel fraction at the start of cruise.

2-g subsonic Maneuver

The subsonic 2-g maneuver is of utmost importance especially in the design mission for Super

Sonix. Inherent in the mission of hub to hub service lies the fact that the aircraft needs to

operate in and out of high-traffic airports. A 2-g maneuver may be required to initiate hard

turns and/or evasive maneuvers to clear traffic, or clear the airspace quickly about these

airports. For the 2-g subsonic maneuver, the equation for supersonic flight is used, without any

wave drag contribution: that is, 𝐶𝐷𝑤=0. The load factor is set to 2, for a 2-g maneuver. The Table

2 below shows flight conditions and aircraft parameters that are used for the 2-g subsonic

maneuver.


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subsonic 2g maneuver, V =422ft/s @ h=10000ft

Altitude 10000 ft

0.001755 slug/ft3

0.79

0.963501

V 422 ft/s

CD,0 0.016

e 0.8

AR 2.1

q 156.27 lb/ft2

n 2

dh/dt 0 ft/s

g 32.17 ft/s2

dV/dt 0 ft/s2

Table 2: Parameters for subsonic 2-g maneuver constraint

An altitude of 10000 ft is chosen to represent the maximum altitude at which such a maneuver

may be required. Flight speed is set at 250 kts, which relates to 422 ft/s True Air Speed (TAS).

Aircraft parameters are, as before, determined based on sizing, and the fuel fraction is derived

using the takeoff and climb portions of the design mission.

An important aspect to be aware of is the maximum lift coefficient required to perform the

maneuver during subsonic flight. As wing loading increases, higher lift is required to perform

the 2-g maneuver. The wing loading can only increase until a point where the CLmax of the lifting

surface is reached, and any further increase in wing loading would lead to the stalling of the

wing. A CLmax, subsonic is assumed to be 1.2, a number that is chosen to be lower than the

maximum lift coefficient at takeoff CLmax,TO while being within reasonable limits. This places an

upper limit on the achievable wing loading, manifesting itself on the constraint diagram as a

vertical line.

Takeoff and Landing Ground roll

Takeoff constraints are implemented through the following equation:

Equation 11

Where STO is the takeoff ground roll distance. The equation is obtained assuming that thrust is

of a much larger magnitude than aerodynamic and ground friction drag. It also takes into

account the performance with one engine inoperative during takeoff. Takeoff velocity VTO is

assumed to be 10% higher than Vstall.

𝑇𝑆𝐿𝑊0

= 1.1 2𝛽2

𝛼𝑔𝜌𝐶𝐿𝑚𝑎𝑥𝑆𝑇𝑂

𝑊0

𝑆


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Landing constraint is dependent on the availability of thrust reversers on the engines. For the

present analysis, thrust reversers were used to aid in braking of the aircraft during its ground

roll.

Equation 12

αrev is the ratio of the reversed thrust to the static sea level thrust, TSL. The friction coefficient, μ

is typically set to be from 0.3 to 0.5 with the brakes being applied. SL , landing pavement length

is mandated to be one third of the available runway pavement, the remaining two thirds for

safety considerations.

Takeoff and landing constraints are imposed at particular airports which Super Sonix serves. At

present this analysis will include extreme conditions at hub airports, two of the most restrictive

of which are Dubai (DXB) and New York (JFK). Tables 3(a) and 3(b) below show the parameters

that are used for operations out of Dubai:

landing ground roll 4374ft @ h = 34ft,43° hot day[DXB]

Altitude 34 ft

ΔT 43 :R

0.002193 slug/ft3

0.939672

rev 0.4

1

CL max land 2

g 32.17 ft/s2

μ 0.3

SL 4374 ft

Table 3(a): Landing constraint parameters for DXB airport Table 3(b): Takeoff constraint parameters for DXB airport

Table 3: Landing and Takeoff parameters for DXB airport

Based on the typical data, rev was assumed to be 0.4. The fuel fraction was taken to be 1 for

these maneuvers. In landing case, this implies that the aircraft can land immediately without

burning any fuel. CLmax,land and CLmax,TO were estimated to be 2 and 1.5 respectively[2]. The

standard atmosphere temperature at sea level is 59:R; however Dubai presents a hot climate of

108:R due to its geographical location[9]. Since there is a significant variation of density at this

temperature (and a corresponding degradation in engine performance), a ΔT from standard

temperature of 43:R is incorporated into this performance constraint.

𝑇𝑆𝐿𝑊0

=𝛽2

𝛼𝑟𝑒𝑣

1.15 2

𝜌𝐶𝐿𝑚𝑎𝑥𝑔𝑆𝐿

𝑊0

𝑆 − 𝜇

takeoff ground roll 10000ft @ h = 34ft,43° hot day[DXB]

Altitude 34 ft

ΔT 43 :R

0.002193 slug/ft3

0.939672

1

CL max,TO 1.5

g 32.17 ft/s2

STO 10000 ft


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The parameters that are used for JFK are as follows:

Table 4(a): Landing constraint parameters for JFK Table 4(b): Takeoff constraint parameters for JFK

Table 4: Takeoff and Landing parameters for JFK

The constraint in this case is the short takeoff and landing ground roll. This represents the

shortest runway at JFK airport[8,10], and meeting this constraint will allow us to utilize all of the

runways not only at JFK but all other hub airports as well.

Second Segment Climb Gradient

Second segment climb gradient is a Federal Aviation Authority (FAA) requirement specified in

the FAR 25.111[11]. It dictates the rate of climb for an aircraft from takeoff at 35 feet to clear

400 foot high obstacles along the climb out flight path as a function of the number of engines it

possesses. The second segment climb gradient is defined by:

Equation 13

Where T is the thrust at takeoff, W is the aircraft weight, and L/D is the lift-to-drag coefficient.

Equation 14

𝐶𝐺𝑅 =𝑇

𝑊−

1

𝐿 𝐷

𝑇𝑆𝐿𝑊0

=𝛽

𝛼 𝑁 − 1

𝑁 𝐶𝐺𝑅 +

1

𝐿 𝐷

landing ground roll 2800ft @ h = 13ft,25° hot day[JFK]

Altitude 13 ft

ΔT 25 :R

0.002267 sl/ft3

1

rev 0.4

1

CL max land 2

g 32.17 ft/s2

μ 0.3

SL 2800 ft

takeoff ground roll 8000ft @ h = 13ft,25° hot day[JFK]

Altitude 13 ft

ΔT 25 :R

0.002267 slug/ft3

1

1

CL max,TO 1.5

g 32.17 ft/s2

STO 8000 ft


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Substituting L/D=CL/CD and their respective expressions, we arrive at the following relationship

of thrust-to-weight ratio and climb gradient.

