October 2012

NASA/CR-2012-217774

Solution-Space Screening of a Hypersonic 

Endurance Demonstrator

Bernd Chudoba, Gary Coleman, Amit Oza, and Lex Gonzalez

University of Texas at Arlington, Arlington, Texas

Paul Czysz

HyperTech Concepts LLC, St. Louis, Missouri
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Hampton, Virginia 23681-2199 under Contract NNL09AA00A

October 2012

NASA/CR-2012-217774

Solution-Space Screening of a Hypersonic 

Endurance Demonstrator

Bernd Chudoba, Gary Coleman, Amit Oza, and Lex Gonzalez

University of Texas at Arlington, Arlington, Texas

Paul Czysz

HyperTech Concepts LLC, St. Louis, Missouri
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

 

Trade names and trademarks are used in this report for identification only. Their usage does not 

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Administration. 
i

FOREWORD

This report summarizes the results of the Solution-Space Screening for a Hypersonic Endurance 

Demonstrator study performed from 14 June 2010 through 31 August 2010 under the National Institute of 

Aerospace (NIA) contract NNL09AA00A, Task Order No. NIA Activity C10-2800-UTA for the National 

Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) by the Aerospace 

Vehicle Design (AVD) Laboratory at the Mechanical and Aerospace Department (MAE) at the The 

University of Texas at Arlington (UTA).

The study was funded by the Systems Analysis and Concepts Directorate (SACD), Vehicle Analysis 

Branch (VAB), with John J. Korte as study monitor and Lawrence L. Green as alternate study monitor.

Bernd Chudoba was the manager of the Solution-Space Screening for a Hypersonic Endurance 

Demonstrator project and Gary Coleman was the deputy manager. The study was conducted within the 

AVD Laboratory at UTA MAE under the direction of Bernd Chudoba.

The support of the following individuals is gratefully acknowledged: Ajay Kumar (NASA LaRC), John J. 

Korte (NASA LaRC), Charles P. Leonard (NASA LaRC), Lawrence L. Green (NASA LaRC), Jeffrey S. 

Robinson (NASA LaRC), Roger A. Lepsch (NASA LaRC), John G. Martin (NASA LaRC), Janet M. 

Ross (NASA LaRC), David E. Glass (NASA LaRC), John R. Olds (SpaceWorks Engineering, Inc.), 

William J.D. Escher (SpaceWorks Engineering, Inc.), Kevin G. Bowcutt (The Boeing Company), 

Heribert Kuczera (EADS Space Transportation), Dietrich E. Koelle (TCS-TransCostSystems), Peter W. 

Sacher (AeroSpace Consulting), Ivan Burdun (Intelonics, Inc.), and Georg Poschmann (Airbus Industrie).

The study deliverables consist of:

 Weekly teleconferences with supporting Microsoft PowerPoint files.

 Two-day UTA MAE AVD Laboratory workshop with supporting Microsoft PowerPoint file (1415 

September 2010).

 One-day NASA LaRC VAB VIP presentation with supporting Microsoft PowerPoint file (21 October 

2010).

 AVD Laboratory final report in Adobe PDF format (01 November 2010).
ii

CONTENTS

FOREWORD i

CONTENTS ii

LIST OF FIGURES iv

LIST OF TABLES vi

NOTATION viii

1 INTRODUCTION.............................................................................................................................1

2 OVERALL STUDY METHODOLOGY..................................................................................................2

3 MISSION REQUIREMENT AND RESEARCH OBJECTIVES....................................................................5

4 DB AND KB REVIEW OF HYPERSONIC DEMONSTRATORS ...............................................................8

4.1 Hypersonic flight vehicle data-base (DB) 9

4.2 Hypersonic flight vehicle knowledge-base (KB) 11

5 OVERALL TRADE SPACE AND REDUCED TRADE SPACE................................................................ 14

6 PARAMETRIC SIZING AND SOLUTION-SPACE SCREENING ............................................................. 17

6.1 AVD sizing process summary 17

6.2. Disciplinary methods overview 18

6.3 Description of solution space visualization 19

6.4 Solution-space screening 24

6.4.1 Air launched, Mach 6, hydrogen fuel 25

6.4.2 Air launched, Mach 8, hydrogen fuel 28

6.4.3 Air launched, Mach 6, kerosene fuel 31

6.4.4 Air launched, Mach 8, kerosene fuel 34

6.4.5 Air launched, Mach 6, dual-fuel 37

6.4.6 Air launched, Mach 8, dual-fuel 40

6.4.7 Expendable booster, Mach 6, hydrogen fuel 43

6.4.8 Expendable booster, Mach 8, hydrogen fuel 46

6.4.9 Expendable booster, Mach 6, kerosene fuel 49

6.4.10 Expendable booster, Mach 8, kerosene fuel 52

7 SOLUTION-SPACE COMPARISON AND BASELINE SELECTION ........................................................ 55

7.1 Solution-space comparison 55
iii

7.2 Design-point comparison 56

7.3 Baseline vehicle selection 58

8 SUMMARY OF CONCLUSIONS AND RECOMMENDATIONS................................................................ 60

8.1 Design lessons learned 60

8.2 Recommendations 61

REFERENCES ................................................................................................................................. 63

APPENDIX A: DISCIPLINARY METHODS LIBRARY .............................................................................. 65

APPENDIX B: TRADE-STUDY ASSUMPTIONS AND DATABASE ............................................................ 75

B-1 Trade 001: Air launched, Mach 6, hydrogen fuel 75

B-2 Trade 002: Air launched, Mach 8, hydrogen fuel 82

B-3 Trade 003: Air launched, Mach 6, kerosene fuel 89

B-4 Trade 004: Air launched, Mach 8, kerosene fuel 96

B-5 Trade 005: Air launched, Mach 6, dual fuel 103

B-6 Trade 006: Air launched, Mach 8, dual fuel 110

B-7 Trade 007: Expendable booster, Mach 6, hydrogen fuel 117

B-8 Trade 008: Expendable booster, Mach 8, hydrogen fuel 123

B-9 Trade 009: Expendable booster, Mach 6, kerosene fuel 129

B-10 Trade 0010: Expendable booster, Mach 8, kerosene fuel 135
iv

LIST OF FIGURES

2-1 Study approach to develop reusable hypersonic endurance test bed. 2

2-2 AVD Laboratory integrated design life-cycle process. 3

2-3 AVD Laboratory integrated VAB-AVD team study approach. 4

3-1 Iterative nature of the mission and objectives and baseline vehicle(s) selection process. 5

3-2 HYFAC project research objectives identification and evaluation participants. 6

3-3 VAB/AVD Laboratory baseline vehicle development sequence. 7

4-1 Dedicated AVD Laboratory data-base (DB) and organization scheme. 9

4-2 Selected AVD Laboratory data-base (DB) entries. 10

4-3 Design lessons-learned of selected design case-studies. 12

4-4 KB development steps resulting in numerical design guidelines. 12

4-5 Integration scheme of data domain, knowledge domain, and process domain. 13

5-1 Reduced trade space explored. 16

6-1 AVDsizing methodology visualized via Nassi-Schneidermann structogram. 17

6-2 Complete hypersonic vehicle convergence at each design point (e.g., Mach 6, 30 minutes,

cruiser configuration). 19

6-3 Solution-space carpet plot of TOGW and Istr for varying vehicle slenderness (τ) and cruise 

time. 20

6-4 Landing and thrust minus drag (T-D) constraints imposed on solution space. For the Mach 6 

demonstrator, the landing constraint is more constraining than T-D. 20

6-5 Superposition of structural indices to provide the final constraint for determining technical 

solution space. 21

6-6 Definition of structural capability indices used for this study. 22

6-7 B-52H underwing mount geometric constraints. 22

6-8 Summary of B747-SCA and B-52H constraints for hypersonic demonstrator study. 23

6-9 Summary of Minotaur I and Taurus XL first-stage constraints for hypersonic demonstrator 

study. 23

6-10 Summary of solution spaces for each design trade explored. 24

6-11 Air-launched Mach 6 hydrogen-fueled trade solution space. 26

6-12 B-52H planform-area constraint and suggested design point for Mach 6 hydrogen-fueled

trade. 26

6-13 Summary of suggested design point for air-launched Mach 6 hydrogen-fueled trade. 27

6-14 Air-launched Mach 8 hydrogen-fueled trade solution space. 29

6-15 B-52H planform-area constraint and suggested design point for Mach 8 hydrogen-fueled

trade. 29

6-16 Summary of suggested design point for air-launched Mach 8 hydrogen-fueled trade. 30

6-17 Air-launched Mach 6 kerosene-fueled trade solution space. 32

6-18 B-52H planform-area constraint and suggested design point for Mach 6 kerosene-fueled

trade. 32

6-19 Summary of suggested design point for air-launched Mach 6 kerosene-fueled trade. 33

6-20 Air-launched, Mach 8 kerosene-fueled trade solution space. 35

6-21 B-52H planform-area constraint and suggested design point for Mach 8 kerosene-fueled

trade. 35

6-22 Summary of suggested design point for air-launched Mach 8 kerosene-fueled trade. 36

6-23 Air-launched Mach 6 dual-fuel trade solution space. 38
v

6-24 B-52H planform area constraint and suggested design point for Mach 6 dual-fuel trade. 38

6-25 Summary of suggested design point for air-launched Mach 6 dual-fuel trade. 39

6-26 Air-launched Mach 8 dual-fuel trade solution space. 41

6-27 B-52H planform-area constraint and suggested design point for Mach 6 dual-fuel trade. 41

6-28 Summary of suggested design point for air-launched Mach 8 dual-fuel trade. 42

6-29 Expendable-booster Mach 6 hydrogen-fueled trade solution space. 44

6-30 Expendable-booster constraints and suggested design point for Mach 6 hydrogen-fueled

trade. 44

6-31 Summary of suggested design point for expendable-booster Mach 6 hydrogen-fueled trade. 45

6-32 Expendable-booster Mach 8 hydrogen-fueled trade solution space. 47

6-33 Expendable-booster constraints and suggested design point for Mach 8 hydrogen-fueled

trade. 47

6-34 Summary of suggested design point for expendable booster Mach 8 hydrogen-fueled trade. 48

6-35 Expendable-booster Mach 6 kerosene-fueled trade solution space. 50

6-36 Expendable-booster constraints and suggested design point for Mach 6 kerosene-fueled

trade. 50

6-37 Summary of suggested design point for expendable-booster Mach 6 kerosene-fueled trade. 51

6-38 Expendable-booster Mach 8 kerosene-fueled trade solution space. 53

6-39 Expendable-booster constraints and suggested design point for Mach 8 kerosene-fueled

trade. 53

6-40 Summary of suggested design point for expendable-booster Mach 8 kerosene-fueled trade. 54

7-1 Hydrogen-fueled vehicles allow for a larger technical solution space compared to kerosenefueled vehicles. 55

7-2 Configuration geometry of proposed hydrogen and kerosene hypersonic baseline vehicle 

designs. 56

7-3 The Mach 8 air-launched vehicle represents the largest operational flexibility, while the 

Mach 6 air-launched vehicle has the larger growth capability and design confidence. 

Because the Mach 6, 30 min and Mach 8, 10 min solution curves over-lay, the Mach 6, 30

min vehicle potentially could perform the Mach 8 mission for 10 min. 58

7-4 The Mach 6 kerosene-fuel expendable-booster trade is the only trade study that allows for 30

minutes cruise endurance. 59

A-2 Wetted area description of a hypersonic air-breather. 67

A-3 Assumed trajectory of the hypersonic cruiser. 69

A-4 Structural index prediction description. 74
vi

LIST OF TABLES

4-1 Past Hypersonic Demonstrator Projects and Programs 8

5-1 Overall Trade-Space Concepts, Categories and Options 15

5-2 Summary of Design Trades Executed 16

6-1 Summary of Disciplinary Methods 18

6-2 Air-Launch Mach 6 Hydrogen-Fuel Mission Summary 25

6-3 Air-Launch, Mach 8, Hydrogen-Fuel Mission Summary 28

6-4 Air-Launch, Mach 6, Kerosene-Fuel Mission Summary 31

6-5 Air-Launch, Mach 8, Kerosene-Fuel Mission Summary 34

6-6 Air-Launch, Mach 6, Dual-Fuel Mission Summary 37

6-7 Air-Launch, Mach 8, Dual-Fuel Mission Summary 40

6-8 Expendable-Booster, Mach 6, Hydrogen-Fuel Mission Summary 43

6-9 Expendable-Booster, Mach 8, Hydrogen-Fuel Mission Summary 46

6-10 Expendable-Booster, Mach 6, Kerosene-Fuel Mission Summary 49

6-11 Expendable-Booster, Mach 8, Kerosene-Fuel Mission Summary 52

7-1 Design Characteristics for Hydrogen-Based Suggested Vehicle Selection 57

7-2 Design Characteristics for Kerosene-Based Suggested Vehicle Selection 57

A-1 Hypersonic Cruiser Planform Description Method 65

A-2 Hypersonic Air-Breather Volume and Wetted Area Estimation 66

A-3 Hypersonic Cruiser Trajectory Determination Method 68

A-4 Convergence Empty Weight Estimation Method 71

A-5 Structural Index Estimation Method 73

B-1.1 Air-Launch Mach 6 Hydrogen-Fuel Mission Summary 75

B-1.2 Summary of Rocket Accelerators Explored 75

B-1.3 Summary of Scramjet Stream Thrust Constants 76

B-1.4 Summary of Weight and Volume Constants 77

B-1.5 Trade-001 Air-Launch, Mach 6, Hydrogen-Fuel Output Database 77

B-2.1 Air-Launch, Mach 8, Hydrogen-Fuel Mission Summary 82

B-2.2 Summary of Rocket Accelerators Explored 82

B-2.3 Summary of Scramjet Stream Thrust Constants 83

B-2.4 Summary of Weight and Volume Constants 84

B-2.5 Trade-002 Air-Launch, Mach 8, Hydrogen-Fuel Output Database 84

B-3.1 Air-Launch, Mach 6, Kerosene-Fuel Mission Summary 89

B-3.2 Summary of Rocket Accelerators Explored 89

B-3.3 Summary of Scramjet Stream Thrust Constants 90

B-3.4 Summary of Weight and Volume Constants 91

B-3.5 Trade-003 Air-Launch, Mach 6, Kerosene-Fuel Output Database 91

B-4.1 Air-Launch, Mach 6, Kerosene-Fuel Mission Summary 96

B-4.2 Summary of Rocket Accelerators Explored 96

B-4.3 Summary of Scramjet Stream Thrust Constants 97

B-4.4 Summary of Weight and Volume Constants 98

B-4.5 Trade-004 Air-Launch, Mach 8, Kerosene-Fuel Output Database 98

B-5.1 Air-Launch, Mach 6, Dual-Fuel Mission Summary 103

B-5.2 Summary of Rocket Accelerators Explored 103
vii

B-5.3 Summary of Scramjet Stream Thrust Constants 104

B-5.4 Summary of Weight and Volume Constants 105

B-5.5 Trade-005 AirLaunch, Dual-Fuel, Mach 6 Output Database 105

B-6.1 Air-Launch, Mach 8, Dual-Fuel Mission Summary 110

B-6.2 Summary of Rocket Accelerators Explored 110

B-6.3 Summary of Scramjet Stream Thrust Constants 111

B-6.4 Summary of Weight and Volume Constants 112

B-6.5 Trade-006 Air-Launch, Mach 8, Dual-Fuel, Output Database 112

B-7.1 Expendable-Booster, Mach 6, Hydrogen-Fuel Mission Summary 117

B-7.2 Summary of Scramjet Stream Thrust Constants 118

B-7.3 Summary of Weight and Volume Constants 118

B-7.4 Trade-007 Expendable-Booster, Mach 6, Hydrogen-Fuel, Output Database 119

B-8.1 Expendable-Booster, Mach 8, Hydrogen-Fuel Mission Summary 123

B-8.2 Summary of Scramjet Stream Thrust Constants 124

B-8.3 Summary of Weight and Volume Constants 124

B-8.4 Trade-008 Expendable-Booster, Mach 8, Hydrogen-Fuel Output Database 125

B-9.1 Expendable-Booster, Mach 6, Kerosene-Fuel Mission Summary 129

B-9.2 Summary of Scramjet Stream Thrust Constants 130

B-9.3 Summary of Weight and Volume Constants 130

B-9.4 Trade-009 Expendable-Booster, Mach 6, Kerosene-Fuel Output Database 131

B-10.1 Expendable-Booster, Mach 8, Kerosene-Fuel Mission Summary 135

B-10.2 Summary of Scramjet Stream Thrust Constants 136

B-10.3 Summary of Weight and Volume Constants 136

B-10.4 Trade-010 Expendable-Booster, Mach 8, Kerosene-Fuel Output Database 137
viii

NOTATION

SYMBOLS

b span

c spatular width

c/s spatula ratio, spatula width to outboard semispan

CD0 zero-lift drag coefficient

Cev expansion component velocity coefficient

Cea expansion angularity coefficient

(Cf/2 Aw/A3)c dimensionless boundary-layer skin-friction quantity

(Cf Aw/A3)b burner effective drag coefficient

CL lift curve slope

CLmax maximum lift coefficient

Cpe expansion specific heat

f fuel-to-air ratio

h vehicle height

hc/lc ratio of external compression height to length

hiso/liso ratio of isolator height to length

hpr fuel heating value

Isp specific impulse

Istr structural index, ratio of structural weight to wetted area

Kstr structural weight shape factor

l vehicle length

L’ induced drag coefficient

L/D Lift-to-Drag ratio

Lcomb length of combustor

lc/lw ratio of external compression length to total vehicle length

Nrkt number of rocket motors

Sfront frontal area

Spln planform area

Swet wetted area

tcruise cruise endurance time

Trkt total thrust from rocket motor

Vfx/V3 ratio of fuel velocity to axial flow velocity

Vf/V3 ratio of fuel velocity to total flow velocity

Vppl propellant volume

Vprop propulsion system volume

Vsys systems volume

Vtotal total volume

Vvoid void volume

W/S wing loading

Wmargin design weight margin

Wppl propellant weight

Wprop total propulsion system weight

Wstr structural weight

Wsys systems weight
ix

GREEK SYMBOLS

c compression system ratio of specific heats

e expansion system ratio of specific heats

1 adiabatic compression efficiency

b burner efficiency

1n first nozzle angle

2n second nozzle angle

 Küchemann’s slenderness parameter

ACRONYMS

AVD Aerospace Vehicle Design Laboratory

CE Configuration Evaluation

c.g. center of gravity

CL Configuration Layout

DB Data-Base

DBS Data-Base System

HL Horizontal Landing

HTO Horizontal Take Off

KB Knowledge-Base

KBS Knowledge-Base System

LaRC Langley Research Center

Li-AL Lithium-Aluminum alloy

M Managerial

MLW Maximum Landing Weight

MSTC Multiple Stages To Cruise

OEW Operating Empty Weight

OWEw Operating Weight Empty from weight budget

OWEv Operating Weight Empty from volume budget

PDE Pulse Detonation Engine

PS Parametric Sizing

RBCC Rocket-Based Combined Cycle

RJ Ramjet

RKT Rocket

RSM Response Surface Method

S Synthesis

SiC/SiCMMC Silicon Carbide/Silicon Carbide Metal Matrix Composite

SEP Société Européenne de Propulsion

SERN Single Expansion Ramp Nozzle

SJ Scramjet

SSTC Single Stage To Cruise

T Technologies

TBCC Turbine-Based Combined Cycle

TD Thrust minus Drag

TJ Turbojet

TOGW Take Off Gross Weight

TPS Thermal Protection System
x

TSTC Two Stage To Cruise

T/W Thrust-to-Weight ratio

VAB Vehicle Analysis Branch

VTO Vertical Take Off
1

1 INTRODUCTION

The Solution Space Screening for a Hypersonic Endurance Demonstrator task has been a two and onehalf month study with the aim:

 to demonstrate the Aerospace Vehicle Design (AVD) Laboratory sizing process applied to a fast 

turnaround project by using a dedicated knowledge-harvesting approach coupled with a unique sizing 

methodology to represent the first step in the conceptual design phase;

 to identify and visualize the solution space available for a hypersonic endurance demonstrator (20 to 

30 minutes) that employs an air-breathing propulsion system;

 to propose prospective baseline vehicle(s) based on (1) available industry capability and (2) highpriority research (technology) required.

 to demonstrate a best-practice product development and technology forecasting environment that 

integrates the key team members, including (1) manager (decision maker), (2) synthesis specialist

(integrator), and (3) technologist (disciplinary researcher).

In an effort to increase the air-breathing endurance capability of current hypersonic research aircraft (i.e., 

X-43, 7 seconds; X-51, 5 minutes), the NASA Langley Research Center (LaRC) Vehicle Analysis Branch

(VAB) has tasked the AVD Laboratory at the University of Texas Arlington (UTA) with exploring the 

technical and operational solution space for a 20 minutes to 30 minutes cruise endurance demonstrator 

operating at Mach 6 to Mach 8. The primary challenge has been to explore that portion of the available 

industry capability that will require future technology complementation, with the aim of arriving at a 

technically feasible demonstrator within a given time frame and budget. Consequently, this study 

necessitated the use of a simulation capability to assess and visualize the physical design drivers and 

sensitivities of the operational and technical domain.

The overall goal of the project has been the development of a concept for an air-breathing hypersonic 

endurance flight vehicle to increase our existing understanding and knowledge-base regarding airbreathing propulsion, associated thermal proctection systems (TPS), and any operational peculiarities of 

long-duration hypersonic flight (e.g., maintenance, turnaround, practical range, etc.).

This report introduces the AVD Laboratory’s product development and technology forecasting 

methodology as applied to the problem introduced above. Because the focus of this activity has been on 

the exploration of the available solution space, a unique screening process has been employed to assess 

the implication of (a) the mission, (b) the baseline vehicle, and (c) the operational scenarios on key 

research objectives to be defined.

This study concludes that an air-launched, liquid-hydrogen-fueled, 30 minutes Mach 6 demonstrator (with 

10 minutes Mach 8 capability) provides the largest feasible solution space of the trades that have been

examined (i.e., largest design margins with lowest technical risk) when compared with a kerosene-fueled

equivalent.
2

2 OVERALL STUDY METHODOLOGY

The study has been organized into three distinct phases with the following individual work elements or 

tasks defined for each phase, see Figure 2-1:

Phase I Preparatory Activities

[DB/KB query, initial trade matrix]

3 weeks

Phase II Configuration and Technology Identification

[parametric solution space exploration]

6 weeks

Phase III Recommendations

[baseline flight vehicle(s) identification]

2 weeks

Fig. 2-1. Study approach to develop reusable hypersonic endurance test bed.

Task 1: Research Strategy Definition

The objective is to formulate, discuss, harmonize, and adopt research ground rules for the 11-week study. 

Bernd Chudoba and Gary Coleman traveled from 14 to 18 June 2010 to NASA LaRC to jointly define the 

research strategy with the VAB team members.

Task 2: Literature Review, DB/KB Query

A primary literature search is conducted to identify relevant past and present data and knowledge that are

related to the planning of a hypersonic endurance demonstrator.

Task 3: Reference Vehicle Definition

The X-15 is selected as the appropriate reference aircraft or analog for the endurance demonstrator.
3

Task 4: Definition of Initial Trade Space Scenarios

The initially wide-open trade space for this study is refined successively and constrained to trades of

immediate relevance for VAB.

Task 5: Parametric Sizing

The AVD Laboratory’s parametric sizing (PS) methodology is executed. The conceptual design consists

of three individual phases executed in sequence: (1) Parametric Sizing (PS), (2) Configuration Layout

(CL), and (3) Configuration Evaluation (CE) (see Figure 2-2).

Task 6: Modification of Trade Space Scenarios

The deliverables generated during the PS phase allow the AVD Laboratory and VAB researchers to 

review and modify the initial trade matrix. The PS phase is the key phase during which the researchers

gain an initial physical understanding of the design problem and sensitivities at hand.

