EXTROVERT ADVANCED C ONCEPT E XPLORATION ADL P-­‐2013050202 Sean C hait, B rett K ubica Georgia I nstitute o f T echnology School o f A erospace E ngineering X-­‐57 CONDOR Runway-­‐Based Space Launch System Aerodynamics May 2, 2013
EXTROVERT ADVANCED CONCEPT EXPLORATION
2
Publishing Information We gratefully acknowledges support under the NASA Innovation in Aerospace Instruction Initiative, NASA Grant No. NNX09AF67G, to develop the techniques that allowed such work to be done in core courses, and the resources used to publish this. Tony Springer is the Technical Monitor. Copyright except where indicated, is held by the authors indicted on the content. Please contact the indicated authors komerath@gatech.edu for information and permission to copy. Disclaimer “Any opinions, findings, and conclusions or recommendations expressed in this material are those of the author(s) and do not necessarily reflect the views of the National Aeronautics and Space Administration.” The X-­‐57 Condor Launch System is the next genera0on in Low Earth Orbit access vehicles. The system is capable of delivering a 100,000kg payload into orbit, return safely to Earth, refuel, and be capable of repea0ng the mission in the same day! Although significant technological advancement is necessary for the system to come to frui0on, its concept sets a basis for a poten0ally high value launch system. Mission Flight Profile Li# [kN] 0 5000 10000 15000 20000 0 10 15 Mach Number 5 20 LiA Required LiA Available Condor Li# Available vs. Required Condor Supersonic/Hypersonic Configura0on Condor Subsonic Configura0on X-57 Condor: Runway-Based Space Launch
System Aerodynamics
Integrative Assignment
AE3021A
Sean Chait
Brett Kubica
Daniel Guggenheim School of Aerospace Engineering
Georgia Institute of Technology
Atlanta GA 30332-0150
Spring 2013
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Contents
1 Summary
9
2 Previous Vehicles
2.1 Horizontal Takeoff Vehicles . . . . . . . . . . . .
2.1.1 HOTOL: Horizontal Takeoff and Landing
2.1.2 Skylon . . . . . . . . . . . . . . . . . . .
2.2 Hypersonic Vehicles . . . . . . . . . . . . . . . .
2.2.1 Blackswift . . . . . . . . . . . . . . . . .
2.2.2 Boeing X-51: Waverider . . . . . . . . .
2.3 Heavy Lift Aircraft . . . . . . . . . . . . . . . .
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3 Conceptual Design
3.1 Flight Regime Sizing . . . . . . . . . .
3.1.1 Rocket Stage . . . . . . . . . .
3.1.2 Supersonic-Hypersonic Stage . .
3.1.3 Subsonic Stage . . . . . . . . .
3.2 Initial Wing Loading Design . . . . . .
3.2.1 Subsonic Stage Aircraft . . . . .
3.2.2 Supersonic-Hypersonic Aircraft
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4 Aerodynamic Analysis
4.1 Low Speed Flight Regime
4.2 Supersonic Flight Regime
4.3 Hypersonic Flight Regime
4.4 Thrust Available . . . . .
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5 Final Condor Design
5.1 Final Configuration . . . . . .
5.1.1 Payload Capabilities .
5.1.2 Subsonic Transport . .
5.1.3 Hypersonic Transport .
5.2 Final Flight Regimes . . . . .
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6 Conclusions
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A Preliminary Sizing MATLAB Scripts
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3
CONTENTS
CONTENTS
4
List of Figures
1.1
1.2
Condor Space Access System . . . . . . . . . . . . . . . . . . . . . . . . .
Condor Launch System Overview . . . . . . . . . . . . . . . . . . . . . .
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2.1
2.2
2.3
2.4
HOTOL Horizontal Takeoff Space
Skylon [5] . . . . . . . . . . . . .
Blackswift [6] . . . . . . . . . . .
Boeing X-51: WaveRider [8] . . .
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4.1
4.2
Current Subsonic Transport Configuration . . . . . . . . . . . . . . . . .
Current Supersonic Transport Configuration . . . . . . . . . . . . . . . .
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5.1
5.2
5.3
5.4
5.5
5.6
5.7
Subsonic Condor Top View . .
Subsonic Condor Side View . .
Subsonic Condor Front View .
Hypersonic Condor Top View .
Hypersonic Condor Side View .
Hypersonic Condor Front View
Condor Flight Profile . . . . . .
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A.1
A.2
A.3
A.4
A.5
A.6
A.7
A.8
Initial Mass Size Script . . . . . . . . . .
Subsonic Wing Loading Script . . . . . .
Supersonic Wing Loading Script . . . . .
Subsonic Aerodynamic Analysis Code (a)
Subsonic Aerodynamic Analysis Code (b)
Supersonic Aerodynamic Analysis Code .
Hypersonic Aerodynamic Analysis Code .
Available Thrust Analysis Code . . . . . .
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Access
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Vehicle [3]
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LIST OF FIGURES
LIST OF FIGURES
6
List of Tables
2.1
Heavy Lift Aircraft Parameters . . . . . . . . . . . . . . . . . . . . . . .
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3.1
3.2
3.3
3.4
3.5
3.6
3.7
3.8
∆V Losses . . . . . . . . . . . . . . .
Rocket Stage Sizing . . . . . . . . . .
Required Stage ∆V 0 s . . . . . . . . .
Supersonic-Hypersonic Stage Sizing .
Subsonic Weight Fuel Estimation [15]
Subsonic Stage Sizing . . . . . . . . .
Subsonic Wing Loading . . . . . . . .
Supersonic Wing Loading . . . . . .
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4.1
4.2
4.3
4.4
4.5
4.6
Zero-Lift Drag Buildup . . . . . . . . . . . . . . .
Subsonic Aerodynamic Performance Estimation .
Supersonic Airfoil Coefficients . . . . . . . . . . .
Supersonic Aerodynamic Performance Estimation
Hypersonic Aerodynamic Performance Estimation
Thrust Available . . . . . . . . . . . . . . . . . .
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5.1
5.2
5.3
5.4
Condor Payload Bay Dimensions
Subsonic Wing Dimensions . . . .
Subsonic Fuselage Dimensions . .
Hypersonic Vehicle Dimensions .
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LIST OF TABLES
LIST OF TABLES
8
Chapter 1
Summary
Modern launch vehicles capable of placing a payload into low earth orbit are expensive,
non-reusable, and take months of preparation. In order to make any large scale space
infrastructure, such as the proposed Space Solar Power (SSP) architecture, viable a new
solution launch system must be developed. This system must have the reusability of a
modern heavy-lift cargo aircraft, capable of loading, refueling, and executing its mission
multiple times a day. The requirement of reusability and quick turnaround time makes
the current configuration of a multi-stage rocket used by most modern launch systems
inadequate. These systems require months of preparation, and are not reusable, driving
up the cost per kilogram of placing a payload into orbit. For a launch system to be able
to perform multiple flights to orbit a day, it is necessary for all stages of the proposed
system to be completely reusable, requiring only the level of maintenance seen by current
cargo aircraft. From these requirements it was determined that a vehicle (or vehicles)
based off of the principle of horizontal runway takeoff and landing was the only viable
option for producing the required launch system.