Equation 15

Equation 15 assumes one engine is inoperative during the climb (N being total number of

engines). It is noted that Δ𝐶𝐷0 represents the additional parasite drag due to deployment of

flaps during takeoff. Typical value for this variable is 0.0065 for plain flaps and 0.0033 for

slotted flaps. The parameters used in the analysis are as follows:

second segment climb gradient of 3% above 34ft,43°

hot day

Altitude 34 ft

ΔT 43 :R

0.002193 slug/ft3

0.939672

1

CL max TO 1.5

g 32.17 ft/s2

N 4

CGR 3 %

CD,0 0.016

AR 2.1

eTO 0.6

Δ CD,0 0.0033

Table 5: Second segment climb constraint parameters

The DXB airport is a constraint here due to its hot climate. As the Super Sonix aircraft has 4

engines, the CGR is mandated to be 3% [11]. The determination of the number of engines will be

discussed in a later part of this report that deals with engine selection analysis. The Oswald

efficiency factor is reduced from the earlier value of 0.8 to 0.6 to account for the deployment of

flaps.

The following figure shows the constraint diagram that is obtained from the above-mentioned

analyses.

𝑇𝑆𝐿𝑊0

=𝛽

𝛼 𝑁 − 1

𝑁 𝐶𝐺𝑅 +

𝐶𝐷0+ Δ𝐶𝐷0

𝐶𝐿𝑇𝑂+

𝐶𝐿𝑇𝑂𝜋𝐴𝑅𝑒𝑇𝑂


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Figure 1: Constraint Diagram

The performance of the Super Sonix aircraft is primarily constrained by the supersonic 1-g

cruise, second segment climb and subsonic 2-g maneuver. The current design point lies at or

above TSL/W0=0.45 and W0/S=70 lb/ft2. Compared with historically averaged data of

TSL/W0=0.38 and W0/S=85 lb/ft2 (from SRR initial sizing analysis), it is observed that these values

are within acceptable limits.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

40 60 80 100 120 140 160

Th

rust

-to

-We

igh

t

Wing Loading (lb/ft2)

1g steady, level flight, M =1.8 @ h=50000ft

subsonic 2g manuever, V =422ft/s @ h=10000ft

takeoff ground roll 10000ft @ h = 34ft,43° hot day[DXB]

landing ground roll 4374ft @ h = 34ft,43° hot day[DXB]

second segment climb gradient of 3% above 34ft,43° hot day

takeoff ground roll 8000ft @ h = 13ft,25° hot day[JFK]

landing ground roll 2800ft @ h = 13ft,25° hot day[JFK]


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Table 6: Assumptions Sensitivity Table

The table above organizes certain key assumptions that are used in the constraint analysis in

order of the sensitivity of the solution on the values. Notably, small variations in CD,0 and α

cause significant changes in the constraint plots. The nominal cruise climb capability, on the

other hand, has little to no appreciable effect on the behavior of the graphs. Further analysis is

required, for example the generation of a carpet plot, using the predicted values of TSL/W0 and

W0/S as starting points to determine actual aircraft performance.

Concept Selection

The concept generation and selection process is a very important part in the design process.

The objective of this process is to develop a concept that meets the customer requirements and

well as the needs of the design mission. In order to create the best concept for the specific

mission, team Super Sonix went through a comprehensive concept selection process outlined

by the flowchart in the Appendix. Each of the steps is discussed in detail.

Brainstorming

To come up with a comprehensive database of design concepts, each member in team Super

Sonix generated ideas and sketches of what they considered to be a good concept for the

mission. The concepts generated from the brainstorming session are included in the appendix.

Various design aspects such as engine type/placement, wing and fuselage shapes, number/size/

location of control surfaces and the cabin layout were described.


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Pugh’s matrix

In order to make a qualitative comparison between various concepts, Pugh’s method was used.

This method is an iterative process where the different concepts were compared with a

predefined datum or baseline concept, for a set of design criteria. For the selection of the Super

Sonix concept, two runs of the Pugh’s matrix had to be made.

In order to generate a good concept, a good set of design criteria needed to be chosen. The list

of design criteria used by team Super Sonix was generated using the Quality Function

Deployment and the mission of the aircraft. A list of ten design criteria was used for the 1st run

of the Pugh’s matrix. Using the various ideas generated from the brainstorming session, 8

different concepts were generated and these were compared to the Concorde. These concepts

are shown in Figure 2.

Figure 2: Concepts used for the Pugh’s method analysis

For the 1st run of the Pugh’ matrix analysis, the Concorde was chosen as the datum because it is

the best commercial supersonic transport aircraft that ever flew. Each of the concepts was

compared to the datum for each design criteria and was evaluated on a three level scale –

positive(+), negative(-) and same(s). After the entire matrix had been traversed, the total

number of positives, negatives and same’s were compiled. Based on the evaluation, three

baseline concepts were selected to undergo the next iteration of the Pugh’s method. These


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were concept 1, concept 5 and concept 6. Table 7 shows the first run and the compiled results

of the Pugh’s matrix.

Table 7: First run of Pugh’s matrix

Hybrid Concepts

From the 1st run of Pugh’s matrix, the positive attributes of the various concepts were

incorporated into the three baseline model to create hybrids. The negative attributes in these

concepts were reduced or eliminated. A high performance aft arrow wing concept retrieved

from [12] was included in this list because of its predicted high performance supersonic

characteristics. Different technologies were incorporated into each of the different concepts so

that these technologies could be compared. A very qualitative assessment of various

technologies was made at this point. Viability and actual performance of the concepts need to

be tested.


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Hybrid concept 1:

Figure 3: Hybrid concept 1

Hybrid concept 1 looked like a scaled down version of the Concorde. The droop nose was

missing and was replaced by a hybrid nose design, which reduced the number of moving parts,

and hence the weight of the aircraft. It also decreased complexity and increased safety of the

aircraft. The sonic boom mitigation was taken care by the nose design which created a series of

weak compression shocks similar to the engine inlet of the SR-71. Canards and horizontal

stabilizers are missing on the aircraft. However, like on the Concorde, inner, middle and outer

elevons are used to control roll, pitch and yaw motion for the aircraft. The concept has a delta

wing with the wing root at mid fuselage. It also has an anhedral angle and droops along the

chord, from leading edge to the trailing edge. The engine location is aft of the delta wing and is

placed under the wing.


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Hybrid concept 2


Hybrid concept 2 has an over wing engine which should help reduce the effect of engine noise.

This is located on the aft end of the delta wing. The wing has no anhedral or dihedral angle on

this concept. Nose shaping, similar to the one on the F-5 shaped sonic boom demonstrator [13],

would be used to reduce the sonic boom overpressure. Canards with a dihedral angle would be

present on the upper section of the fuselage.


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Hybrid concept 3


The 3rd hybrid concept made use of the Gulfstream / NASA Quiet Spike™ technology for sonic

boom mitigation. A major difference with this concept is that it has under wing inlets for the

engine and over wing outlets, which is similar to the YF-23. Canards with an anhedral angle are

its primary control surfaces. This concept has a delta wing mounted on the bottom section of

the fuselage. The main wing has a dihedral angle.


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Aft arrow wing concept

Figure 6: Aft Arrow wing concept [12]

The aft arrow wing concept [12] was included in the list of hybrid concepts for the next run of the

Pugh’s matrix because of its expected performance. The aircraft geometry helps to reduce the

sonic signature of the aircraft. Having swept canards also helps bring the sonic signature from

an N-wave to a Plateau wave form. Another major difference in the aircraft geometry is the

inclusion of two vertical tails on either wing. The concept has a highly swept delta wing, with a

dihedral angle, on the bottom of the fuselage.