Task 7: Selection of Baseline Aircraft/System Architecture

After sufficiently exploring and visualizing the available solution space for the endurance demonstrator, 

the design team is in the position to select baseline flight vehicle parameters such as launch type, fuel 

type, size, and operational requirements.

Fig. 2-2. AVD Laboratory integrated design life-cycle process.
4

Task 8: Research Aircraft Proposal Writing and Plan for Future Work

Once the technical solution-space topography for a hypersonic demonstrator has been established, a 

benefit and uncertainty requirement-topography must be imposed. Next, the flight demonstrator research 

objectives will be related to industry capability available and future technology required.

A central requirement for the AVD Laboratory team has been to work in partnership with the NASA 

VAB team via visits, weekly progress telecons, email communication, and presentation and report 

deliverables. Accepting the novelty of the design task, the aim of the AVD team is to generate 

deliverables emphasizing transparency, reproducibility, and physical correctness. This study approach, 

which is shown in Figure 2-3, details the three mindsets at work throughout the project life cycle: (a) 

Managerial (M), (b) Synthesis (S), and (c) Technology (T). This integration scheme maximizes the 

interaction between the VAB and the AVD Laboratory along the three principal mindsets at work.

Figure 2-3 further addresses the implications of the overall project time constraints. The actual trade 

matrix executed is limited to the study of the most important operational requirements and flight vehicle 

design parameters. Before a baseline vehicle selection can be reliably made, a comprehensive set of 

constraints & requirements representing the M, S, and T mindsets are explored via trade studies.

Please note the step called Objectives Matching shown in Figure 2-3. The following chapter will address 

its meaning.

Fig. 2-3. AVD Laboratory integrated VAB-AVD team study approach.
5

3 MISSION REQUIREMENTS AND RESEARCH OBJECTIVE

The overall objective of this study is to explore and visualize the technical solution space for a hypersonic 

endurance demonstrator.

The NASA VAB operational and technology requirements for this demonstrator are:

 scramjet test vehicle

 reusable

 unmanned

 multiple aircraft (at least three test articles)

 entry into service circa 2020



To evaluate the technical feasibility of such a research vehicle, the following mission requirements are 

selected by NASA VAB:

 design speed: Mach 6 to 8 (possibly Mach 12)

 maximum endurance: 20 to 30 minutes

 payload: test instrumentation

 fuel selection: hydrogen or kerosene

 operation: straight line or point-to-point

The broad direction specified by VAB in June 2010 translates into a large n-dimensional design trade 

space. Please note that the VAB-defined design mission is considered a starting point only, thus the 

mission itself is a variable. Since the targeted flight regime is novel terrain for the designer, it is essential 

to trade flight vehicles capable of satisfying alternative missions. Clearly, the sizing exposure will 

iteratively enable the designer to define and justify a feasible baseline mission and baseline vehicle

combination, see Figure 3-1.

Fig. 3-1. Iterative nature of the mission & objectives & baseline vehicle(s) selection process.
6

Figure 3-1 illustrates the iterative nature of the mission selection process. The unknown-terrain nature of a 

20 to 30 minutes air-breathing demonstrator requires a modification of the traditionally utilized product 

development procedures. As shown in this figure, the AVD Laboratory screening & sizing methodology 

is the primary tool utilized to arrive at a (a) baseline mission which harmonizes with (b) the overall 

research objectives and (c) the baseline vehicle.

The sizing team is tasked to execute alternative missions resulting in prospective baseline vehicle(s). 

Throughout the sizing phase, the involved mindsets (managerial (M), synthesis (S), technology (T)) are 

successively gaining physical insight into the characteristic of the product. Consequently, true product 

understanding is evolving while the solution space alternatives are perturbed. The mission-trading needs 

to happen during the parametric sizing (PS) phase, an essential task before a baseline objectives catalogue 

can be formally defined. Clearly, the traditional notion of pre-defining the mission and objectives is not 

feasible with a product of such novel characteristics. The screening & sizing approach becomes the 

enabling means to arrive at a balanced set of (a) mission, (b) objectives, and (c) baseline vehicle(s).

Between July 1969 and June 1970, the McDonnell Aircraft Company had been tasked by NASA (NASA 

Contract NAS2-5458) to conduct a comprehensive Hypersonic Research Facilities Study (HYFAC), see 

Reference 1. The objective of this study has been to assess research and development requirements for (a) 

flight facilities (demonstrator) and (b) ground facilities (e.g., wind tunnels) towards air-breathing 

operational hypersonic aircraft that satisfy multiple future operational missions. Overall, the study 

provides the required characteristics for flight-test research facilities and ground-test research facilities. In 

analogy to the present study, the McDonnell Aircraft Company make use of a dedicated sizing 

methodology as the primary means for the numerical design solution space identification. Noteworthy is 

the time and people effort invested by multiple participants to identify and evaluate the research 

objectives for future hypersonic missions and associated hardware, see Figure 3-2.

Fig. 3-2. HyFAC project research objectives identification and evaluation participants.
7

The McDonnell Aircraft Company HyFAC study highlights the significance of the objectives matching 

process, a process of significance for identifying and balancing the triangle relation between the (a) 

baseline mission, (b) research objectives, and (c) baseline vehicle(s). Due to project time constraints, the 

present research undertaking excludes the research objectives development and matching step. Figure 3-3 

illustrates the finally implemented baseline vehicle development sequence for the present study by 

omitting the objectives matching step shown in Figure 3-1. It is recommended to formally complement 

the existing study at a later step by including the objectives matching logic as an essential ingredient

supporting decision-making.

Fig. 3-3. VAB/AVD Laboratory baseline vehicle development sequence.
8

4 DATA-BASE AND KNOWLEDGE-BASE REVIEW OF HYPERSONIC DEMONSTRATORS

A key component enabling the development of hypersonic flight test vehicles is effective management of 

the knowledge-generation and knowledge-preservation activity. As illustrated before, the research 

approach implemented places emphasis on elevating the understanding with regards to project aims and 

objectives, overall resulting in an informed and structured approach. In the present context, the research 

challenge is best formulated with the question: How to efficiently synchronize the understanding 

available with the understanding required to specify a feasible air-breathing hypersonic demonstrator 

with the technical resources, team support and time available? Due to the limited timeframe available, the 

DB and KB assistances have become indispensable to expedite the learning process.

The scope and complexity of the present research undertaking is seen as catalyst opportunity, which 

translates into a chance to evaluate past and present data and knowledge for its utilization in the context of 

a technically demanding demonstrator with not seen-before performance capability. Table 4-1 lists highspeed flight vehicles of direct relevance in the context of a future endurance testbed.

Table 4-1. Past Hypersonic Demonstrator Projects and Programs

Start Date End Date Project/Program Organization Description

1952 1968 X-15 North American/NASA/USAF Mach 6 to 8 rocket powered hypersonic research vehicle. 3 test 

vehicles, 199 flights

1957 1959 Griffon 02 Nord Aviation Manned ramjet demonstrator

1962 1971 D-21 Lockheed Mach 4 ram-jet UAV launched from the SR-71

1964 1965 MHCV Lockheed Manned Hypersonic Cruise Vehicle, some description of a 

demonstrator

1967 1968 UHTV Vought Universal Hypersonic Test Vehicle, flexible and modular hypersonic 

test vehicle 

1967 1969 X-15 Delta North American/NASA Delta wing X-15

1969 1970 HYFAC MAC/NASA HYpersonic FACilities study, 32 rocket/air-breather configurations 

explored

1969 1969 X-15 SERJ Marquardt Super Charged Ejector Ramjet (RJ) X-15

1969 1969 X-15 Scram Boeing Scramjet (SJ) X-15

1970 1972 IGV MAC/USAF Incremental growth vehicle

1972 1972 PPD Scramjet Test 

Vehicle

Propulsion Performance Demonstrator, vertical takeoff cone with 

four scramjets around its periphery; rocket acceleration to test speed

1975 1977 X-24C NHFRF Lockheed/NASA National Hypersonic Flight Research Facility, B-52 launched, Mach4.8 70,000 lbs vehicle; envisioned as a X-15 type flight operation 

1976 1980 ASALM Martin Hydrocarbon fuel air-launched cruise missile

1980 1981 SLRV Shuttle Launch Research Vehicle, Mach 8 aerodynamic configuration 

demonstrator

1985 1985 RSFTP Ramjet/Scramjet Flight Test Program, M 4-7 F-15 launched vehicle 

1989 1990 HYPAC MBB Sänger demonstrator study

1990 1995 BMFT MBB/UK/UT/Dornier/MTU Hypersonic technology program, HYTEX and RADUGA D2

1996 2004 X-43A NASA LaRC/NASA Dryden Scaled hypersonic scramjet demonstrator

1999 1999 SSTO 

Demonstrator

Hyper Tec RBCC hypersonic demonstrators based on HYFAC Studies

1999 1999 Trailblazer NASA Glenn Modification of the NASA wing body to include RBCC and TBCC

2000 2002 X-43B NASA LaRC/NASA Dryden Reusable combined cycle demonstrator

2001 2002 X-43C NASA LaRC/NASA Dryden Hydrocarbon variant of the X-43A, RJ/SJ

2002 Present HYFLY Boeing/DARPA Mach 6 ramjet powered cruise missile demonstrator

2003 Present X-51A Boeing Scramjet propulsion research vehicle

2005 2007 X-43D NASA LaRC/NASA Dryden HYFLITE III, M 12 variant of the X-43A

2007 2007 HyCAUSE DARPA/ADST 2-stage sounding rocket for hypersonic propulsion demonstration

2007 2008 Falcon HTV-3X Lockheed/DARPA TBCC hydrocarbon hypersonic demonstrator
9

The following two sub-chapters present the flight vehicle conceptual design data-base (DB) and 

knowledge-base (KB) as developed and utilized for the present research undertaking. The main flight 

vehicle research & design work is directly benefitting from this dedicated DB & KB foundation.

4.1 HYPERSONIC FLIGHT VEHICLE DATA-BASE (DB)

The first step in efficiently utilizing existing high-speed aircraft design knowledge has been a systematic

literature survey, which in itself has been an ongoing effort throughout the existence of the AVD 

Laboratory and of course during the current research period. Source for accessing normal and radical 

design data and knowledge have been (a) public domain literature, (b) institution and company internal 

sources, and (c) expert advice. For efficient handling of design related data and information, a dedicated 

computer-based aircraft conceptual design data-base (DB) has been set up, see Figure 4-1. Reference 2 

presents the literature DB file-structure. This system handles disciplinary and inter-disciplinary literature 

relevant for conceptual design (methodologies, flight mechanics, aerodynamics, etc.), interviewprotocols, flight vehicle case study information (descriptive-, historical-, numerical information on 

conventional and unconventional flight vehicle configurations), simulation and flight test information, etc. 

The overall requirement for the creation of the DB has been simplicity in construction, maintenance, and 

operation, to comply with the underlying time constraints.

A detailed description of the DB is beyond the scope of the present discussion. The system has become a 

steadily growing, comprehensive, and effective working tool. Clearly, the quality of such system is only 

as good as the degree of completeness, actuality, and familiarity by the user. The DB has matured to be 

the central instrument for managing aircraft design data and information. However, the true potential of 

this system for utilizing design data and information has been opened up by proceeding as follows:

Fig. 4-1. Dedicated AVD Laboratory DB and organization scheme.
10

1. availability of a reference list containing meaningful entries; (DB)

2. availability of these references as a hardcopy on the table; (DB)

3. utilization of time to absorb the data & information; (DB)

4. review, select, classify, subtract, and document the data & information provided; (DB)

5. extraction, combination and utilization of data & information in a pre-defined manner. (KB)

The first four steps are handled within the DB. The DB has been put to use to provide in an intermediate 

step (step four) suitably selected, structured, and condensed flight vehicle conceptual design data and 

information. The research goal, to develop an air-breathing hypersonic demonstrator requires to account 

for as many design-related interactions as necessary, since the rationale for the evolution of aircraft is 

diverse as a quick browse through aviation history reveals. The aircraft design disciplines identified 

relevant and the representative case studies of design ingenuity selected, see Table 4-1, both elements 

need to be appreciated mutually, to efficiently serve the design understanding where innovation provided 

answers to otherwise troublesome problems. The updated DB embodies a technology-baseline attained, 

which is considered state-of-the-art for the current research undertaking.

Figure 4-2 summarizes four particular entries in the DB: (a) digital DB of past hypersonic demonstrator 

projects, (b) digital DB for rocket engines, (c) digital DB for carrier aircraft, and (c) digital DB for past 

hypersonic vehicle design solutions (visual/geometry evidence).

Fig. 4-2. Selected AVD Laboratory data-base (DB) entries.
11

Summarizing, all four DB development steps have been followed and completed to a satisfactory degree 

within the time span allocated. The knowledge-base (KB) step five has been organized outside the DB. 

Clearly, it is the process of knowledge extraction, knowledge compilation, and knowledge provision into 

an organized and concise format which finally makes relevant high-speed aircraft conceptual design 

knowledge available ‘at the fingertips’ for problem solving activities. For this purpose, a simplified 

knowledge-base (KB) has been constructed as detailed in the following sub-chapter.

4.2 HYPERSONIC FLIGHT VEHICLE KNOWLEDGE-BASE (KB)

The aircraft conceptual design knowledge-base (KB), as advanced and utilized for the present research 

undertaking, has to be considered an early development-version of a fully operational design knowledgebased system (KBS). Without reiterating the capability of exemplary KBSs, the KB system utilized here 

is a ‘manual’ system in contrast to the ideally automated KBS. However, independent on the degree of 

automation, both systems have in common that knowledge itself is the focus and that the knowledge 

acquisition activity is recognized as being one of the most problematic areas of KBS development. 

Clearly, it is the knowledge collecting, knowledge management and knowledge utilization activity, where 

the priorities for the present flight vehicle conceptual design KB have been laid due to time constraints 

imposed.

The primary objective of developing the dedicated hypersonic aircraft conceptual design KB has been, to 

make relevant normal and radical design knowledge effortlessly available. The particular strength of the 

system manifests, in that it enables the user to advance his/her understanding with respect to the variety of 

legacy high-speed aircraft and launch vehicle configurations by identifying aircraft configuration 

commonalties and peculiarities. This feature has been empowered by placing particular emphasis on 

consistently grouped flight vehicle configuration-specific design knowledge. As a result, design detail, for 

example longitudinal stability, can be compared between the range of aircraft configurations. This 

approach finally enables a reliable and trust-worthy generic aircraft configuration parameter identification 

process.

The hypersonic flight vehicle conceptual design KB for fixed-wing and lifting-body designs is subdivided 

into two main sections:

(a) Longitudinal Motion

(b) Lateral/Directional Motion

Each motion is subdivided into:

- Flight Character (Design Constraining Flight Conditions: trim, control, stability)

- Aerodynamic Character (Stability and Control Derivatives: u, u/t, w(), w/t (/t), ...)

- Flow Character (Flow Phenomena: tuck, pitch-up, non-linearity, ...)

- Additional Grounds (Landing gear location, geometry limitations, c.g. range, ...)

Figure 4-3 overviews the lessons-learned section as described above. This section clearly emphasizes on 

physical understanding and design related decision-making of relevant aircraft case studies.

Figure 4-4 introduces the steps required to arrive at knowledge-derived numerical design guidelines. At 

first, intimate technical understanding of pertinent design case studies enables the identification of gross 

design-drivers and variables with significant impact on the overall design. Those gross design drivers then

form the basis for the underlying sizing relations in the sizing methodology. The resulting numerical

design guidelines represent a true continuum of the pertinent design characteristic in contrast to the 

narrow exposure of typical point-design characteristics.
12

Fig. 4-3. Design lessons-learned of selected design case-studies.

Fig. 4-4. KB development steps resulting in numerical design guidelines.
13

The ‘living-character’ of the DB and KB is ensured by permitting unconstrained data & knowledge 

entries as gained during the iterative design life-cycle, see Figure 3-3.

In summary, the dedicated hypersonic vehicle DB and KB have both matured towards fully integrated 

design support domains. The AVD Laboratory is routinely utilizing the project-specific DB and KB in 

concert with the process domain (sizing methodology), see Figure 4-5.

Fig. 4-5. Integration scheme of data domain, knowledge domain, and process domain.
14

5 OVERALL TRADE-SPACE AND REDUCED TRADE-SPACE

The challenge of designing a 20 to 30 minutes hypersonic endurance demonstrator is embodied in the 

fundamentally unknown vehicle solution space and solution topography. Based on the best understanding 

available at the outset, it is required to define an initial or ‘start’ trade-space by taking relevant constraints

and requirements into account.

It is to be expected that this initial trade-space, with associated constraints & requirements, will naturally 

mature during the configuration exploration phase. The configuration exploration phase is tasked to 

identify two primary solution-space areas of significance: (a) the solution space area based on presently 

available industry capability, and (b) the solution space area requiring prospective future technologies. 

Dependent on the establishment of overall project objectives (technology development, low-cost & risk 

demonstrator, etc.), the physical understanding generated will help to refine the initial trade-space scope.

Clearly, the early identification of the correct trade-space and technology combinations requires using 

logic, organization and transparency before any baseline design can be selected. This approach will 

provide the greatest insight into the design problem within the time assigned.

The process of rectifying thus reducing the theoretical trade-space available consists of: (a) Formulate a 

classification scheme for the design options available. (b) Focus the DB/KB development and team 

learning on relevant design trade-studies. (c) Harmonize pre-selected trades with VAB’s team’s long-term 

research objectives.

Table 5-1 presents the overall trade-space adopted classification scheme addressing (1) mission concept, 

(2) staging configuration, (3) operations concept, and (4) hardware concept. If all of the options shown in

this general trade-space Table 5-1 would be executed, the total number of trades would exceed 90,000+

cases.

Applying the DB/KB lessons-learned and harmonization with VAB’s research objects further allows

reducing and focusing the trade-space:

1. Mission Concepts: Mach 6 and Mach 8 design trades are given priority; point to point and flyback options are explored. Mach 12 has been eliminated.

2. Staging Configurations and Operational Concepts: HyFAC (Reference 3) determined that airlaunch and vertical take-off provide the largest research value for a hypersonic demonstrator 

relative to horizontal takeoff and single-stage vehicles. Air-launch and vertical takeoff with a 

booster allow for smaller and lighter demonstrators which can focus on testing the high-speed 

regime. Consequently, the trades selected will focuse on air-launch and vertical takeoff options.

3. Hardware Concepts: Alternative vehicle concepts have been grouped as follows:

a. Lifting body - for this speed range, the lifting body provides improved volumetric efficiency 

over wing bodies; therefore, the lifting body has been selected as the sole volume supply 

option (Reference 3, 4).

b. Off-the-shelf accelerator rocket – the off-the-shelf rocket motor (low risk item) is selected to 

accelerate the ramjet to start Mach number.
15

c. Dual-mode ramjet cruise engine - the dual mode ramjet/scramjet is selected to allow for 

testing of both modes with a single vehicle.

d. Fuel selection limited to liquid hydrogen and kerosene - the fuel selection is determined by 

the operational vehicle envisioned; for possible reusable TSTC launch vehicles, hydrogen 

appears to be the most likely choice. Kerosene appears to be an operationally practical option 

for a military hypersonic point-to-point vehicle. Consequently, both options (hydrogen and 

kerosene) are explored.

Table 5-1. Overall Trade-Space Concepts, Categories and Options

CONCEPT/CONFIGURATION CATEGORIES TOTAL TRADE OPTIONS SELECTED TRADES

Mission Concept Mach number and duration design Mach 6 design Mach 6

design Mach 8 design Mach 8

design Mach 12

test duration 0 to 30 minutes

test range options point-to-point point-to-point

fly-back fly-back

Staging Configuration SSTC integrated booster, propellant and oxidizer 

tanks 

TSTC air launch air launch

expendable booster expendable booster

oxidizer drop tanks

MSTC any combination of TSTC options

Operations Concept launch HTO

VTO

recovery HL

Hardware Concept lift & volume supply lifting body lifting body

wing body

propulsion concept: 

(accelerator engine)

RKT RKT

TJ

RBCC

PDE

propulsion concept:

(cruise engine) 

SJ

dual mode RJ/SJ dual mode RJ/SJ

RKT

fuel selection hydrogen hydrogen

methane

kerosene kerosene

primary & secondary 

controls

aerodynamic

mix mix

The above reasoning is reducing the overall trade-space to 10 trade studies, consisting of a constant test 

vehicle concept (lifting body, dual mode ramjet/scramjet, horizontal landing) with varying (a) design 

Mach number, (b) endurance, and (c) launch concept. The reduced trade-space is introduced with Table 5-

2 and Figure 5-1.
16

Table 5-2. Summary of Design Trades Executed

MISSION STAGING 

CONFIGURATION

OPERATIONS

CONCEPT

HARDWARE

CONCEPT

Atmospheric Test Range 

Options TSTC Launch Fuel Selection

Trade 

#

design 

Mach 6 

design 

Mach 8 

test

duration

point-topoint air 

launch

expendable 

booster

HTO VTO hydrogen kerosene

dual 

fuel 

1 x 0 - 30 min x x x x

2 x 0 - 30 min x x x x

3 x 0 - 30 min x x x x

4 x 0 - 30 min x x x x

5 x 0 - 30 min x x x x

6 x 0 - 30 min x x x x

7 x 0 - 30 min x x x x

8 x 0 - 30 min x x x x

9 x 0 - 30 min x x x x

10 x 0 - 30 min x x x x

Fig. 5-1. Reduced trade-space explored. Air Launch Expendable booster

Mach 6

Mach 8

Mach 6

Mach 8

Hydrogen Dual Fuel Kerosene
17

6 PARAMETRIC SIZING AND SOLUTIONS SPACE SCREENING

For each individual trade study, the total system design solution space is identified and visualized with the 

AVD Laboratory parametric sizing program AVDsizing. This ‘best practice’ sizing approach has been 

developed through a thorough review of parametric sizing processes and methods from the 1960s to 

present for subsonic to hypersonic vehicles, see Reference 5. With this framework in place, the available 

solution space is identified considering both technical and operational constraints.

6.1 AVD SIZING PROCESS SUMMARY

AVDsizing is a constant mission sizing process capable of first-order solution space screening of a wide 

variety of conventional and unconventional vehicle configurations. Solution space screening implies an 

overall focus on visualizing multi-disciplinary design interactions and trends. AVDsizing is based on the 

Hypersonic Convergence sizing approach for transonic to hypersonic vehicle applications as developed at 

formerly McDonnell Aircraft Company between 1970 and 1990, see Reference 6. The modular process 

implemented with AVDsizing relies upon a robust disciplinary methods library for analysis and a unique 

multi-disciplinary analysis (MDA) sizing logic and software kernel enabling data storage, design 

iterations, and process convergence. The integration of the disciplinary methods library and the generic 

multi-disciplinary sizing logic enables the consistent evaluation and comparison of radically different 

flight vehicles, see References 7, 8. The flight vehicle configuration independent implementation of 

AVDsizing allows for rapid parametric exploration of the complete flight vehicle system via a convergence 

check to mission. Figure 6-1 visualizes the top level sizing process implemented.

At the heart of the process is the weight and balance budget. The results from the geometry, performance 

constraint and trajectory modules (weight ratio, required T/W ratio, and vehicle geometry) are provided to 

a weight & volume available and required logic. For a given vehicle slenderness parameter ( 

Fig. 6-1. AVDsizing methodology visualized via Nassi-Schneidermann structogram.

AVDsizing

Weight budget: compute OWEw Volume budget: compute OWEv Iterate Spln until OWEw and OWEv converge

Iterate for each  specified Iterate over any independent design variable

Geometry Constraint Analysis: T/W=f(W/S) Trajectory: ff=f(trajectory,aero,propulsion)

Fundamental Sizing Steps OEW esitmation Sizing Logic Trajectory Anlaysis Convergence Logic Constraint Anlaysis

Constraints

Take-off

Approach speed

(W/S)TO

Feasible solution space Cruise

Aborted Landing OEI 2nd Segment Climb OEI Current Design Point Current (W/S)TO

Required (T/W)TO

Trajectory
18

 

 ⁄ ), the planform area is iterated through the total design process until weight & volume available 

equal weight & volume required.

