EXTROVERT      

ADVANCED  CONCEPT  EXPLORATION  ADL  P-­‐2013050202  

 Sean  Chait,  Brett  Kubica  

 Georgia  Institute  of  Technology    School  of  Aerospace  Engineering  

  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.”  

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X-57 Condor: Runway-Based Space LaunchSystem Aerodynamics

Integrative Assignment

AE3021A

Sean ChaitBrett Kubica

Daniel Guggenheim School of Aerospace EngineeringGeorgia Institute of Technology

Atlanta GA 30332-0150

Spring 2013

2

Contents

1 Summary 9

2 Previous Vehicles 132.1 Horizontal Takeoff Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . 13

2.1.1 HOTOL: Horizontal Takeoff and Landing . . . . . . . . . . . . . . 132.1.2 Skylon . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

2.2 Hypersonic Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142.2.1 Blackswift . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142.2.2 Boeing X-51: Waverider . . . . . . . . . . . . . . . . . . . . . . . 15

2.3 Heavy Lift Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

3 Conceptual Design 173.1 Flight Regime Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

3.1.1 Rocket Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 183.1.2 Supersonic-Hypersonic Stage . . . . . . . . . . . . . . . . . . . . . 193.1.3 Subsonic Stage . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

3.2 Initial Wing Loading Design . . . . . . . . . . . . . . . . . . . . . . . . . 213.2.1 Subsonic Stage Aircraft . . . . . . . . . . . . . . . . . . . . . . . . 213.2.2 Supersonic-Hypersonic Aircraft . . . . . . . . . . . . . . . . . . . 21

4 Aerodynamic Analysis 234.1 Low Speed Flight Regime . . . . . . . . . . . . . . . . . . . . . . . . . . 234.2 Supersonic Flight Regime . . . . . . . . . . . . . . . . . . . . . . . . . . 254.3 Hypersonic Flight Regime . . . . . . . . . . . . . . . . . . . . . . . . . . 274.4 Thrust Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5 Final Condor Design 295.1 Final Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

5.1.1 Payload Capabilities . . . . . . . . . . . . . . . . . . . . . . . . . 295.1.2 Subsonic Transport . . . . . . . . . . . . . . . . . . . . . . . . . . 295.1.3 Hypersonic Transport . . . . . . . . . . . . . . . . . . . . . . . . . 31

5.2 Final Flight Regimes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32

6 Conclusions 35

A Preliminary Sizing MATLAB Scripts 37

3

CONTENTS CONTENTS

4

List of Figures

1.1 Condor Space Access System . . . . . . . . . . . . . . . . . . . . . . . . . 91.2 Condor Launch System Overview . . . . . . . . . . . . . . . . . . . . . . 11

2.1 HOTOL Horizontal Takeoff Space Access Vehicle [3] . . . . . . . . . . . . 132.2 Skylon [5] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142.3 Blackswift [6] . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152.4 Boeing X-51: WaveRider [8] . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.1 Current Subsonic Transport Configuration . . . . . . . . . . . . . . . . . 244.2 Current Supersonic Transport Configuration . . . . . . . . . . . . . . . . 25

5.1 Subsonic Condor Top View . . . . . . . . . . . . . . . . . . . . . . . . . 305.2 Subsonic Condor Side View . . . . . . . . . . . . . . . . . . . . . . . . . 315.3 Subsonic Condor Front View . . . . . . . . . . . . . . . . . . . . . . . . 315.4 Hypersonic Condor Top View . . . . . . . . . . . . . . . . . . . . . . . . 315.5 Hypersonic Condor Side View . . . . . . . . . . . . . . . . . . . . . . . . 325.6 Hypersonic Condor Front View . . . . . . . . . . . . . . . . . . . . . . . 325.7 Condor Flight Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

A.1 Initial Mass Size Script . . . . . . . . . . . . . . . . . . . . . . . . . . . 38A.2 Subsonic Wing Loading Script . . . . . . . . . . . . . . . . . . . . . . . 39A.3 Supersonic Wing Loading Script . . . . . . . . . . . . . . . . . . . . . . 39A.4 Subsonic Aerodynamic Analysis Code (a) . . . . . . . . . . . . . . . . . 40A.5 Subsonic Aerodynamic Analysis Code (b) . . . . . . . . . . . . . . . . . 41A.6 Supersonic Aerodynamic Analysis Code . . . . . . . . . . . . . . . . . . 42A.7 Hypersonic Aerodynamic Analysis Code . . . . . . . . . . . . . . . . . . 43A.8 Available Thrust Analysis Code . . . . . . . . . . . . . . . . . . . . . . . 44