Significant research into previous attempts at a horizontal take-off/horizontal landing
space launch system gave rise to the concepts used in the preliminary design of the current
launch system concept under consideration, the Condor. The Condor will consist of two
main stages, each with their own vehicle. The first is a large blended wing body aircraft
which will be carrying the second stage, a blunt body hypersonic craft. The first stage
will be responsible for runway based take-off and getting the entire system to an altitude
Figure 1.1: Condor Space Access System
9
CHAPTER 1. SUMMARY
and Mach number such that the second stage can be deployed, accelerate through the
sound barrier, and eventually reach orbit. After separation, the first stage will return,
under its own power, to the launch facility. The second stage will use its highly advanced
LACE engines, first in an air breathing configuration until hypersonic speeds are reached,
then in rocket configuration to insert the payload into orbit. Once the payload has been
successfully jettisoned, the stage two vehicle will re-enter the Earth’s atmosphere and land
at its designated launch facility. The entire Condor system will then be quickly refueled
and loaded for another launch opportunity. To allow for the uninterrupted delivery of
cargo into orbit, launch facilities will located around the world at various locations to
prevent possible interruptions due to weather and other atmospheric conditions.
After an extensive research effort, which gave rise to a preliminary conceptual design
for the Condor, further analysis was conducted to refine and determine the feasibility
of the design. Basic mass sizing and wing loading techniques were used along with a
preliminary flight profile to determine a set of geometric of parameters that defined the
launch system. From here a modeling software was used to solidify the initial design and
produce all of the information to run a thorough aerodynamic analysis. This analysis
revealed that the current vehicle and flight profile were not sufficient to accomplish the
goals of the mission. Modifications were made to both the first and second stage vehicle
designs along with the intended mission profile in order to ensure the entire Condor
launch system would be able to function efficiently. Final aerodynamic analysis showed
that the launch system framework described in this report is theoretically possible in
terms of aerodynamics although significant advances in structural design, materials, and
hypersonic control systems are necessary. This is to ensure that the Condor is fully able
to cope with the extreme conditions it will face while still performing to full mission
success. Although out of reach of modern technology, the Condor may very well be the
basis for a future large payload space launch system.
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CHAPTER 1. SUMMARY
Figure 1.2: Condor Launch System Overview
11
CHAPTER 1. SUMMARY
12
Chapter 2
Previous Vehicles
In the pursuit of a new design, it is wise to consider approaches that others have tried (and
oftentimes failed) to determine the most feasible approach to meet your mission criteria.
Complete design concepts may have already been developed and analyzed, thus saving the
time of performing a comparable analysis to determine the feasibility of a method. From
the successes and failures of these previous attempts, it is possible to develop a sound,
new vehicle concept that may very well perform to the desired specifications. It is for
this reason that a significant research effort into horizontal takeoff and landing (HOTOL)
orbital vehicles and hypersonic vehicles was undertaken prior to conceptual design of our
craft. The purpose of this research was two-fold. First, to examine the aerodynamic
and payload capabilities of previous design attempts and second, to compare different
propulsion systems and their optimal flight regimes. By learning about previous and
current attempts at HOTOL vehicles, we were able to gather enough information to
create a feasible preliminary design for a heavy-lift HOTOL vehicle to bring our payload
into Low Earth Orbit (LEO).
2.1 Horizontal Takeoff Vehicles
2.1.1 HOTOL: Horizontal Takeoff and Landing
Developer: Rolls Royce and British Aerospace
Payload: 7,000kg
Propulsion System: Rolls Royce RB545 air-breathing rocket engine
Description:
The HOTOL was an attempt at a single-stage-to-orbit vehicle developed by Rolls
Royce and British Aerospace between 1982 and 1986. The concept involved a ”space
Figure 2.1: HOTOL Horizontal Takeoff Space Access Vehicle [3]
13
2.2. HYPERSONIC VEHICLES
CHAPTER 2. PREVIOUS VEHICLES
plane” which would takeoff horizontally from a runway with the aid of a rocket propelled
trolley. The craft’s air breathing rocket engine (the RB545 Swallow) would then take
over and behave like a turbojet until approximately Mach 5. At this point the engine
would transfer to a pure rocket mode for the remainder of the climb to orbit [1].
Issues during the air breathing phase of operations involving the center of gravity
and center of pressure resulted in the need for major redesigns. As a result the payload
fraction of the craft was drastically reduced thus decreasing the economy of the craft.
Funding for the project was ceased by the British government in the mid-80’s [2].
2.1.2 Skylon
Figure 2.2: Skylon [5]
Developer: Reactions Engine Limited
Payload: 15,000 kg
Propulsion System: SABRE (Synergistic Air-Breathing Rocket Engine)
Description:
The Skylon space plane is a single-stage-to-orbit craft currently under development
by Reaction Engines Limited and funded by both the British government and European
Space Agency. The vehicle is based off of the HOTOL design, as several of its key designers
were members of the original team. An air-breathing rocket engine, with many similarities
to a LACE system, is used. The primary difference between the SABRE engine and the
conventional LACE design is that the air is not liquefied within the engine, increasing
efficiency (the engine has an atmospheric ISP of 3500s). During the flight regime up to
Mach 5.4, the cooled, highly compressed air is fed into a rocket combustion chamber,
where it is ignited with liquid oxygen. This allows for high thrust throughout the entire
flight. After this point, the air intake is closed off and stored liquid oxygen is used for
the remainder of the flight [4] .
The hopes of the project is to be able to launch a 15,000kg payload at a cost of a little
over a thousand dollars per kilo. To date, the project has passed all design and testing
reviews, garnering it continuous funding.
2.2 Hypersonic Vehicles
2.2.1 Blackswift
Developer: Lockheed Martin, Boeing, and the USAF
Propulsion System: Combination turbine engine/ramjet
14
CHAPTER 2. PREVIOUS VEHICLES
2.2. HYPERSONIC VEHICLES
Figure 2.3: Blackswift [6]
Description:
The Blackswift is a hypersonic concept aircraft developed by Lockheed Martin, Boeing
and the USAF. Under the Air Force’s Falcon project, expectations are that the craft
will be able to function as a hypersonic cruise vehicle, capable of delivering its payload
anywhere on the planet within a few hours. Unlike many other hypersonic concepts, the
craft is being designed to take off and land on a conventional runway under its own power.