6.2 DISCIPLINARY METHODS LIBRARY OVERVIEW

The following methods are utilized from the disciplinary methods library for this hypersonic demonstrator 

study, see Reference 5. The methods selected are of consistent first-order nature, including empirical, 

semi-empirical and reduced-order analytical types. Table 6-1 summarizes the disciplinary methods used

for this study. Selected methods are further documented in Appendix A.

Table 6-1. Summary of Disciplinary Methods

DISCIPLINE METHOD TITLE DESCRIPTION REFERENCE

Geometry Planform

Vehicle length, span and spatular width for current 

planform area based on constant leading edge sweep and 

c/s.

Czysz [6]

Bottom Surface Total volume and dimensions determined from nondimensional engine constants.Appendix A

Top Surface

Total volume, dimensions and wetted area computed for 

a compound elliptical cross-section. Top surface height 

determined from specified slenderness parameter.

Appendix A

Aerodynamics Drag Polar

McDonnell-Douglas empirical correlations (circa 1970) 

based on vehicle slenderness, frontal area and wetted 

area with spatular corrections from Pike.

HyFAC [3]

Pike [10]

 

 

Lift-Curve Slope McDonnell-Douglas empirical correlations (circa 1970) 

of all-body hypersonic vehicles.

HyFAC [3]

 

Maximum Lift (low speed) FDL-7 wind tunnel data. FDL-7 report

Propulsion Scramjet - Modified 1-D 

Cycle Analysis

1-D stream thrust analysis with corrections inlet spillage 

drag. RSM from Bradford used for truncated SERN 

nozzle performance. 

Heiser and Pratt

[12], Bradford

[13]

Ramjet – Marquardt Data Representative data from Marquardt study (circa 1960). Marquardt [14]

Rocket – Pratt & Whitney 

Method

Analytic off-design performance estimation of rocket 

thrust and Isp based on ideal rocket equation.

Czysz [6]

Performance Landing Wing loading requirement for given stall speed and 

maximum trimmed lift coefficient.

Coleman [5]

Trajectory

2-D energy integration method (altitude and velocity), 

constant q trajectory to cruise velocity, cruise climb, 

maximum L/D descent.

Appendix A

Stability and Control Trim effects Engine cowl location effect on trim drag.

HyFAC [3] Czysz 

[6]

Weight and Volume Hypersonic Convergence 

Weight and Volume Budget

Empirical weight and volume estimation of structure, 

systems, payload and propellant.

Appendix A
19

6.3 DESCRIPTION OF SOLUTION SPACE VISUALIZATION

The overall product solution space consists of individually converged total flight vehicle design points. 

For a fixed vehicle slenderness parameter (), the complete weight breakdown and trajectory are 

computed for every individual vehicle planform iteration. The process is repeated until the weight and 

volume required meet the weight and volume available, see Figure 6-2.

A vehicle geometry solution space contour or topography is determined by varying the vehicle 

slenderness and re-converging each design point. The operational mission solution space is created by

varying cruise time and re-converging each solution contour. The result is a continuous carpet plot 

comparing individually converged flight vehicle solutions based on structural index, Istr, and TOGW, see 

Figure 6-3. The structural index, Istr, is a metric of the structural efficiency of the concept, and is defined 

as structural weight per unit wetted area. This parameter will be further discussed when addressing the 

description of the solution space constraints.

Solution Space Constraint Description: Having generated a carpet plot consisting of individually 

converged flight vehicles of varying vehicle slenderness () and cruise time, the next step is to 

superimpose the aborted landing constraint, the thrust minus drag (T-D) constraint and the structural 

technology level available (Istr). The landing constraint is computed from the prescribed approach speed, 

which translates to the required 1g stall speed and required stall wing loading. Additionally, mapping the 

required wing loading to the TOGW and Istr, the T-D constraint can be added to the solutions, see Figure 

6-4.

The T-D constraint represents the highest allowable which will still have positive acceleration during 

the ascent portion of the trajectory. If the vehicle is stouter (reduced planform area and increased vehicle 

height), then this limits the wave drag increase and the reduced capture area results in negative thrust, see 

Figure 6-4.

Fig. 6-2. Each design point represents a converged complete hypersonic vehicle (Example: Mach6, 30 minutes, cruiser configuration).

AVDsizing

Weight budget: compute OWEw Volume budget: compute OWEv Iterate Spln until OWEw and OWEv converge

Iterate for each  specified Iterate over any independent design variable

Geometry Constraint Analysis: T/W=f(W/S) Trajectory: ff=f(trajectory,aero,propulsion)

Fundamental Sizing Steps OEW esitmation Sizing L

Trajectory Anlaysis Convergence Logic Constraint Anlaysis

Constraints

Take-off

Feasible 

sp

Cruise

Aborted Landing OEI 2nd Segment Climb OEI 

Curren

Required (T/W)TO

Trajectory

Geometry Weight Volume

 0.125 TOGW (kg) 22136 Vtotal (m3) 131.3

Spln (m2) 103.31 Wppl (kg) 10047 Vppl (m3) 75.4

Swet (m2) 244.63 OEW (kg) 12090

b (m) 9.57 Wmargin 1099

c (m) 1.91 Wsys (kg) 3834 Vsys (m3) 10.3

l (m) 18.00 Wstr (kg) 4736 Vvoid (m3) 26.3

h (m) 2.84 Wprop (kg) 2420 Vprop (m3

) 19.3

Aero and propulsion results across trajectory Geometry, Weight and Volume Results constant 1.5

pln

 

S

Vtotal 
20

Fig. 6-3. Solution space carpet plot of TOGW and Istr for varying vehicle slenderness (τ) and cruise 

time.

Fig. 6-4. Landing and T-D constraints imposed on the solution space. For the Mach 6 

demonstrator, the landing constraint is more constraining than T-D.

0

5000

10000

15000

20000

25000

30000

35000

40000

0 1 0 2 0 30 40

TOGW

(kg)

Istr=Wstr/Swet

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical 

Challenge

0.075

0.125

Geometry Weight Volume

 0.125 TOGW (kg) 22136 Vtotal (m3) 131.3

Spln (m2) 103.31 Wppl (kg) 10047 Vppl (m3) 75.4

Swet (m2) 244.63 OEW (kg) 12090 b (m) 9.57 Wmargin 1099

c (m) 1.91 Wsys (kg) 3834 Vsys (m3) 10.3

l (m) 18.00 Wstr (kg) 4736 Vvoid (m3) 26.3 h (m) 2.84 Wprop (kg) 2420 Vprop (m3

) 19.3

Aero and propulsion results across trajectory Geometry, Weight and Volume Results

Weight budget: compute OWEw Volume budget: compute OWEv Iterate Spln until OWEw and OWEv converge

Iterate for each  specified Iterate over each cruise time Geometry Constraint Analysis: T/W=f(W/S) Trajectory: ff=f(trajectory,aero,propulsion)

0

5000

10000

15000

20000

25000

30000

35000

40000

0 1 0 2 0 3 0 4 0

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical 

Challenge

T-D

constraint

0.075

0.125
21

The final constraints of relevance for identifying the solution space include: (a) launch vehicle load 

capability, (b) geometry limits for the carrier (air-launch) aircraft, and (c) expendable booster staging 

options. Options for the air-launched carrier vehicle are the B747-100SCA and B-52H; both options have 

been explored as possibilities. The B-52H employs an under wing mount constrained by: (a) the 

maximum load of the pylon, and (b) the geometric boundaries between the fuselage and inboard engine, 

the test vehicle wing and engine exhaust plum. The X-24C was intended to be the largest vehicle to 

possibly fit under the B-52H wing mount. Therefore, the X-24C’s TOGW, length and width represent a 

guide for the maximum capability of the B-52H air-launcher for this investigation, see Figure 6-7. The 

B747-SCA is a modified B747-100 designed to carry the Space Shuttle Orbiter. For this study, the OEW, 

length and span of the Space Shuttle Orbiter are used as a guide for the maximum air-lift capability of the 

B747-SCA, see Figure 6-8.

Figure 6-4 represents the structural weight per wetted area required to converge the configuration to each 

specific slenderness value (). When superimposing relevant material and structural concept technology 

levels onto the vehicle structural index carpet plot, the left boundaries of the solution space are

determined. For vehicle slenderness parameters which require structural indices beyond this limit, the 

structural and shingle material are not feasible, see Figure 6-5. Figure 6-6 documents the structural 

indices utilized to derive the technology solution space boundaries pertinent to the flight mission.

Fig. 6-5. Superposition of structural indices provides the final constraint to determine the technical 

solution space.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

40000

0.00 10.00 20.00 30.00 40.00

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical 

Challenge

T-D

constraint

0.075

0.125
22

Fig. 6-6. Definition of structural capability indices used for this study. (Ref 6)

Fig. 6-7. B-52H under-wing mount geometric constraints. (Ref 15)

Integral Hydrogen Fuel tank

Refractory 

shingles

Advanced 

schingles 

Refractory 

shingles

Advanced 

schingles 

kg/m2 kg/m2 kg/m2 kg/m2

Shingle 6.30 4.88 6.30 4.88

Insulation 4.39 4.39 4.39 4.39

Structure 7.08 7.08 3.61 3.61

Foam insulation 1.61 1.61 1.61 1.61

total 19.38 17.97 15.92 14.50

Integral Kerosene fuel tank

Refractory 

shingles

Advanced 

schingles 

Refractory 

shingles

Advanced 

schingles 

kg/m2 kg/m2 kg/m2 kg/m2

Shingle 6.30 4.88 6.30 4.88

Insulation 4.39 4.39 4.39 4.39

Structure 7.08 7.08 3.61 3.61

Foam insulation 0.00 0.00 0.00 0.00

total 17.77 16.36 14.31 12.89

Li-AL structure Composite strucutre

Li-AL structure Composite strucutre

Integral fuel Tank Structural constant (MAC circa 1970) Break-down of industrial capability index
23

Fig. 6-8. Summary of B747-SCA and B-52H constraints for the hypersonic demonstrator study. (Ref 

15,16)

Fig. 6-9. Summary of Minotaur I and Taurus XL 1st stage constraints for the hypersonic demonstrator 

study. (Ref 17)

B-52H

Design Payload: X-24C

Payload = 31,751 kg

Spanlimit = 7.3 m

Lengthlimit = 17.3 m

B747-100SCA

Design Payload: Space Shuttle

Payload =78,000 kg

Spanlimit = 23.79 m

Lengthlimit = 37.24 m

Castor 120

Taurus XL – 1

st Stage

Payload = 19,046 kg

Velsep = 2,209 m/sec

Altsep = 41 km

M55A1

Minotaur I – 1

st Stage

Payload =13,116 kg

Velsep = 1,510 m/sec

Altsep = 33 km

Payload equals 

payload to orbit plus 

weight of upper stages

Velocity and altitude 

referenced to first 

stage separation point

Both boosters are 

currently in production
24

When considering expendable boosters as the launch method for the hypersonic demonstrator, the 

boosters are found to fit the hypersonic demonstrator options as the 2nd-stage of either the Minotaur I or 

Taurus XL launch vehicles. These representative boosters are selected based on their maximum payload 

weight, separation velocity and separation altitude, see Figure 6-9. The maximum payload weight 

capacity of the booster 1

st stage is taken to be the maximum payload to orbit, plus the weight of the upper 

stages.

During the screening process, each solution space is bounded by operational factors and technology 

factors for landing, T-D, and structural index. Next, the carrier/launch vehicle constraints are examined to 

determine the appropriate air-launch vehicle options for each trade.

6.4 SOLUTION SPACE SCREENING

The selection of the trade-space and the accompanying trade-matrix results in a solution space screening 

activity overall consisting of two (2) launch options, two (2) cruise Mach numbers, and three (3) fuel 

combinations. The solution space deliverables for each option are visualized relative to each other with 

Figure 6-10. For each trade, the cruise time will be increased from 0 min to 30 min in increments of 10 

min while vehicle slenderness is varied, generating the distinct solution space carpet plot. Since Figure 6-

10 compares discrete flight vehicle types (launch method, Mach number, fuel), note that the ten (10) 

identified and visualized trade solution spaces demonstrate regions of operational and technical feasibility 

with a varying TOGW y-axis scale. In total, 237 flight vehicle design solutions have been converged.

The remainder of this chapter is structured to address each specific trade-study in terms of: (1) tradesummary, (2) mission summary, (3) solutions space visualization, (4) carrier vehicle constraints, (5) 

additional sub-trades explored during the study, and (6) selected vehicle baseline design point.

Fig. 6-10. Relative comparison of solution spaces for each design trade explored.

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.0425

0.0625

0.0825

tcruise 

Increasing Technical Challenge

T-D

constraint

0

2000

4000

6000

8000

10000

12000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.05 0.0725 0.095 tcruise



Increasing Technical Challenge

T-D

constraint

0.1175

0.14 Air Launch Expendable booster

Mach 6

Mach 8

Mach 6

Mach 8

Hydrogen Dual Fuel Kerosene

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint 0 min

10 min

20 min

30 min

0.1 0.1275

tcruise 0.155

Increasing Technical Challenge

T-D

constraint

0.1825

0.21

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.12

0.17

0.195

tcruise

Increasing Technical Challenge

T-D

constraint

0.145

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint 0 min

10 min

20 min

30 min

0.02 0.04250.0538



Increasing Technical Challenge

T-D

constraint

0.0313

0.065

tcruise

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint 0 min

10 min

20 min

30 min

0.05

0.10

0.125



Increasing Technical Challenge

T-D

constraint

tcruise

0.075

0

10000

20000

30000

40000

50000

60000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical Challenge

T-D

constraint

0

5000

10000

15000

20000

25000

30000

35000

0.00 10.00 20.00 30.00

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.02 0.052 0.084

tcruise



Increasing Technical Challenge

T-D

constraint

0.036 0.068

0.10

0

5000

10000

15000

20000

25000

30000

35000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint 0 min

10 min

20 min

30 min

0.05 0.10

0.15

tcruise



Increasing Technical Challenge

T-D

constraint

0.075

0.125

0

5000

10000

15000

20000

25000

30000

35000

40000

0.00 10.00 20.00 30.00 40.00

TOGW

(kg)

Istr=Wstr/Swet

B-52 under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical Challenge

T-D

constraint

0.075 0.125
25

6.4.1 AIR-LAUNCHED, MACH 6, HYDROGEN FUEL

Trade Summary: This trade explores the Mach 6 cruise mission with varying endurance time. The 

vehicle is air-launched at 36,000 ft, 0.8M via carrier aircraft and rocket accelerated to ramjet start at 

3.0M. Hydrogen fuel is utilized for the rocket and ramjet/scramjet. This trade determines that (a) the 

mission is feasible with current industrial capability, (b) the vehicle can be air-launched from the B747-

100SCA, and (c) the RL-10-5A liquid hydrogen rocket is an appropriate accelerator motor.

Mission Summary: Table 6-2 summarizes the mission and operational constraints for the air-launched, 

Mach 6, hydrogen trade study.

Table 6-2. Air-Launch, Mach-6, Hydrogen Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection LH, LOX

 Fuel density 74.63kg/m3(4.65 lbs/ft3)

 Oxidizer density 1,287 kg/m3( 803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-11 presents the solution space for the hydrogen fueled, Mach 6, 

air-launch trade. The solution space is bounded by the landing constraint, composite and aluminum 

structural constraints, and the 0 to 30 minutes cruise requirements. This trade is feasible with either 

structural or TPS material. The design point selected has an endurance of 30 minutes. It is composed of an 

aluminum structure and refractory metal TPS. This design point yields the largest design margin from the 

structural, landing and T-D constraints.

Carrier Constraints: While the B-52H could handle the weight of the demonstrator vehicle, it cannot 

accommodate the planform under the wing (constrained by the distance between the fuselage and the 

inboard engine). Figure 6-12 shows the B-52H’s span and length constraints compared to the constant -

contour () of varying cruise time. From this comparison it becomes clear that only the 10 

minutes cruise vehicle can be accommodated under the B-52H wing.
26

Fig. 6-11. Air-launched Mach 6 hydrogen-fueled trade solution space.

Fig. 6-12. B-52H planform area constraint and suggested design point, Mach 6, hydrogen-fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

40000

0.00 10.00 20.00 30.00 40.00

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical 

Challenge

T-D

constraint

0.075

0.125

Carrier Constraint

• Insufficient span and length clearance 

for B-52H

• B747-100 required

0

2

4

6

8

10

12

14

16

18

20

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

10 min cruise required 

for B-52 under wing 

planform constraint 

(X-24c planform)

 = 0.125
27

Additional Trade Studies and Sensitivity Analysis: In addition, this trade study explores utilizing the 

RL-10A-5 (Reference 18) and Vulcain (Reference 17) rocket motors as well as varying the spatular ratio 

(c/s) of the lifting body. From this analysis it is determined that the RL-10 provides sufficient thrust for 

this vehicle, resulting in a lighter engine and TOGW vehicle relative to the Vulcain accelerated variant. A 

spatular ratio (c/s) of 0.5 provides the best balance between scramjet capture area and drag due to wetted 

area. As long as the spatular ratio is contained between 0.25 and 0.75, the parameter has a second order 

effect on vehicle size. Later studies can determine an optimum c/s ratio.

Suggested Design Point Summary: The selected design point has a cruise endurance of 30 minutes, 

seven (7) RL-10A-5 rocket motors, it is composed of aluminum structure, refractory metal TPS and is airlaunched from the B747-100SCA. This design point provides adequate margin from the structural, 

landing and T-D constraints. The vehicle is summarized in Figure 6-13.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 6 cruise mission is feasible with LH2 fuel, aluminum structure and metal TPS.

 The RL-10A-5 is a satisfactory of-the-shelf rocket for this mission.

 A c/s of 0.5 provides a good balance between engine capture area and wetted area. Spatular ratio 

has a 2nd order effect on vehicle size and is therefore fixed at c/s = 0.5 for the remaining tradestudies.

 The B-52H could accommodate the vehicle from a load perspective, but the vehicle would not fit 

between the fuselage and inboard engine. Therefore, the B747-100SCA is the preferred launch 

vehicle.

Fig. 6-13. Summary of suggested design point for the air-launched, Mach 6 hydrogen-fuel trade.

Design Summary

tcruise 30 min

Down range 4,060 km 2,190 nm

TOGW 22,136 kg 48,802 lbs

Wppl 10,047 kg 22,149 lbs

OEW 12,090 kg 26,653 lbs

 0.125

Spln 103.3 m

2

1,112 ft2

B 9.57 m 31 ft

L 18.00 m 59 ft

L/D cruise 2.46

Isp cruise(s) 2,613 s

Trkt 453 kN 102 klbs

Nrkt (64.7 kN each) 7

Carrier Constraint

• Insufficient span and length clearance 

for B-52B

• B747-100 required

Suggested Design Point

• 30 min cruise

•  = 0.125

• B747-100 Launch Vehicle

• RL-10 Rocket motor

• Li-Al structure

• SiC or Refractory TPS

0

2

4

6

8

10

12

14

16

18

20

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

10 min cruise required 

for B-52 under wing 

planform constraint 

(X-24c planform)

 = 0.125
28

6.4.2 AIR-LAUNCHED, MACH 8, HYDROGEN FUEL

Trade Summary: This trade explores the Mach 8 cruise mission with varying cruise endurance. The 

vehicle is air-launched at 36,000 ft, 0.8M via carrier aircraft and rocket accelerated to ramjet start at 

3.0M. Hydrogen fuel is utilized for the rocket and ramjet/scramjet. This trade determines that (a) the 

mission is feasible with current industrial capability, (b) the vehicle can be air-launched from the B747-

100SCA, and (c) the Vulcain liquid hydrogen rocket is an appropriate accelerator motor.

Mission Summary: Table 6-3 summarizes the mission and operational constraints for the air-launched, 

Mach 8, hydrogen trade study.

Table 6-3. Air-Launch, Mach 8, Hydrogen-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection LH, LOX

 Fuel density 74.63kg/m3(4.65 lbs/ft3)

 Oxidizer density 1,287 kg/m3( 803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-14 shows the solution space for the hydrogen fueled, Mach 8, 

air-launch trade. The solution space is constrained by the landing constraint, the composite and aluminum 

structural constraints and the 0 to 30 minutes cruise requirement. This trade demonstrates that the vehicle 

is feasible with either composite or aluminum structure from a cruise time of 0.0 minutes to 30 minutes. 

The design point is selected at the intersection of the landing constraint and the 30 minutes solution curve, 

providing a design margin for the structural (Istr) and propulsion (T-D) constraints.

Note that the slope of the 30 minutes cruise curve increases from = 0.125 to 0.15. This increase in slope 

indicates that the vehicle size is increasingly sensitivity to changes in structural weight. The aluminum 

structure is selected to provide a conservative weight estimate; however, a composite structure vehicle 

will be easier to converge in future studies.

Carrier Constraints: The B-52H can support the weight of the demonstrator vehicle for 0 to 20 minutes

cruise vehicles. However, it cannot accommodate the vehicle under its wing, see Figure 6-15. Plotting the 

span and vehicle length of a constant slenderness parameter () with varying cruise time, it 

becomes clear that even the 0 minutes cruise time vehicle violates the planform requirement. Thus, the 

B747-100SCA must be utilized for the Mach-8 mission.
29

Fig. 6-14. Hydrogen-fuel, Mach 8 air-launch trade solution space.

Fig. 6-15. B-52H planform area constraint and suggested design point, Mach 8, hydrogen-fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

10000

20000

30000

40000

50000

60000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05

0.10

0.15

tcruise



Increasing Technical 

Challenge

T-D

constraint

• Carrier Constraint

• Insufficient span and length clearance 

for B-52H

• B747-100 required

0

5

10

15

20

25

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

10min cruise required 

for B-52 under wing 

planform constraint 

(X-24c planform)

 = 0.15
30

Additional Trade Studies and Sensitivity Analysis: For this cruise mission, the RL-10A-5 and Vulcain 

rocket motors are explored. From this analysis it has been determined that the RL-10A-5 would require an 

excessive number of engines to provide sufficient thrust for the 30 minutes vehicle, requiring twelve (12) 

rocket motors. Therefore, the Vulcain engine is preferred even though it possesses more thrust then 

required for the design.

Suggested Design Point Summary: The selected design point has an endurance of 30 minutes, one (1) 

Vulcain rocket motor, aluminum structure, refractory metal TPS, and it is air-launched from the B747-

100SCA. This design point provides adequate margin from the structural and T-D constraints. The vehicle 

is summarized with Figure 6-16.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 8 cruise mission is feasible with LH2 fuel, aluminum structure and refractory metal 

TPS.

 The Vulcain rocket is a satisfactory of-the-shelf rocket for this mission; however, a LH2 rocket 

between the RL-10A-5 and Vulcain thrust classes is preferred.

 The B-52H cannot support the 30 minutes Mach 8 vehicle nor geometrically accommodate the 

vehicle under the wing. Therefore, the B747-100SCA is the preferred air-launch vehicle.

Fig. 6-16. Summary of suggested design point for the air-launched, Mach 8 hydrogen-fuel trade.