5

LIST OF FIGURES LIST OF FIGURES

6

List of Tables

2.1 Heavy Lift Aircraft Parameters . . . . . . . . . . . . . . . . . . . . . . . 16

3.1 ∆V Losses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 183.2 Rocket Stage Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193.3 Required Stage ∆V ′s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193.4 Supersonic-Hypersonic Stage Sizing . . . . . . . . . . . . . . . . . . . . . 203.5 Subsonic Weight Fuel Estimation [15] . . . . . . . . . . . . . . . . . . . 213.6 Subsonic Stage Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213.7 Subsonic Wing Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . 223.8 Supersonic Wing Loading . . . . . . . . . . . . . . . . . . . . . . . . . . 22

4.1 Zero-Lift Drag Buildup . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244.2 Subsonic Aerodynamic Performance Estimation . . . . . . . . . . . . . . 254.3 Supersonic Airfoil Coefficients . . . . . . . . . . . . . . . . . . . . . . . . 274.4 Supersonic Aerodynamic Performance Estimation . . . . . . . . . . . . . 274.5 Hypersonic Aerodynamic Performance Estimation . . . . . . . . . . . . 284.6 Thrust Available . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

5.1 Condor Payload Bay Dimensions . . . . . . . . . . . . . . . . . . . . . . 295.2 Subsonic Wing Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . 305.3 Subsonic Fuselage Dimensions . . . . . . . . . . . . . . . . . . . . . . . . 305.4 Hypersonic Vehicle Dimensions . . . . . . . . . . . . . . . . . . . . . . . 31

7

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 spaceinfrastructure, such as the proposed Space Solar Power (SSP) architecture, viable a newsolution launch system must be developed. This system must have the reusability of amodern heavy-lift cargo aircraft, capable of loading, refueling, and executing its missionmultiple times a day. The requirement of reusability and quick turnaround time makesthe current configuration of a multi-stage rocket used by most modern launch systemsinadequate. These systems require months of preparation, and are not reusable, drivingup the cost per kilogram of placing a payload into orbit. For a launch system to be ableto perform multiple flights to orbit a day, it is necessary for all stages of the proposedsystem to be completely reusable, requiring only the level of maintenance seen by currentcargo 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 viableoption for producing the required launch system.

Significant research into previous attempts at a horizontal take-off/horizontal landingspace launch system gave rise to the concepts used in the preliminary design of the currentlaunch system concept under consideration, the Condor. The Condor will consist of twomain stages, each with their own vehicle. The first is a large blended wing body aircraftwhich will be carrying the second stage, a blunt body hypersonic craft. The first stagewill 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 thesound 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 advancedLACE 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 beensuccessfully jettisoned, the stage two vehicle will re-enter the Earth’s atmosphere and landat its designated launch facility. The entire Condor system will then be quickly refueledand loaded for another launch opportunity. To allow for the uninterrupted delivery ofcargo into orbit, launch facilities will located around the world at various locations toprevent possible interruptions due to weather and other atmospheric conditions.

After an extensive research effort, which gave rise to a preliminary conceptual designfor the Condor, further analysis was conducted to refine and determine the feasibilityof the design. Basic mass sizing and wing loading techniques were used along with apreliminary flight profile to determine a set of geometric of parameters that defined thelaunch system. From here a modeling software was used to solidify the initial design andproduce all of the information to run a thorough aerodynamic analysis. This analysisrevealed that the current vehicle and flight profile were not sufficient to accomplish thegoals of the mission. Modifications were made to both the first and second stage vehicledesigns along with the intended mission profile in order to ensure the entire Condorlaunch system would be able to function efficiently. Final aerodynamic analysis showedthat the launch system framework described in this report is theoretically possible interms of aerodynamics although significant advances in structural design, materials, andhypersonic control systems are necessary. This is to ensure that the Condor is fully ableto cope with the extreme conditions it will face while still performing to full missionsuccess. Although out of reach of modern technology, the Condor may very well be thebasis for a future large payload space launch system.