This was to be achieved through a combination turbine engine and ramjet propulsion
system. The turbine engine will be used to get Blackswift to speeds approaching Mach 3
and then the ramjet will accelerate the vehicle to Mach 6. Currently the project is not in
development as funding for the project was drastically cut causing the project’s ultimate
cancellation [6] .
2.2.2 Boeing X-51: Waverider
Figure 2.4: Boeing X-51: WaveRider [8]
Developer: Boeing
Payload: 270 kg
Propulsion System: Pratt Whitney SJX61
Description:
The Boeing X-51 WaveRider is an unmanned hypersonic demonstration aircraft being
developed by the United States Air Force Research Lab. The term ”WaveRider” comes
from the unmanned crafts use of shockwaves to produce additional lift. The X-51 is a ride
along craft, which is attached to the wing of a B-52 and carried to an altitude of 50,000
feet before deployment. After separation, a solid rocket booster ignites to accelerate the
craft to approximately Mach 4.5, at which time a hydrocarbon-fueled scramjet is ignited,
further accelerating the craft to a theoretical max speed of greater than Mach 6 [9].
The WaveRider has been subject to several setbacks due to failures after the ignition
of the scramjet during the transition between the start up fuel and conventional JP-7 jet
15
2.3. HEAVY LIFT AIRCRAFT
CHAPTER 2. PREVIOUS VEHICLES
fuel, an attribute that makes the X-51 unique in hypersonic craft. These failures, causing
the premature end of hypersonic flight, coupled with other aerodynamic control issues
make the current reliability of the craft low. Also current design allows for a payload of
only 270kg, far below the range of any practical transport application [7] .
2.3 Heavy Lift Aircraft
Research was also conducted into current subsonic heavy lift aircraft. This was done to
gage the current capabilities of heavy lift systems in the subsonic regime so as to determine
the best course of action for designing the subsonic portion of our launch system. From
this research, an understanding of payload fraction and the sizing of necessary lifting
surfaces for aircraft performing in this flight regime was developed. Due to the scale of
the payload that must be placed into orbit (100,000 kg), the largest and most powerful
of modern heavy-lift aircraft were examined.
Aircraft
C-5 Galaxy [10]
C-17 Globemaster [11]
An-225 Mriya [12]
An-124 Ruslan [13]
A-380-800 [14]
Max Takeoff-Weight
381,000 kg
265,350 kg
640,000 kg
405,000 kg
590,000 kg
Max Payload
122,470 kg
77,519 kg
250,500 kg
150,000 kg
149,800 kg
Table 2.1: Heavy Lift Aircraft Parameters
16
Wing Span
75.31 m
51.75 m
88.4 m
73.3 m
79.75 m
Wing
576
353
905
628
845
Area
m2
m2
m2
m2
m2
Chapter 3
Conceptual Design
After examining previous attempts at horizontal take-off and landing spacecraft systems,
it was determined that the SSTO framework would not be feasible for delivering a 100,000
kg payload to orbit while utilizing completely reusable components. An aircraft that is
capable of producing the lift necessary to take-off and pass through the subsonic flight
regime would be so large that any attempt to reach supersonic, let alone hypersonic
speeds, would meet insurmountable heating, loading, and thrust issues that would end in
the craft’s failure. It was for this reason that a three stage, two craft system was decided
upon.
Our design for a heavy-lift, horizontal takeoff and landing, hypersonic vehicle consists
of a smaller body designed for supersonic/hypersonic flight, which carries the payload,
and a large detachable flying wing for subsonic flight. The smaller body will have a
blunt nose to decrease heating in hypersonic flight and a delta wing with a high sweep
angle and low aspect ratio. We will use the SABRE hybrid rocket engines as our main
powerplant. The SABRE is an offshoot of the LACE engine concept. It uses a helium
loop to precool incoming air and turn the turbine. It can compress air from ambient to
140 atmospheres. The engine uses liquid hydrogen as fuel and will close off its inlet at
high speed and altitude, becoming exclusively a rocket engine [4].
The flying wing will be attached on takeoff and used to take the second stage vehicle
to an altitude and velocity such that, once detached, it will be able to sustain flight.
The vehicle will take-off under the power of the turbofan engines of the flying wing stage
and be responsible for bringing the entire system to altitude. As the vehicle approaches
the supersonic flight regime, the large wing will detach and return to the airfield. The
vehicle then relies solely on the SABRE engines, which adjust their inlets with increasing
Mach number and altitude. At approximately Mach 5.14 or 28.5km the inlet seals and
the SABRE becomes a hydrogen-fueled rocket. When the vehicle reaches LEO (approximately Mach 25 at 150km) it will jettison its payload and return to a designated airfield
under little or no power from its engines.
From this baseline concept, an initial flight plan and aircraft conceptual design was
developed using different mass and wing area sizing techniques. These baseline values
were then utilized in the aerodynamic analysis and based on the results of that analysis,
adjusted accordingly such that the Condor is able to successfully complete its mission.
3.1 Flight Regime Sizing
The preliminary mission profile will be broken into five different stages of flight:
17
3.1. FLIGHT REGIME SIZING
CHAPTER 3. CONCEPTUAL DESIGN
Stage 1: Takeoff to Mach 0.8 at 10,000 meters (approximately 241 meters per second)
and aerial refuel
Stage 2: Airbreathing phase of the SABRE engines from Mach 0.8 at 10,000 meters to
Mach 5.5 at 20,000 meters
Stage 3: Rocket phase of the SABRE engines from Mach 5.5 at 20,000 meters to Low
Earth Orbit (LEO), which occurs at 150,000 meters and Mach 25 (approximately 7,780
meters per second)
Stage 4: Jettison of payload
Stage 5: Vehicle re-entry and landing
The initial mass sizing of the heavy-lift, horizontal takeoff and landing craft was
conducted in three stages based upon our expected flight regimes. The flight regimes
are defined as the takeoff/transonic phase, the supersonic/hypersonic phase, and the
rocket stage. The rocket stage will encompass orbit insertion, jettison of payload, and
re-entry. These regimes are based upon the staging of the vehicle as well as the state of
the engine during those phases. Derivation of the subsequent phases and their respective
mass requirements are described below. A comprehensive MATLAB script was developed
for all calculations and to provide a framework for design iteration.
3.1.1 Rocket Stage
The rocket stage is the final stage of prior to insertion into low Earth orbit and will commence when the vehicle has been accelerated to a speed of Mach 5.5. At this point in time
the SABRE engines will convert from a hybrid airbreathing rocket engine to a convential
rocket engine running on liquid hydrogren and oxygen. Since the vehicle will already be
travelling at Mach 5.5, the ∆V requirements of this phase can be determined using the
necessary velocity requirements of entering orbit and the velocity already achieved during
the supersonic-hypersonic phase. While determining these total ∆V requirements of this
stage, other considerations must also be made. To improve efficiency, the launch site of
the craft and maneuvers of the rocket phase will be utilized to obtain an extra ∆V from
rotation of the Earth, thereby reducing fuel required to enter LEO. However, losses due
to gravity, drag, and maneuvering the craft in atmosphere must also be accounted for.