Design Summary

tcruise 30 min

Down range 6,300 km 3,402 nm

TOGW 40,900 kg 90,170 lbs

Wppl 20,821 kg 45,903 lbs

OEW 20,079 kg 44,267 lbs

 0.15

Spln 161.2 m

2

1,735 ft2

b 11.95 m 39 ft

l 22.48 m 74 ft

L/D cruise 2.31

Isp cruise(s) 2,246 s

Trkt 1,015 kN 228 klbs

Nrkt (1015 kN each) 1

• Suggested Design Point

• 30 min cruise

•  = 0.15

• B747-100 Launch Vehicle

• Vulcain Rocket

• Li-Al structure

• SiC/SiCMMC TPS
31

6.4.3 AIR-LAUNCHED, MACH 6, KEROSENE FUEL

Trade Summary: This trade explores the Mach 6 cruise mission with varying endurance time. The 

vehicle is air-launched at 36,000 ft, 0.8M via carrier aircraft and rocket accelerated to ramjet start at 

3.0M. Kerosene fuel is utilized for the rocket and ramjet/scramjet. It is determined that the 30 minutes

mission is not feasible with current industrial capability. The 20 minutes cruise represents the endurance 

limit bounded by the intersection of the landing constraint and composite structure constraint. This flight 

vehicle can be air-launched from the B-52H, it utilizes the Merlin kerosene rocket for acceleration to 

ramjet start. However, an off-the-shelf rocket may present an integration problem due to the high 

slenderness of kerosene vehicles, reducing the upper surface height.

Mission Summary: Table 6-4 summarizes the mission and operational constraints for the air-launched, 

Mach 6, kerosene trade study.

Table 6-4. Air-Launch, Mach 6, Kerosene-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection LH, LOX

 Fuel density 820.0 kg/m3(51.2 lbs/ft3)

 Oxidizer density 1,287 kg/m3(803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-17 shows the solution space for the kerosene-fueled, Mach 6, 

air-launched trade. The solution space is constrained by the landing constraint, the composite structural 

constraint, and the 0 to 20 minutes cruise requirement. This trade demonstrates that the kerosene vehicle 

is more severely constrained via the landing field length compared to the hydrogen vehicle due to the 

increased fuel density, which translates into higher wing loadings. In case the landing constraint is 

relaxed, a 30 minutes composite vehicle is technically feasible. The selected design point is at the 

intersection of the landing constraint, composite structure constraint, and 20 minutes cruise solution 

curve. This design point represents a maximum endurance vehicle while still allowing for: (a) the aborted 

landing condition, (b) a reasonable propulsion margin (distance from T-D), and (c) a limited structural 

margin. In order to increase the structural design margin, the cruise time must be reduced to 10 minutes.
32

Fig. 6-17. Kerosene-fuel Mach-6 air-launched trade solution space.

Fig. 6-18. B-52 planform area constraint and suggested design point, Mach 6, kerosene-fuel trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

0.00 10.00 20.00 30.00

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.02 0.052 0.084

tcruise



Increasing Technical 

Challenge

T-D

constraint

0.036 0.068

0.10

• Carrier Constraint

• 20 min cruise could fit B-52 under wing 

mount.

0

2

4

6

8

10

12

14

16

18

20

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

20 min cruise required 

for B-52 under wing 

planform constraint 

(X-24c planform)

 = 0.068
33

Carrier Constraints: The B-52H can support the weight of all vehicles converged in Figure 6-17. The 20 

minutes cruise vehicle complies with the planform constraints for the B-52H wing mount, see Figure 6-

18. Therefore, the B-52H is selected as the launch vehicle for this trade.

Additional Trade Studies and Sensitivity Analysis: None; the Merlin engine is representative of the 

thrust class required.

Suggested Design Point Summary: The selected design point for this trade consists of air-launch from 

the B-52H and acceleration to ramjet start with the Merlin rocket motor. However, due to the slenderness 

required for the vehicle, the rocket accelerator integration will not geometrically fit into the upper surface 

of the vehicle. This integration will require further design studies if a kerosene Mach 6, air-launched 

demonstrator is selected. The vehicle is summarized in Figure 6-19.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 6 kerosene mission is feasible for 20 minutes cruise endurance with a composite 

structure. If the landing constraint is relaxed, the endurance can be increased to 30 minutes or 

the 20 minutes vehicle could be constructed of aluminum.

 The Merlin rocket has satisfactory thrust performance; however, an off-the-shelf rocket will 

present integration problems due to the required slenderness of the kerosene endurance vehicle.

 The B-52H could support the selected 20 minutes Mach 6 vehicle.

Fig. 6-19. Summary of suggested design point for the air-launched, Mach 6 kerosene-fuel trade.

Design Summary

tcruise 20 min

Down range 3,480 km 1,880 nm

TOGW 14,191 kg 31,287 lbs

Wppl 7,715 kg 17,009 lbs

OEW 6,476 kg 14,277 lbs

 0.07

Spln 58.4 m

2

628 ft2

b 7.19 m 24 ft

l 13.53 m 44 ft

L/D cruise 3.79

Isp cruise(s) 943 s

Trkt 512 kN 115 klbs

Nrkt (512 kN each) 1

• Suggested Design Point

• 20 min cruise

•  = 0.07

• B-52 Launch Vehicle

• Merlin Rocket motor

• Composite structure

• Si/SiCMMCTPS
34

6.4.4 AIR-LAUNCHED, MACH 8, KEROSENE FUEL

Trade Summary: This trade explores the air-launched, Mach 8 kerosene cruise mission with varying 

cruise endurance. The vehicle is air-launched at 36,000 ft, 0.8M via carrier aircraft and rocket accelerated 

to ramjet start at 3.0M. Kerosene fuel is utilized for the rocket and ramjet/scramjet. It is determined that 

the 30 minutes mission is not feasible with the current industrial capability. The 4.5 minutes cruise 

represents the endurance limit bounded by the intersection of the landing constraint and composite 

structure constraint. The 4.5 minutes cruise vehicle can be air-launched from the B-52H; it utilizes the 

Merlin kerosene rocket for acceleration to ramjet start. However, an off-the-shelf rocket may present an 

integration problem due to the required low slenderness of kerosene vehicles.

Mission Summary: Table 6-5 summarizes the mission and operational constraints for the air-launched, 

Mach 8, kerosene trade study.

Table 6-5. Air-Launch, Mach 8, Kerosene-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1, LOX

 Fuel density 820.0 kg/m3(51.2 lbs/ft3)

 Oxidizer density 1,287 kg/m3(803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-20 shows the solution space for the kerosene-fueled, Mach 8, 

air-launched trade. The solution space is bounded by the landing constraint, composite structural 

constraint, and 0 to 30 minutes cruise contour. This demonstrates that the kerosene vehicle is more 

severely constrained by the landing field length relative to hydrogen due to the increased fuel density 

which translates into higher wing loadings. The increase in cruise Mach number from 6 to 8 shifts the 

solution space up and to the left, leaving only a small feasible region with a maximum endurance of 4.5 

minutes. In case the landing constraint is relaxed, a 10 minutes cruise composite vehicle is technically 

feasible. However, such a vehicle will be dangerously close to the thrust minus drag (T-D) constraint.

The selected design point is at the intersection of the landing and composite structure constraint, 

representing the maximum endurance based on the current industry capability available.
35

Fig. 6-20. Kerosene-fuel Mach 8 air-launched trade solution space.

Fig. 6-21. B-52H planform area constraint and suggested design point, Mach 8, kerosene-fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 2 0 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint 0 min

10 min

20 min

30 min

0.02 0.04250.0538



Increasing Technical 

Challenge

T-D

constraint

0.0313

0.065

tcruise

• Carrier Constraint

• Insufficient span and length clearance 

for B-52H

• B747-100 required

0

2

4

6

8

10

12

14

16

18

20

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

Vehicle too large to fit 

under B-52 wing

 = 0.065
36

Carrier Constraints: As with the Mach 8 hydrogen vehicle, none of the vehicles will fit geometrically 

under the B-52 wing. Therefore, the B747-100SCA is the required carrier aircraft, see Figure 6-21.

Additional Trade Studies and Sensitivity Analysis: None; the Merlin engine is representative of the 

thrust class required.

Suggested Design Point Summary: The selected design point for this mission will consist of an airlaunch from the B747-100SCA, and acceleration to ramjet start with the Merlin rocket motor. However, 

due to slenderness requirements, the rocket accelerator integration will geometrically not fit into the upper 

surface of the vehicle. The integration will require further research if a kerosene Mach 8, air-launched 

demonstrator is selected. The vehicle is summarized in Figure 6-22.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 8 cruise mission is feasible with kerosene for a maximum endurance of 4.5 minutes. If 

the landing constraint is relaxed, the endurance can be increased to 10 minutes. This represents 

the smallest solution space of all trades explored.

 The Merlin rocket has satisfactory thrust performance; however, an off-the-shelf rocket will 

present integration problems due to the required slenderness of the kerosene endurance vehicles.

 The B747-100SCA is required for air-launch.

Fig. 6-22. Summary of suggested design point for the air-launched, Mach 8 kerosene-fuel trade.

Design Summary

tcruise 4.5 min

Down range 3,270 km 1,770 nm

TOGW 19,013 kg 41,917 lbs

Wppl 10,627 kg 23,429 lbs

OEW 8,386 kg 18,488 lbs

 0.0675

Spln 76.7 m

2

826 ft2

b 8.24 m 27 ft

l 15.51 m 51 ft

L/D cruise 3.39

Isp cruise(s) 753 s

Trkt 512 kN 115 klbs

Nrkt (512 kN each) 1

• Suggested Design Point

• 20 min cruise

•  = 0.07

• B-52 Launch Vehicle

• Merlin Rocket motor

• Composite structure

• Si/SiCMMCTPS
37

6.4.5 AIR-LAUNCHED, MACH 6, DUAL-FUEL

Trade Summary: The Mach 8 hydrogen fueled demonstrator shows that the 1,000 kN thrust class 

Vulcain rocket is feasible; however, only a 600 kN rocket is required. Given that the Merlin rocket is in 

the 500 kN thrust class and the X-24C utilized a kerosene rocket and hydrogen scramjet (Reference 15), it 

has been decided to add a dual-fuel option to the air-launch studies.

This trade explores the Mach 6 cruise mission with varying endurance time. The vehicle is air-launched at 

36,000 ft, 0.8M via carrier aircraft and rocket accelerated to ramjet start at 3.0M. Kerosene is utilized for 

the rocket and hydrogen for the ramjet/scramjet. This trade determines that (a) the mission is feasible with 

the current industrial capability, (b) the vehicle can be air-launched from the B747-100SCA, and (c) the 

Merlin kerosene rocket is an appropriate accelerator motor.

Mission Summary: Table 6-6 summarizes the mission and operational constraints for the air-launched, 

Mach 6, dual-fuel trade study.

Table 6-6. Air-Launch, Mach 6, Dual-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection RP-1, LH, LOX

 Kerosene fuel density 820.0 kg/m3(51.2 lbs/ft3)

 Hydrogen fuel density 74.63kg/m3(4.65 lbs/ft3)

 Oxidizer density 1,287 kg/m3(803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-23 shows the solution space for the dual-fueled, Mach 6, airlaunched trade. The solution space is constrained by the landing constraint, the composite and aluminum 

structural constraints, and the 0 to 30 minutes cruise requirements. This vehicle is feasible with composite 

structure and either TPS. The design point selected has an endurance of 30 minutes at Mach 6 and is 

composed of a composite structure and refractory metal TPS. This design point provides the required 

endurance with some margin for both structural (Istr) and propulsion (T-D) technology constraints. While 

the dual-fuel variant is technically feasible, it requires a lighter structure compared to the equivalent 

hydrogen vehicle due to the reduced Isp of the kerosene rocket.

Carrier Constraints: Similar to the Mach 6 hydrogen vehicle, the B-52H can support the weight of the 

vehicle but not the geometry, see Figure 6-24.
38

Fig. 6-23. Dual-fuel Mach 6 air-launched trade solution space.

Fig. 6-24. B-52H planform area constraint and suggested design point, Mach 6, dual-fuel trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

0 1 0 2 0 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint

0 min

10 min

20 min

30 min

0.05 0.10

0.15

tcruise



Increasing Technical 

Challenge

T-D

constraint

0.075

0.125

• Carrier Constraint

• 15 min cruise could fit B-52 under wing 

mount.

0

2

4

6

8

10

12

14

16

18

20

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

10 to 15 min cruise 

required for B-52 under 

wing planform 

constraint (X-24c 

planform)

 = 0.15
39

Additional Trade Studies and Sensitivity Analysis: None; the Merlin engine is representative for the 

thrust class required.

Suggested Design Point Summary: The selected design point for this mission will consist of air-launch 

from the B747-100SCA and acceleration to ramjet start with the Merlin rocket motor. This vehicle will 

require composite structure to compensate for the reduced Isp of kerosene. The vehicle is summarized in 

Figure 6-25.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 6 cruise mission is feasible with the dual-fuel option for 30 minutes endurance. 

However, a composite structure is required to compensate for the reduced Isp and heavier kerosene 

rocket fuel.

 The Merlin rocket has satisfactory performance as an off-the-shelf rocket. The dual-fuel option 

will not have the same integration issues as the kerosene-only vehicles due to the decrease in 

slenderness.

 The 30 minutes vehicle requires the B-747-100SCA for air launch.

Fig. 6-25. Summary of suggested design point for the air-launched, Mach 6 dual-fuel trade.

Design Summary

tcruise 30 min

Down range 4,226 km 2,282 nm

TOGW 19,606 kg 43,224 lbs

Wppl 10,126 kg 22,324 lbs

OEW 9,480 kg 20,900 lbs

 0.15

Spln 84.7 m

2

912 ft2

B 8.66 m 28 ft

L 16.30 m 53 ft

L/D cruise 2.03

Isp cruise(s) 2,619 s

Trkt 512 kN 115 klbs

Nrkt (512 kN each) 1

• Suggested Design Point

• 30 min cruise

•  = 0.15

• B747-100 Launch Vehicle

• Merlin Rocket motor

• Composite structure

• SiC or refractory metal 

TPS
40

6.4.6 AIR-LAUNCHED, MACH 8, DUAL-FUEL

Trade Summary: The Mach 8 hydrogen fueled demonstrator, using the 1,000 kN thrust class Vulcain 

rocket, shows feasibility. However, only a 600 kN rocket is required. Given that the Merlin rocket is in 

the 500 kN thrust class and the X-24C has been utilizing a kerosene rocket and hydrogen scramjet 

(Reference 15), it has been decided to add a dual-fuel option to the air-launch studies.

This trade explores the Mach 8 cruise mission with varying endurance time. The vehicle is air-launched at 

36,000 ft, 0.8M via carrier aircraft and rocket accelerated to ramjet start at 3.0M. Kerosene is utilized for 

the rocket and hydrogen for the ramjet/scramjet. This trade determines that (a) the mission is feasible with 

the current industrial capability, (b) the vehicle can be air-launched from the B747-100SCA, and (c) the 

Merlin rocket is an appropriate thrust class accelerator motor.

Mission Summary: Table 6-7 summarizes the mission and operational constraints for the air-launched, 

Mach 6, dual-fuel trade study.

Table 6-7. Air-Launch, Mach 8, Dual-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1, LH, LOX

 Kerosene fuel density 820.0 kg/m3(51.2 lbs/ft3)

 Hydrogen fuel density 74.63kg/m3(4.65 lbs/ft3)

 Oxidizer density 1,287 kg/m3(803.34 lbs/ft3) 

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-26 shows the solution space for the dual-fuel, Mach 8, airlaunched trade. The solution space is constrained by the landing constraint, composite structure, and the 0 

to 30 minutes solution requirements. Interestingly, the thrust minus drag (T-D), landing and composite 

structure constraints all coalesce at a single point on the 30 minutes solution curve, representing a zero 

margin design point. To allow for a propulsion margin, a 25 minutes cruise time design point is selected. 

In the case of the dual-fuel Mach 8 vehicle, the increased weight of kerosene and reduced Isp results in 

reduced cruise endurance compared to the all-hydrogen vehicle alternative.
41

Fig. 6-26. Dual-fuel, Mach-8 air-launched trade solution space.

Fig. 6-27. B-52H planform area constraint and suggested design point, Mach 8, dual-fuel trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Landing

constraint 0 min

10 min

20 min

30 min

0.05

0.10

0.125



Increasing Technical 

Challenge

T-D

constraint

tcruise

0.075

• Carrier Constraint

• Insufficient span and length clearance 

for B-52H

• B747-100 required

0

5

10

15

20

25

0 5 10 15 20

l (m)

b (m)

B-52 

under wing

0 min

10 min

20 min

30 min

0min cruise required 

for B-52 under wing 

planform constraint 

(X-24c planform)

 = 0.125
42

Carrier Constraints: While the B-52H could support the weight of the vehicle, however, its planform 

size is too large for all cruise endurance points at Mach 8, see Figure 6-27.

Additional Trade Studies and Sensitivity Analysis: None; the Merlin engine is representative for the

thrust classes required.

Suggested Design Point Summary: The selected design point for this mission consists of air-launch 

from the B747-100SCA and acceleration to ramjet start with the Merlin rocket motor. Due to the increase 

in fuel density (increasing wing loading) and the reduced Isp of the kerosene rocket, the cruise time must 

be reduced to 25 minutes to provide some propulsion margin. The vehicle is summarized in Figure 6-28.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 8 cruise mission is feasible with the dual-fuel option for 30 minutes endurance; 

however, the 30 minutes option represents a zero margin design point. Therefore the cruise 

endurance is reduced to 25 minutes.

 The Merlin rocket has satisfactory performance for an off-the-shelf rocket. The dual-fuel option 

will not have the same integration issues as the kerosene-only vehicles due to the decrease in 

slenderness.

 All cruise times violate the B-52 under wing geometry constraint, requiring the B-747-100SCA 

carrier aircraft for air launch.

Fig. 6-28. Summary of suggested design point for the air-launched, Mach 8 dual-fuel trade.

Design Summary

tcruise 25 min

Down range 5,560 km 3,000 nm

TOGW 25,635 kg 56,516 lbs

Wppl 13,667 kg 30,130 lbs

OEW 11,968 kg 26,385 lbs

 0.13

Spln 114.4 m

2

1,231 ft2

B 10.06 m 33 ft

L 18.94 m 62 ft

L/D cruise 2.28

Isp cruise(s) 2,246 s

Trkt 512 kN 115 klbs

Nrkt (512 kN each) 1

• Suggested Design Point

• 25 min cruise

•  = 0.13

• B747-100 Launch Vehicle

• Merlin Rocket

• Composite structure

• SiC/SiCMMC TPS
43

6.4.7 EXPENDABLE BOOSTER, MACH 6, HYDROGEN FUEL

Trade Summary: This trade explores the Mach 6 cruise mission with varying cruise endurance. The 

vehicle is vertically launched on top of an expendable booster and accelerated to ramjet start at 3.0M. 

Hydrogen fuel is used for the ramjet/scramjet. The trade study determines that this mission is feasible 

with the current industrial capability available. The vehicle is within the weight, separation velocity, and 

separation altitude constraints for the Castor 120 expendable booster.

Mission Summary: Table 6-8 summarizes the mission and operational constraints for the expendable 

booster, Mach 6, hydrogen-fuel trade study.

Table 6-8. Expendable Booster, Mach 6, Hydrogen Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection LH

 Hydrogen fuel density 74.63kg/m3(4.65 lbs/ft3)

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-29 shows the solution space for the hydrogen-fueled, Mach 6, 

expendable booster trade. The solution space is constrained by the landing constraint, composite structure 

constraint and 0 to 30 minutes cruise requirement. The landing and thrust minus drag (T-D) constraints 

cross at the 10 minutes cruise solution, meaning that for cruise times below 10 minutes the landing 

constraint is dominant, whereby above 10 minutes the T-D constraint dominates. This switch in dominant 

constraints is due the increase fuel weight of the 20 to 30 minutes endurance vehicles, leading to 

increased thrust requirements.

Expendable Booster Constraints: All Mach 6 hydrogen vehicles meet the weight, separation velocity, 

and separation altitude constraints of the M55A1 and Castor 120 expendable boosters, see Figure 6-30.

Additional Trade Studies and Sensitivity Analysis: None.
44

Fig. 6-29. Hydrogen-fuel, Mach 6 expendable booster trade solution space.

Fig. 6-30. Expendable booster constraints and suggested design point, Mach 6, hydrogen- fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.12

0.17

0.195

tcruise

Increasing Technical 

Challenge

T-D

constraint

0.145

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 500 1000 1500 2000 2500

Weight, kg

Velocity at Transition, m/sec

Design Point

M55A1 Expendable Booster

Castor 120 Expendable Booster
45

Suggested Design Point Summary: The selected design point for this mission consists of vertical launch 

using the M55A1 expendable booster for acceleration to ramjet start at Mach 3. A composite structure has 

been chosen for the design point, although aluminum is technically possible. This choice has been made 

in order to increase the propulsion margin relative to the thrust minus drag (T-D) constraint. The vehicle 

is summarized in Figure 6-31.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade:

 The Mach 6 cruise mission is feasible with the hydrogen fuel option for 30 minutes endurance.

 The M55A1 expendable booster meets the requirements for the Mach 6 mission.

 For increasing cruise durations, the thrust minus drag (T-D) becomes more constraining relative 

to the aborted landing constraint.

Fig. 6-31. Summary of suggested design point for the expendable booster, Mach 6 hydrogen- fuel 

trade.

Design Summary

tcruise 30 min

Down range 4,120 km 2,224 nm

TOGW 25,635 kg 25,364 lbs

Wppl 3,757 kg 8,283 lbs

OEW 7,709 kg 16,995 lbs

 0.175

Spln 63.5 m

2

683.5 ft2

B 7.50 m 25 ft

L 14.1 m 46 ft

L/D cruise 1.88

Isp cruise(s) 2,600 s

• Suggested Design Point

• 30 min cruise

•  = 0.175

• M55A1 Expendable 

Booster

• Composite structure

• SiC/SiCMMC or Refractory 

TPS
46

6.4.8 EXPENDABLE BOOSTER, MACH 8, HYDROGEN FUEL

Trade Summary: This trade explores the Mach 8 cruise mission with varying endurance time. The 

vehicle is vertically launched on top of an expendable booster and accelerated to ramjet start at 3.0M. 

Hydrogen fuel is used for the ramjet/scramjet. It is determined that this mission is feasible with the 

current industrial capability. The vehicle is within the weight, separation velocity, and separation altitude 

constraints for the Castor 120 expendable booster.

Mission Summary: Table 6-9 summarizes the mission and operational constraints for the expendable 

booster, Mach 8, hydrogen fuel trade study.

Table 6-9. Expendable Booster, Mach 8, Hydrogen Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection LH

 Hydrogen fuel density 74.63kg/m3(4.65 lbs/ft3)

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-32 shows the solution space for the hydrogen fueled, Mach 8, 

expendable booster trade. The solution space is constrained by the landing constraint, composite structure 

constraint and 0 to 30 minutes solution curves. The demonstrator vehicle is feasible with either composite 

or aluminum structure for cruise times from 0.0 to 30 minutes. The design point is selected at the 

intersection of the landing constraint and aluminum structure constraint. This has been done to provide 

sufficient design margin for the structural and propulsion (T-D) constraints.

Expendable Booster Constraints: The selected design point meets the weight, separation velocity, and 

separation altitude constraints of the Castor 120 expendable booster. However, a more powerful booster is 

required if TOGW is expected to increase, see Figure 6-33.

Additional Trade Studies and Sensitivity Analysis: None.
47

Fig. 6-32. Hydrogen-fuel, Mach 8 expendable booster trade solution space.

Fig. 6-33. Expendable booster constraints and suggested design point, Mach 8, hydrogen-fuel trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

5000

10000

15000

20000

25000

30000

35000

40000

0 10 2 0 3 0 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint 0 min

10 min

20 min

30 min

0.1 0.1275

tcruise 0.155

Increasing Technical 

Challenge

T-D

constraint

0.1825

0.21

0

5000

10000

15000

20000

25000

0 500 1000 1500 2000 2500

Weight, kg

Velocity at Transition, m/sec

M55A1 Expendable Booster

Castor 120 Expendable Booster
48

Suggested Design Point Summary: The selected design point for this mission consists of vertical launch 

using the Castor 120 expendable booster for acceleration to ramjet start at Mach 3. The design point is 

selected at the intersection of the landing constraint and aluminum structure constraint. Note that for 

slenderness parameters greater than the design point, the gradient of TOGW with respect to the structural 

index is increasing almost asymptotically. Note that a small increase in material weight will spiral into a 

large increase in vehicles size. The vehicle is summarized in Figure 6-34.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 8 cruise mission is feasible with the hydrogen fuel option for 30 minutes endurance.