10

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 (andoftentimes failed) to determine the most feasible approach to meet your mission criteria.Complete design concepts may have already been developed and analyzed, thus saving thetime of performing a comparable analysis to determine the feasibility of a method. Fromthe 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 forthis reason that a significant research effort into horizontal takeoff and landing (HOTOL)orbital vehicles and hypersonic vehicles was undertaken prior to conceptual design of ourcraft. The purpose of this research was two-fold. First, to examine the aerodynamicand payload capabilities of previous design attempts and second, to compare differentpropulsion systems and their optimal flight regimes. By learning about previous andcurrent attempts at HOTOL vehicles, we were able to gather enough information tocreate a feasible preliminary design for a heavy-lift HOTOL vehicle to bring our payloadinto 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 propelledtrolley. The craft’s air breathing rocket engine (the RB545 Swallow) would then takeover and behave like a turbojet until approximately Mach 5. At this point the enginewould 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 gravityand center of pressure resulted in the need for major redesigns. As a result the payloadfraction 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 EuropeanSpace Agency. The vehicle is based off of the HOTOL design, as several of its key designerswere members of the original team. An air-breathing rocket engine, with many similaritiesto a LACE system, is used. The primary difference between the SABRE engine and theconventional LACE design is that the air is not liquefied within the engine, increasingefficiency (the engine has an atmospheric ISP of 3500s). During the flight regime up toMach 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 entireflight. After this point, the air intake is closed off and stored liquid oxygen is used forthe 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 littleover a thousand dollars per kilo. To date, the project has passed all design and testingreviews, 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 craftwill be able to function as a hypersonic cruise vehicle, capable of delivering its payloadanywhere on the planet within a few hours. Unlike many other hypersonic concepts, thecraft 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 propulsionsystem. The turbine engine will be used to get Blackswift to speeds approaching Mach 3and then the ramjet will accelerate the vehicle to Mach 6. Currently the project is not indevelopment as funding for the project was drastically cut causing the project’s ultimatecancellation [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” comesfrom the unmanned crafts use of shockwaves to produce additional lift. The X-51 is a ridealong craft, which is attached to the wing of a B-52 and carried to an altitude of 50,000feet before deployment. After separation, a solid rocket booster ignites to accelerate thecraft 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 ignitionof 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, causingthe premature end of hypersonic flight, coupled with other aerodynamic control issuesmake the current reliability of the craft low. Also current design allows for a payload ofonly 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 togage the current capabilities of heavy lift systems in the subsonic regime so as to determinethe best course of action for designing the subsonic portion of our launch system. Fromthis research, an understanding of payload fraction and the sizing of necessary liftingsurfaces for aircraft performing in this flight regime was developed. Due to the scale ofthe payload that must be placed into orbit (100,000 kg), the largest and most powerfulof modern heavy-lift aircraft were examined.

Aircraft Max Takeoff-Weight Max Payload Wing Span Wing AreaC-5 Galaxy [10] 381,000 kg 122,470 kg 75.31 m 576 m2

C-17 Globemaster [11] 265,350 kg 77,519 kg 51.75 m 353 m2

An-225 Mriya [12] 640,000 kg 250,500 kg 88.4 m 905 m2

An-124 Ruslan [13] 405,000 kg 150,000 kg 73.3 m 628 m2

A-380-800 [14] 590,000 kg 149,800 kg 79.75 m 845 m2

Table 2.1: Heavy Lift Aircraft Parameters

16

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 100,000kg payload to orbit while utilizing completely reusable components. An aircraft that iscapable of producing the lift necessary to take-off and pass through the subsonic flightregime would be so large that any attempt to reach supersonic, let alone hypersonicspeeds, would meet insurmountable heating, loading, and thrust issues that would end inthe craft’s failure. It was for this reason that a three stage, two craft system was decidedupon.

Our design for a heavy-lift, horizontal takeoff and landing, hypersonic vehicle consistsof 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 ablunt nose to decrease heating in hypersonic flight and a delta wing with a high sweepangle and low aspect ratio. We will use the SABRE hybrid rocket engines as our mainpowerplant. The SABRE is an offshoot of the LACE engine concept. It uses a heliumloop to precool incoming air and turn the turbine. It can compress air from ambient to140 atmospheres. The engine uses liquid hydrogen as fuel and will close off its inlet athigh speed and altitude, becoming exclusively a rocket engine [4].

The flying wing will be attached on takeoff and used to take the second stage vehicleto 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 stageand be responsible for bringing the entire system to altitude. As the vehicle approachesthe supersonic flight regime, the large wing will detach and return to the airfield. Thevehicle then relies solely on the SABRE engines, which adjust their inlets with increasingMach number and altitude. At approximately Mach 5.14 or 28.5km the inlet seals andthe SABRE becomes a hydrogen-fueled rocket. When the vehicle reaches LEO (approxi-mately Mach 25 at 150km) it will jettison its payload and return to a designated airfieldunder little or no power from its engines.