Table 3.1 details those values used which are based off of typical industry standards for
initial sizing.
∆Vlosses
∆Vgravity
∆Vdrag
∆Vsteering
∆Vrotation
Magnitude
1000 m/s
50 m/s
100 m/s
400 m/s
Table 3.1: ∆V Losses
The mass ratio of the rocket stage was determined using the rocket equation and the
aforementioned losses. When in pure rocket flight the SABRE engines operate at an Isp of
460s, a very high value even for liquid rockets [16]. A high (for a rocket) structural ratio
of 0.10 was assumed due to the added complexity of the SABRE engines, the large fuel
18
CHAPTER 3. CONCEPTUAL DESIGN
Entrance Mach
5.5
10
15
Mass Ratio
4.497
3.276
2.305
3.1. FLIGHT REGIME SIZING
Structural Mass [kg]
63,500
33,800
17,000
Propellant [kg]
572,000
305,000
153,000
Total Mass [kg]
735,000
438,000
270,000
Table 3.2: Rocket Stage Sizing
tanks required to store liquid hydrogen and structural/control considerations for hypersonic maneuvering. The full payload of 100,000kg was also assumed.
∆VLEO = Isp g0 ln(M R) − ∆Vgravity − ∆Vdrag − ∆Vsteering + ∆Vrotation
M R = e(∆VLEO +∆Vgravity +∆Vdrag +∆Vsteering −∆Vrotation )/(Isp g0 )
Implementation of the modified form of the rocket equation in MATLAB (See Appendix A) created a powerful design iteration tool to be used to determine the point at
which the engines will be converted from an airbreathing to a pure rocket mode. Analysis (seen in Table 2.2) showed that the frontier of Mach 15, the limit of hypersonic,
airbreathing flight, yeilds the minimum mass of the final rocket stage, as a large portion
of the velocity required to achieve orbit has already been acquired in the supersonichypersonic phase. However, aerodynamic analysis will show that although this creates a
lowest mass vehicle, the engines will not have sufficient thrust (due to engine lapse) to
maintain flight. It was therefore determined that an entrance in the rocket phase would
begin at Mach 5.5, increasing our initial mass drastically but this would grant us enough
thrust to insert into LEO successfully. The mass of the final, rocket stage will drive
the mass and fuel requirements of all prior stages (since they must carry it to speed),
therefore it minimization is essential. As will be described in Section 3.1.2, the efficiency
of the SABRE engines during its airbreathing operations is drastically higher than that
during rocket operations (due to the use of air as an oxidizer) making it desirable to have
as much velocity change as possible during the this earlier stage of flight.
Initial sizing of the rocket staging and the determination of the threshold which will
mark entrance into ”rocket” mode therefore make it possible to calculate the remaining
required ∆V 0 sor the remaining two stages based on a transition between subsonic and
supersonic modes at Mach 0.8 and hypersonic and rocket phases at Mach 5.5.
Mach Regime
0-0.80
0.80-5.5
5.5-25.0
∆Vrequired [m/s]
241
1,410
6,784
Table 3.3: Required Stage ∆V 0 s
3.1.2 Supersonic-Hypersonic Stage
After the entrance point into the rocket stage was determined and initial sizing determined, it is possible to next determine the mass of the supersonic-hypersonic phase of
19
3.1. FLIGHT REGIME SIZING
Entrance Mach
0.8
Exit Mach
5.5
CHAPTER 3. CONCEPTUAL DESIGN
Mass Ratio
1.041
Propellant [kg]
29,900
Total Mass [kg]
765,000
Table 3.4: Supersonic-Hypersonic Stage Sizing
operations. A single phase was decided upon for travel between transonic speeds and the
end of the hypersonic regime due to the extremely efficient capabilities of the SABRE
engines. Unlike traditional supersonic and hypersonic engines such as ramjets and scramjets which are only efficient once they approach their optimal Mach range, the SABRE
engines perform at the same efficiency throughout flight operations. Due to the use of air
as an oxidizer, the engines run at an Isp f 3600s, a figure that rivals many electric propulsion systems [16]. The dual capabilities of the engines also negate the need for a physical
stage separation of the craft during the transition from hypersonic to rocket flight, rather
a simple ”mode” change of the engines. For this reason the only extra structural mass
needed in this phase is that required to carry the fuel required, all other structural mass
was already accounted for in the rocket stage. Also due to the high efficiency of the
engines during airbreathing operations, a significantly smaller amount of fuel is needed
during this phase of operations than the rocket stage, even though the total velocity
gained during hypersonic activities is greater than that of the rocket burn.
3.1.3 Subsonic Stage
Upon completion of the initial sizing of the stages necessary to accelerate the craft and
payload from high transonic speeds to Low Earth Orbit, it is possible to determine the
size of the initial vehicle required to take the payload from takeoff to hypersonic vehicle
deployment. The transition from the subsonic to supersonic-hypersonic phases will mark
the only physical seperation of staged craft during flight operations. The logic for this
design decision lies in the aerodynamic differences between subsonic and super/hypersonic
flight. An aircraft with a significant amount of surface area is required to get such a
massive payload off the ground and up to typical cruise conditions. However such an
aircraft would not be able to withstand the forces placed on it during supersonic and
hypersonic flight, let alone re-entry. Therefore the ”payload” of the subsonic, heavy lift
aircraft will be the hypersonic vehicle, designed to have enough thrust to power through
the Mach barrier and quickly up to high Mach numbers, as well as an aerodynamic design
that tends itself to high speed flight.
For optimal lift and structural weight considerations, a blended wing body configuration was chosen for the subsonic phase of flight. These aircraft have an empty weight
fraction of below 50% and are ideal when such a high payload is being carried. For our
considerations an empty weight fraction of 0.50 was assumed. The current design iteration also utilizes six GE 90-115 Turbofan engines, the same engine used on the Boeing
777 [15]. Using the specific fuel consumption of the engines, an estimated time of takeoff
to hypersonic vehicle detachment of 45 minutes, and the average weight of JP-7 jet fuel,
initial aerodynamic design considerations, an approximate weight of fuel can be generated
for sizing purposes. The values used can be found in Table 3.5.