Although it helped to increase options for the final concept, choosing between various

technologies on a qualitative basis was difficult. For this reason, further research was done into

each of the design criteria used for comparison of aircraft concepts. This included making a

detailed description of each design criteria and the systems it related to. A table of the design

criteria along with its detailed description is in the appendix.


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The 2nd run of the Pugh’s matrix used the aft arrow wing concept as the datum. The three

hybrid concepts were compared to each of the design criteria (using the detailed description

table). Like the previous run, the positives, negatives and the same’s were compiled.

2nd run AFT ARROW

WING CONCEPT HYBRID CONCEPT 1 HYBRID CONCEPT 2 HYBRID CONCEPT 3

D

A

T

U

M

SONIC BOOM - - - SUBSONIC NOISE s + s

CONTROL SURFACES - s s TURN AROUND TIME s + +

AIRPORT COMPATIBLE + + - SAFETY s + -

EASE OF MANUFACTURE - s - EMPTY WEIGHT + + s

COST - - -

+ 2 5 1

- 4 2 5

s 3 2 2

Table 8: Second run of Pugh’s matrix

From the 2nd run of the Pugh’s matrix, it was possible to see how the hybrid concept 2 had

better design characteristics for the mission when compared to the other three concepts. For

the final design, the positive attributes of this concept were enhanced and the negatives were

reduced. Using this information and the positive aspects of the other concepts, the final hybrid

conceptual design was created.


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Figure 7: 3-view and isometric drawing of final design

Figure 7 shows the 3-view and isometric drawing of the final design concept. The final concept

selected from the Pugh’s matrix analysis had a blunt nose design. This creates a bow shock in

front of the aircraft, which will eliminate shock effects on the wings of the aircraft and also

reduces the effect of coalescent shocks along the bottom of the aircraft. The aircraft concept

had a double delta wing with aft mounted engines. The concept also had swept canards like the

aft arrow wing concept.

Certain aircraft characteristics, such as the engine placement (over/under wing) and vertical

stabilizers- one on fuselage or two on wings had not been decided. Further research needed to

be done before a definitive choice was made. The effects of dihedral or anhedral angles on the

wings and canards needed to be examined as well. Door locations on the aircraft are yet to be

decided.


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Advanced Technologies

The Super Sonix aircraft will need to balance three major design qualities to be successful.

These qualities are: low sonic Boom, efficient supersonic cruise, and advance engine design to

reduce noise and pollution. There are many ways that each design quality can be met. It is

important to note that not all design qualities can be met to 100%, this is primarily due to

mission requirements and a coupled relationship between the different design solutions. What

follows is a description of the different ways that each design quality will be met.

Boom shaping technologies

A modern supersonic aircraft will need to not only be able to fly fast, but also be able to fly fast

quietly. Quiet supersonic flight is a challenge because of the many different technologies that

must be used in tandem to reduce an aircraft’s sonic boom signature. There are two different

types of sonic boom signatures, an N-wave and a plateau wave. An N-wave sonic boom

signature is created when aircraft develops lift rapidly. This can be in the form of high aspect

ratio wings with low sweep. A plateau wave is created when lift is developed gradually and or a

blunted nose configuration is used. The plateau wave is more desirable because the sonic boom

energy is spread out over a longer period of time, and thus there is a lower overpressure and a

smaller sound signature. Thus, it can be concluded that the objective of the Super Sonix aircraft

is to produce plateau wave sonic boom signatures with significantly lower overpressures. The

methods in which this objective will be achieved are presented below for each design

configuration.

Nose Design

Blunt Nose

The implementation of a blunt nose on the Super Sonix aircraft will result in a bow shock in

front of the aircraft. Normally this would not be desired; however the objective of producing

this bow shockwave is to reduce coalescence shockwaves produced under the aircraft. These

coalescence shockwaves are stronger and produce the N-wave pressure signatures seen on

earlier supersonic aircraft. Using a blunt nose configuration results in a plateau wave signature,

thus lowering the sonic boom effects. The effects of this technology on the overpressure

produced by the aircraft are presented below in Figure 8.


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Figure 8: Blunt nose effect on boom overpressure [23]

It is important to note that implementing a blunt nose configuration on the Super Sonix aircraft

will result in a large increase in wave drag. The wave drag penalty is one of the main reasons

why this has not been implemented in previous designs. Thus, another big challenge of the

Super Sonix aircraft is to reduce wave drag in other parts of the aircraft by careful fuselage and

tail design. These concepts are presented in more detail later.

Nose Keel

Another nose design that may reduce sonic boom sound signature includes installing a keel at

the front of the aircraft. The keel has a channel down the centerline of the device that allows

air to flow in-between two panels. Then high voltage electricity is arched across the air to ionize

the incoming air. This process produces a plasma out in front of the aircraft that when

implemented will increase the effective length of the aircraft [20]. While this nose design results

in a significant overpressure drop across the aircraft, there are many environmental issues

associated with this design. Creating plasma at high altitudes results in the production ozone

which is a greenhouse gas. Given the altitude that this gas is produced at, it will have much

more severe environmental consequences than if the same device was implemented on sea

level. Furthermore, there are issues associated with how one might


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Dihedral Angle

Using dihedral angle on any of the lifting surfaces on the aircraft can have a profound impact on

the sonic boom signature [21]. Tilting the lifting surface upward a set angle results in lift produce

that is on multiple planes. This has the effect of increasing the effective length of the aircraft,

thus smoothing out the effective area distribution. An example of dihedral angle is shown in

Figure 9.

Figure 9: Dihedral angle design

In Figure 9, the angle gamma is the amount that the wings are tilted upward. Full

implementation of this ideal over an entire lifting surface may not be feasible; however,

sections of the main wing can have dihedral angle. This will help with increasing the effective

length of the aircraft while slowly building a lift producing surface. While using a dihedral angle

on the wings may have several advantages for supersonic aircraft, there are several issues

associated with using it in the subsonic regime. During subsonic flight, wings with dihedral angle

will experience wash out due to lateral flow instabilities in the airflow. Furthermore, increasing

dihedral angle reduces planform area of the wing and introduces structural issues in wingbox

design.

Effective Area distribution

Aerodynamic Cowling

Carefully shaping the area distribution is a primary design consideration for the Super Sonix

aircraft. Research has found that for sonic boom minimization it is important to have a rapid

buildup of volume at the nose of the aircraft [17]. This idea was introduced earlier in the blunt

nose design section. A nose design that involves a rapid buildup of volume at the nose of the

aircraft will produce significant wave drag, and thus the rest of the aircraft must compensate by


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having a low wave drag design. These ideas are implemented by using a varying diameter

fuselage, which if designed improperly can lead to severe weight penalties. The solution to this

problem is to use a classic cylindrical fuselage design, while varying the effective diameter of

the body by using cowling held in place by light-weight trusses. The spaces between the

fuselage cylinder and the outer cowling truss will be filled with avionics equipment, fuel,

insulation, or some form of payload. When the fuselage travels over the wing structure, the

fuselage cowling will shrink to fit the fuselage cylinder, thus smoothing out the area distribution

of the aircraft. An example of this is presented below in Figure 10.