From this baseline concept, an initial flight plan and aircraft conceptual design wasdeveloped using different mass and wing area sizing techniques. These baseline valueswere 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 toMach 5.5 at 20,000 meters

Stage 3: Rocket phase of the SABRE engines from Mach 5.5 at 20,000 meters to LowEarth Orbit (LEO), which occurs at 150,000 meters and Mach 25 (approximately 7,780meters 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 wasconducted in three stages based upon our expected flight regimes. The flight regimesare defined as the takeoff/transonic phase, the supersonic/hypersonic phase, and therocket stage. The rocket stage will encompass orbit insertion, jettison of payload, andre-entry. These regimes are based upon the staging of the vehicle as well as the state ofthe engine during those phases. Derivation of the subsequent phases and their respectivemass requirements are described below. A comprehensive MATLAB script was developedfor 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 com-mence when the vehicle has been accelerated to a speed of Mach 5.5. At this point in timethe SABRE engines will convert from a hybrid airbreathing rocket engine to a conventialrocket engine running on liquid hydrogren and oxygen. Since the vehicle will already betravelling at Mach 5.5, the ∆V requirements of this phase can be determined using thenecessary velocity requirements of entering orbit and the velocity already achieved duringthe supersonic-hypersonic phase. While determining these total ∆V requirements of thisstage, other considerations must also be made. To improve efficiency, the launch site ofthe craft and maneuvers of the rocket phase will be utilized to obtain an extra ∆V fromrotation of the Earth, thereby reducing fuel required to enter LEO. However, losses dueto 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 forinitial sizing.

∆Vlosses Magnitude∆Vgravity 1000 m/s∆Vdrag 50 m/s

∆Vsteering 100 m/s∆Vrotation 400 m/s

Table 3.1: ∆V Losses

The mass ratio of the rocket stage was determined using the rocket equation and theaforementioned losses. When in pure rocket flight the SABRE engines operate at an Ispof460s, a very high value even for liquid rockets [16]. A high (for a rocket) structural ratioof 0.10 was assumed due to the added complexity of the SABRE engines, the large fuel

18

CHAPTER 3. CONCEPTUAL DESIGN 3.1. FLIGHT REGIME SIZING

Entrance Mach Mass Ratio Structural Mass [kg] Propellant [kg] Total Mass [kg]5.5 4.497 63,500 572,000 735,00010 3.276 33,800 305,000 438,00015 2.305 17,000 153,000 270,000

Table 3.2: Rocket Stage Sizing

tanks required to store liquid hydrogen and structural/control considerations for hyper-sonic maneuvering. The full payload of 100,000kg was also assumed.

∆VLEO = Ispg0ln(MR)−∆Vgravity −∆Vdrag −∆Vsteering + ∆Vrotation

MR = e(∆VLEO+∆Vgravity+∆Vdrag+∆Vsteering−∆Vrotation)/(Ispg0)

Implementation of the modified form of the rocket equation in MATLAB (See Ap-pendix A) created a powerful design iteration tool to be used to determine the point atwhich the engines will be converted from an airbreathing to a pure rocket mode. Anal-ysis (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 portionof the velocity required to achieve orbit has already been acquired in the supersonic-hypersonic phase. However, aerodynamic analysis will show that although this creates alowest mass vehicle, the engines will not have sufficient thrust (due to engine lapse) tomaintain flight. It was therefore determined that an entrance in the rocket phase wouldbegin at Mach 5.5, increasing our initial mass drastically but this would grant us enoughthrust to insert into LEO successfully. The mass of the final, rocket stage will drivethe 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 efficiencyof the SABRE engines during its airbreathing operations is drastically higher than thatduring rocket operations (due to the use of air as an oxidizer) making it desirable to haveas 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 willmark entrance into ”rocket” mode therefore make it possible to calculate the remainingrequired ∆V ′sor the remaining two stages based on a transition between subsonic andsupersonic modes at Mach 0.8 and hypersonic and rocket phases at Mach 5.5.

Mach Regime ∆Vrequired[m/s]0-0.80 241

0.80-5.5 1,4105.5-25.0 6,784

Table 3.3: Required Stage ∆V ′s

3.1.2 Supersonic-Hypersonic Stage

After the entrance point into the rocket stage was determined and initial sizing deter-mined, it is possible to next determine the mass of the supersonic-hypersonic phase of

19

3.1. FLIGHT REGIME SIZING CHAPTER 3. CONCEPTUAL DESIGN

Entrance Mach Exit Mach Mass Ratio Propellant [kg] Total Mass [kg]0.8 5.5 1.041 29,900 765,000