20
CHAPTER 3. CONCEPTUAL DESIGN
3.2. INITIAL WING LOADING DESIGN
Sizing Parameter
SFCcruise
SFCsealevel
SFCAverage
Average Thrust During Takeoff
Time to Separation
Weight of Fuel Per Engine
Weight of Fuel Total
Value
1.47E-5 kg/s/N
0.92E-5 kg/s/N
1.195E-5 kg/s/N
400,000 N
45 min
17136 kg
77436 kg
Table 3.5: Subsonic Weight Fuel Estimation [15]
We /W0
0.50
Structural Mass [kg]
843,000
Propellant [kg]
77436
Total Mass [kg]
1,680,000
Table 3.6: Subsonic Stage Sizing
3.2 Initial Wing Loading Design
Wing loading is an important performance parameter. An aircraft with lighter wing loading will result in an increased climb rate and a decrease in thrust required at cruise. An
aircraft with higher wing loading is better suited for high speed flight, due to the smaller
wing and decreased drag. However, the higher wing loading means faster takeoff/landing
distance and velocity, because the wing will need to generate the same amount of lift
as the larger wing. A seperate analysis was run for both the subsonic and supersonichypersonic craft to determine their respective wing loading.
3.2.1 Subsonic Stage Aircraft
To estimate wing loading in the subsonic regime, we set several constants. The altitude
is set at sea level, Mach number set at 0.15, and angle of attack set at 7 degrees. The
variables are weight, aspect ratio, sweep, and lift-curve slope. In the process of estimating
wing loading, we ran a function in Matlab to return a wing span and area to help size
our wings at the same time. After several iterations we selected a wing with an aspect
ratio of 8, sweep of 25 degrees, lift-curve slope of 5, span of 121 meters, and area of 1,804
square meters. This configuration yielded a wing loading of 931kg/m2 for a 1,680,000 kg
aircraft.
3.2.2 Supersonic-Hypersonic Aircraft
To estimate wing loading in the supersonic to hypersonic regime, we set weight, aspect
ratio, and wing length (root chord) as our variables. We decided to use wing length as a
variable, because we assumed that our supersonic stage would utilize a delta wing. Held
constant is altitude at 10,000m, angle of attack at 2 degrees, and Mach number at 1.5.
The Matlab function returns the same values as the subsonic estimate. After iterations,
we selected a delta wing with an aspect ratio of 0.5, wing length of 50 meters, span of
13.4 meters, and area of 335.2 square meters. The wing loading is 1193.5kg/m2 Our
only concern with this configuration is its ability to produce the required lift at transonic
speeds after the subsonic wing detaches. This will be determined, and the wing adjusted,
after aerodynamic analysis.
21
3.2. INITIAL WING LOADING DESIGN
Iteration
Input
Weight [kg]
Aspect Ratio
Sweep [deg]
dCl/dα
Output
Wing Span [m]
Wing Area [m2 ]
Wing Loading [kg/m2 ]
CHAPTER 3. CONCEPTUAL DESIGN
1
2
3
4
5
6
1,680,000
7
25
6
1,680,000
8
25
6
1,680,000
7
30
6
1,680,000
7
25
5
1,680,000
8
25
5
1,680,000
8
25
6
106
1,575
1,066
112
1,514
1087
258
9,373
179.2
114
1,837
914.6
121
1,804
931.1
112
1,542
1,089
Table 3.7: Subsonic Wing Loading
Iteration
Input
Weight [kg]
Aspect Ratio
Wing Length [m]
Output
Wing Span [m]
Wing Area [m2 ]
Wing Loading [kg/m2 ]
1
2
3
4
5
6
765,000
2
50
765,000
1.5
50
765,000
1
50
765,000
0.5
50
765,000
0.5
75
765,000
1
75
6.4
160
4,774
8.5
214
3,580
12.8
320
2,387
13.4
335
1,193
17.1
641
1,193
8.5
320
2,387
Table 3.8: Supersonic Wing Loading
22
Chapter 4
Aerodynamic Analysis
To further refine the design of the launch system, a baseline aerodynamic analysis must
be done to determine the current benefits and deficiencies of the proposed aircraft. Prior
to analysis, a firm geometric design of the aircraft was developed. Using the open source
software OpenVSP, an aircraft was designed based off of the general parameters of wing
span, area, aspect ratio, etc previously determined during the initial mass and wing loading sizing of the stages. For initial analysis purposes a NACA 0012 airfoil was assumed
for the subsonic phase until investigation suggests a preferred alternative. The flying
wing which constitutes the detachable subsonic stage was modeled with the supersonichypersonic hybrid blunt body delta wing craft attached to its bottom. The size of the
fuselage of the supersonic-hypersonic stage was modeled after the necessary space requirements of the USAF C-5 heavy lift aircraft to account for a large payload. Additional space
was added for engines, fuel, and cockpit, however the space designated for this is subject
to growth as further research and analysis is conducted. Once all of these parameters
were determined and successfully modeled in OpenVSP, the software’s features allowed
for simple extraction of geometric data for use in aerodynamic analysis.
The resulting baseline design was utilized in initial vehicle aerodynamic analysis. As
the analysis was conducted and deficiencies identified, modifications were made to the
aircraft design and flight profile accordingly. Initial analysis, the resulting changes, and
the final aerodynamic capabilities of the craft are described below.
4.1 Low Speed Flight Regime
The blended-wing heavy lift subsonic transport plays the vital role of taking the payload
(i.e. the supersonic-hypersonic transport) from the runway to a speed on the border of the
transonic flight regime at which point the second stage can begin operations. The second
stage is being designed for optimal performance in the supersonic and hypersonic flight
regimes which tends it towards physical characteristics that make it highly impractical
(if not completely impossible) to produce sufficient lift during subsonic flight to take off.
It is for this reason that the subsonic transport phase must produce a large amount of
lift to get its payload to operational speeds and is this optimized for such.
In order to analyze the necessary thrust requirements of the aircraft, a baseline drag
profile must be developed. For the purposes of this portion of analysis an emperical
buildup based off of methods presented in Aircraft Performance and Design by John Anderson [18]. The following formulas were developed using regressions of common transport
aircraft and are a function of the typical aircraft parameters seen below. These results
23
4.1. LOW SPEED FLIGHT REGIME
CHAPTER 4. AERODYNAMIC ANALYSIS
Figure 4.1: Current Subsonic Transport Configuration
can then be easily plugged into the standard lift and drag equations to begin analysis of
the aircraft’s performance. For the purposes of this analysis we are first going assume
worst case scenario of completely turbulent flow and second assume three sections for
the aircraft: the subsonic flying wing (modeled as a wing), the hypersonic delta wing
(modeled as a wing), and a hypersonic blunt body (modelled as a fueslage). Further
refinement of design will later include details such as engines, extra fuel tanks, interface
mechanisms, etc and will add a significant amount of zero-lift drag. For this reason, as
well as to account for parasite drag, the calculated Cdo will be multiplied by three. Also
for the purposes of this assignment, it is being assumed that a highly advanced airfoil is
being used such that it has a critical Mach number 0.81 and will therefore allow the craft
to operate at Mach 0.9 without encountering any supersonic flow.