Figure 10: Non-Axisymmetrical fuselage shape [22]

From Figure 10 it can be seen that the supersonic aircraft concept has a carefully shaped

fuselage to reduce the drag and sonic boom signature.

Engine Nacelle Placement

The placement of the engine nacelles can contribute significantly to the overall area of the

aircraft. To help smooth the area distribution with the engine nacelles, one can add carefully

shaped cowlings to the engines. If the engines are placed underneath the fuselage and a part of

the tail of the engine is stuck out behind the wing, the cowling on the engine could be larger

close to the end of the wing, to help smooth the effective area of the aircraft.

Canards

Another component of the Super Sonix aircraft is the canards located at the nose of the aircraft.

These lifting devices help add longitudinal stability to the aircraft and are the primary elevator

control surfaces of the aircraft. For these devices to be efficient on the Super Sonix aircraft,

they will need to have large dihedral angle, such that the lifting area contribution is gradual and

the sonic boom signature is carefully crafted.


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Efficient supersonic cruise

The majority of a supersonic aircraft’s design mission is spent in the supersonic flight

configuration. This means that the Super Sonix aircraft needs to be efficient in the supersonic

flight regime. The key design features that enable an aircraft to perform efficiently in the

supersonic flight regime are: high aspect ratio low sweep wings resulting in natural laminar

supersonic flow. It is unlikely that a natural laminar supersonic flow wing can be created,

however getting the flow over the wing to be close to laminar will result in significant drag

reduction. This wing design is in direct contradiction to what is required for low sonic boom

signature. For low sonic boom sound signature the Super Sonix aircraft needs to develop lift

more gradually; thus highly swept wings and dihedral angle are major design requirements. The

wing design for efficient supersonic flight and low sonic boom signature are fundamentally

different, and thus a balance needs to be reached where the Super Sonix aircraft can fly quietly

at supersonic speeds.

Active Flow Management

Because the Super Sonix aircraft will have a hybrid wing design that balances supersonic flight

efficiency and quiet supersonic flight operations, some design solutions may not be efficient in

during the entire operating mission of the aircraft. For that reason, active flow control will be

used to help keep the flow over the wings attached to the body, and thus improve lift. An

example of where this technology might be needed is during takeoff and landing.

Engine Technology

Correctly pairing an engine to an aircraft is one of the most important steps in the

conceptualization process. For this purpose, a database of over 200 engines was created;

outlining details about each engine and the aircraft that it was used in. The engines were then

filtered into two categories, subsonic and supersonic engines. While some engines that operate

in the subsonic regime can fly supersonic, the majority of the engines put into the subsonic

category were high-bypass turbo fans. This is primarily due to the size of the engine and design

of the engine inlet. Therefore, all of the engines considered for the Super Sonix aircraft were

medium bypass turbofans, low bypass turbofans, and turbojets. The engines from this set were

then filtered for the proper thrust profile that fit the preliminary sizing studies. Figure 11 shows

the flow chart detailing the engine selection process. The engine results from this process are

listed in Table 9.


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Figure 11: Engine Database

Manufacturer Designation Dry

Thrust Weight, dry (lb)

Aviadvigatel (Solovyev)

D-30F6 20944 5,326

GE F101-102 17390 4,460

GE/RR F136 24210 4,300

P&W F135 23852 3,750

P&W PW1000G 33393 8,885

Samara NK-25 41900 6,283

Samara NK-321 30843 7,588

Table 9: Viable supersonic engines [5]

Out of the engines listed, the Samara NK-321 is thought to be the best engine that meets many

of the requirements. The reason is due to the fact that the Samara NK-321 has supersonic

capabilities, they have also been used in commercial supersonic transport with the TU-144.

From the sizing code, it was determined that the aircraft will require 107,724 lbf of thrust. With

the thrust rated at 30,843 lbf, the Super Sonix aircraft will have a 4-engine configuration.

Apart from NK-321, another engine that is under consideration for Super Sonix aircraft is QSST

engine. The engine is capable of producing 33,000 lbf of thurst and will be available by 2010.

With the latest, state of the art technology and significant reduction of noise, this engine is a

good option and fits perfectly into Super Sonix engine requirements. As of now, the


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development of the engine is contracted to General Electric, Pratt & Whitney, and Rolls-Royce

and detail specifications of the engine are currently not available.

There are a few aspects of note in regards to the Samara NK-321. Firstly, the Samara is a

Russian engine and there have been problems in the past exporting TU-144 for the SST studies

because the engines are export controlled. Secondly, the Samara engine is also very noisy.

Furthermore Samara engine is quite old, entering service in 1987. Figure 12 below illustrates

the Thrust to Weight ratio of the 7 engines against the manufactured year. It is noted that there

are several engines that have yet to be produced (P&W 1000G, P&W F135 and the GE F136).

Therefore, even though some of those engines are expected to produce compatible thrust, only

the Samara NK-321 will be used as a baseline engine with additional technological

improvement as of today.

Figure 12: Thrust-to-weight ratios for viable engines

Modifications for Samara NK-321 Engine

The first technological concept in regards to reducing noise is employing the variable engine

cycle [3, 4], an engine that is designed to operate efficiently under mixed flight conditions. One

concept is the tandem fan engine where the engine has two fans, both mounted on the low

0

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1975 1980 1985 1990 1995 2000 2005 2010 2015

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Samara NK-321

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P&W 1000G

Aviadvigatel D-30F6

GE F136


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pressure shaft with an axial gap between the units. In normal flight, the engine is in 'series'

mode, meaning that the flow leaving the front fan passes directly to the second fan. This works

like a conventional turbofan. But during takeoff, climb out, final descent and approach, the

front of the fan is allowed to discharge automatically through a nozzle on the underside of the

nacelle. This will allow air to enter the rear fan and will flow to the rest of the engine. This is

called the 'parallel' mode. In parallel mode, it increases airflow of the engine at thrust which

leads to a lower jet velocity and quieter engine. One downside of the variable engine cycle is

the complexity of it, which can lead to difficulty in manufacturing and also servicing.

As has been discussed above, the Samara NK-321 engine is very noisy. Papamoschou and

Debiasi [14] did a study in order to reduce engine noise and in order for the NK-321 to be a

viable base engine, this study has to be taken into account. It was found that employing an

eccentric exhaust configuration will be beneficial to reducing engine noise. They studied the

effects of engine configurations have in regards to noise. Figure 13 shows the 4 configurations

that they propose would help with engine noise. The first one, B03-MIX was the baseline

engine, a military turbofan engine with BPR of 0.3 at mach 1.6 and height of 16,000 m. The

notation MIX stands for mixed flow engine. The second configuration, B16-MIX, is a mixed flow

turbofan. It was suggested an increased bypass ratio from 0.3 in the baseline engine to 1.6. This

configuration will yield less noise but will not have an efficient supersonic performance. The

third configuration, B16-SEP-COAX, is an engine with bypass ratio of 1.6. It has separate flow

and has a coaxial configuration. A separate flow is normally better than mixed flow because

the bypass exhaust design will reduce the mach wave emission from the core and will therefore

produce less noise. This configuration is normally used in current subsonic aircrafts. The last

configuration, also said to be the best solution to the noise versus performance tradeoff

problem, according to Papamoschou and Debiasi, is the eccentric engine configuration. This

configuration works by maximizing the theory of mach wave elimination discussed above. The

eccentric configuration minimizes the mach wave and therefore produces less noise.