Table 3.4: Supersonic-Hypersonic Stage Sizing

operations. A single phase was decided upon for travel between transonic speeds and theend of the hypersonic regime due to the extremely efficient capabilities of the SABREengines. Unlike traditional supersonic and hypersonic engines such as ramjets and scram-jets which are only efficient once they approach their optimal Mach range, the SABREengines perform at the same efficiency throughout flight operations. Due to the use of airas an oxidizer, the engines run at an Ispf 3600s, a figure that rivals many electric propul-sion systems [16]. The dual capabilities of the engines also negate the need for a physicalstage separation of the craft during the transition from hypersonic to rocket flight, rathera simple ”mode” change of the engines. For this reason the only extra structural massneeded in this phase is that required to carry the fuel required, all other structural masswas already accounted for in the rocket stage. Also due to the high efficiency of theengines during airbreathing operations, a significantly smaller amount of fuel is neededduring this phase of operations than the rocket stage, even though the total velocitygained 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 andpayload from high transonic speeds to Low Earth Orbit, it is possible to determine thesize of the initial vehicle required to take the payload from takeoff to hypersonic vehicledeployment. The transition from the subsonic to supersonic-hypersonic phases will markthe only physical seperation of staged craft during flight operations. The logic for thisdesign decision lies in the aerodynamic differences between subsonic and super/hypersonicflight. An aircraft with a significant amount of surface area is required to get such amassive payload off the ground and up to typical cruise conditions. However such anaircraft would not be able to withstand the forces placed on it during supersonic andhypersonic flight, let alone re-entry. Therefore the ”payload” of the subsonic, heavy liftaircraft will be the hypersonic vehicle, designed to have enough thrust to power throughthe Mach barrier and quickly up to high Mach numbers, as well as an aerodynamic designthat tends itself to high speed flight.

For optimal lift and structural weight considerations, a blended wing body configu-ration was chosen for the subsonic phase of flight. These aircraft have an empty weightfraction of below 50% and are ideal when such a high payload is being carried. For ourconsiderations an empty weight fraction of 0.50 was assumed. The current design itera-tion also utilizes six GE 90-115 Turbofan engines, the same engine used on the Boeing777 [15]. Using the specific fuel consumption of the engines, an estimated time of takeoffto 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 generatedfor 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 ValueSFCcruise 1.47E-5 kg/s/NSFCsealevel 0.92E-5 kg/s/NSFCAverage 1.195E-5 kg/s/N

Average Thrust During Takeoff 400,000 NTime to Separation 45 min

Weight of Fuel Per Engine 17136 kgWeight of Fuel Total 77436 kg

Table 3.5: Subsonic Weight Fuel Estimation [15]

We/W0 Structural Mass [kg] Propellant [kg] Total Mass [kg]0.50 843,000 77436 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 load-ing will result in an increased climb rate and a decrease in thrust required at cruise. Anaircraft with higher wing loading is better suited for high speed flight, due to the smallerwing and decreased drag. However, the higher wing loading means faster takeoff/landingdistance and velocity, because the wing will need to generate the same amount of liftas the larger wing. A seperate analysis was run for both the subsonic and supersonic-hypersonic 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 altitudeis set at sea level, Mach number set at 0.15, and angle of attack set at 7 degrees. Thevariables are weight, aspect ratio, sweep, and lift-curve slope. In the process of estimatingwing loading, we ran a function in Matlab to return a wing span and area to help sizeour wings at the same time. After several iterations we selected a wing with an aspectratio of 8, sweep of 25 degrees, lift-curve slope of 5, span of 121 meters, and area of 1,804square meters. This configuration yielded a wing loading of 931kg/m2for a 1,680,000 kgaircraft.

3.2.2 Supersonic-Hypersonic Aircraft

To estimate wing loading in the supersonic to hypersonic regime, we set weight, aspectratio, and wing length (root chord) as our variables. We decided to use wing length as avariable, because we assumed that our supersonic stage would utilize a delta wing. Heldconstant 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 of13.4 meters, and area of 335.2 square meters. The wing loading is 1193.5kg/m2 Ouronly concern with this configuration is its ability to produce the required lift at transonicspeeds after the subsonic wing detaches. This will be determined, and the wing adjusted,after aerodynamic analysis.

21

3.2. INITIAL WING LOADING DESIGN CHAPTER 3. CONCEPTUAL DESIGN

Iteration 1 2 3 4 5 6Input

Weight [kg] 1,680,000 1,680,000 1,680,000 1,680,000 1,680,000 1,680,000Aspect Ratio 7 8 7 7 8 8Sweep [deg] 25 25 30 25 25 25

dCl/dα 6 6 6 5 5 6Output

Wing Span [m] 106 112 258 114 121 112Wing Area [m2] 1,575 1,514 9,373 1,837 1,804 1,542

Wing Loading [kg/m2] 1,066 1087 179.2 914.6 931.1 1,089

Table 3.7: Subsonic Wing Loading

Iteration 1 2 3 4 5 6Input

Weight [kg] 765,000 765,000 765,000 765,000 765,000 765,000Aspect Ratio 2 1.5 1 0.5 0.5 1