F FW ing = 1 + 2(t/c) + 60(t/c)4
f
60
F FF uselage = 1 + 3 +
f
400
ρ∞ V∞ c
Rec =
µ∞
0.074
Cf turbulent =
Re0.2
c
X
Sweti
CDo =
C fi · (
) · F Fi · Qi
Srefi
Component
Subsonic Blended Wing
Hypersonic Wing
Hypersonic Blunt Body
Total Zero-Lift Drag
Cf
0.0020
0.0017
0.0015
Swet
3800
649.4
790.2
FF [m2 ]
1.252
1.252
1.079
(4.1)
(4.2)
(4.3)
(4.4)
(4.5)
Q
1.3
1.1
1.1
CDo
0.0064
0.0001
0.0001
0.0238
Table 4.1: Zero-Lift Drag Buildup
After calculation of zero-lift drag, it is possible to obtain an estimate for induced drag
(based on the desired angle of attack) and thus calculate the desired drag created by the
aircraft. Lift calculations are also attainable using the lift curve slope previously derived.
24
CHAPTER 4. AERODYNAMIC ANALYSIS 4.2. SUPERSONIC FLIGHT REGIME
Mach
0.9
Alt [m]
10,000
α[deg]
3
b [m]
88.4
Lift [kN]
17,200
Drag [kN]
977
Tavail [kN]
1350
Table 4.2: Subsonic Aerodynamic Performance Estimation
e0 = 1.78(1 − 0.045AR0.68 ) − 0.64
(4.6)
K = (πARe0 )−1
(4.7)
Cd = Cdo + KCL2
(4.8)
L = 1/2ρV∞2 SCL
(4.9)
D = 1/2ρV∞2 SCD
(4.10)
Using the aircraft parameters obtained during the initial conceptual sizing of the
aircraft, it was found that although the aircraft was able to produce enough lift and
thrust to take-off from a runway unassisted, it was not able to produce enough lift to
reach an altitude of 10,000 meters at Mach 0.8. The analysis was repeated for different
geometric configurations using the current flight profile where it was determined that a
subsonic aircraft of nearly double the wing area, in the current configuration, would be
necessary to maintain level flight. Modification of the flight regimes was then considered.
After analysis, it was determined that only a 5% increase in wing area would be necessary
if the aircraft was travelling at Mach 0.9 at an altitude of 10,000 meters. The higher
entrance Mach number of the second stage was then analyzed for performance and it was
determined that both the subsonic and supersonic aricraft would be able to function under
the new flight regime. Therefore, the suggested changes to the geometric configuration
of the subsonic aircraft and the changes to the overall flight regime were put in place.
4.2 Supersonic Flight Regime
Figure 4.2: Current Supersonic Transport Configuration
In the first iteration of the supersonic case, we utilized slender wing theory to estimate
lift and induced drag on the wing. This assumption can be used up to about Mach 5,
because of the high sweep of the delta wing. Maximum wing sweep is 80 degrees. No part
of the wing is outside of the Mach cone until the aircraft surpasses Mach 5 by the equation:
25
4.2. SUPERSONIC FLIGHT REGIME CHAPTER 4. AERODYNAMIC ANALYSIS
1
)
(4.11)
M
To analyze the aerodynamic forces at supersonic speeds the aspect ratio must first be
corrected for compressible flow,
√
β = M2 − 1
(4.12)
µ = arcsin(
AR =
ARo
β
(4.13)
Then, using slender wing theory, the coefficient of lift and wave drag can be determined,
π
(4.14)
CL,wing = ARα
2
CD,wave =
4α2
β
(4.15)
The skin friction coefficient can be calculated after adjusting for compressible, turbulent flow conditions,
T∗
2
= 1 + 0.1198M∞
(4.16)
T∞
Cf = 0.295
µ∗
T∗ 3
=(
)2
µ∞
T∞
(4.17)
T∞ µ∞
T∞
log(Re∞ ∗ ∗ )−2.45
∗
T
T µ
(4.18)
Now the total drag coefficient can be found,
CD = CD,wave + Cf
(4.19)
Total lift generated by the Condor is calculated by finding the total lift of the wing
and using an adjustment estimation to account for the loss of lift due to the fuselage
interference [17],
1
q∞ = V∞2
2
(4.20)
Lwing = q∞ SCL,wing
(4.21)
rf uselage 2 rf uselage 4
+
)
(4.22)
b
b
Total drag acting of the Condor in flight is calculated assuming a flat plate for the
fuselage. The flat plate is equal to the largest cross-sectional area of the fuselage, therefore
overestimating drag due to the fuselage,
Ltotal = Lwing (1 −
Dtotal = q∞ (Swing + Sf uselage )CD,total
(4.23)
Finally, thrust required is equal to total drag, unless the wings do not generate enough
lift. If the wings do not produce sufficient lift, then the thrust required is equal to the
total drag plus the thrust vector in the lift direction needed to compensate for the lack
of lift,
26
CHAPTER 4. AERODYNAMIC ANALYSIS 4.3. HYPERSONIC FLIGHT REGIME
Treq = D +
W −L
sin(α)
(4.24)
A series of altitudes, angles of attack, and Mach numbers were run to determine the
aerodynamic coefficients and forces on the aircraft during the supersonic phase. Based on
initial analysis, wing span needed to be increased for the supersonic phase. The second
stage of the Condor’s wingspan was adjusted from an original 15 meters to a much larger
36 meters. Although this change drastically increased drag and the forces on the aircraft,
it was necessary to generate sufficient lift for mission completion. The added thrust
requirements caused by the increase in drag were well within the pre-existing thrust
margin of the SABRE engines.
Mach
1.5
2
3
4
5.5
Alt [m]
10,000
10,000
10,000
15,000
20,000
α[deg]
8
6
4
4
6
CL
0.235
0.114
0.047
0.034
0.037
CD,wave
0.0697
0.0253
0.0069
0.0050
0.0081
Cf
0.00020
0.00018
0.00016
0.00015
0.00014
CD,total
0.0700
0.0260
0.0071
0.0052
0.0083
Table 4.3: Supersonic Airfoil Coefficients
Mach
1.5
2
3
4
5.5
Alt [m]
10,000
10,000
10,000
15,000
20,000
α[deg]
8
6
4
4
4
Lift [kN]
7,578
10,104
15,156
12,706
10,964
Drag [kN]
2,548
1,652
1,028
634
395
Table 4.4: Supersonic Aerodynamic Performance Estimation
4.3 Hypersonic Flight Regime
In the hypersonic regime (above Mach 5), Modified Newtonian Aerodynamics can be
applied. Here we get:
Cp = Cpo,min sin2 θ
(4.25)
Assuming γ = 1.4 and as Mach number approach infinity, the equation becomes:
Cp = 1.84sin2 θ
(4.26)
We assume that the Cp on the sides of the wing and body not directly interacting
with the flow are zero.
