THE DECISION MAKER’S GUIDE TO 

ROBUST, RELIABLE AND INEXPENSIVE 

ACCESS TO SPACE 

by 

Gary N. Henry, Lieutenant Colonel, USAF 

July 2004 

Occasional Paper No. 37 

Center for Strategy and Technology 

Air War College 

Air University 

Maxwell Air Force Base, Alabama 

The Decision-Maker’s Guide to Robust, Reliable, and 

Inexpensive Access to Space 

Gary N. Henry, Lieutenant Colonel, U. S. Air Force 

July 2004 

The Occasional papers series was established by the Center for Strategy 

and Technology as a forum for research on topics that reflect long-term 

strategic thinking about technology and its implications for U.S. national 

security. Copies of No. 37 in this series are available from the Center for 

Strategy and Technology; Air War College; 325 Chennault Circle; 

Maxwell AFB, Alabama 36112. The fax number is (334) 953-6158; 

phone (334) 953-6460 

Occasional Paper No. 37 

Center for Strategy and Technology 

Air University 

Maxwell Air Force Base, Alabama 36112 

Contents 

Page

Disclaimer.................................................................................................. iii 

Author..........................................................................................................v 

Illustrations .................................................................................................vi 

Tables....................................................................................................... viii 

Preface .........................................................................................................x 

Abstract.................................................................................................... xiii 

I. Introduction .............................................................................................1 

II. Background ............................................................................................9 

III. Defining the Problem...........................................................................31 

IV. The Search for Solutions ....................................................................69 

V. Conclusions and Recommendations ....................................................99 

Notes........................................................................................................109 

Disclaimer 

The views expressed in this academic research paper are those of

the author(s) and do not reflect the official policy or position of the Air 

War College, the U.S. government or the Department of Defense. In 

accordance with Air Force Instruction 51-303, it is not copyrighted, but is 

the property of the United States government. 

iii 
iv 
Author 

Lieutenant Colonel Gary N. Henry graduated from the United 

States Air Force Academy with academic distinction in 1984. Colonel 

Henry’s varied career includes assignments in laboratory research, 

academic instruction, developmental flight test, and system acquisition. 

He began his career conducting high-energy laser vulnerability research 

under the auspices of the Strategic Defense Initiative from 1984-87 and 

then served as an Assistant Professor of Astronautical Engineering at the

USAF Academy from 1989-93. While there he co-edited the textbook 

Space Propulsion Analysis and Design and received the 1993 USAF

Science and Engineering Award for successfully launching the DOD’s 

first land-based hybrid sounding rocket. He was then selected to attend 

USAF Test Pilot School and subsequently assigned to the 418th Flight Test 

Squadron as Chief of the Flight Dynamics Branch from 1994-1996 where 

he directed numerous MC-130H Combat Talon and AC-130U Gunship 

test missions. Colonel Henry later became the Chief of the Airborne Laser 

(ABL) System Program Office Air Vehicle Integration and Test IPT at 

Kirtland AFB, New Mexico, from 1998-1999, where he was responsible 

for preparing the ABL weapon system for first flight. In 2000, Colonel 

Henry was assigned as the Lead ABL Program Element Monitor in the

Office of the Assistant Secretary of the Air Force for Acquisition at the 

Pentagon. He is currently the Chief, Advanced Space Systems Division, 

Space Superiority Materiel Wing, Space and Missile Systems Center, at 

Los Angeles AFB, California, and will assume command of the 1st Air and 

Space Test Squadron, 30th Space Wing, Vandenberg AFB, California, in 

the summer of 2004. Colonel Henry’s academic credentials include a

Bachelor of Science in astronautical engineering from the USAF 

Academy, a Master of Science in aeronautical and astronautical 

engineering from Stanford University, and a Master of Strategic Studies 

from the Air War College. 

v 
Illustrations 

Page 

Figure 1. Robust Reliable Inexpensive Space Access (R2

ISA) 

“Trade Space”.....................................................................................6 

Figure 2. R2

ISA Linkages to Vehicle Design and Space 

Architecture ........................................................................................7 