Wing Length [m] 50 50 50 50 75 75Output

Wing Span [m] 6.4 8.5 12.8 13.4 17.1 8.5Wing Area [m2] 160 214 320 335 641 320

Wing Loading [kg/m2] 4,774 3,580 2,387 1,193 1,193 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 mustbe done to determine the current benefits and deficiencies of the proposed aircraft. Priorto analysis, a firm geometric design of the aircraft was developed. Using the open sourcesoftware OpenVSP, an aircraft was designed based off of the general parameters of wingspan, area, aspect ratio, etc previously determined during the initial mass and wing load-ing sizing of the stages. For initial analysis purposes a NACA 0012 airfoil was assumedfor the subsonic phase until investigation suggests a preferred alternative. The flyingwing which constitutes the detachable subsonic stage was modeled with the supersonic-hypersonic hybrid blunt body delta wing craft attached to its bottom. The size of thefuselage of the supersonic-hypersonic stage was modeled after the necessary space require-ments of the USAF C-5 heavy lift aircraft to account for a large payload. Additional spacewas added for engines, fuel, and cockpit, however the space designated for this is subjectto growth as further research and analysis is conducted. Once all of these parameterswere determined and successfully modeled in OpenVSP, the software’s features allowedfor simple extraction of geometric data for use in aerodynamic analysis.

The resulting baseline design was utilized in initial vehicle aerodynamic analysis. Asthe analysis was conducted and deficiencies identified, modifications were made to theaircraft design and flight profile accordingly. Initial analysis, the resulting changes, andthe 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 thetransonic flight regime at which point the second stage can begin operations. The secondstage is being designed for optimal performance in the supersonic and hypersonic flightregimes 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 oflift 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 dragprofile must be developed. For the purposes of this portion of analysis an empericalbuildup based off of methods presented in Aircraft Performance and Design by John An-derson [18]. The following formulas were developed using regressions of common transportaircraft 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 ofthe aircraft’s performance. For the purposes of this analysis we are first going assumeworst case scenario of completely turbulent flow and second assume three sections forthe 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). Furtherrefinement of design will later include details such as engines, extra fuel tanks, interfacemechanisms, etc and will add a significant amount of zero-lift drag. For this reason, aswell as to account for parasite drag, the calculated Cdowill be multiplied by three. Alsofor the purposes of this assignment, it is being assumed that a highly advanced airfoil isbeing used such that it has a critical Mach number 0.81 and will therefore allow the craftto operate at Mach 0.9 without encountering any supersonic flow.

FFWing = 1 + 2(t/c) + 60(t/c)4 (4.1)

FFFuselage = 1 +60

f 3+

f

400(4.2)

Rec =ρ∞V∞c

µ∞(4.3)

Cfturbulent =0.074

Re0.2c

(4.4)

CDo =∑

Cfi · (Sweti

Srefi

) · FFi ·Qi (4.5)

Component Cf Swet FF [m2] Q CDo

Subsonic Blended Wing 0.0020 3800 1.252 1.3 0.0064Hypersonic Wing 0.0017 649.4 1.252 1.1 0.0001

Hypersonic Blunt Body 0.0015 790.2 1.079 1.1 0.0001Total Zero-Lift Drag 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 theaircraft. Lift calculations are also attainable using the lift curve slope previously derived.

24

CHAPTER 4. AERODYNAMIC ANALYSIS 4.2. SUPERSONIC FLIGHT REGIME

Mach Alt [m] α[deg] b [m] Lift [kN] Drag [kN] Tavail[kN]0.9 10,000 3 88.4 17,200 977 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 +KC2L (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 theaircraft, it was found that although the aircraft was able to produce enough lift andthrust to take-off from a runway unassisted, it was not able to produce enough lift toreach an altitude of 10,000 meters at Mach 0.8. The analysis was repeated for differentgeometric configurations using the current flight profile where it was determined that asubsonic aircraft of nearly double the wing area, in the current configuration, would benecessary 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 necessaryif the aircraft was travelling at Mach 0.9 at an altitude of 10,000 meters. The higherentrance Mach number of the second stage was then analyzed for performance and it wasdetermined that both the subsonic and supersonic aricraft would be able to function underthe new flight regime. Therefore, the suggested changes to the geometric configurationof 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 estimatelift 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 partof 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

µ = arcsin(1

M) (4.11)

To analyze the aerodynamic forces at supersonic speeds the aspect ratio must first becorrected for compressible flow,

β =√M2 − 1 (4.12)

AR =ARo

β(4.13)

Then, using slender wing theory, the coefficient of lift and wave drag can be deter-mined,

CL,wing =π

2ARα (4.14)

CD,wave =4α2

β(4.15)

The skin friction coefficient can be calculated after adjusting for compressible, turbu-lent flow conditions,

T ∗

T∞= 1 + 0.1198M2

∞ (4.16)

µ∗

µ∞= (

T ∗

T∞)32 (4.17)