At Mach 10 and 15 we are using the thrust vector in the lift direction to compensate
for a lack of lift. We could extend the wings at high altitude to generate more lift to give
the aircraft more thrust for climbing.
27
4.4. THRUST AVAILABLE
Mach
5.5
10
15
Alt [m]
25,000
40,000
50,000
CHAPTER 4. AERODYNAMIC ANALYSIS
α[deg]
10
17
20
Lift [kN]
12,290
5,746
4,185
Drag [kN]
4,032
2,089
1,757
Treq [kN]
4,032
7,073
10,581
Table 4.5: Hypersonic Aerodynamic Performance Estimation
4.4 Thrust Available
Thrust in an air-breathing engine is proportional to the density at altitude over the density
at sea level. We assume air-breathing to 20km. Above 20km the SABRE switches over to
rocket propulsion. It is assumed that thrust is constant, because the density is so small
and pressure is close to vacuum.
Altitude [m]
10,000
20,000
30,000
30,000
Thrust Available [kN]
2,646.4
569
117 (airbreathing)
11,760 (rocket)
Table 4.6: Thrust Available
From this analysis, we need to start the rocket at a lower altitude than initially
thought. The aircraft can get to Mach 5.5 at 20,000 meters before switching over to
rockets for the climb to LEO.
28
Chapter 5
Final Condor Design
The Condor takes off with hydrogen and oxygen tanks empty. It climbs to 10,000 meters
and Mach 0.9 under the power of six GE-90’s, where it levels off in order to fill the
hydrogen tank and some of the oxygen tank. When aerial refuel is complete, the high
aspect ratio wing is jettisoned and returns to the airport of origin and the SABRE engines
engage. The Condor then goes into a shallow dive to break the sound barrier and reach
Mach 1.5 to generate sufficient lift to climb. The aircraft then climbs to 25,000 meters
and Mach 5.5, at which point the SABRE engines switch over to liquid rocket. The
hypersonic vehicle climbs to 50,000 meters and Mach 15 then executes a vertical, nose
up maneuver to climb to 150,000 meters and Mach 25 entering Low Earth Orbit. The
payload is jettisoned in LEO. The Condor then begins its descent back to Earth. The low
aspect ratio wing extends outward to increase wing area on landing. A large parachute
deploys after touchdown to help slow the massive vehicle to a stop.
5.1 Final Configuration
5.1.1 Payload Capabilities
The current design of the Condor launch system has the capability of placing a 100,000
kg payload into low Earth orbit. The geometric dimensions of the cargo bay were roughly
based off of that of one of the largest air transports in operation, the C-5 Galaxy. The
cargo bay is enclosed within the hypersonic stage vehicle, as it is utilized through all
stages of flight.
Length [m]
25
Max Width [m]
5
Max Height [m]
5
Max Payload Volume [m3 ]
625
Table 5.1: Condor Payload Bay Dimensions
5.1.2 Subsonic Transport
The final parameters of the subsonic transport stage were initially developed during the
conceptual design phase described in Chapter 3 but were later refined based on the
aerodynamic analysis described in Chapter 4, where they were adjusted such that the
aircraft would be able to perform to its desired performance. The fuselage dimensions
described below were derived to encompass both the cargo bay and the necessary volume
29
5.1. FINAL CONFIGURATION
CHAPTER 5. FINAL CONDOR DESIGN
for avionics, cockpit, control systems, fuel etc. This fuselage is also the fuselage for the
hypsersonic stage vehicle but must be considered whenever examining the subsonic phase
of flight as it will be attached during this time.
Wing Span [m]
126
Sweep [deg]
25
Wing Area [m2 ]
1900
Table 5.2: Subsonic Wing Dimensions
The sizing for the fuselage was conducted under two considerations. First the necessary cargo bay volume capacity and second the volume required to fit the liquid oxygen
and hydrogen fuel used during the majority of the mission. Initial mass sizing produced
an estimated required fuel mass, then the density of liquid oxygen and hydrogen was used
to calculate the necessary storage volume. The majority of the LO2 and H2 will be held
within storage compartments that surround the cargo bay, which will aid in the cooling
of the fuselage as it experiences the high speed flow.
Fuselage Length [m]
50
Max Height [m]
10
Max Width [m]
15
Table 5.3: Subsonic Fuselage Dimensions
Figure 5.1: Subsonic Condor Top View
30
Fuel Capacity [m3 ]
4,063
CHAPTER 5. FINAL CONDOR DESIGN
5.1. FINAL CONFIGURATION
Figure 5.2: Subsonic Condor Side View
Figure 5.3: Subsonic Condor Front View
5.1.3 Hypersonic Transport
Similar to the process undertaken in the design of the subsonic transport, initial sizing of
the hypersonic vehicle occured during the conceptual design phase and was refined after
aerodynamic analysis. The fuselage is the same as that described in the previous section.
Fuselage Length [m]
50
Wing Span [m]
36
Sweep [deg]
73
Table 5.4: Hypersonic Vehicle Dimensions
Figure 5.4: Hypersonic Condor Top View
31
Wing Area [m2 ]
888
5.2. FINAL FLIGHT REGIMES
CHAPTER 5. FINAL CONDOR DESIGN
Figure 5.5: Hypersonic Condor Side View
Figure 5.6: Hypersonic Condor Front View
5.2 Final Flight Regimes
After initial conceptual design and multiple modifications during aerodynamic analysis,
the following flight profile was decided upon:
Stage 1: Take-off to Mach 0.9 at 10,000m
Stage 2: Airbreathing phase of SABRE engines from Mach 0.9 at 10,000m to Mach 5.5
at 25,000m
Stage 3: Rocket phase of SABRE engines from Mach 5.5 at 25,000m to Mach 25 in Low
Earth Orbit
Stage 4: Payload deployment
Stage 5: Vehicle re-entry and landing
32
CHAPTER 5. FINAL CONDOR DESIGN
33
5.2. FINAL FLIGHT REGIMES
5.2. FINAL FLIGHT REGIMES
CHAPTER 5. FINAL CONDOR DESIGN
34
Chapter 6
Conclusions
The Condor is a horizontal take-off/landing vehicle designed to deliver up to 100,000
kilograms to Low Earth Orbit (LEO). It is a fully reusable aircraft that can reach orbit
from any latitude of plus or minus 40 degrees. The Condor operates in two different stages:
a subsonic and hypersonic stage. The subsonic stage consists of the hypersonic stage with
a high aspect ratio wing attached. It is powered by six GE-90 engines in subsonic flight.