 The Castor 120 expendable booster meets the requirements for the Mach 8 mission design point 

although any increase in TOGW will require the use of a more powerful expendable booster.

 The large gradients in the hydrogen solution curves lead to large changes in TOGW with small 

changes in material weight.

Fig. 6-34. Summary of suggested design point for the expendable booster, Mach 8, hydrogen-fuel 

trade.

Design Summary

tcruise 30 min

Down range 6,000 km 3,239 nm

TOGW 19,577 kg 43,160 lbs

Wppl 7,423 kg 16,365 lbs

OEW 12,153 kg 26,793 lbs

 0.1825

Spln 95.67 m

2

1,230 ft2

b 9.20 m 30 ft

l 17.32 m 57 ft

L/D cruise 1.98

Isp cruise(s) 2,248 s

• Suggested Design Point

• 30 min cruise

•  = 0.1825

• Castor 120 Expendable 

Booster

• Li-Al structure

• SiC/SiCMMC or Refactory

TPS
49

6.4.9 EXPENDABLE BOOSTER, MACH 6, KEROSENE FUEL

Trade Summary: This trade explores the Mach 6 cruise mission with varying endurance time. The 

vehicle is vertically launched on top of an expendable booster and accelerated to ramjet start at Mach 3. 

Kerosene fuel is used for the ramjet/scramjet. It is determined that this mission is feasible with current 

industrial capability. The vehicle is within the weight, separation velocity, and separation altitude 

constraints for the Castor 120 expendable booster.

Mission Summary: Table 6-10 summarizes the mission and operational constraints for the expendable 

booster, Mach 6, kerosene-fuel trade study.

Table 6-10. Expendable Booster, Mach 6, Kerosene-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection RP-1

Kerosene-fuel density 820.0 kg/m3(51.2 lbs/ft3)

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-35 shows the solution space for the kerosene-fueled, Mach 6, 

expendable booster trade. The solution space is constrained by the landing constraint, composite structure 

constraint, and the 0 to 30 minutes solution curves. The solution space demonstrates that the vehicle is 

feasible with either composite or aluminum structure for cruise times between 0 to 30 minutes. The 

design point is selected at the intersection of the landing constraint and aluminum structure constraint. 

This selection provides a healthy design margin from the structural and propulsion (T-D) constraints.

Expendable Booster Constraints: All Mach 6 kerosene vehicles meet the weight, separation velocity, 

and separation altitude constraints of the M55A1 and Castor 120 expendable boosters, see Figure 6-36.

Additional Trade Studies and Sensitivity Analysis: None.
50

Fig. 6-35. Kerosene-fuel, Mach 6 expendable booster trade solution space.

Fig. 6.36. Expendable booster constraints and suggested design point, Mach 6, kerosene-fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

2000

4000

6000

8000

10000

12000

0 1 0 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.05 0.0725 0.095 tcruise



Increasing Technical 

Challenge

T-D

constraint

0.1175

0.14

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 500 1000 1500 2000 2500

Weight, kg

Velocity at Transition, m/sec

Design Point

M55A1 Expendable Booster

Castor 120 Expendable Booster
51

Suggested Design Point Summary: The selected design point for this mission consists of vertical launch 

using the M55A1 expendable booster for acceleration to ramjet start at Mach 3. The design point is 

selected at the intersection of the landing constraint and aluminum structure constraint. The vehicle is 

summarized in Figure 6-37.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 6 cruise mission is feasible with the kerosene-fuel option for 30 minutes endurance.

 The N55A1 expendable booster meets the requirements for the Mach 6 mission.

Fig. 6.37. Summary of suggested design point for the expendable booster, Mach 6 kerosene-fuel trade.

Design Summary

tcruise 30 min

Down range 4,523 km 2,442 nm

TOGW 8,345 kg 18,398 lbs

Wppl 3,536 kg 7,796 lbs

OEW 4,809 kg 10,602 lbs

 0.085

Spln 34.75 m

2

374 ft2

B 5.55 m 18 ft

L 10.44 m 34 ft

L/D cruise 4.08

Isp cruise(s) 970 s

• Suggested Design Point

• 30 min cruise

•  = 0.085

• M55A1 Expendable 

Booster

• Composite structure

• SiC/SiCMMC TPS
52

6.4.10 EXPENDABLE BOOSTER, MACH 8, KEROSENE FUEL

Trade Summary: This trade explores the Mach 8 cruise mission with varying endurance time. The 

vehicle is vertically launched on top of an expendable booster and accelerated to ramjet start at 3.0M. 

Kerosene fuel is used for the ramjet/scramjet. It is determined that the mission is feasible with current 

industrial capability, though with a reduction in cruise time to 20 minutes. The vehicle is within the 

weight, separation velocity, and separation altitude constraints for the Castor 120 expendable booster.

Mission Summary: Table 6-11 summarizes the mission and operational constraints for the expendable 

booster, Mach 8, kerosene-fuel trade study.

Table 6-11. Expendable Booster, Mach 8, Kerosene-Fuel Mission Summary

Mission Requirements

Endurance 0, 10, 20 and 30 min

Payload 0 kg (0 lbs)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1

 Kerosene fuel density 820.0 kg/m3(51.2 lbs/ft3)

Operational Constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Solution Space Visualization: Figure 6-38 shows the solution space for the kerosene-fueled, Mach 8, 

expendable booster trade. The solution space is constrained by the landing constraint, composite structure 

constraint and the 0 to 30 minutes solution curve. The solution space is severely constrained by the 

aborted landing constraint. This is due to the increased fuel density translating into higher vehicle wing 

loadings. In case the landing constraint is relaxed, a 30 minutes composite vehicle is technically feasible. 

This point is at the intersection of the structural capability and the thrust minus drag (T-D) constraints, 

resulting in a zero design margin for the structural weight and thrust available. Consequently, the design 

point selected is reduced to 20 minutes endurance.

Expendable Booster Constraints: All Mach 8 kerosene vehicles meet the weight, separation velocity, 

and separation altitude constraints of the M55A1 and Castor 120 expendable boosters, see Figure 6-39.

Additional Trade Studies and Sensitivity Analysis: None.
53

Fig. 6-38. Kerosene-fuel, Mach 8, expendable booster trade solution space.

Fig. 6-39. Expendable booster constraints and suggested design point, Mach 8, kerosene-fuel 

trade.

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 1 0 20 3 0 40

TOGW

(kg)

Istr=Wstr/Swet

Landing

constraint

0 min

10 min

20 min

30 min

0.0425

0.0625

0.0825

tcruise



Increasing Technical 

Challenge

T-D

constraint

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 500 1000 1500 2000 2500

Weight, kg

Velocity at Transition, m/sec

M55A1 Expendable Booster

Castor 120 Expendable Booster

Design Point
54

Suggested Design Point Summary: The selected design point for this mission consists of vertical launch 

using the M55A1 expendable booster for acceleration to ramjet start at Mach 3. The design point is 

selected at the intersection of the 20 minutes solution curve and the composite structure constraint. In case 

the aborted landing constraint is relaxed, the thrust minus drag constraint for the 30 minutes cruise vehicle 

is feasible with a minimal thrust margin. Reducing the endurance to 20 minutes allows for a sufficient 

propulsion margin that meets the aborted launch constraint. The vehicle is summarized in Figure 6-40.

Summary of Conclusions and Key Points: Below are the primary conclusions from this trade.

 The Mach 8 cruise mission is infeasible with the kerosene-fuel option for 30 minutes endurance; 

however, the 20 minutes option can be accomplished.

 The M55A1 expendable booster meets the requirements for the Mach 8 mission.

Fig. 6-40. Summary of suggested design point for the expendable booster, Mach 8, kerosene-fuel trade.

Design Summary

tcruise 20 min

Down range 5,640 km 3,045 nm

TOGW 12,027 kg 26,515 lbs

Wppl 6,074 kg 13,391 lbs

OEW 5,953 kg 13,124 lbs

 0.075

Spln 51.28 m

2

552 ft2

b 6.74 m 22 ft

l 12.68 m 42 ft

L/D cruise 3.92

Isp cruise(s) 732 s

• Suggested Design Point

• 20 min cruise

•  = 0.075

• M55A1 Expendable 

Booster

• Composite structure

• SiC/SiCMMC TPS
55

7 SOLUTION SPACE COMPARISON AND BASELINE SELECTION

Using the results of this study, two endurance hypersonic demonstrators have been identified as 

prospective baseline vehicles for research and development, concept formulation and definition, and 

system development efforts. It has been determined that the goal of first flight within the 10 to 20 year 

time span can be achieved with reasonable confidence using mostly existing industrial capability. 

Required technology development efforts would primarily focus on scramjet engine requirements for (a) a 

hydrogen-based, and/or (b) a kerosene-based operational infrastructure.

In summary, the current research undertaking has covered and delivered sensitivity trends for launch and 

staging options, accelerator motor selection, ramjet/scramjet fuel selection, material concept and 

configuration arrangement, all measured against the operational mission (i.e. cruise time, speed 

requirement). Considering the broadness of these engineering options evaluated, the value of parametric 

sizing (PS) on physical understanding and system-level decision-making has been demonstrated. Clearly, 

parametric sizing utilizes the first principles mindset and tools to answer how changes within the mission, 

operational scenario and overall research objectives influence the design ‘hardware’ requirements, thus 

the decision-making process. The recommendations and conclusions of the solution space trade analysis 

follow.

7.1 SOLUTION SPACE COMPARISON

A. Design-Level Summary

A direct comparison of the hydrogen and kerosene demonstrator trade space illustrates that hydrogen 

vehicles have a larger feasible design space relative to kerosene equivalents, see Figure 7-1. Comparing 

kerosene vehicles relative to hydrogen vehicles, the kerosene designs show larger sensitivity to landing 

constraints due to increased vehicle density (which increases wing loading) and the requirement for a 

lighter structure to compensate for reduced fuel Isp values. Comparing hydrogen vehicles relative to

Fig. 7-1. Hydrogen-fueled vehicles allow for a larger technical solution space compared to kerosenefueled vehicles.

Hydrogen, Insulated integral fuel tank Kerosene, Non-insulated fuel tank

0

10000

20000

30000

40000

50000

60000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

M8, AL, LH2

Increasing Technical 

Challenge

M6, AL, LH2

M6, EB, LH2

M8, EB, LH2

0

10000

20000

30000

40000

50000

60000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Increasing Technical 

Challenge

M8, AL, RP-1

M6, AL, RP-1

M8, EB, RP-1

Increasing Technical 

Challenge

M6, EB, RP-1

Structural Technology Levels (Istr= Wstr/Swet)

Li - AL Structure - Refractory shingles 

Li - AL Structure – SEP SiC/SiCMMC shingles

Composite Structure - Refractory shingles 

Composite Structure - SEP SiC/SiCMMC shingles
56

kerosene vehicles, the trade-off between fuel weight density and energy density characteristics yields a 

higher total system benefit for hydrogen.

B. Mission-Level Summary

In order to explore the hypersonic design relationships at mission level, Figure 7-1 superimposes the outer 

contours of the hydrogen and kerosene solution spaces. Both design spaces, with decreasing maximum 

TOGW, include (a) M=8 Air-Launch, (b) M=6 Air-Launch, (c) M=8 Expendable Booster, and (d) M=6 

Expendable Booster. This discussion centers on the cruise time constraint equal to 30 min (positive curve 

at the top of the trade space). For the hydrogen-based demonstrators, the individual solution spaces offer a 

vehicle point-design each that meets the operational limit while having the largest structural technology 

margin compared to kerosene equivalents. The M=8 Air-Launch option could be considered the higher 

risk solution for the 30 minutes cruise mission. For the kerosene-based demonstrators, only the M=6 

Expendable Booster trade offers a feasible 30 minutes endurance solution. The remaining trades do not 

present feasible solutions for the 30 minutes demonstrator due to structural constraints. This shows that 

overall vehicle feasibility is dependent on not-yet-available structural industry capability, thus requiring 

future structures technology developments.

7.2 DESIGN POINT COMPARISON

The following discussion reviews the converged baseline vehicle design points selected from the 

hypersonic flight vehicle design solution space screening activity presented in Chapter 6. For more 

information regarding the demonstrator selection for individual hydrogen- and kerosene-fuel trades, 

please refer to the earlier sections. Figure 7-2 presents the short-list overview of prospective baseline 

vehicle configuration-, speed- and fuel combinations. Table 7-1 and Table 7-2 are summarizing the 

general ‘parametric’ design characteristics for the feasible baseline vehicle options utilizing either 

hydrogen or kerosene fuel.

Fig. 7-2. Configuration geometry of proposed hydrogen and kerosene hypersonic baseline vehicle 

designs.

Air Launch Expendable booster

Mach 6

Mach 8

Mach 6

Mach 8

Hydrogen Kerosene
57

Table 7-1. Design Characteristics for Hydrogen-Based Suggested Vehicle Selection

Table 7-2. Design Characteristics for Kerosene-Based Suggested Vehicle Selection

tcruise 30 min 30 min 30 min 30 min

Down 

range

4060 km 2190 n m 6300 km 3402 n m 4120 km 2224 n m 6000 km 3239 n m

TOGW 22136 kg 48802 lbs 40900 kg 90170 lbs 25635 kg 25364 lbs 19577 kg 43160 lbs

Wppl 10047 kg 22149 lbs 20821 kg 45903 lbs 3757 kg 8283 lbs 7423 kg 16365 lbs

OEW 12090 kg 26653 lbs 20079 kg 44267 lbs 7709 kg 16995 lbs 12153 kg 26793 lbs

 0.125 0.15 0.175 0.1825

Spln 103.3 m

2 1112 ft2 161.2 m2 1735 ft2 63.5 m2 683.5 ft2 95.67 m2 1230 ft2

B 9.57 m 31 ft 11.95 m 39 ft 7.5 m 25 ft 9.2 m 30 ft

L 18 m 59 ft 22.48 m 74 ft 14.1 m 46 ft 17.32 m 57 ft

L/D cruise 2.46 2.31 1.88 1.98

Isp 

cruise (s) 2613 s 2246 s 2600 s 2248 s

Trkt 453 kN 102 klbs 1015 kN 228 klbs

Nrkt 7 at 64.7kN each 1 at 1015 kN each

Mach 6, Air-Launch, LH2 Mach 8, Air-Launch, LH2 Mach 6, Expendable Booster, LH2 Mach 8, Expendable Booster, LH2

tcruise 20 min 4.5 min 30 min 20 min

Down 

range

3480 km 1880 n m 3270 km 1770 n m 4523 km 2442 n m 5640 km 3045 n m

TOGW 14191 kg 31287 lbs 19013 kg 41917 lbs 8345 kg 18398 lbs 12027 kg 26515 lbs

Wppl 7715 kg 17009 lbs 10627 kg 23429 lbs 3536 kg 7796 lbs 6074 kg 13391 lbs

OEW 6476 kg 14277 lbs 8386 kg 18488 lbs 4809 kg 10602 lbs 5953 kg 13124 lbs

 0.07 0.0675 0.085 0.075

Spln 58.4 m

2 628 ft2 76.7 m2 826 ft2 34.75 m2 374 ft2 51.28 m2 552 ft2

b 7.19 m 24 ft 8.24 m 27 ft 5.55 m 18 ft 6.74 m 22 ft

l 13.53 m 44 ft 15.51 m 51 ft 10.44 m 34 ft 12.68 m 42 ft

L/D cruise 3.79 3.39 4.08 3.92

Isp 

cruise (s) 943 s 753 s 970 s 732 s

Trkt 512 kN 115 klbs 512 kN 115 klbs

Nrkt 1 at 512 kN each 1 at 512 kN each

Mach 6, Air-Launch, RP-1 Mach 8, Air-Launch, RP-1 Mach 6, Expendable Booster, RP-1 Mach 8, Expendable Booster, RP-1
58

7.3 BASELINE VEHICLE SELECTION

While feasible options for both, the hydrogen-fueled and kerosene-fueled vehicles, exist, the selection of 

the fuel type alone is not a sufficient indicator for demonstrator feasibility. The selection criteria for the 

fuel type are primarily determined by the required operational vehicle characteristics, in this case being a 

robust air-breathing propulsion system flying test bed. Clearly, additional criteria are needed to measure 

the risk and benefit merits of this demonstrator vehicle. At this point we ask the simple question: “If a 

hydrogen fueled scramjet is required, what demonstrator is recommended?” and “If a kerosene-fueled 

scramjet is required, what demonstrator is recommended?”

For each fuel requirement, trade-studies will have to address the following four qualitative metrics:

1. Versatility Which vehicle represents the largest flexibility of its operational capability?

2. Growth Capability Which vehicle is the least sensitive to scale? In other words, which vehicle is 

least sensitive to changes in structural capability which are assumed for this study?

3. Design Confidence Which vehicle has the largest technology margins and allows for a design point 

which has sufficient margin in terms of structural technology, T-D and landing distance?

4. Limitations Which vehicle has any perceived limitations that would hinder development?

If hydrogen scramjet testing is required, assessment results are presented with Figure 7-3:

Fig. 7-3. The Mach 8 air-launched case represents the largest operational flexibility while Mach 6 

air-launched has larger growth capability and design confidence. Since M6 30 min and M8 

10 min solution curves overlay, it appears that the M6 30 min vehicle could perform the 

Mach 8 mission for 10 minutes.

• Selection

• Mach 6, 30 min cruise, air launch 

with Mach 8, 10 min cruise capability

(the solutions spaces overlap)

• B747-100 air-launch

(planform constrained on the B-52) 

• RL-10 rocket accelerator

• Li-Al structure

• Selection Criteria

• Versatility

(largest flexibility in operation)

• Mach 8, air-launch, 30 min

• Growth capability in concept

(sensitivity to increase in scale)

• Mach 6, air-launch, 30 min

• Design confidence

(largest technology margins)

• Mach 6, air-launch, 30 min

• Limitations

(Hydrogen demonstrator)

• B747-100 carrier required for 

air-launch

• Operational limitation of 

Hydrogen (relative to Kerosene)

0

10000

20000

30000

40000

50000

60000

0 10 20 30 40

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

M8, AL, LH2

Increasing Technical 

Challenge

M6, AL, LH2
59

Observing that the Mach 6, 30 minutes vehicle can perform the Mach 8 mission for 10 minutes, this 

scenario provides a compromise which will allow for both, the endurance and speed requirements to be 

accomplished at a lower risk option compared to the Mach 8, 30 minutes vehicle. Consequently, the 

selection of this particular baseline design provides a superior design margin and a concept less sensitive 

to structural and propulsion technology requirements.

If kerosene scramjet testing is required, assessment results are presented with Figure 7-4:

Given the increased density of kerosene (which increases W/S and causes the landing constraint to 

increase) accompanied with a reduced energy density, the required structural technology must increase to 

compensate. This leaves the Mach 6, 30 minutes vehicle as the only viable technical option for kerosene 

scramjets. Furthermore, it is important to note that the Mach 6, 30 minutes solution overlays with the 

Mach 8, 0 minutes cruise time solution. Consequently, the Mach 6, 30 minutes research vehicle can 

accelerate to Mach 8, but it will not have sufficient fuel for 30 minutes but 10 minutes cruise endurance.

Fig. 7-4. The Mach 6 kerosene-fuel expendable booster trade is the only trade-study which allows for 

30 minutes cruise endurance.

Both research demonstrators represent attractive options, each offering the capability to explore advanced 

propulsion design concepts.

• Selection

• Mach 6, 30 min cruise

• Composite structure

(possible aluminum) 

• M55A1 Expendable Booster

• Selection Criteria

• Versatility

(largest flexibility in operation)

• Mach 6, expendable, 30 min

• Growth capability in concept

(sensitivity to increase in scale)

• Mach 6, expendable, 30 min

• Design confidence

(largest technology margins)

• Mach 6, expendable, 30 min

• Limitations

(Kerosene demonstrator)

• High level of structural 

technology required

• Operation of Kerosene at high 

temperatures

0

10000

20000

30000

40000

50000

60000

0 10 20 30

TOGW

(kg)

Istr=Wstr/Swet

B-52 

under wing

Increasing Technical 

Challenge

M8, AL, RP-1

M6, AL, RP-1

M8, EB, RP-1

Increasing Technical 

Challenge

M6, EB, RP-1
60

8 SUMMARY OF CONCLUSIONS AND RECOMMENDATIONS

This report documents a parametric sizing (PS) study performed to develop a program strategy for (a) 

research and development (R&D), and (b) procurement of a feasible next-generation hypersonic airbreathing endurance demonstrator. Overall project focus has been on complementing technical and 

managerial decision-making during the earliest conceptual design phase towards minimization of 

operational, technical, and managerial risks.

The first segment in the course of the conceptual design phase is parametric sizing (PS), the second 

segment performs configuration layout (CL), and the third segment is the configuration evaluation (CE) 

stage. The early, thus critical, PS segment requires a systematic approach, enabling the generation of 

physically correct understanding and knowledge of the solution space available before the subsequent CD 

stages (CL and CE) are locking in on a baseline system in this very solution space.

In the context of the present research undertaking, the AVD Laboratory team has utilized a dedicated 

parametric sizing (PS) tool to measure sensitivities and classical figures-of-merit for the manager [M], 

synthesis specialist [S], and technologist [T]. The systematic approach applied (screening & sizing) is 

iteratively harmonizing the relationships amongst: (a) mission selection, (b) research & technology 

objectives definition, and (c) baseline vehicle(s) characterization. The above outlined process arrives at a 

justification package able to characterize and defend the suggested baseline hypersonic vehicle design

selected.

8.1 DESIGN LESSONS LEARNED

In addition to the primary flight vehicle system recommendations communicated in Chapter 7.3, several 

design lessons have been learned through the course of this project which are worthy of note.

 Increasing cruise time from 0 to 30 minutes increases vehicle size and technology requirements 

(30 minutes cruise hypersonic demonstrator appears possible).

 LH2 fuel allows for a larger technical solution space relative to the kerosene option.

 Air-launch from the B-52 is limited due to under-wing geometry (planform) constraints rather 

than under-wing load limitations.

 Selection of scramjet fuel is not driven by technical feasibility of the demonstrator test-bed, but by 

requirements specified for the operational aircraft (range and payload requirements, 

infrastructure).

 Air-launch and expendable booster launch are both viable options with LH2.

 Launch arrangement should be based on flight rate requirement and associated operational cost.

 Off-the-shelf accelerator rocket motors are available, thereby reducing overall development 

program costs and initial program risks.

 Landing constraints, driven by the abort mission, tend to constrain the solution space.
61

 Dual fuel option marginally decreases size of vehicle, relative to the 30 minutes LH2 variant.

 A reduced cruise time Mach 8 mission could represent an off-design point for the Mach 6 

demonstrator (Merlin thrust class rocket is no longer required).

 A 30 minutes turning cruise flight has minimum effect on vehicle size due to operation at higher 

L/D at large turning radius and low load factor.

The study results generated within the available time frame conclude with the recommendations outlined 

in Chapter 8.2. It is felt that the recommendations require attention before a selection of confidence can 

be made for a baseline vehicle and the resulting moving forward with the design.

8.2 RECOMMENDATIONS

The parametric sizing results clearly indicate that the design of a hypersonic endurance demonstrator is 

far from trivial. Although the parametric sizing (PS) phase is considered not complete at this point, the 

results generated allow the decision-makers (manager, synthesis specialist, and technologist) to plan 

ahead and proceed with some degree of confidence. Clearly, more research is required for selecting a 

baseline hypersonic demonstrator concept.

Remaining Top-Level Questions

The remaining top-level questions at (a) synthesis level, (b) managerial level, and (c) technology level

are:

A. Synthesis Level

• What future scenarios and operational systems warrant hydrogen and/or kerosene scramjet 

research?