The comparison between the four was based on the assumption that all engines produce the

same supersonic cruise thrust, and have the same core characteristics. A solution that

Papamopschou and Debiasi proposed to reduce noise was to have an intermediate bypass ratio

in order to satisfy the need for a high bypass ratio on landing/takeoff, which would lead to

quieter noise and also a low bypass ratio during supersonic cruise for efficient cruise.


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Figure 13: Proposed engine exhaust configurations [14]

The main source of the engine noise is Mach wave radiation. Mach waves are shock waves that

are generated by supersonic motion of turbulent eddies relative to the air. It can be assumed

that the velocity of the eddies will be close to 80% of the jet velocity engine. The problem with

such a high eddy is that it will create a system of shocks that will produce noise that can be

heard form a very far distance. A solution to this problem is known as the mach wave

elimination method. This works by preventing the formation of shocks by making sure the

eddies are subsonic. To make the eddies subsonic, a secondary flow was created to surround

the jet so it will reduce the velocities of the eddies.


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Figure 14: Directivity of Overall sound pressure level [14]

Figure 14 shows the difference of the Overall Sound Pressure Level for the baseline engine with

the three derivatives. It is seen that the sound levels are decreased appreciably with the new

derivatives. Figure 15 below is a comparison of the Perceived Noise Level (PNL) time histories

between the different configurations on a given flight path. All engines produced the same

thrust of 126 KN. As can be seen from the plot, the eccentric exhaust configuration will be 13dB

quieter than the baseline engine, while the mixed and coaxial configurations are 5 and 6 dB

quieter than that of the baseline.


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Figure 15: Perceived Noise Level [14]

There are two ways to go about selecting the engine. The first will be to use the Samara NK-321

as a base engine and improve that engine with the technology discussed above. The result will

be the Samara engine with less noise and that will be good given that the Samara has been

tried and tested in a supersonic commercial aircraft. Another route would be to consider an

engine still being developed today that will be available in 2020. The QSST engine discussed

earlier would be a good option assuming it will produce a similar Thrust to what is required.

Details for the QSST engine are unavailable at the moment so it is not possible to make a

selection basing on so little information today.


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Cabin Layout

Several factors contribute to the cabin layout of the Super Sonix aircraft concept. Passengers

desire a comfortable, spacious seating environment, but it is important to meet aircraft design

constraints and FAA requirements as well. Passengers will be paying a hefty sum of money to

fly supersonic on this aircraft, but it is debatable to what extent a passenger will hold seating

layout a priority in their travel with such short flight times.

Business analysis thus far has shown that the target customer base values time a bit more than

comfort, but at the same time people who can afford to pay for time savings like this generally

may expect more from their seating. There is a delicate balance between passenger comfort,

safety, efficiency, and aircraft size and shape.

Figure 16: Passenger Seating Options (pitch)

0

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120

Flat Bed Lie Flat Recline Suite

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Figure 17: Passenger Seating Options (width)

After analyzing historical data (Figures 15 and 16 above), it is apparent that typical subsonic

transports run by the airlines have a maximum seat pitch in first class of about 40 inches,

though it is generally much less. Larger specialty aircraft with long flight times may have seat

pitches between 55 to even 100 inches or more with fully reclining “sleeper” seats or suites.

Generally, supersonic flights will be no longer than 4 hours, so traditional reclining seats appear

to be the most reasonable option. For coach, the maximum seat pitch is around 36 inches but

the same idea still applies. These rough numbers give a general idea of how long the seating

section would be. Given the overall aircraft length from sizing, the length seems practical.

Figure 18: Passenger Seating vs. Overall a/c Length

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Flat Bed Lie Flat Recline Suite

Seat

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90 ft


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As shown in Figure 17 above, the seats would take up 90 feet of the 180 feet of the aircraft

length. Because of aerodynamics requirements the aircraft length was extended to 180 feet,

giving a cabin that is just slightly more than half the length of our aircraft. With all this extra

length in the fuselage, the positions of major aircraft components such as fuel tanks and

landing gear can be adjusted to meet the c.g. requirements outlined in this report.

Figure 19: Seating Layout with Galleys and Restrooms

The current seating configuration gives first class passengers 7 rows of 2 seats, and gives coach

passengers 15 rows of 3 seats for an overall total of 59 passengers. Two flight attendants will

have folding “jump seats” available in the cabin between the two classes.

Other parameters, such as aisle and seat width, are also obtained using the method described

above. The first class aisle is 28 inches wide, while the coach aisle is 20 inches wide. First class

and coach seats are also 28 and 20 inches wide respectively and also represent the upper limits

for seating in a typical subsonic transport such as a Boeing 737 or Airbus 320. The seat width

also takes into account the extra fuselage width available with only 2 seats per row in first class.

Overhead compartments, though not yet modeled in CATIA, will fit easily into the 9 foot

diameter inner-cabin cylinder.

Placing the floor 2.5 feet from the bottom of the cylinder allows for a 6.5 foot aisle height. The

inner cabin (inner fuselage) is designed as a basic cylinder for easy cabin pressurization. Though

this inner section is cylindrical, the idea is that outer fuselage geometry can be built around this

cylinder to allow for area rule compliance as well as aerodynamic shaping. Having this inner

cylinder should allow more room for adjustability, and it is desirable from a structural

standpoint.

Also worth mention, the CATIA model of the cabin is easily adjustable for changes in any one of

these parameters. A change in one number in the model will adjust the entire seating


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arrangement for the aircraft. The seating is modeled separately so that it can be adjusted and

imported into the main product file.

Figure 20: Seating Dimensions, Coach (L) First Class (R)


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Sizing Studies

Sizing Approach

The current sizing code being utilized is similar to the previous code, but now has more

components integrated. For example, in addition to the iterative weight calculator, the sizing

code now includes a center of gravity calculator and a boom overpressure calculator. An

important consequence of this integration is the ability to quickly alter all of the estimations.

For example, a change in the takeoff weight of the aircraft will result in a change in both the

boom overpressure and the center of gravity. This is represented as a flowchart in Figure 20.

This sizing code results in a slight increase in the estimated takeoff gross weight. It is now

expected to be around 284,000 lb. The results from the current sizing are summarized in Table

10.

Figure 21: Sizing Approach flowchart


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Table 10: Current Sizing Attributes

Component Weights

Aircraft weight statement should list all possible weight within an aircraft. The sum of weight

statement should be equal to the gross takeoff weight. The importance of estimating aircraft

component weight cannot be understated as it is critical to the center of gravity calculation,

structural design, and aerodynamic design. Well estimated aircraft component weights will help

better placement of the wing, horizontal tail, canards, fuel, and engine. The baseline aircraft

used as a comparison was the Concord. However because information could not be found on

the Concord, a database of 16 small and large commercial aircraft component weights was

created.