Figure 3. Civil and Military Requirements Comparison ..........................17 

Figure 4. Shuttle Spacelift Performance —Vision vs. Reality ................24 

Figure 5. Specific Orbital Energy.............................................................33 

Figure 6. Two Stage Vehicle Mass Contributors......................................38 

Figure 7. Feasible Regions for Launch Systems ......................................42 

Figure 8. Staging Impact Upon Vehicle Mass..........................................43 

Figure 9. X-33-Based SSTO RLV Design Concepts................................50 

Figure 10. Parametric Analysis of a Single-Stage-to-Orbit Vehicle ........53 

Figure 11. ELV Space Launch Cost Trendlines.......................................58 

Figure 12. Labor Intensity of Aerospace Systems....................................61 

Figure 13. Fiscal Year 2002 Space Shuttle Budget Summary..................65 

Figure 14. Time Allocation at the Space Shuttle Orbital 

Processing Facility............................................................................66 

Figure 15. Space Launch Design Linkages to R2

ISA................................70 

Figure 16. Space Shuttle Mass Distribution .............................................73 

Figure 17. Airbreathing versus Rocket Performance ...............................74 

vi 
Figure 18. Notional Hypersonic Corridor.................................................80 

Figure 19. Rocket/Air-Breathing Propulsion 

Options/Characteristics.....................................................................82 

Figure 20. Construction of a Body “Waverider” Producing a 

Conical Shock...................................................................................85 

Figure 21. Waveriders Optimized for Different Mach Numbers 

and Conditions..................................................................................86 

Figure 22. Boeing FASST Concept..........................................................87 

Figure 23. R2

ISA and the Space Transportation Architecture: 

Vision versus Reality........................................................................95 

vii 
Tables

 Page 

Table 1. Commercial Space Transportation Study Attributes and 

Requirements....................................................................................20 

Table 2. RLV “Technical/Programmatic” Development Lessons 

Learned .............................................................................................28 

Table 3. RLV “Environmental” Development Lessons Learned .............29 

Table 4. Propulsion Technology Performance Summary.........................36 

Table 5. Key Propulsion Performance Characteristics.............................37 

Table 6. Mass Characteristics of Aerospace Systems ..............................39 

Table 7. Velocity Budgets to Low-Earth Orbits for Selected 

Launch Vehicles...............................................................................40 

Table 8. Aircraft to Launch Vehicle Comparison ....................................44 

Table 9. Properties of Candidate RLV Propellants ..................................46 

Table 10. Selected Temperature Distributions .........................................48 

Table 11. X-33 Based Single-Stage-to-Orbit Design Concepts 

Description........................................................................................51 

Table 12. Launch Vehicle Costs...............................................................59 

Table 13. Transportation Vehicle Hardware Costs per Pound .................60 

Table 14. Simplified Expendable Launch Vehicle Cost Excursions........64 

Table 15. X-43 Series Hypersonic Tests ..................................................83 

viii 
ix 
Preface 

I naively began my research thinking that our failure as a nation to

achieve truly robust, reliable, and inexpensive space access (R2

ISA) was

simply the result of several misplaced technology bets, poor timing, and 

bad luck. I was wrong. I soon discovered that one cannot adequately 

address this complex topic without entering the realm of public policy,

economics, and programmatics, as well as the technical complexity 

inherent with R2

ISA. 

Although I am critical of NASA at times in this paper, without 

question, NASA is blessed with the services of some of the most brilliant 

and dedicated scientists and engineers this nation has to offer. However, if

NASA was an organization in crisis before the tragic loss of Columbia, it 

will soon be under siege. This is wrong. Decisions spanning more than 

thirty years by the Executive Branch, Congress, the Department of 

Defense, and the entire aerospace community contributed to this tragedy. 

The loss of Columbia brings some difficult issues to the forefront. 