• Is the flight vehicle capability targeted satisfying the program objectives in terms of time and 

resources available?

• What is the required demonstrator capability able to accommodate a wide range of test conditions 

contributing to general hypersonic research?

• What is the required demonstrator capability able to accommodate a wide range of test conditions 

contributing to specific hypersonic research?

B. Managerial Level

• What is the sensitivity characterizing expendable booster cost and air-launch cost?

• Does a hydrogen, Mach 8 and 30 minutes demonstrator warrant the increased technology & cost 

requirement relative to the Mach 6 and 30 minutes, Mach 8 and 10 minutes vehicle?

• What effect will a RBCC, such as an ejector ramjet, have on the vehicle and its technology 

requirements?

• What are the maximum allowable down-range and cross-range requirements?
62

C. Technology Level

• Are primary disciplinary and multi-disciplinary technology parameters sufficiently represented 

throughout the design life-cycle?

• Can operational vehicle and demonstrator vehicle (a) technology, (b) operational mission & flight 

test program, and (c) vehicle utilization be predicted?

• What technology breakthroughs are necessary or desirable for each of the final baseline 

demonstrator vehicle types?

Recommendations for Future Work

With these questions in mind, the following steps are necessary to complete the conceptual design study:

A. Complete requirements and objectives research

• Expand requirements & objectives definition activity for proposed operational hypersonic 

applications; particular interest should be given to refining endurance and fuel selection.

• Expand survey of hypersonic technologies (ground & in-flight) which support near-term

experimental validation and verification towards an operational system.

• Expand survey of hypersonic technologies (ground & in-flight) which require longer-term

experimental validation and verification towards an operational system.

B. Expand demonstrator parametric sizing (PS) study

• Expand kerosene trades with wing-body combinations with various abort, emergency, or failure 

scenarios.

• Expand fuel trades to include natural gas as a middle ground between the performance of 

hydrogen and available infrastructure of kerosene.

• Explore impact of RBCC and/or TBCC demonstration capability to support future launch and 

point-to-point vehicle programs.

• Compare demonstrator vehicle solution spaces based on development and operational cost 

metrics.

• Explore the solution space of operational vehicle concepts (vehicles with payload) while the 

technology demonstrator will be designed to validate a suite of technologies to directly satisfy 

those operational missions.

C. Complete conceptual design: configuration layout (CL) and configuration evaluation (CE) stages

• CL and CE will validate and refine the initial operational and technical assumptions made during 

parametric sizing step.

• Development of baseline demonstrator vehicle (conceptual design & safety assessment).

• Identification of associated operational vehicle(s) (conceptual design & safety assessment).
63

REFERENCES

1. Pirrello, C.J. and Czysz, P.A., “Hypersonic Research Facilities Study: Research Requirements 

Analysis and Facility Potential,” Volume 4, Part 3, Phase III Final Studies, National Aeronautics 

and Space Administration, NASA CR-114329, 1970.

2. Chudoba, B., “Stability and Control of Conventional and Unconventional Aircraft Configurations, A 

Generic Approach,” 1

st Edition, Books on Demand, 2001.

3. Pirrello, C.J. and Czysz, P.A., “Hypersonic Research Facilities Study: Summary,” Volume 1, 

National Aeronautics and Space Administration, NASA-CR 114322, 1970.

4. Brewer, G.D., “Hydrogen Aircraft Technology,” CRC Press, Boca Raton, Florida, 1991.

5. Coleman, G., “Aircraft Conceptual Design: An Adaptable Parametric Sizing Methodology,” Ph.D. 

Dissertation, Department of Aeronautical Engineering, The University of Texas at Arlington, 

Arlington, Texas, 2010.

6. Czysz, P.A., “Hypersonic Convergences,” Volume 1, Air Force Research Laboratory, WrightPatterson Air Force Base, AFRL-VA-WP-TR-2004-3114, 2004.

7. Chudoba B. and Coleman, G., “Development of Advanced Commercial Transport Configurations (N

+ 3 and Beyond), Through the Assessment of Past, Present, and Future, Technologies,” Presentation 

given to the Aeronautics Systems Analysis Branch at NASA Langley Research Center, Hampton, 

Virginia, 24-25 June, 2009.

8. Ingenito, A., Gulli, S. and Bruno, C., “Sizing of a TBCC Powered Hypersonic Vehicle,” Journal of 

Aircraft, TBD [Submitted].

9. Pirrello, C.J. and Czysz, P.A., “Hypersonic Research Facilities Study: Flight Vehicle Synthesis,”

Volume 2, Part 2, Phase I Preliminary Studies, National Aeronautics and Space Administration, 

NASA CR-114324, 1970.

10. Pike, J., “Minimum Drag Bodies of a Given Length and Base using Newtonian Theory,” AIAA 

Journal, Vol. 15, No. 6, 1977, pp. 769770.

11. Warneke, C.H. and Kinroth, G.D., “Lifting Reentry Vehicle Preliminary Designs for FDL-7MC and 

FDL-5MA Configurations, Volume IV: Configuration Evolution, Design Approach Investigation and 

Supporting Analyses,” Technical Report AFFDL-TR-68-97, Air Force Flight Dynamics Laboratory, 

Wright-Patterson Air Force Base, Ohio, 1969.

12. Heiser, W.H. and Pratt, D.T., “Hypersonic Airbreathing Propulsion,” AIAA Educational Series, 

Washington, DC, 1994.

13. Bradford, J.E., “A Technique for Rapid Prediction of Aft-body Nozzle Performance for Hypersonic 

Launch Vehicle Design,” Ph.D. Dissertation, School of Aeronautical Engineering, Georgia Institute 

of Technology, Atlanta, Georgia, 2001.
64

14. Flornes, B.J. and Escher, J.D., “A Study of Composite Propulsion Systems for Advanced Launch 

Vehicle Applications,” Volume 2 Main Technical Report, Report 25, The Marquardt Corporation, 

Van Nuys, California, 1967.

15. Combs, H.G., “Configuration Development Study of the X-24C Hypersonic Research Airplane, 

Executive Summary,” NASA CR-145274, NASA Langley Research Center, 1977.

16. Klijin, M.S., et al, “Selection of Carrier Aircraft and a Launch Method for Air Launching Space 

Vehicles,” AIAA-2008-7835, 2008.

17. Isakowitz, S.J., Hopkins, J.B. and Hopkins, J.P., “International Reference Guide to Space Launch 

Systems,” 4

th Edition, American Institute of Aeronautics and Astronautics, Reston, Virginia, 2004.

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accessed June 2010.
65

Appendices

Appendix A. Disciplinary Methods Library

1. Geometry

Table A-1. Hypersonic Cruiser Planform Description Method

Method overview

Discipline

Geometry 

Design phase

Sizing

Method title

Hypersonic cruiser 

planform description 

Categorization 

Semi-empirical

Author

Czysz/Coleman

Reference: (Modified from) Czysz, P.A., “Hypersonic Convergences,” Volume 1, Air Force Research 

Laboratory, Wright-Patterson Air Force Base, AFRL-VA-WP-TR-2004-3114, 2004.

Brief description

Planform description of hypersonic cruisers with delta-wing planform with a given spatular ratio (c/s).

Assumptions

Simplified geometry

Applicability

General glider and air breather configuration

Execution of method

Input: Spln, c/s, LE

Analysis description

c s

S

l

p LE

1 /

tan ln







LE s  l / tan

 c  c /ss

b  2s  c

Output: l, s, c, b

Experience

Triangular Configuration Spatular Configuration
66

Accuracy

Dependent on assumed values

General comments

Use the figure provided for guidance for K0

Table A-2. Hypersonic Air-Breather Volume and Wetted Area Estimation

Method overview

Discipline

Geometry

Design phase

Sizing

Method title

Hypersonic air-breather 

volume and wetted area

Categorization

Semi-empirical

Author

Czysz/Coleman

Reference: (Modified from) Czysz, P.A., “Hypersonic Convergences,” Volume 1, Air Force Research 

Laboratory, Wright Patterson Air Force Base, AFRL-VA-WP-TR-2004-3114, 2004.

Brief description

Volume and wetted area are computed based on planform parameters, vehicle slenderness () ,and engine 

design cowl location. 

Assumptions

Simplified geometry

Applicability

General glider and air breather configuration

Execution of method

Input: Spln, t, c/s, Lw, LE, , 

s ,Lc

/Lw, hc/Lc, 1, ht,/Liso, Lcomb

Analysis description

Compute underside geometry:

  

 

 

 

 

Iterate ht until Liso convergence:

 

 

Call one-dimensional stream thrust analysis 

at cruise condition to compute contraction 

ratios:

 

 

 

 

 

 

 

 

 

 

 

 

 

 

 

Compute minimum  (flat top):

 

  ( 

 

 

 )(

 

 

 

 

 

 )

 

 

Compute height of upper surface (flat top)

 

 

 

 ( 

 

 

 )

Compute wetted area per planform area:

 

 

 

 

 

 

 

  

 √ 

 

 √ 

 

 

 

Estimate frontal area and capture area:

 

 

 

 

  

Output: , , , , , , , , , 
67

Experience

Accuracy

Dependent on assumed values

General comments

Wetted are based on MAC pointed-nose configuration, 

expanded by spatula.

Further description:

The parameterization of the volume (Figure A-1) and wetted area (Figure A-2) for a hypersonic air-breather follows.

Fig. A-1. Volume parameterization of a hypersonic air-breather.

Fig. A-2. Wetted area description of a hypersonic air-breather.

2. Performance

Trajectory thrust requirement and fuel requirement

Lw Lc htop

hc ht Liso Lcomb scowl

c

sout
68

Table A-3. Hypersonic Cruiser Trajectory Determination Method

Method overview

Discipline

Propulsion 

Design phase

Sizing

Method title

Hypersonic cruiser trajectory 

Categorization 

Numerical

Author

HYFAC

Reference: Czysz, P.A., “Hypersonic Convergence,” Air Force Research Laboratory, Wright Patterson Air Force 

Base, AFRL-VA-WP-TR-2004-3114, 2004.

Brief description

From an assumed segmented trajectory, an energy integration is performed to compute the required fuel weight. 

From the computed drag and propulsion-system performance data, the thrust that is required at sea level is computed

at each step. The largest thrust requirement is utilized for the acceleration.

Assumptions

Step climb up to transonic acceleration.

Constant altitude transonic acceleration.

Constant dynamic pressure climb to cruise altitude.

Cruise-climb (constant CL) and max L/D descent.

Applicability

Air-breathing hypersonic or supersonic cruisers or firststage launchers.

Execution of method

Input 

Trajectory, CD0, L’, T/Tsl , nmax, Isp at each step

Analysis description

At each point, the following equation is utilized to compute the total fuel burn and thrust requirement (see the 

additional information following this table in the further description section).

Each segment is then integrated based on constant, altitude, velocity, or dynamic pressure.

The total fuel fraction is then summed for weight and volume convergence.

The largest thrust-to-weight ratio is used for engine weight estimation.

Output: WR, (T/W)TO

Experience

Accuracy

Depends on aero and propulsion system accuracy

General comments

This type of trajectory tends to yield the lowest thrust requirement because 

of the constant altitude transonic acceleration. Transonic acceleration is 

typically what sets the thrust requirement for the vehicle.
69

Further description:

Assumed trajectory:

(1) Climb to 10,000 ft, (2) constant altitude acceleration to 0.8 M, (3) constant Mach climb to 12,000 ft, (4) 

constant altitude acceleration through the transonic region to maximum dynamic pressure, (5) constant dynamic 

pressure climb to cruise altitude, (6) cruise-climb to altitude, (7) maximum L/D descent, and (8) landing (see 

Figure A-3 below).

Fig. A-3. Assumed trajectory of the hypersonic cruiser.

At each integration step () (each segment of the trajectory is broken down by predefined step size), compute

the following:

Gravity relief:

 

 

 

 

 

 

Aerodynamic efficiency:

 

 

 

 

 

 

 ̅

 

 

 

 

 

 

Acceleration available:
70

 (

 

 

)

 

 

 

 

Energy at step i:

 

 

 

 

 

 

 

Derivatives:

 ̇

 

 

 

 

 ̇

 

  

 

 

 

 

Then,

 

 

 

 

 

 

 

 
71

3. Weight and Balance

Empty Weight and Volume Formulation

Table A-4. Convergence Empty Weight Estimation Method

Method overview

Discipline

Weight 

estimation

Design phase

Parametric sizing

Method title

Convergence empty-weight 

estimation

Categorization

Empirical

Author

Coleman/

Czysz

Reference: Dissertation

Brief description

A modification of the hypersonic convergence method for estimating the converged empty weight based on volume 

and mass. This method has been modified to allow for the incorporation of additional methods for structural, 

propulsion, systems, and operational item weights beyond what are presented in hypersonic convergence.

Assumptions

Wing area is not constant

Applicability

Any aircraft or launcher configuration. Applicability 

depends on the methods used for the structural, 

propulsion, and systems weight.

Execution of method

Input: WR, T/W, Wpay, Wcrew, Vpay, Vcrew

Analysis description

Solve the following system for Spln and OWE:

Weight budget: 

   

 

pay

sys TW

str sys oper TW pay crw W

f T W WR E

W C W T W WR E W W

OWE

a



 

   





/ /

/ /

1 max

1

max



Volume budget:

 

k T W  WR

S k k V V V

OWE

ve

WR

pln vv vs fix pay crew

fuel max

1

1.5

/

1



     









Use the additional methods for Wstr, Wsys,, fsys, Woper and ETW

Output: OEW, TOGW, OWE, Spln

Experience
72

Accuracy

Depends upon additional methods

Time to calculate

Depends on structural weight 

estimation

General comments

Works well for any configuration. Is at the 

heart of AVDsizing. The convergence logic 

will take the output and feed it back 

through the geometry trajectory and 

constraints until convergence

Further description

Additional weight relationships:

 ton/ton

 ton

 ton/person

 ton/person

 

 

 

( ) kg thrust/kg weight

Additional volumetric relationships:

 kg/m3

 ( ) m

3

/person

 m

3

/person

 m

3

/m3

 ( ) m

3

/person

 m

3

/person

 m

3

/m3

 m

3

 m

3

/person

 ; m3/ton thrust

 (

 

 )
73

4. Structural weight

Table A-5. Structural Index Estimation Method

Method overview

Discipline

Structure 

Design phase

Sizing

Method title

Structural index 

Categorization

Empirical

Author

Czysz

Reference: Czysz, P.A., “Hypersonic Convergence,” Volume 1, Air Force Research Laboratory, Wright Patterson 

Air Force Base, AFRL-VA-WP-TR-2004-3114, 2004.

Brief description

Empirical methods for structural weight derived as a function of planform area, OEW, and the shape factor Kstr. Kstr

represents the structural efficiency that is required for given hypersonic vehicle slenderness ().

Assumptions

Blended-body or wing-body hypersonic cruiser or launch 

vehicle

Integrated thermal projection and structural sandwich

Applicability

Both passive and actively cooled structures

Hypersonic cruisers and launch vehicles

Execution of method

Input: , Spln, Swet, OEW

Analysis description

Compute structural weight and structural index required for a given , Spln, Swet, and OEW:

Wstr I

str

 Swet Kstr S OEW 0.138

pln

Output: Wstr, Istr

Experience

Accuracy

Has worked well for a variety of hypersonic cruisers 

projects at MAC

Proves valid for the Sanger II

General comments

Due to the transition from hot to cold structure, the structural index does 

not need to be greater than 18 kg/m2. The rule of thumb at MAC is 21 

kg/m2for demonstrators (with cheap and heavier materials) and 18 kg/m2 

for operational vehicles.
74

Further description:

The structural index is selected from Figure A-4 based on the predicted maximum.

Fig. A-4. Structural index prediction description.
75

Appendix B. Trade-Study Assumptions and Database

B-1 TRADE 001: AIR-LAUNCHED, MACH 6, HYDROGEN FUEL

Mission summary: Table B-1.1 summarizes the mission constants for this trade study.

Table B-1.1. Air-Launched, Mach 6, Hydrogen Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection LH, LOX

Fuel density 74.63kg/m3(4.65 lb/ft3)

Oxidizer density 1287 kg/m3(803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7)[11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[5]

Propulsion 

Rocket engine

Using the constants that are given in Table B-1.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-1.2. Summary of Rocket Accelerators Explored

RL-10A-5, P&W Vulcain, SEP

ETW – Engine thrust-to-weight ratio 46.12 61.36

ISPvac – Vacuum ISP 373.0 s 440

Tvac – Vacuum thrust 64.7 kN (14,500 lb) 1015 kN (lb)

 – Nozzle-expansion ratio 4.0 45.0

Pc – combustion-chamber pressure 38.6 atm 102.0 atm

O/F – oxidizer-to-fuel ratio 6.0 5.6
76

Dual-Mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that were assumed in the stream thrust 

analysis for this trade are summarized below in Table B-1.3.

Table B-1.3. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.65

hc/lc 0.088

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
77

Weight, volume, and balance

Hypersonic convergence weight and volume formulation[6]

Table B-1.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 46.12 or 61.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000133 or 0.000088 m3/kg

Kve_DMR – dual mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 65 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design point database file: Table B-1.5 summarizes the design-point data collected by AVDsizing.

Table B-1.5. Trade 001, Air-Launched, Mach 6, Hydrogen-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 6

(NA) D_MVIHN 0

(NA) D_WR 1
78

Table B.1.5. Continued

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 4520

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic 

Acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 74.63

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.125

Planform area SPLN 103.30815

Ratio of wetted area to planform area AKW 2.36792

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.1612

Table B.1.5. Continued
79

Total vehicle length AL_TOTAL 18.00049

Span BPLN 9.56531

Height above centerline BASE_HEIGHT 2.83738

Spatular width CSPAT 1.91306

Capture area ACAP 5.62205

Length of external compression ALC 11.70032

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.43662

Width of cowl at inlet W3 3.34968

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00858

Height of cowl to length of isolator ALI_L 0.0858

Height at combustor exit to nozzle length H_LN 0.05581

Stall/approach 

performance

Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 82389762.44

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01106

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.40469

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.54888

Trajectory summary Climb range CLRANGE 180394.6293

Cruise range CRRANGE 3236961.266

Decent range DERANGE 853093.8265

Total range RANGE_TOTAL 4270449.722

Climb time T_CLIMB 2.68974

Cruise time T_CRUISE 30

Descent time T_DESCENT 22.32068

Total flight time T_FLT 55.01042

Mach number at max rocket T/W AMACH_TJ 1.38

Maximum rocket T/W TW_TJ_TM 1.78764

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 0.847

Maximum capture area of scramjet (NA) AC_W_MAX 5.62205

(NA) AMACH_ACAP 0.81

Table B.1.5. Continued
80

(NA) QBAR_ACAP 10171.29338

(NA) ALT_ACAP 11000

(NA) CFN_ACAP 0

Average cruise L/D LD_CRUISE 2.46135

Average cruise Isp ISP_CRUISE 2613.15135

Minimum acceleration during scramjet mode ANMIN 0.41159

Oxidizer-to-fuel ratio OF_TRAJ 0.86403

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 22136.37226

Propellant weight WPPL 10046.82982

Total fuel weight WFUEL 5357.2487

Fuel 1 weight WFUEL1 5357.2487

Weight of oxidizer Wox 4689.58113

Manufacturer’s zero-fuel weight AMZFW 12089.54244

Operating weight empty OWE 12089.54244

Operating empty weight OEW 12089.54244

Weight margin WMARGIN 1099.04931

Operational items weight WOPER 0

Systems weight WSYS 3834.32679

Structural weight WSTR 4735.68883

Rocket propulsion system weight WP_TJ 858.02115

Ram/scramjet weight WP_SC 1562.45636

Total fuel fraction FF_TOTAL 0.45386

Structural weight fraction WSTR_TOGW 0.21393

Total weight ratio WR 1.83103

Total volume V_TOTAL 131.25381

Fixed systems volume V_FIX 5

Total systems volume V_SYS 10.25015

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 75.42793

Total fuel volume V_FUELI1 71.78412

Total oxidizer volume V_OX 3.64381

Total propellant density PPL_DEN 133.19775

Total fuel density FUEL_DEN 74.63

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 5.26307

Table B.1.5. Concluded
81

Total ram/scramjet volume VENG_SC 14.06211

Total engine volume VENG 19.32517

Total void volume VVOID 26.25076

Convergence check Operating-weight-empty weight budget OWE_W 12089.54244

Operating-weight-empty volume budget OWE_V 12089.51592

Planform area SPLN 103.30815

Capture-area required AC_RE 5.62205

Capture-area available AC_AV 5.62205

Structural index AISTR 19.35895

Propulsion index AIP 160.2794

T/W rocket TW_TJ_MAX 1.78764

T/W scramjet TW_SC_MAX 0.847

Wing loading WS 214.27517
82

B-2 TRADE 002: AIR-LAUNCHED, MACH 8, HYDROGEN FUEL

Mission summary: Table B-2.1 summarizes the mission constants for this trade study.

Table B-2.1. Air-Launched, Mach 8, Hydrogen-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection LH, LOX

 Fuel density 74.63kg/m3(4.65 lb/ft3)

 Oxidizer density 1287 kg/m3(803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7) [11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[5].

Propulsion 

Rocket engine

Using the constants that are given in Table B-2.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-2.2. Summary of Rocket Accelerators Explored

RL-10A-5, P&W Vulcain, SEP

ETW – Engine thrust-to-weight ratio 46.12 61.36

ISPvac – Vacuum ISP 373.0 s 440

Tvac – Vacuum thrust 64.7 kN (14,500 lb) 1015 kN (lb)

 – Nozzle-expansion ratio 4.0 45.0

Pc – combustion-chamber pressure 38.6 atm 102.0 atm

O/F – oxidizerto-fuel ratio 6.0 5.6
83

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized in Table B-2.3.

Table B-2.3. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.65

hc/lc 0.088

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
84

Weight, volume, and balance

The hypersonic convergence weight and volume formulation is summarized in Table B-2.4.

Table B-2.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 46.12 or 61.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000133 or 0.000088 m3/kg

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrw – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 65 percent of the body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-2.5 summarizes the design-point data that were collected by AVDsizing.