Cf = 0.295T∞T ∗

log(Re∞T∞T ∗

µ∞µ∗

)−2.45 (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 wingand using an adjustment estimation to account for the loss of lift due to the fuselageinterference [17],

q∞ =1

2V 2∞ (4.20)

Lwing = q∞SCL,wing (4.21)

Ltotal = Lwing(1−rfuselage

b

2

+rfuselage

b

4

) (4.22)

Total drag acting of the Condor in flight is calculated assuming a flat plate for thefuselage. The flat plate is equal to the largest cross-sectional area of the fuselage, thereforeoverestimating drag due to the fuselage,

Dtotal = q∞(Swing + Sfuselage)CD,total (4.23)

Finally, thrust required is equal to total drag, unless the wings do not generate enoughlift. If the wings do not produce sufficient lift, then the thrust required is equal to thetotal drag plus the thrust vector in the lift direction needed to compensate for the lackof lift,

26

CHAPTER 4. AERODYNAMIC ANALYSIS 4.3. HYPERSONIC FLIGHT REGIME

Treq = D +W − Lsin(α)

(4.24)

A series of altitudes, angles of attack, and Mach numbers were run to determine theaerodynamic coefficients and forces on the aircraft during the supersonic phase. Based oninitial analysis, wing span needed to be increased for the supersonic phase. The secondstage of the Condor’s wingspan was adjusted from an original 15 meters to a much larger36 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 thrustrequirements caused by the increase in drag were well within the pre-existing thrustmargin of the SABRE engines.

Mach Alt [m] α[deg] CL CD,wave Cf CD,total

1.5 10,000 8 0.235 0.0697 0.00020 0.07002 10,000 6 0.114 0.0253 0.00018 0.02603 10,000 4 0.047 0.0069 0.00016 0.00714 15,000 4 0.034 0.0050 0.00015 0.0052

5.5 20,000 6 0.037 0.0081 0.00014 0.0083

Table 4.3: Supersonic Airfoil Coefficients

Mach Alt [m] α[deg] Lift [kN] Drag [kN]1.5 10,000 8 7,578 2,5482 10,000 6 10,104 1,6523 10,000 4 15,156 1,0284 15,000 4 12,706 634

5.5 20,000 4 10,964 395

Table 4.4: Supersonic Aerodynamic Performance Estimation

4.3 Hypersonic Flight Regime

In the hypersonic regime (above Mach 5), Modified Newtonian Aerodynamics can beapplied. Here we get:

Cp = Cpo,minsin2θ (4.25)

Assuming γ = 1.4 and as Mach number approach infinity, the equation becomes:

Cp = 1.84sin2θ (4.26)

We assume that the Cp on the sides of the wing and body not directly interactingwith the flow are zero.

At Mach 10 and 15 we are using the thrust vector in the lift direction to compensatefor a lack of lift. We could extend the wings at high altitude to generate more lift to givethe aircraft more thrust for climbing.

27

4.4. THRUST AVAILABLE CHAPTER 4. AERODYNAMIC ANALYSIS

Mach Alt [m] α[deg] Lift [kN] Drag [kN] Treq[kN]5.5 25,000 10 12,290 4,032 4,03210 40,000 17 5,746 2,089 7,07315 50,000 20 4,185 1,757 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 densityat sea level. We assume air-breathing to 20km. Above 20km the SABRE switches over torocket propulsion. It is assumed that thrust is constant, because the density is so smalland pressure is close to vacuum.

Altitude [m] Thrust Available [kN]10,000 2,646.420,000 56930,000 117 (airbreathing)30,000 11,760 (rocket)

Table 4.6: Thrust Available

From this analysis, we need to start the rocket at a lower altitude than initiallythought. The aircraft can get to Mach 5.5 at 20,000 meters before switching over torockets 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 metersand Mach 0.9 under the power of six GE-90’s, where it levels off in order to fill thehydrogen tank and some of the oxygen tank. When aerial refuel is complete, the highaspect ratio wing is jettisoned and returns to the airport of origin and the SABRE enginesengage. The Condor then goes into a shallow dive to break the sound barrier and reachMach 1.5 to generate sufficient lift to climb. The aircraft then climbs to 25,000 metersand Mach 5.5, at which point the SABRE engines switch over to liquid rocket. Thehypersonic vehicle climbs to 50,000 meters and Mach 15 then executes a vertical, noseup maneuver to climb to 150,000 meters and Mach 25 entering Low Earth Orbit. Thepayload is jettisoned in LEO. The Condor then begins its descent back to Earth. The lowaspect ratio wing extends outward to increase wing area on landing. A large parachutedeploys 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 100,000kg payload into low Earth orbit. The geometric dimensions of the cargo bay were roughlybased off of that of one of the largest air transports in operation, the C-5 Galaxy. Thecargo bay is enclosed within the hypersonic stage vehicle, as it is utilized through allstages of flight.