The Condor can take-off fully loaded with fuel and payload, but is designed to takeoff with liquid hydrogen and oxygen tanks empty to enable faster climbing and mission
completion. This will also decrease the amount of JP-7 fuel burned in each mission,
which makes for a more environmentally friendly design. After an aerial refuel of the
liquid hydrogen and oxygen tanks, the high aspect ratio detaches and returns to the
airfield of origin. The hypersonic vehicle consists of a delta wing and four SABRE hybrid
rocket engines. This vehicle is then used to accelerate the payload to velocity and altitude
required for LEO.
The initial two-stage concept for the Condor was chosen, because of the enormous
payload that it is required to carry and the velocities that it must carry the payload to.
A large, high aspect ratio wing works great for generating lift at low speeds, but will
produce terribly high drag forces in supersonic and hypersonic regimes. Therefore, it
was determined that the large wing would detach prior to supersonic flight. The SABRE
hybrid rocket engines were selected as supersonic propulsion, because of their versatility.
They are air-breathing to Mach 5.5 and go to pure rocket propulsion for acceleration
beyond that to the Mach 25 required for LEO.
Sizing based on initial weight estimations was grossly underestimated. After several
iterations of aerodynamic analysis in the subsonic, supersonic, and hypersonic regimes it
was determined that the large and delta wings would need to increase substantially in
wingspan and wing area. Also, the volume required to hold the liquid rocket fuel was
overlooked in the early stages of design. The volume of the fuselage of the Condor more
than doubled to hold the hydrogen and oxygen fuels. This drastically increased our drag
and thrust required for the mission.
Aerodynamically, there were still a few lift issues for the Condor even after increasing
the wing sizes. It was determined that the transition velocity between stages would need
to be increased from Mach .8 to Mach .9 in order to maintain sufficient lift. Also, from
the supersonic analysis there is insufficient lift in the low supersonic range (below Mach
1.5). To compensate for this problem, the Condor will go into a shallow dive to accelerate
to Mach 1.5. After reaching this velocity the Condor will generate enough lift to allow it
to continue its ascent to LEO.
35
CHAPTER 6. CONCLUSIONS
From this analysis, it has been determined that this design is feasible, though not
at this point in time. The materials needed to manufacture the wings and body of this
aircraft will need to withstand immense aerodynamic forces and incredible temperatures.
The SABRE engine is still in the design and testing phase of development. Also, this
design requires a high-speed refueling tanker to fuel the Condor at speeds approaching
Mach .9. Currently, no such tanker exists. The runways for take-off and landing will
have to be massive in comparison to current airfields. In the relatively near future, when
these necessities are available, the Condor will be a truly viable design.
36
Appendix A
Preliminary Sizing MATLAB Scripts
Initial sizing design interation calculation were done via MATLAB. All scripts used can
be found below.
37
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.1: Initial Mass Size Script
38
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.2: Subsonic Wing Loading Script
Figure A.3: Supersonic Wing Loading Script
39
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.4: Subsonic Aerodynamic Analysis Code (a)
40
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.5: Subsonic Aerodynamic Analysis Code (b)
41
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.6: Supersonic Aerodynamic Analysis Code
42
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.7: Hypersonic Aerodynamic Analysis Code
43
APPENDIX A. PRELIMINARY SIZING MATLAB SCRIPTS
Figure A.8: Available Thrust Analysis Code
44
Bibliography
[1] ”HOTOL.”
Encyclopedia
Astronautica.
http://www.astronautix.com/lvs/hotol.htm
N.p..
Web.
6
Mar
2013.
[2] ”HOTOL.” Wikipedia. N.p.,
24
http://en.wikipedia.org/wiki/HOTOL
2013.
Web.
6
Mar
2013.
02
[3] HOTOL Over France. N.d. International Space Art NetworkWeb. 6 Mar 2013.
http://api.ning.com/files
[4] Amos, Jonathan. ”Skylon spaceplane engine concept achieves key milestone.”
BBC pag. Web. 6 Mar. 2013. Skylon spaceplhttp://www.bbc.co.uk/news/scienceenvironment-20510112ane engine .
[5] Skylon
front
view.
N.d.
WikipediaWeb.
6
Mar
2013.
http://en.wikipedia.org/wiki/File:Skylonf rontv iew.jpgP age, Lewis.”DARP Areleases0 Blackswif
//www.theregister.co.uk/2008/03
[6] Blackswift. N.d. Air Force Times Web. 6 Mar 2013. http://www.airforcetimes.com/
[7] Rosenburg, Zach. ”Second X-51 hypersonic flight ends prematurely.” Flight n.d., n. pag.
Web. 6 Mar. 2013. http://www.flightglobal.com/news/articles/second-x-51-hypersonicflight-ends-prematurely-358056
[8] Waverider. N.d. GixmagWeb. 6 Mar 2013. http://images.gizmag.com/hero/x-51awaverider.jpg
[9] Warwick, Graham. ”X-51A to demonstrate first practical scramjet .” Flight n.d., n. pag.
Web. 6 Mar. 2013. http://www.flightglobal.com/news/articles/x-51a-to-demonstratefirst-practical-scramjet-215592
[10] ”C-5 A/B/C Galaxy
C-5M Super Galaxy.” The official website of the
U.S. Air Force. 16 Aug. 2012. Air Mobility Command. 6 Mar. 2013
http://www.af.mil/information/factsheets/factsheet.asp?id=84
[11] ”C-17 Globemaster III.” Boeing. 06 Mar. 2013 http://www.boeing.com/defensespace/military/c17/
[12] ”AN-225 Mriya / Super Heavy Transport.” AN-225 Mriya. Antonov Company. 06 Mar.
2013 http://www.antonov.com/aircraft/transport-aircraft/an-225-mriya
[13] ”AN-124-100 Ruslan / Ruslan heavy transport and its modifications.” AN-124-100
Ruslan. Antonov Company. 06 Mar. 2013 http://www.antonov.com/aircraft/transportaircraft/an-124-100-ruslan
45
BIBLIOGRAPHY
BIBLIOGRAPHY
[14] ”A380-800.” A380 aircraft:
Technology and innovation,
range,
specifications, cabin space
comfort. 25 Feb. 2013. Airbus. 06 Mar. 2013
http://www.airbus.com/aircraftfamilies/passengeraircraft/a380family/a380-800/
[15] ”Model
GE90-115B.”
GE
Aviation.
N.p..
Web.
13
http://www.geaviation.com/engines/commercial/ge90/ge90-115b.html
Mar
2013.
[16] ”SABRE (Rocket Engine).” http://en.wikipedia.org/wiki/SABRE
[17] Ashley Landahl, “Aerodynamics of Wings and Bodies,” Addison-Wesley Publishing
Company, Inc., 1965.
[18] Anderson, John. Aircraft Performance and Design. 1. McGraw-Hill, Print, 1998.
46