Table B-2.5. Trade 002, Air-Launched, Mach 8, Hydrogen-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 8

(NA) D_MVIHN 0

(NA) D_WR 1
85

Table B.2.5. Continued

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 4520

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to 

transonicacceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 74.63

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.15

Planform area SPLN 161.1835

Ratio of wetted area to planform area AKW 2.42036

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.19721

Table B.2.5. Continued
86

Total vehicle length AL_TOTAL 22.4842

Span BPLN 11.94791

Height above centerline BASE_HEIGHT 3.97159

Spatular width CSPAT 2.38958

Capture area ACAP 8.77163

Length of external compression ALC 14.61473

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.79446

Width of cowl at inlet W3 4.18405

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00617

Height of cowl to length of isolator ALI_L 0.06173

Height at combustor exit to nozzle length H_LN 0.03318

Stall/approach 

performance

Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 102912054.3

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01196

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.17812

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.4934

Trajectory summary Climb range CLRANGE 646761.7275

Cruise range CRRANGE 4353951.878

Decent range DERANGE 1299474.693

Total range RANGE_TOTAL 6300188.299

Climb time T_CLIMB 6.3516

Cruise time T_CRUISE 30

Descent time T_DESCENT 24.89358

Total flight time T_FLT 61.24517

Mach number at max rocket T/W AMACH_TJ 0.93

Maximum rocket T/W TW_TJ_TM 2.5297

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 1.20697

Maximum capture area of scramjet (NA) AC_W_MAX 8.77163

Table B.2.5. Continued

(NA) AMACH_ACAP 0.81
87

(NA) QBAR_ACAP 10171.29338

(NA) ALT_ACAP 11000

(NA) CFN_ACAP 0

Average cruise L/D LD_CRUISE 2.30527

Average cruise Isp ISP_CRUISE 2246.44966

Minimum acceleration during scramjet mode ANMIN 0.2189

Oxidizer-to-fuel ratio OF_TRAJ 0.57576

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 40900.46186

Propellant weight WPPL 20821.12118

Total fuel weight WFUEL 13149.3266

Fuel 1 weight WFUEL1 13149.3266

Weight of oxidizer Wox 7671.79458

Manufacturer’s zero-fuel weight AMZFW 20079.34068

Operating weight empty OWE 20079.34068

Operating empty weight OEW 20079.34068

Weight margin WMARGIN 1825.39461

Operational items weight WOPER 0

Systems weight WSYS 5112.69451

Structural weight WSTR 7341.23756

Rocket propulsion system weight WP_TJ 1686.21222

Ram/scramjet weight WP_SC 4113.80178

Total fuel fraction FF_TOTAL 0.50907

Structural weight fraction WSTR_TOGW 0.17949

Total weight ratio WR 2.03694

Total volume V_TOTAL 306.95318

Fixed systems volume V_FIX 5

Total systems volume V_SYS 17.27813

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 182.15457

Total fuel volume V_FUELI1 176.19358

Total oxidizer volume V_OX 5.96099

Total propellant density PPL_DEN 114.30469

Total fuel density FUEL_DEN 74.63

Total oxidizer density OX_DEN 1287

Table B.2.5. Concluded

Total rocket volume VENG_TJ 9.10501
88

Total ram/scramjet volume VENG_SC 37.02422

Total engine volume VENG 46.12922

Total void volume VVOID 61.39064

Convergence check Operating-weight-empty weight budget OWE_W 20079.34068

Operating-weight-empty volume budget OWE_V 20079.39617

Planform area SPLN 161.1835

Capture-area required AC_RE 8.77163

Capture-area available AC_AV 8.77163

Structural index AISTR 18.81777

Propulsion index AIP 110.23243

T/W rocket TW_TJ_MAX 2.5297

T/W scramjet TW_SC_MAX 1.20697

Wing loading WS 253.75092
89

B-3 TRADE 003: AIR-LAUNCHED, MACH 6, KEROSENE FUEL

Mission summary: Table B-3.1 summarizes the mission constants for this trade study.

Table B-3.1. Air-Launched, Mach 6, Kerosene-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1, LOX

Fuel density 820.0 kg/m3(51.2 lb/ft3)

Oxidizer density 1287 kg/m3( 803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7) [11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6].

Propulsion 

Rocket engine

Using the constants that are given in Table B-3.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-3.2. Summary of Rocket Accelerators Explored

Merlin , Space X

ETW – Engine thrust-to-weight ratio 96.0

ISPvac – Vacuum ISP 304.0 s

Tvac – Vacuum thrust 512.0 kN (115.0 lb)

 – Nozzle-expansion ratio 14.0

Pc – combustion-chamber pressure 60.69 atm

O/F – oxidizer-to-fuel ratio 2.17
90

Dua- mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that were assumed in the stream thrust 

analysis for this trade are summarized in Table B-3.3.

Table B-3.3. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 43380.0

f (stoichiometric) 0.0680

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.40

Cev 0.99

Cpe 1.51

Cea 0.98

c 1.362

e 1.28

Geometric constants

lc/lw 0.50

hc/lc 0.067

hiso/liso 0.1

Lcomb 2.0 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
91

Weight, volume, and balance

Hypersonic convergence weight and volume formulation.

[6]

Table B-3.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 96.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000088 m

3

/kg

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 50 percent of body 

length to keep the trim drag manageable.

[6]

.

Cost

No cost model was utilized.

Design-point database file: Table B-3.5 summarizes the design-point data that were collected by AVDsizing.

Table B-3.5. Trade 003, Air-Launched, Mach 6, Kerosene-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 20

Design Mach number D_MACH 6

(NA) D_MVIHN 0

(NA) D_WR 1

Table B.3.5. Continued
92

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 4572

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic 

acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.07

Planform area SPLN 58.38376

Ratio of wetted area to planform area AKW 2.30811

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.09352

Table B.3.5. Continued

Total vehicle length AL_TOTAL 13.53204
93

Span BPLN 7.19081

Height above centerline BASE_HEIGHT 1.41398

Spatular width CSPAT 1.43816

Capture area ACAP 2.19091

Length of external compression ALC 6.76602

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 0.83076

Width of cowl at inlet W3 2.26892

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00803

Height of cowl to length of isolator ALI_L 0.08034

Height at combustor exit to nozzle length H_LN 0.04212

Stall/approach 

performance

Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 61937258.73

Mach number for approach AMACH 0.1884

Landing gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for 

flaps DE_LG 0

Form drag CD0 0.00953

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.90312

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.45864

Trajectory summary Climb range CLRANGE 304462.8653

Cruise range CRRANGE 2158897.648

Decent range DERANGE 1078335.405

Total range RANGE_TOTAL 3541695.918

Climb time T_CLIMB 3.92479

Cruise time T_CRUISE 20

Descent time T_DESCENT 26.46614

Total flight time T_FLT 50.39093

Mach number at max rocket T/W AMACH_TJ 0.81

Maximum rocket T/W TW_TJ_TM 3.34752

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 0.54832

Table B.3.5. Continued

Maximum capture area of scramjet (NA) AC_W_MAX 2.19091

(NA) AMACH_ACAP 0.81

(NA) QBAR_ACAP 10171.29338
94

(NA) ALT_ACAP 11000

(NA) CFN_ACAP 0

Average cruise L/D LD_CRUISE 3.79014

Average cruise Isp ISP_CRUISE 942.59695

Minimum acceleration during scramjet mode ANMIN 0.19489

Oxidizer-to-fuel ratio OF_TRAJ 0.47237

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 14191.36724

Propellant weight WPPL 7715.23442

Total fuel weight WFUEL 5217.89052

Fuel 1 weight WFUEL1 5217.89052

Weight of oxidizer Wox 2497.3439

Manufacturer’s zero-fuel weight AMZFW 6476.13282

Operating weight empty OWE 6476.13282

Operating empty weight OEW 6476.13282

Weight margin WMARGIN 588.73935

Operational items weight WOPER 0

Systems weight WSYS 2936.18125

Structural weight WSTR 1759.09396

Rocket propulsion system weight WP_TJ 543.66565

Ram/scramjet weight WP_SC 648.45261

Total fuel fraction FF_TOTAL 0.54366

Structural weight fraction WSTR_TOGW 0.12396

Total weight ratio WR 2.19133

Total volume V_TOTAL 31.22742

Fixed systems volume V_FIX 5

Total systems volume V_SYS 6.2491

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 8.30372

Total fuel volume V_FUELI1 6.36328

Total oxidizer volume V_OX 1.94044

Total propellant density PPL_DEN 929.12997

Table B.3.5. Concluded

Total fuel density FUEL_DEN 820

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 4.59289

Total ram/scramjet volume VENG_SC 5.83607
95

Total engine volume VENG 10.42896

Total void volume VVOID 6.24548

Convergence check Operating-weight-empty weight budget OWE_W 6476.13282

Operating-weight-empty volume budget OWE_V 6476.18811

Planform area SPLN 58.38376

Capture-area required AC_RE 2.19091

Capture-area available AC_AV 2.19091

Structural index AISTR 13.05393

Propulsion index AIP 779.90749

T/W rocket TW_TJ_MAX 3.67772

T/W scramjet TW_SC_MAX 0.54832

Wing loading WS 243.07046
96

B-4 TRADE 004: AIR-LAUNCHED, MACH 8, KEROSENE FUEL

Mission summary: Table B-4.1 summarizes the mission constants for this trade study.

Table B-4.1. Air-Launched, Mach 6, Kerosene-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1, LOX

 Fuel density 820.0 kg/m3(51.2 lb/ft3)

 Oxidizer density 1287 kg/m3(803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7) [11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion 

Rocket engine

Using the constants that are given in Table B-4.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-4.2. Summary of Rocket Accelerators Explored

Merlin, Space X

ETW – Engine thrust-to-weight ratio 96.0

ISPvac – Vacuum ISP 304.0 s

Tvac – Vacuum thrust 512.0 kN (115.0 lb)

 – Nozzle-expansion ratio 14.0

Pc – combustion-chamber pressure 60.69 atm

O/F – oxidizer-to-fuel ratio 2.17
97

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that were assumed in the stream thrust 

analysis for this trade are summarized in Table B-4.3.

Table B-4.3. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 43380.0

f (stoichiometric) 0.0680

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.40

Cev 0.99

Cpe 1.51

Cea 0.98

c 1.362

e 1.28

Geometric constants

lc/lw 0.50

hc/lc 0.067

hiso/liso 0.1

Lcomb 2.0 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
98

Weight, volume, and balance

Hypersonic convergence weight and volume formulation.

[6]

Table B-4.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 96.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000088 m

3

/kg

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed wascomputed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 50 percent of the body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-Point database file: Table B-4.5 summarizes the design-point data that were collected by AVDsizing.

Table B-4.5. Trade 004, Air-Launched, Mach 8, Kerosene-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 4.5

Design Mach number D_MACH 8

(NA) D_MVIHN 0

(NA) D_WR 1
99

Table B.4.5. Continued

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.0675

Planform area SPLN 76.73385

Ratio of wetted area to planform area AKW 2.30252

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.08529

Total vehicle length AL_TOTAL 15.51353
100

Table B.4.5. Continued

Span BPLN 8.24376

Height above centerline BASE_HEIGHT 1.6733

Spatular width CSPAT 1.64875

Capture area ACAP 3.44037

Length of external compression ALC 8.84271

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.08575

Width of cowl at inlet W3 2.7345

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00692

Height of cowl to length of isolator ALI_L 0.06922

Height at combustor exit to nozzle length H_LN 0.04097

Stall/approach 

performance Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 71006721

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.00945

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.92578

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.44328

Trajectory summary Climb range CLRANGE 788626.134

Cruise range CRRANGE 652138.1791

Decent range DERANGE 1826877.811

Total range RANGE_TOTAL 3267642.124

Climb time T_CLIMB 7.34318

Cruise time T_CRUISE 4.5

Descent time T_DESCENT 32.73813

Total flight time T_FLT 44.58131

Mach number at max rocket T/W AMACH_TJ 1.2

Maximum rocket T/W TW_TJ_TM 2.74501

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 0.9407

Maximum capture area of scramjet (NA) AC_W_MAX 3.44037

(NA) AMACH_ACAP 0.81
101

Table B.4.5. Continued

(NA) QBAR_ACAP 10171.29338

(NA) ALT_ACAP 11000

(NA) CFN_ACAP 0

Average cruise L/D LD_CRUISE 3.38824

Average cruise Isp ISP_CRUISE 752.86825

Minimum acceleration during scramjet mode ANMIN 0.13563

Oxidizer-to-fuel ratio OF_TRAJ 0.45572

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 19013.29831

Propellant weight WPPL 10627.20547

Total fuel weight WFUEL 7271.3604

Fuel 1 weight WFUEL1 7271.3604

Weight of oxidizer Wox 3355.84507

Manufacturer’s zero-fuel weight AMZFW 8386.09284

Operating weight empty OWE 8386.09284

Operating empty weight OEW 8386.09284

Weight margin WMARGIN 762.37208

Operational items weight WOPER 0

Systems weight WSYS 3241.77485

Structural weight WSTR 2347.78996

Rocket propulsion system weight WP_TJ 543.66424

Ram/scramjet weight WP_SC 1490.49171

Total fuel fraction FF_TOTAL 0.55894

Structural weight fraction WSTR_TOGW 0.12348

Total weight ratio WR 2.26724

Total volume V_TOTAL 45.37162

Fixed systems volume V_FIX 5

Total systems volume V_SYS 6.81486

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 11.47501

Total fuel volume V_FUELI1 8.86751

Total oxidizer volume V_OX 2.60749

Total propellant density PPL_DEN 926.11757

Total fuel density FUEL_DEN 820

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 4.59288
102

Table B.4.5. Concluded

Total ram/scramjet volume VENG_SC 13.41443

Total engine volume VENG 18.0073

Total void volume VVOID 9.07432

Convergence check Operating-weight-empty weight budget OWE_W 8386.09284

Operating-weight-empty volume budget OWE_V 8386.12797

Planform area SPLN 76.73385

Capture-area required AC_RE 3.44037

Capture-area available AC_AV 3.44037

Structural index AISTR 13.2883

Propulsion index AIP 730.81375

T/W rocket TW_TJ_MAX 2.74501

T/W scramjet TW_SC_MAX 0.9407

Wing loading WS 247.7824
103

B-5 TRADE 005: AIR-LAUNCHED, MACH 6, DUAL FUEL

Mission summary: Table B-5.1 summarizes the mission constants for this trade study.

Table B-5.1. Air-Launched, Mach 6, Dual-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection RP-1, LH, LOX

 Kerosene-fuel density 820.0 kg/m3(51.2 lb/ft3)

 Hydrogen-fuel density 74.63kg/m3(4.65 lb/ft3)

 Oxidizer density 1287 kg/m3(803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7) [11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion

Rocket engine

Using the constants that are given in Table B-5.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-5.2. Summary of Rocket Accelerators Explored

Merlin, Space X

ETW – Engine thrust-to-weight ratio 96.0

ISPvac – Vacuum ISP 304.0 s

Tvac – Vacuum thrust 512.0 kN (115.0 lb)

 – Nozzle-expansion ratio 14.0

Pc – combustion-chamber pressure 60.69 atm

O/F – oxidizer-to-fuel ratio 2.17
104

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized in Table B-5.3.

Table B-5.3 Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.65

hc/lc 0.088

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
105

Weight, volume, and balance

Hypersonic convergence weight and volume formulation[6]

Table B-5.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 96.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000133 or 0.000088 m

3

/kg

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m

3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 65 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-5.5. summarizes the design-point data that were collected by AVDsizing.

Table B-5.5. Trade 005, Air-Launched, Mach 6, Dual-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 6

(NA) D_MVIHN 0

(NA) D_WR 1
106

Table B.5.5. Continued

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic 

acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 2

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.15

Planform area SPLN 84.74165

Ratio of wetted area to planform area AKW 2.41685

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.19496

Table B.5.5. Continued
107

Total vehicle length AL_TOTAL 16.30293

Span BPLN 8.66324

Height above centerline BASE_HEIGHT 2.86038

Spatular width CSPAT 1.73265

Capture area ACAP 4.61165

Length of external compression ALC 10.5969

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.30114

Width of cowl at inlet W3 3.03378

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00858

Height of cowl to length of isolator ALI_L 0.0858

Height at combustor exit to nozzle length H_LN 0.05694

Stall/approach 

performance

Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 74619857.12

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01195

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.17812

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.48596

Trajectory summary Climb range CLRANGE 213817.6566

Cruise range CRRANGE 3238239.258

Decent range DERANGE 773931.1643

Total range RANGE_TOTAL 4225988.079

Climb time T_CLIMB 2.90413

Cruise time T_CRUISE 30

Descent time T_DESCENT 21.20696

Total flight time T_FLT 54.11109

Mach number at max rocket T/W AMACH_TJ 2.62051

Maximum rocket T/W TW_TJ_TM 2.66202

(NA) Amach_sc 0

Maximum scramjet T/W TW_SC_TM 0.90956

Maximum capture area of scramjet (NA) AC_W_MAX 4.61165

(NA) Amach_acap 0.81

Table B.5.5. Continued
108

(NA) Qbar_acap 10171.29338

(NA) Alt_acap 11000

(NA) Cfn_acap 0

Average cruise L/D LD_CRUISE 2.02834

Average cruise Isp ISP_CRUISE 2618.93452

Minimum acceleration during scramjet mode ANMIN 0.27328

Oxidizer-to-fuel ratio OF_TRAJ 0.60278

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 19606.07443

Propellant weight WPPL 10125.86402

Total fuel weight WFUEL 6287.96357

Fuel 1 weight WFUEL1 1746.66415

Fuel 2 weight WFUEL2 4541.29942

Weight of oxidizer Wox 3837.90045

Manufacturer’s zero-fuel weight AMZFW 9480.21041

Operating weight empty OWE 9480.21041

Operating empty weight OEW 9480.21041

Weight margin WMARGIN 861.83731

Operational items weight WOPER 0

Systems weight WSYS 3416.83367

Structural weight WSTR 3171.79376

Rocket propulsion system weight WP_TJ 543.66375

Ram/scramjet weight WP_SC 1486.08192

Total fuel fraction FF_TOTAL 0.51647

Structural weight fraction WSTR_TOGW 0.16178

Total weight ratio WR 2.06811

Total volume V_TOTAL 117.01368

Fixed systems volume V_FIX 5

Total systems volume V_sys 9.68055

Total payload volume V_pay 0

Total crew volume V_crew 0

Total propellant volume V_ppl 65.96299

Fuel 1 volume V_fueli1 2.13008

Fuel 2 volume V_fueli2 60.85086

Total oxidizer volume V_ox 2.98205

Total propellant density Ppl_den 153.50827

Total fuel density Fuel_den 820

Table B.5.5. Concluded
109

Total oxidizer density Ox_den 1287

Total rocket volume Veng_tj 4.59287

Total ram/scramjet volume Veng_sc 13.37474

Total engine volume Veng 17.96761

Total void volume Vvoid 23.40274

Convergence check Operating-weight-empty weight budget OWE_W 9480.21041

Operating-weight-empty volume budget OWE_V 9480.18806

Planform area Spln 84.74165

Capture-area required Ac_re 4.61165

Capture-area available Ac_av 4.61165

Structural index Aistr 15.48668

Propulsion index Aip 143.72015

T/W rocket TW_TJ_MAX 2.66202

T/W scramjet TW_SC_MAX 0.90956

Wing loading Ws 231.3629
110

B-6 TRADE 006: AIR-LAUNCHED, MACH 8, DUAL FUEL

Mission summary: Table B-6.1 summarizes the mission constants for this trade study.

Table B.6.1. Air-Launched, Mach 8, Dual-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 11,000 m (36,000 ft)

Launch velocity 0.8 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection RP-1, LH, LOX

 Kerosene-fuel density 820.0 kg/m3(51.2 lb/ft3)

 Hydrogen-fuel density 74.63kg/m3(4.65 lb/ft3)

 Oxidizer density 1287 kg/m3(803.34 lb/ft3) 

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7)[11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion 

Rocket engine

From the constants that are given in Table B-6.2, the sizing process determines the number of rockets required,

which yields a minimum vehicle TOGW. Atmospheric losses were accounted for by using the P&W method.

Table B-6.2. Summary of Rocket Accelerators Explored

Merlin, Space X

ETW – Engine thrust-to-weight ratio 96.0

ISPvac – Vacuum ISP 304.0 s

Tvac – Vacuum thrust 512.0 kN (115.0 lb)

 – Nozzle-expansion ratio 14.0

Pc – combustion-chamber pressure 60.69 atm

O/F – oxidizer-to-fuel ratio 2.17
111

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized in Table B-6.3.

Table B-6.3. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.65

hc/lc 0.088

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Lw Lc htop

hc ht Liso Lcomp
112

Weight, volume, and balance

Hypersonic convergence weight and volume formulation[6]

Table B-6.4. Summary of Weight and Volume Constants

Weight

ETW_rkt – Rocket thrust-to-weight ratio (T/Weng) 96.0 kg/kg

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_rkt – rocket volume per kg thrust (VRKT/T) 0.000133 or 0.000088 m

3

/kg

Kve_DMR – dual mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m

3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrw – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 65 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-6.5 summarizes the design-point data that were collected by AVDsizing.

Table B-6.5. Trade 006, Air-Launched, Mach 8, Dual-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 25

Design Mach number D_MACH 8

(NA) D_MVIHN 0

(NA) D_WR 1
113

Table B.6.5. Continued

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic 

acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 11000

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Air-breathing transition Mach number AMACH_TJS 3

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 2

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.13

Planform area SPLN 114.37446

Ratio of wetted area to planform area AKW 2.37932

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.1696

Table B.6.5. Continued
114

Total vehicle length AL_TOTAL 18.94007

Span BPLN 10.06459

Height above centerline BASE_HEIGHT 3.06947

Spatular width CSPAT 2.01292

Capture area ACAP 6.22428

Length of external compression ALC 12.31105

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.51161

Width of cowl at inlet W3 3.52452

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00617

Height of cowl to length of isolator ALI_L 0.06173

Height at combustor exit to nozzle length H_LN 0.03403

Stall/approach 

performance

Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 86690295.96

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01124

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.35937

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.46921

Trajectory summary Climb range CLRANGE 519231.8808

Cruise range CRRANGE 3627110.709

Decent range DERANGE 1409570.9

Total range RANGE_TOTAL 5555913.49

Climb time T_CLIMB 5.29807

Cruise time T_CRUISE 25

Descent time T_DESCENT 27.39591

Total flight time T_FLT 57.69398

Mach number at max rocket T/W AMACH_TJ 1.23

Maximum rocket T/W TW_TJ_TM 2.03595

NA AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 1.15596

Maximum capture area of scramjet (NA) AC_W_MAX 6.22428

(NA) AMACH_ACAP 0.81

Table B.6.5. Continued
115

(NA) QBAR_ACAP 10171.29338

(NA) ALT_ACAP 11000

(NA) CFN_ACAP 0

Average cruise L/D LD_CRUISE 2.27915

Average cruise Isp ISP_CRUISE 2246.37844

Minimum acceleration during scramjet mode ANMIN 0.29657

Oxidizer-to-fuel ratio OF_TRAJ 0.603

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 25635.0344

Propellant weight WPPL 13666.90797

Total fuel weight WFUEL 8488.3351

Fuel 1 weight WFUEL1 2358.72421

Fuel 2 weight WFUEL2 6129.6109

Weight of oxidizer Wox 5178.57286

Manufacturer’s zero-fuel weight AMZFW 11968.12644

Operating weight empty OWE 11968.12644

Operating empty weight OEW 11968.12644

Weight margin WMARGIN 1088.01149

Operational items weight WOPER 0

Systems weight WSYS 3814.90023

Structural weight WSTR 4052.12375

Rocket propulsion system weight WP_TJ 543.66377

Ram/scramjet weight WP_SC 2469.42719

Total fuel fraction FF_TOTAL 0.53313

Structural weight fraction WSTR_TOGW 0.15807

Total weight ratio WR 2.14194

Total volume V_TOTAL 159.01458

Fixed systems volume V_FIX 5

Total systems volume V_SYS 11.36058

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 89.03358

Fuel 1 volume V_FUELI1 2.87649

Fuel 2 volume V_FUELI2 82.13334

Total oxidizer volume V_OX 4.02376

Total propellant density PPL_DEN 153.50284

Total fuel density FUEL_DEN 820

Table B.6.5. Concluded
116

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 4.59287

Total ram/scramjet volume VENG_SC 22.22484

Total engine volume VENG 26.81772

Total void volume VVOID 31.80292

Convergence check Operating-weight-empty weight budget OWE_W 11968.12644

Operating-weight-empty volume budget OWE_V 11968.10381

Planform area SPLN 114.37446

Capture-area required AC_RE 6.22428

Capture-area available AC_AV 6.22428

Structural index AISTR 14.89021

Propulsion index AIP 134.42261

T/W rocket TW_TJ_MAX 2.03595

T/W scramjet TW_SC_MAX 1.15596

Wing loading WS 224.1325
117

B-7 TRADE 007: EXPENDABLE BOOSTER, MACH 6, HYDROGEN FUEL

Mission summary: Table B-7.1 summarizes the mission constants for this trade study.

Table B-7.1. Expendable-Booster, Mach 6, Hydrogen-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 3.0 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection LH

 Hydrogen-fuel density 74.63kg/m3(4.65 lb/ft3)

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7) [11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike 

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized in Table B-7.2.
118

Table B-7.2. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.60

hc/lc 0.08

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Weight, volume, and balance

Hypersonic convergence weight and volume formulation.

[6]

Table B-7.3. Summary of Weight and Volume Constants

Weight

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrw – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Lw Lc htop

hc ht Liso Lcomp
119

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 60 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-7.4 summarizes the design-point data that were collected by AVDsizing.