Length [m] Max Width [m] Max Height [m] Max Payload Volume [m3]25 5 5 625

Table 5.1: Condor Payload Bay Dimensions

5.1.2 Subsonic Transport

The final parameters of the subsonic transport stage were initially developed during theconceptual design phase described in Chapter 3 but were later refined based on theaerodynamic analysis described in Chapter 4, where they were adjusted such that theaircraft would be able to perform to its desired performance. The fuselage dimensionsdescribed 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 thehypsersonic stage vehicle but must be considered whenever examining the subsonic phaseof flight as it will be attached during this time.

Wing Span [m] Sweep [deg] Wing Area [m2]126 25 1900

Table 5.2: Subsonic Wing Dimensions

The sizing for the fuselage was conducted under two considerations. First the neces-sary cargo bay volume capacity and second the volume required to fit the liquid oxygenand hydrogen fuel used during the majority of the mission. Initial mass sizing producedan estimated required fuel mass, then the density of liquid oxygen and hydrogen was usedto calculate the necessary storage volume. The majority of the LO2 and H2 will be heldwithin storage compartments that surround the cargo bay, which will aid in the coolingof the fuselage as it experiences the high speed flow.

Fuselage Length [m] Max Height [m] Max Width [m] Fuel Capacity [m3]50 10 15 4,063

Table 5.3: Subsonic Fuselage Dimensions

Figure 5.1: Subsonic Condor Top View

30

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 ofthe hypersonic vehicle occured during the conceptual design phase and was refined afteraerodynamic analysis. The fuselage is the same as that described in the previous section.

Fuselage Length [m] Wing Span [m] Sweep [deg] Wing Area [m2]50 36 73 888

Table 5.4: Hypersonic Vehicle Dimensions

Figure 5.4: Hypersonic Condor Top View

31

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 5.5at 25,000m

Stage 3: Rocket phase of SABRE engines from Mach 5.5 at 25,000m to Mach 25 in LowEarth Orbit

Stage 4: Payload deployment

Stage 5: Vehicle re-entry and landing

32

CHAPTER 5. FINAL CONDOR DESIGN 5.2. FINAL FLIGHT REGIMES

Figure 5.7: Condor Flight Profile

33

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 100,000kilograms to Low Earth Orbit (LEO). It is a fully reusable aircraft that can reach orbitfrom 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 witha 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 take-off with liquid hydrogen and oxygen tanks empty to enable faster climbing and missioncompletion. 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 theliquid hydrogen and oxygen tanks, the high aspect ratio detaches and returns to theairfield of origin. The hypersonic vehicle consists of a delta wing and four SABRE hybridrocket engines. This vehicle is then used to accelerate the payload to velocity and altituderequired for LEO.

The initial two-stage concept for the Condor was chosen, because of the enormouspayload 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 willproduce terribly high drag forces in supersonic and hypersonic regimes. Therefore, itwas determined that the large wing would detach prior to supersonic flight. The SABREhybrid 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 accelerationbeyond that to the Mach 25 required for LEO.

Sizing based on initial weight estimations was grossly underestimated. After severaliterations of aerodynamic analysis in the subsonic, supersonic, and hypersonic regimes itwas determined that the large and delta wings would need to increase substantially inwingspan and wing area. Also, the volume required to hold the liquid rocket fuel wasoverlooked in the early stages of design. The volume of the fuselage of the Condor morethan doubled to hold the hydrogen and oxygen fuels. This drastically increased our dragand thrust required for the mission.

Aerodynamically, there were still a few lift issues for the Condor even after increasingthe wing sizes. It was determined that the transition velocity between stages would needto be increased from Mach .8 to Mach .9 in order to maintain sufficient lift. Also, fromthe supersonic analysis there is insufficient lift in the low supersonic range (below Mach1.5). To compensate for this problem, the Condor will go into a shallow dive to accelerateto Mach 1.5. After reaching this velocity the Condor will generate enough lift to allow itto continue its ascent to LEO.

35

CHAPTER 6. CONCLUSIONS

From this analysis, it has been determined that this design is feasible, though notat this point in time. The materials needed to manufacture the wings and body of thisaircraft will need to withstand immense aerodynamic forces and incredible temperatures.The SABRE engine is still in the design and testing phase of development. Also, thisdesign requires a high-speed refueling tanker to fuel the Condor at speeds approachingMach .9. Currently, no such tanker exists. The runways for take-off and landing willhave to be massive in comparison to current airfields. In the relatively near future, whenthese 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 canbe 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

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46