List of aircrafts in the database includes small commercial aircraft of: Citation 500, MDAT-30,

MDAT-50, F-28, MDAT-70, DC-9-10, 737-200, and 727-100. For large aircraft database it

includes: 727-200, 707-320, Dc-8-62, DC-10-10, L-1011, DC-10-40, 747-400, and SCAT-15. Total

of 16 aircrafts, including mass fraction given: wing system, tail system, body system, alighting

gear system, nacelle system, propulsion system (less dry Engine), flight control system, auxiliary

power system, instrument system, hydraulic and pneumatic system, electrical system, avionics

systems, furnishing and equipment system, air-conditioning system, anti-icing system, and load

and handling system.


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Figure 22: Component weight mass fraction

0

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This aircraft component weight database was set to provide foresight on different component

mass fraction of small versus large commercial aircraft, while providing a range that the

component weight fraction should be within.

Major concern was the compatibility of the aircraft within the database to the Super Sonix.

None of the aircraft in the data base have the ability to fly at Mach 1.6-1.8. The idea of the

database is to create a general range with average and standard deviation of the component

weight fraction with different aircraft.

Figure 21 above show the individual component weight mass fraction in the form of average,

standard deviation above and standard deviation below. This provides insight on the different

range of component mass fraction for the small and large commercial aircraft. The combined

average (third figure from top) for the mass fraction of the two aircraft size is also presented in

the form of average, and plus, minus one STD. Specific values of combined averages can be

found in the appendix. The average of the two is shown again on the very bottom of Figure 21.

Using the database of 16 aircrafts, the combined average was chosen to be used as the

component mass fraction. This provides an estimate for 87% of the aircraft weight. The table

above shown a 95.8% estimate on component weight however, some components of the mass

fraction will not be considered (ex. Horizontal tail) because the design does not include this

component. The table above provides the estimate of each component weight of the Super

Sonix; this is used to calculate the center of gravity and later used to provide insight on aircraft

stability dynamic and control.

Center of Gravity

In order to determine the center of gravity, an excel spreadsheet was created and is shown in

Figure 22. This spreadsheet utilizes the derived component weights of our aircraft. The

weights can either be divided up and entered individually (passengers and seats), or the total

weight of a part can be placed at the center of gravity for that specific part (engines). The

current center of gravity is 97 feet from the tip of the nose. This is roughly 54% of the total

length of the aircraft (180 ft). It should be noted that this estimation does not include all of the

component weights. However, the spreadsheet is currently calculating the center of gravity

using 87% of the gross takeoff weight. All of the major component weights such as wings,

fuselage, fuel, canard, tail, engines and passengers are included in the 87%. Components such

as baggage, air conditioning, and the anti-icing system are not included as of yet due to their

placement not being accurate enough yet. The center of gravity is expected to change slightly

when these components are added. However, it is not expected to change significantly

because the location of most of the missing components is expected to be close to the center of

gravity.


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Figure 23: Center of Gravity Calculator

Shock Prediction

The prediction of the sonic boom overpressure is of a great importance. It is vital for a sound

argument as to the feasibility of allowing supersonic flights over land. The resolution of the

boom overpressure prediction problem involves a fundamental understanding of sonic boom

pressure signatures as well as the comprehension of boom prediction methods and

implementations. Sonic boom signatures are discussed in an earlier section of the report:

Predominant far-field signatures (i.e. the signatures seen on the ground) include N-wave and

plateau shaped characteristics. As noted earlier, the plateau and N-wave signatures have very

different magnitudes and are determined by specific geometry. In order to provide a full

computation of the boom overpressure, two methods are considered to calculate the

overpressure for the N-wave and plateau signatures respectively.

N-wave

N-wave boom overpressure prediction is based on the study performed by H.W. Carlson [16].

Much of the detail required for successful implementation of the “simplified” model is not

available at this preliminary stage of design of the Super Sonix aircraft, therefore the “super

simplified” model is used to calculate bow shock overpressure. In this model, charts and

approximations are used in place of some of the more detailed equations to determine the

shape factor KS and the pressure amplification factor Kp.


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The equation for the overpressure is given as:

Equation 16

The variables are defined as follows:

Variable Definition

M Cruise Mach number W Aircraft weight (MTOW is used for this prediction) l Aircraft Characteristic length (Fuselage length) pv Atmospheric pressure at altitude pg Ground atmospheric pressure (assumed to be Sea level pressure) h Altitude of aircraft above ground (assumed to be Cruise altitude) hv Altitude of aircraft above sea level (assumed to be Cruise altitude) Kp Pressure amplification factor Ks Shape factor (assumed to be similar to Concorde)

Table 11: Definition of variables for overpressure calculation

To use the charts a lift parameter KL is calculated based on the following equation:

Equation 17

Based on this, a shape factor is obtained from Figure 23. It is noted that the curve

corresponding to the heavy bomber shape (highlighted on the figure) is assumed to be the

closest planform shape to Super Sonix. It is used in conjunction with the cruise altitude to

determine the pressure amplification factor.

Figure 24: Shape factor and pressure amplification factor charts

The “super simplified” analysis only depends on certain macro properties of the aircraft, which

are determined from initial sizing studies conducted earlier. The aircraft weight is at the top of

cruise, and is therefore the MTOW (W0) multiplied by the fuel fractions of segments 1 and 2 in


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the design mission profile. The following table displays the input parameter values as well as

the final result:

N-wave signature

M 1.8 W 270000 lb h 50000 ft l 180 ft Pg 2116.2 lb/ft2

Pv 242.35 lb/ft2

b 1.50 Kl 0.015

Ks 0.085 Kr 1.9 Kp 1.13

ΔPmax 1.89 lb/ft2

Table 12: Input Parameters for N-wave overpressure calculation

It is seen that the overpressure is far above the target limit of 0.3 lb/ft2, a fact that should not

seem surprising as N-wave signatures do indeed result in high overpressures. To validate the

model, the parameters for the Concorde are used and an overpressure of 2.02 lb/ft2, and

matches very closely with the actual overpressure of 2 lb/ft2 [17].

Plateau wave

Calculations for the plateau wave are based on work done by R. Seebass and A.R. George [15].

The formal derivation of the equations involves complex Whitham F-functions and only the

resulting equations are presented as follows.

The main equation for prediction shock overpressure pso is shown below:

Equation 18


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It relies on several auxiliary equations that are computed as follows:

Equation 19: Auxiliary Equations for Plateau wave overpressure calculation

In these equations, the parameter H represents a characteristic altitude to act as a basis for

computations. This height is taken to be H=25000ft, a value that has historically been found to

provide accurate results [15]. characterizes the rate of advance of the signal as it propagates

through a real atmosphere. This signal advance is highly nonlinear and therefore must be

referenced to the advance in a homogeneous atmosphere, where it is well understood.

symbolizes this ratio of real advance to homogeneous advance. The input parameters and

computed overpressure are tabulated below:

Plateau Boom Signature

M 1.8

W 270000 lb

h 50000 ft

l 180 ft

Pg 2116 lb/ft2

beta 1.50

alfa 0.74

gamma 1.2

k 2.15

W~ 1.16

Pso 0.70 lb/ft2

Table 13: Input Parameters for Plateau wave overpressure calculation

For the same aircraft parameters and flight conditions, it is seen that the plateau wave

overpressure is predicted to be almost 2.5 times lower, a fact that follows directly from the

discussion earlier. Even though the overpressure of 0.70 lb/ft2 is still higher than the

requirement, it is re-assuring that the move toward a blunt nose profile leading to a plateau

shock signature does indeed have the potential to reduce the boom overpressure considerably.