Table B-7.4. Trade 007, Expendable-Booster, Mach 6, Hydrogen-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 6

(NA) D_MVIHN 0

(NA) D_WR 1

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 17260

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Table B.7.4. Continued
120

Air-breathing transition Mach number AMACH_TJS 2.5

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 74.63

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.175

Planform area SPLN 63.4503

Ratio of wetted area to planform area AKW 2.48815

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.23607

Total vehicle length AL_TOTAL 14.10698

Span BPLN 7.49633

Height above centerline BASE_HEIGHT 2.69141

Spatular width CSPAT 1.49927

Capture area ACAP 3.05364

Length of external compression ALC 8.46419

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.03927

Width of cowl at inlet W3 2.53854

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00885

Height of cowl to length of isolator ALI_L 0.0885

Height at combustor exit to nozzle length H_LN 0.04976

Stall/approach 

performance Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 64568809.22

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01306

Table B.7.4. Continued
121

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 4.95156

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.67345

Trajectory summary Climb range CLRANGE 170096.1216

Cruise range CRRANGE 3238877.436

Decent range DERANGE 709365.5088

Total range RANGE_TOTAL 4118339.066

Climb time T_CLIMB 2.07588

Cruise time T_CRUISE 30

Descent time T_DESCENT 19.55827

Total flight time T_FLT 51.63415

Mach number at max rocket T/W AMACH_TJ 0

Maximum rocket T/W TW_TJ_TM 0

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 1.37328

Maximum capture area of scramjet (NA) AC_W_MAX 3.05364

(NA) AMACH_ACAP 3

(NA) QBAR_ACAP 53177.74012

(NA) ALT_ACAP 17260

(NA) CFN_ACAP 0.76981

Average cruise L/D LD_CRUISE 1.88078

Average cruise Isp ISP_CRUISE 2600.27308

Minimum acceleration during scramjet mode ANMIN 0.33083

Oxidizer-to-fuel ratio OF_TRAJ 0

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 11504.86122

Propellant weight WPPL 3795.67775

Total fuel weight WFUEL 3756.93813

Fuel 1 weight WFUEL1 3756.93813

Weight of oxidizer Wox 38.73962

Manufacturer’s zero-fuel weight AMZFW 7709.18347

Operating weight empty OWE 7709.18347

Operating empty weight OEW 7709.18347

Weight margin WMARGIN 700.83486

Table B.7.4. Concluded
122

Operational items weight WOPER 0

Systems weight WSYS 3133.46936

Structural weight WSTR 2558.26102

Rocket propulsion system weight WP_TJ 0

Ram/scramjet weight WP_SC 1316.61824

Total fuel fraction FF_TOTAL 0.32992

Structural weight fraction WSTR_TOGW 0.22236

Total weight ratio WR 1.49236

Total volume V_TOTAL 88.4481

Fixed systems volume V_FIX 5

Total systems volume V_SYS 8.53792

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 50.37096

Total fuel volume V_FUELI1 50.34086

Total oxidizer volume V_OX 0.0301

Total propellant density PPL_DEN 75.35449

Total fuel density FUEL_DEN 74.63

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 0

Total ram/scramjet volume VENG_SC 11.84956

Total engine volume VENG 11.84956

Total void volume VVOID 17.68962

Convergence check Operating-weight-empty weight budget OWE_W 7709.18347

Operating-weight-empty volume budget OWE_V 7709.188

Planform area SPLN 63.4503

Capture-area required AC_RE 3.05364

Capture-area available AC_AV 3.05364

Structural index AISTR 16.20449

Propulsion index AIP 153.04818

T/W rocket TW_TJ_MAX 0

T/W scramjet TW_SC_MAX 1.37328

Wing loading WS 181.32085
123

B-8 TRADE 008: EXPENDABLE BOOSTER, MACH 8, HYDROGEN FUEL

Mission summary: Table B-8.1 summarizes the mission constants for this trade study.

Table B-8.1. Expendable-Booster, Mach 8, Hydrogen-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 3.0 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,400 ft)

Propellant selection LH

 Hydrogen-fuel density 74.63kg/m3(4.65 lb/ft3)

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7)

[11]

HyFAC database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion 

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized in Table B-8.2.
124

Table B-8.2. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 119954.0

f (stoichiometric) 0.0291

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.10

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.65

hc/lc 0.09

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Weight, volume, and balance

Hypersonic convergence weight and volume formulation[6]

Table B-8.3. Summary of Weight and Volume Constants

Weight

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Lw Lc htop

hc ht Liso Lcomp
125

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 60 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-8.4 summarizes the design-point data that were collected by AVDsizing.

Table B-8.4. Trade 008, Expendable-Booster, Mach 8, Hydrogen-Fuel, Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 8

(NA) D_MVIHN 0

(NA) D_WR 1

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 17260

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Table B.8.4. Continued
126

Air-breathing transition Mach number AMACH_TJS 2.5

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 74.63

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.1825

Planform area SPLN 95.66821

Ratio of wetted area to planform area AKW 2.4918

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.24083

Total vehicle length AL_TOTAL 17.32212

Span BPLN 9.20482

Height above centerline BASE_HEIGHT 3.45871

Spatular width CSPAT 1.84096

Capture area ACAP 5.20628

Length of external compression ALC 11.25938

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 1.38248

Width of cowl at inlet W3 3.22344

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00617

Height of cowl to length of isolator ALI_L 0.06173

Height at combustor exit to nozzle length H_LN 0.03455

Stall/approach 

performance Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 79284769.75

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.01333

Table B.8.4. Continued
127

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 4.88359

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.62393

Trajectory summary Climb range CLRANGE 482063.403

Cruise range CRRANGE 4355546.828

Decent range DERANGE 1161834.763

Total range RANGE_TOTAL 5999444.993

Climb time T_CLIMB 4.47102

Cruise time T_CRUISE 30

Descent time T_DESCENT 22.87244

Total flight time T_FLT 57.34346

Mach number at max rocket T/W AMACH_TJ 0

Maximum rocket T/W TW_TJ_TM 0

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 1.77687

Maximum capture area of scramjet (NA) AC_W_MAX 5.20628

(NA) AMACH_ACAP 3

(NA) QBAR_ACAP 53177.74012

(NA) ALT_ACAP 17260

(NA) CFN_ACAP 0.76981

Average cruise L/D LD_CRUISE 1.98418

Average cruise Isp ISP_CRUISE 2248.34719

Minimum acceleration during scramjet mode ANMIN 0.30152

Oxidizer-to-fuel ratio OF_TRAJ 0

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 19576.79851

Propellant weight WPPL 7423.31196

Total fuel weight WFUEL 7362.23917

Fuel 1 weight WFUEL1 7362.23917

Weight of oxidizer Wox 61.0728

Manufacturer’s zero-fuel weight AMZFW 12153.48654

Operating weight empty OWE 12153.48654

Operating empty weight OEW 12153.48654

Weight margin WMARGIN 1104.86241

Table B.8.4. Concluded
128

Operational items weight WOPER 0

Systems weight WSYS 3844.55785

Structural weight WSTR 4305.28354

Rocket propulsion system weight WP_TJ 0

Ram/scramjet weight WP_SC 2898.78275

Total fuel fraction FF_TOTAL 0.37919

Structural weight fraction WSTR_TOGW 0.21992

Total weight ratio WR 1.6108

Total volume V_TOTAL 170.77108

Fixed systems volume V_FIX 5

Total systems volume V_SYS 11.83084

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 98.69732

Total fuel volume V_FUELI1 98.64986

Total oxidizer volume V_OX 0.04745

Total propellant density PPL_DEN 75.21291

Total fuel density FUEL_DEN 74.63

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 0

Total ram/scramjet volume VENG_SC 26.08904

Total engine volume VENG 26.08904

Total void volume VVOID 34.15422

Convergence check Operating-weight-empty weight budget OWE_W 12153.48654

Operating-weight-empty volume budget OWE_V 12153.45331

Planform area SPLN 95.66821

Capture-area required AC_RE 5.20628

Capture-area available AC_AV 5.20628

Structural index AISTR 18.06016

Propulsion index AIP 123.13898

T/W rocket TW_TJ_MAX 0

T/W scramjet TW_SC_MAX 1.77687

Wing loading WS 204.63223
129

B-9 TRADE 009: EXPENDABLE BOOSTER, MACH 6, KEROSENE FUEL

Mission summary: Table B-9.1 summarizes the mission constants for this trade study.

Table B-9.1. Expendable-Booster, Mach 6, Kerosene-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 3.0 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 26.2 km (86,000 ft)

Propellant selection RP-1

 Kerosene-fuel density 820.0 kg/m3(51.2 lb/ft3)

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7)[11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion 

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized below in Table B-9.2.
130

TableB-9.2. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 43380.0

f (stoichiometric) 0.0680

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.20

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.50

hc/lc 0.07

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Weight, volume, and balance

Hypersonic convergence weight and volume formulation[6]

Table B-9.3. Summary of Weight and Volume Constants

Weight

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Lw Lc htop

hc ht Liso Lcomp
131

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 60 percent of body 

length to keep the trim drag manageable.[6]

Cost

No cost model was utilized.

Design-point database file: Table B-9.4 summarizes the design-point data that were collected by AVDsizing.

Table B-9.4. Trade 009, Expendable Booster, Mach 6, Kerosene–Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 30

Design Mach number D_MACH 6

(NA) D_MVIHN 0

(NA) D_WR 1

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 17260

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Table B.9.4. Continued

Initial descent range (NA) ALT_DE 0
132

Air-breathing transition Mach number AMACH_TJS 2.5

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.085

Planform area SPLN 34.74619

Ratio of wetted area to planform area AKW 2.32289

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.11371

Total vehicle length AL_TOTAL 10.43928

Span BPLN 5.54735

Height above centerline BASE_HEIGHT 1.20217

Spatular width CSPAT 1.10947

Capture area ACAP 1.30401

Length of external compression ALC 5.21964

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 0.64089

Width of cowl at inlet W3 1.75036

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00803

Height of cowl to length of isolator ALI_L 0.08035

Height at combustor exit to nozzle length H_LN 0.05021

Stall/approach 

performance Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 47781475.36

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Table B.9.4. Continued

Form drag CD0 0.00993
133

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.76719

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.57912

Trajectory summary Climb range CLRANGE 273006.9779

Cruise range CRRANGE 3240789.329

Decent range DERANGE 1009672.294

Total range RANGE_TOTAL 4523468.601

Climb time T_CLIMB 3.31144

Cruise time T_CRUISE 30

Descent time T_DESCENT 23.99917

Total flight time T_FLT 57.31061

Mach number at max rocket T/W AMACH_TJ 0

Maximum rocket T/W TW_TJ_TM 0

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 0.62777

Maximum capture area of scramjet (NA) AC_W_MAX 1.30401

(NA) AMACH_ACAP 3

(NA) QBAR_ACAP 53177.74012

(NA) ALT_ACAP 17260

(NA) CFN_ACAP 0.70822

Average cruise L/D LD_CRUISE 4.07714

Average cruise Isp ISP_CRUISE 969.91017

Minimum acceleration during scramjet mode ANMIN 0.21849

Oxidizer-to-fuel ratio OF_TRAJ 0

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 8344.88894

Propellant weight WPPL 3536.36068

Total fuel weight WFUEL 3512.19722

Fuel 1 weight WFUEL1 3512.19722

Weight of oxidizer Wox 24.16346

Manufacturer’s zero-fuel weight AMZFW 4808.52825

Operating weight empty OWE 4808.52825

Operating empty weight OEW 4808.52825

Table B.9.4. Concluded

Weight margin WMARGIN 437.13893
134

Operational items weight WOPER 0

Systems weight WSYS 2669.36452

Structural weight WSTR 1265.47176

Rocket propulsion system weight WP_TJ 0

Ram/scramjet weight WP_SC 436.55305

Total fuel fraction FF_TOTAL 0.42378

Structural weight fraction WSTR_TOGW 0.15165

Total weight ratio WR 1.73544

Total volume V_TOTAL 17.40924

Fixed systems volume V_FIX 5

Total systems volume V_SYS 5.69637

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 4.30194

Total fuel volume V_FUELI1 4.28317

Total oxidizer volume V_OX 0.01878

Total propellant density PPL_DEN 822.03813

Total fuel density FUEL_DEN 820

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 0

Total ram/scramjet volume VENG_SC 3.92898

Total engine volume VENG 3.92898

Total void volume VVOID 3.48185

Convergence check Operating-weight-empty weight budget OWE_W 4808.52825

Operating-weight-empty volume budget OWE_V 4808.58679

Planform area SPLN 34.74619

Capture-area required AC_RE 1.30401

Capture-area available AC_AV 1.30401

Structural index AISTR 15.67892

Propulsion index AIP 1117.75748

T/W rocket TW_TJ_MAX 0

T/W scramjet TW_SC_MAX 0.62777

Wing loading WS 240.16701
135

B-10 TRADE 010: EXPENDABLE BOOSTER, MACH 8, KEROSENE FUEL

Mission summary: Table B-10.1 summarizes the mission constants for this trade study.

Table B-10.1 Expendable-Booster, Mach 8, Kerosene-Fuel Mission Summary

Mission requirements

Endurance 0, 10, 20, and 30 min

Payload 0 kg (0 lb)

Launch altitude 17,260 m (56,630 ft)

Launch velocity 3.0 M

Max dynamic pressure 53.6 kPa (1,120 psf)

Cruise altitude 30.0 km (98,425 ft)

Propellant selection RP-1

 Kerosene-fuel density 820.0 kg/m3(51.2 lb/ft3)

Operational constraints

Takeoff field length 4,572.0 m (15,000 ft)

Landing field length 4,572.0 m (15,000 ft)

 MLW/TOGW 1.0

Maximum axial acceleration 3.0 g

Relevant method assumptions and constants: The geometry, aerodynamic, structural, propulsion weight and 

volume, performance, stability and control, and cost methods and constants are summarized below.

Geometry

The blended-body configuration, as defined by using hypersonic convergence.

[6]

Aerodynamics

CLmax = 0.50 (FDL-7)[11]

Hyfac database, MAC circa 1970[3]

Spatular corrections from Pike

Structure and thermal protection

Structural shape factor Kstr from MAC[6]

Propulsion 

Dual-mode ram-scramjet

The dual-mode ram-scramjet for this trade is a composite of Marquardt ramjet data[14] from Mach 3 to 6 and onedimensional stream thrust analysis[12] from Mach 6 to 8. The constants that are assumed in the stream thrust analysis 

for this trade are summarized below in Table B-10.2.
136

Table B-10.2. Summary of Scramjet Stream Thrust Constants

Cycle constants Value

Hpr (kJ/kg) 43380.0

f (stoichiometric) 0.0680

Vfx/V3 0.50

Vf/V3 0.50

1 0.95

b 0.90

(

 

 

 

 

)

 

0.01

( 

 

 

)

 

0.20

Cev 0.99

Cpe 1.59

Cea 1.00

c 1.362

e 1.22

Geometric constants

lc/lw 0.50

hc/lc 0.07

hiso/liso 0.1

Lcomb 0.762 m

Shock on lip Mach number 8.0

1n 22.0

2n 9.0

Weight, volume, and balance

Hypersonic convergence weight and volume formulation.

[6]

Table B-10.3. Summary of Weight and Volume Constants

Weight

ETW_DMR – Dual-mode ramjet thrust-to-weight ratio (T/Weng) 12.0 kg/kg

 – empty weight margin (OEW/OEW) 0.10

Fsys – variable systems weight (Wsys/OEW) 0.16 kg/kg

Cun – fixed unmanned systems weight 1,900 kg

Wcrew – weight of crew per person (Wcrew/person) 0.0

fprv – crew provision weight per person (Wprv/person) 0.0

Volume

Kve_DMR – dual-mode ramjet per kg thrust (VDMR/T) 0.00075 m3/kg

Kvv – void volume coefficient (Vvoid/Vtotal) 0.20 m3/m3

Kvs – systems volume coefficient (Vsys/Vtotal) 0.02 m3/m3

Vun – fixed unmanned system volume 5.0 m3

Vpcrew – variable crew volume coefficient (Vcrew/person) 0.0

Fcrew – fixed crew volume coefficient 0.0

Performance

The energy-integration method was used to compute the trajectory.

The required approach speed was computed from the assumed CLmax.

Lw Lc htop

hc ht Liso Lcomp
137

Stability and control

No direct computation of stability and control; the scramjet cowl location was constrained to 60 percent of body 

length to keep the trim drag manageable.

[6]

Cost

No cost model was utilized.

Design-point database file: Table B-10.4 summarizes the design-point data that were collected by AVDsizing.

Table B-10.4. Trade 010, Expendable-Booster, Mach 8, Kerosene-Fuel Output Database

Category Description Variable Value

Mission input check Number of design passengers APAXD 0

Maximum number of passengers APAXMAX 0

Number of crew members CREW 0

Weight per passenger WPAX 100

Weight per crew member WCREW 129

Cargo weight WCARGO 0

Cruise switch (0 range, 1 endurance) NCRUISE 1

Design range or endurance D_RANGE 20

Design Mach number D_MACH 8

(NA) D_MVIHN 0

(NA) D_WR 1

Takeoff field length (NA) TOFL 3337.56

Altitude at takeoff ALT_TO 0

Landing field length SLAND 2400

Altitude at landing ALT_LAND 0

(NA) MP_TO 0

(NA) MP_LAND 0

(NA) MP_TRAJ 0

(NA) NTRAJ_ST 0

Maximum axial load factor AN_MAX 3

Cruise normal load factor AN_NORM 1

Altitude step for climb out (NA) ASTEP_CO 10

Velocity at climb out (NA) V_CLIMBOUT 180

Altitude for initial climb (NA) ALT_IC 3048

Acceleration Mach step AMSTEP_AC 0.01

Altitude step for climb to transonic acceleration ASTEP_AC 10

Altitude for transonic acceleration ALT_TC 17260

Mach step for transonic acceleration AMSTEP_TA 0.01

Altitude step for constant q climb ASTEP_QC 10

Initial descent range (NA) ALT_DE 0

Table B.10.4. Continued
138

Air-breathing transition Mach number AMACH_TJS 2.5

Range step for cruise DDCRUISE 0.5

Number of fuels NFUEL 1

Fuel density FUEL_DEN 820

Oxidizer density OX_DEN 1287

Takeoff climb gradient (NA) TO_CGR 0.024

Takeoff climb gradient (NA) TO_OEI 1

Landing climb gradient (NA) ALAND_CGR 0.021

Landing climb gradient (NA) ALAND_OEI 1

Landing weight ratio ALAND_WR 1

Altitude for reserve mission (NA) ALTRES 3048

Range for reserve mission (NA) R_MACH 0

Endurance for reserve mission (NA) TIMERES 120

Configuration (1-lifting body) NBASE 1

Geometry Slenderness parameter TAU 0.075

Planform area SPLN 51.27501

Ratio of wetted area to planform area AKW 2.31308

Spatular width to wing semispan CS_SPAT 0.5

Ratio of frontal area to planform area SF_SREF 0.10127

Total vehicle length AL_TOTAL 12.68148

Span BPLN 6.73883

Height above centerline BASE_HEIGHT 1.37701

Spatular width CSPAT 1.34777

Capture area ACAP 1.92415

Length of external compression ALC 6.34074

Length of cowl ALLC 1.44862

Width of cowl outside spatular WCOUT 0.77855

Width of cowl at inlet W3 2.12631

Height of cowl HCOWL 0.01222

Height of throat HCOWLI 0.00618

Height of cowl to length of isolator ALI_L 0.06179

Height at combustor exit to nozzle length H_LN 0.0297

Stall/approach 

performance Velocity for approach calculation VREL 64.11653

Reynolds number REYNOLDS 58044199.41

Mach number for approach AMACH 0.1884

Landing-gear drag DCD_LG 0.0015

Oswald’s efficiency factor correction for flaps DE_LG 0

Form drag CD0 0.00966

Table B.10.4. Continued
139

Induced drag factor ALIND 0.7

Maximum L/D ALDMAX 5.85781

Approach speed VA_LC 83.34419

Stall speed VS_LC 64.11092

Wing loading at stall WS_STALL 243.87905

Weight ratio WR_TJ 0.49743

Trajectory summary Climb range CLRANGE 943176.4315

Cruise range CRRANGE 2904215.76

Decent range DERANGE 1792538.799

Total range RANGE_TOTAL 5639930.99

Climb time T_CLIMB 8.39916

Cruise time T_CRUISE 20

Descent time T_DESCENT 31.7583

Total flight time T_FLT 60.15745

Mach number at max rocket T/W AMACH_TJ 0

Maximum rocket T/W TW_TJ_TM 0

(NA) AMACH_SC 0

Maximum scramjet T/W TW_SC_TM 0.94606

Maximum capture area of scramjet (NA) AC_W_MAX 1.92415

(NA) AMACH_ACAP 3

(NA) QBAR_ACAP 53177.74012

(NA) ALT_ACAP 17260

(NA) CFN_ACAP 0.70822

Average cruise L/D LD_CRUISE 3.92361

Average cruise Isp ISP_CRUISE 731.50305

Minimum acceleration during scramjet mode ANMIN 0.11526

Oxidizer-to-fuel ratio OF_TRAJ 0

Weight and volume Weight of crew WCRW 0

Weight of design payload WPAY_D 0

Weight of max payload WPAY_MAX 0

Takeoff gross weight TOGW 12027.0251

Propellant weight WPPL 6074.29065

Total fuel weight WFUEL 6044.37741

Fuel 1 weight WFUEL1 6044.37741

Weight of oxidizer Wox 29.91324

Manufacturer’s zero-fuel weight AMZFW 5952.73445

Operating weight empty OWE 5952.73445

Operating empty weight OEW 5952.73445

Weight margin WMARGIN 541.15768

Table B.10.4. Concluded
140

Operational items weight WOPER 0

Systems weight WSYS 2852.43751

Structural weight WSTR 1610.94574

Rocket propulsion system weight WP_TJ 0

Ram/scramjet weight WP_SC 948.19352

Total fuel fraction FF_TOTAL 0.50505

Structural weight fraction WSTR_TOGW 0.13394

Total weight ratio WR 2.02042

Total volume V_TOTAL 27.53721

Fixed systems volume V_FIX 5

Total systems volume V_SYS 6.10149

Total payload volume V_PAY 0

Total crew volume V_CREW 0

Total propellant volume V_PPL 7.39443

Total fuel volume V_FUELI1 7.37119

Total oxidizer volume V_OX 0.02324

Total propellant density PPL_DEN 821.4679

Total fuel density FUEL_DEN 820

Total oxidizer density OX_DEN 1287

Total rocket volume VENG_TJ 0

Total ram/scramjet volume VENG_SC 8.53374

Total engine volume VENG 8.53374

Total void volume VVOID 5.50744

Convergence check Operating-weight-empty weight budget OWE_W 5952.73445

Operating-weight-empty volume budget OWE_V 5952.77207

Planform area SPLN 51.27501

Capture-area required AC_RE 1.92415

Capture-area available AC_AV 1.92415

Structural index AISTR 13.58262

Propulsion index AIP 805.02902

T/W rocket TW_TJ_MAX 0

T/W scramjet TW_SC_MAX 0.94606

Wing loading WS 234.5592

---------- o0o ----------
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Contractor Report

 4. TITLE AND SUBTITLE

Solution-Space Screening of a Hypersonic Endurance Demonstrator 

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NNL09AA00A

 6. AUTHOR(S)

Chudoba, Bernd; Coleman, Gary; Oza, Amit; Gonzalez, Lex; Czysz, Paul

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NASA Langley Research Center 

Hampton, Virginia 23681-2199 

 

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This report documents a parametric sizing study performed to develop a program strategy for research and development and 

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complementing technical and managerial decision-making during the earliest conceptual design phase towards minimization of 

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Hypersonic aircraft; Hypersonics; Life cycle analysis; Vehicle design

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