It is important to mention that these prediction methods are still unrefined, and more accurate

models and programs (such as NASA’s PBOOM program) is to be used in future as the Super

Sonix aircraft become more defined in its characteristics.


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Summary

Compliance Matrix

The compliance matrix was prepared to check the design progress. Recent version of the matrix

can be found in the appendix. Current concept generated has met several requirements

including take off field length and still air range. However, current concept has not yet met the

cabin volume per passenger, cruise efficiency and the sonic boom overpressure. The cabin is to

have at least 50 pax/(ft3/pax); however our current design has 43.1. This is due to having 59

people in such a small cabin. Rearranging the cabin layout slightly is considered to efficiently

use all the volume from the fuselage. The cruise efficiency was roughly calculated by counting

the number of passengers, total fuel weight and the total range. The threshold is 0.3 lb fuel/pax

mi, but the current value came out to be 0.6, which is the twice of the threshold. We are

studying to see if there is a way to improve efficiency. The threshold sonic boom overpressure

is 0.3 lb/ft2. The design needs to meet the threshold value however; the rough calculation from

volume came out to be 0.70 lb/ft2. This value is more than twice of the target value. Currently it

seems that meeting the target value is very difficult. More trade off studies will be completed

to reduce the overpressure produced by the Super Sonix aircraft.

There are some requirements still unmet. Some requirements such as door height and stall

speed have not been found. The level of details needed is higher than the current stage, more

detail specifications studies will be conducted later in the project.


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Walkaround and CAD Model

Figure 25: Walkaround chart

Figure 24 depicts a walkaround chart that is created for the Super Sonix aircraft. CATIA is used

to design the airplane for visual aid and analysis. Previously discussed technologies and

concepts were considered when designing this model. To start, the fuselage was designed to

have different cross sectional area along its length. The fuselage part connected to the nose has

diameter of 10 ft. The diameter gradually increases to 13 ft to about half way of the cabin. The

fuselage then decreases to 9 ft on the back of the cabin section. Tail fuselage is about 60 ft long.

The diameter then quickly decreases to 1 ft, forming concaved shape. Overall, this aircraft is

designed to encompass most of its volume up front in order to reduce sonic boom over

pressure.

The wings consist of two parts, the strakes and main delta wings. Currently, the airfoil selection

hasn’t been done yet, therefore the wing and the strake is assumed to have symmetric airfoil.

This will later be change when further analysis is conducted and the best airfoil for the aircraft

has been selected. The wing has total span of 80 ft and mounted with dihedral angle for


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efficiency during supersonic cruise. The location where the strake and wing meets is arbitrarily

selected since a lot of information about strakes and wings are currently not available. The

strakes are added to produce more lift during subsonic flight and to move the center of gravity

forward. The strakes are also mounted on dihedral angle. Future consideration for this design is

to have all the connections between fuselage, strakes and wings to be smooth. It is important

to have the entire cross sectional area smoothen so drag can be reduced.

The canards are implemented on the airplane as control surface, and to help move the center

of gravity forward. Currently, the canards are mounted on top of the fuselage due to problem

with accessing the door when the gate is used in the airport. The location of canards and

strakes on the fuselage caused the area around door to be small, risking possible collision

between the airport gate and the canard. Further research about the flow effect from the

canard to the wing and the engine inlet will be conducted.

Vertical tail is very similar to Concorde’s. Vertical wing also has symmetric airfoil. The tail is

drawn so that it is swept. Further study is needed find the effect of having vertical tail during

supersonic flight. There are few other concepts, such as having two vertical tails on the both

side of the tip of the wing.

Figure 26: 3-View of Super Sonix Aircraft


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So far, airplane has been modeled to reflect researches currently done. For future work, the

doors are going to be added. The location of door is already decided in the cabin layout section,

so it will be drawn later. The landing gears are in consideration also. The location to put landing

gears based on center of gravity analysis needs to be addressed. The nose is also being

considered to redesign to have more flat bottom. This means, instead of having a symmetric

elliptical cone, it will have asymmetric elliptical cone. This will reduce drag during flight. Finally,

the number of engines and the location engines are in consideration. Further studies needs to

be conducted to find the effect of having the engine on the wing and under the wing.

Next Steps

Given the current status of this project, the next steps will involve optimization. The system

definition review has allowed the Super Sonix aircraft to have its primary characteristics

defined. However at this point it is unknown what relative sizes of each surface or best mass

position will result in the correct configuration to satisfy the design mission. Thus the next steps

of the Super Sonix project is to iterate though sizing variables to reach a final system that

satisfies the design mission. The manner in which this will be completed is outlined in the next

steps flow chart present in the appendix. Some key design qualities that will need to be

modified include: aerodynamic shaping of the aircraft, detailed dynamics and control

calculations, and structural analysis and design.

In conclusion, the Super Sonix aircraft has the key qualities that will allow it to satisfy the

opportunity description. The seating arrangement is sufficient for airlines to make money, the

aerodynamics has qualities that will allow the aircraft to travel supersonic over land, and the

overall design of the aircraft is simple enough to allow for reasonable maintenance. While the

system has been defined, it has not been perfected. In preparation for the system

characteristics review the Super Sonix aircraft will undergo many controlled changes, in an

attempt to find the most efficient design.


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References

[1]. Raymer, D. P., “Aircraft design: A Conceptual Approach”, 4th Edition, AIAA Education

Series 2006

[2]. “Fun Aerospace Questions”,

http://www.adl.gatech.edu/classes/dci/intro/faq_f00.html#CLMax

[3]. Sweetman, B., “USAF Funds Leap-Ahead Jet Technology”

http://www.aviationweek.com/aw/blogs/defense/index.jsp?plckController=Blog&pl

ckScript=blogScript&plckElementId=blogDest&plckBlogPage=BlogViewPost&plckPos

tId=Blog%3A27ec4a53-dcc8-42d0-bd3a-01329aef79a7Post%3A73d6b7cf-a2e8-4efb-

9984-f7494c0747e5, 8/15/2007

[4]. Quiet Supersonic Platform (QSP)

http://www.globalsecurity.org/military/systems/aircraft/qsp.htm

[5]. Meier, N., “Jet Engine Specifications Database”, http://www.jet-engine.net/

[6]. “Tupolev Tu-160 BlackJack”,

http://www.aircraftguru.com/aircraft/aircraft-information.php?craftid=132

[7]. “Fighter Gallery”, http://www.aerospaceweb.org/aircraft/fighter/

[8]. TravelSmart Ltd, “New York Airport Climate and Weather”,

http://new-york-jfk.airports-guides.com/jfk_climate.html

[9]. TravelSmart Ltd, “Dubai Climate and Weather”,

http://www.dubai-dxb.airports-guides.com/dxb_climate.html

[10]. Jeppesen, “Kennedy Intl Airport Chart”, 20 May 2005

[11]. FAA regulations FAR http://ecfr.gpoaccess.gov/cgi/t/text/text-

idx?c=ecfr&sid=be6ffa6d87343c75e7d49e911c95ab06&rgn=div5&view=text&node=

14:1.0.1.3.11&idno=14#14:1.0.1.3.11.2.155

[12]. Carlson, H.W et al, “Application of Sonic-boom Minimization Concepts in Supersonic

Transport Design”, NASA TP D-7218, Langley Research Center, Hampton, VA, June

1973

[13]. Morgenstern, J.M. et al, “F-5 Shaped Sonic Boom Demonstrator’s Persistence of

Boom Shaping Reduction through Turbulence”, Lockheed Martin Aeronautics

Company, Palmdale, CA 93599; AIAA-2005-0012; 43rd AIAA Aerospace Sciences

Meeting and Exhibit - Reno, NV

http://www.adl.gatech.edu/classes/dci/intro/faq_f00.html#CLMax

http://www.aviationweek.com/aw/blogs/defense/index.jsp?plckController=Blog&plckScript=blogScript&plckElementId=blogDest&plckBlogPage=BlogViewPost&plckPostId=Blog%3a27ec4a53-dcc8-42d0-bd3a-01329aef79a7Post%3a73d6b7cf-a2e8-4efb-9984-f7494c0747e5&plckCommentSortOrder=TimeStampAscending

http://www.aviationweek.com/aw/blogs/defense/index.jsp?plckController=Blog&plckScript=blogScript&plckElementId=blogDest&plckBlogPage=BlogViewPost&plckPostId=Blog%3A27ec4a53-dcc8-42d0-bd3a-01329aef79a7Post%3A73d6b7cf-a2e8-4efb-9984-f7494c0747e5




http://www.globalsecurity.org/military/systems/aircraft/qsp.htm

http://www.jet-engine.net/

http://www.aircraftguru.com/aircraft/aircraft-information.php?craftid=132

http://www.aerospaceweb.org/aircraft/fighter/

http://new-york-jfk.airports-guides.com/jfk_climate.html

http://www.dubai-dxb.airports-guides.com/dxb_climate.html

http://ecfr.gpoaccess.gov/cgi/t/text/text-idx?c=ecfr&sid=be6ffa6d87343c75e7d49e911c95ab06&rgn=div5&view=text&node=14:1.0.1.3.11&idno=14#14:1.0.1.3.11.2.155




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[14]. Papamoschou, D. Debiasi, M., “Conceptual Development of Quiet Turbofan Engines

for Supersonic Aircraft”, Journal of Propulsion and Power Vol. 19, No. 2, March–April

2003

[15]. Seebass, R. et al, “Sonic Boom Minimization Revisited”, AIAA-98-2956, 1998

[16]. Carlson, H., W., “Simplified Sonic Boom Prediction”, NASA TP 1122, Langley Research

Center, Hampton, VA, 1978

[17]. Aronstein, D.,C. et al, “Two Supersonic Business Aircraft Conceptual Designs, with

and Without Sonic Boom Constraint”, Journal of Aircraft, Vol. 42 No. 3, May-June

2005

[18]. Jane’s All the World’s Aircraft Engine Database,

http://jawa.janes.com/public/jawa/index.shtml

[19]. Sakata, K., “Supersonic Experimental Airplane (NEXST) for Next Generation SST

Technology” AIAA-2002-0527; 39th Aerospace Sciences meeting & Exhibit - Reno,

Nevada, 2001

[20]. Marconi, F., et al, “Sonic Boom Alleviation using Keel Configuration” AIAA 2002-

0149; 40th Aerospace Sciences Meeting and Exhibit – Reno, Nevada, 2002

[21]. Kroo, I., Sturdza, P., Tracy, R., Chase, J., “Natural Laminar Flow for Quiet and Efficient

Supersonic Aircraft”, AIAA 2002-0146; 40th Aerospace Sciences Meeting and Exhibit –

Reno, Nevada, 2002

[22]. Makino, Y., Suzuki, K., Noguchi, M., Yoshida, K., “Non-Axisymmetrical Fuselage Shape

Modification for Drag Reduction of Low Sonic-Boom Airplane”, AIAA 2003-557; 41st

Aerospace Sciences Meeting and Exhibit – Reno, Nevada, 2003

[23]. Kroo, I. “Unconventional Configurations for Efficient Supersonic Flight”, VKI Lecture

Series on Innovative Configurations and Advanced Concepts for Future Civil Aircraft,

2005

http://jawa.janes.com/public/jawa/index.shtml


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Appendix

Progress Update Chart


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Concept Selection Process


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Initial Concept Drawings


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Detailed Design Criteria

Design Mission Profile

DESIGN CRITERIA DETAILED DESCRIPTION

SONIC BOOM geometry altitude weight wetted area aspect ratio frontal area induced dragcontrol surface

effects

SUBSONIC NOISE engine placement noise mitigation type of engine

CONTROL SURFACES

types of control surfaces static stability number

location of control sufaces

airport compatibility complexity

size of control surfaces weight

TURN AROUND TIME

# of passenger doors # of service doors preflight checks

AIRPORT COMPATIBLE

location of control sufaces location of doors geometry fuel bay location

engine placement

SAFETY engine placement fuel bay location landing gearvertical stabilizer

osciallationsemergency

exits stability debris preflight

EASE OF MANUFACTURE moving parts materials used geometry costs

EMPTY WEIGHT # of engines materials used # of landing gear wing size# of moving

parts

COST procurement operation manufacture maintenance crew


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Historical Seating Data

Combined Mass fraction Averages

Combined Mass Fraction Average Aircraft Weight

Wing System 0.105531226 29970.86824

Tail System 0.023156588 6576.470953

Body System 0.10165365 28869.63663

Alighting Gear System 0.041065154 11662.50363

nacelle System 0.018617868 5287.474416

Propulsion System (less dry Engine) 0.02045158 5808.24884

Flight Control System 0.014202833 4033.604595

Auxiliary Power System 0.003965779 1126.281171

Instrument System 0.004442474 1261.662724

Hydraulic and Pneumatic Sytem 0.007352966 2088.242424

electrical System 0.015134974 4298.33256

Avionic Systems 0.010457658 2969.974953

Furnishing and Equipment System 0.068736975 19521.3009

Air Conditioning System 0.008229526 2337.185405

Anti-icing system 0.004181392 1187.515336

Load and Handling System 0.000184082 52.27922527

Empty Weight 0.448109156 127263.0004

Dry Engine Weight 0.062638655 17789.37796

Empy Weight (Maximum Empty) 0.510747811 145052.3784

0.958112537 272103.9604


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Initial Sizing results


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Requirements Compliance Matrix


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Next Step Flowchart

Aircraft Sizing

Wing

Airfoil

Database

Supersonic

Boom Code

Aerodynamic

Analysis

Dynamic

Stability

Control

Analysis

Aircraft Exterior

Shaping (cowling)

Component

Weight

Structural

Analysis

Component Size

Preliminary

Aerodynamic

Engine

Database

CFD

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