NASA made a decision in late 2002 to defer pursuit of a secondgeneration reusable launch vehicle until at least 2009, and it instituted a 

shuttle life extension program to 2020. While I personally believed this to 

be a mistake prior to the loss of Columbia, I can now say so with absolute 

certainty. Keeping the shuttle flying until 2020 will almost certainly mean 

the loss of another vehicle and crew, as well as an abrupt and permanent

grounding of the remaining shuttle fleet. Failure, due to a “lack of 

resources,” to immediately begin the pursuit of the design and build-out of 

a new space transportation architecture that fails to include a solution for

the shuttle not only puts a price tag on human life but seriously 

jeopardizes the future of manned space flight. 

Recent Space Launch Initiative shuttle replacement cost estimates 

as high as $30-35 billion proved unaffordable for NASA. Further, the 

NASA staff was unable to make the business case to its own administrator 

for the long promised $1000 per pound to low-earth orbit. Meanwhile, the 

Department of Defense (DOD), unable to ride NASA’s coattails, struggles 

to find solutions to its aspirations for an affordable, flexible, and 

operationally robust space launch system. Significant strides towards 

defining force enhancement, counterspace, and force application 

requirements are encouraging, but an enabling space launch system 

remains elusive. 

x 
The commercial space industry is on life-support with global 

capacity far outstripping demand for expensive and cumbersome launch 

services. The DOD and NASA pursuit of independent space launch

solutions is inefficient and unaffordable. I believe the answer to this crisis 

demands a national solution to earth-to-orbit operations that is responsive 

to civil, military, and commercial requirements. 

Is a national solution possible? Does it even make sense?

Potential answers to these vexing questions were painfully slow in 

coming. This paper suggests that there are indeed national solutions that 

can deliver R2

ISA. There are also quicker, riskier, technology-leveraged 

alternatives than those suggested here that could work. However, as a 

nation we’ve been down that road before – placing several multi-billion 

dollar bets that failed to deliver. I am convinced that the approach 

promulgated here—followed to its logical conclusion—is guaranteed to

deliver R2

ISA. There are likely to be similar solutions that can deliver

R2

ISA even faster and cheaper; it is time for all stakeholders to focus our 

effort to defining those solutions. 

One observation I’ve made is that it would be very helpful if the 

aerospace community could reach a reasonable consensus

recommendation regarding a prudent road ahead. This would not only 

help lift the fog that has virtually immobilized us for more than thirty 

years, but also inspire some sorely needed confidence from American 

leadership and her people. Unfortunately, without a clear mandate from 

our senior leadership, such consensus is unlikely. Hence, the decisions of 

our most senior leadership within the executive, congress, DOD, and 

NASA will provide the direction and resources to make R2

ISA a—or not. 

This paper is a modest, unbiased attempt to help prepare them for that 

task. 

My sincere thanks to the Center for Strategy and Technology and 

the unwavering support of Dr. Grant Hammond, Mr. Ted Hailes, and 

particularly, Col John Geis, who painstakingly reviewed numerous drafts 

of this paper and provided invaluable editorial comment. I am also 

sincerely grateful to Dr. Lanny Jines of the Air Force Research Laboratory 

and Dennis Bushnell of NASA Langley Research Center for providing 

personal insight as well as generous access to their respective 

organizations. Thanks to Dr. Kevin Bowcutt, Ramon Chase, Bill 

Claybaugh, Dr. William Heiser, Dr. Harry Karasapolous, Dr. John Olds, 

Russel Rhoades, Ming Tang, Neil Woodward, Edgar Zapata and others for 

challenging the thesis of this paper and providing invaluable advice and 

xi 
assistance. Finally, I’d like to thank my beautiful wife Heather for her

unselfish support as well as my two-year old son Patrick who never quite 

figured out where daddy went all those weekends. 

Despite bleak characterizations of past failure, there is great hope 

for the future. As I researched this paper, I was struck by the sheer 

magnitude and diversity of the technical and programmatic discourse 

surrounding space exploration and exploitation. There is no shortage of 

brilliant people or great ideas. Once R2

ISA is achieved, I am convinced 

that we will witness a renaissance in space not unlike that experienced in 

aviation during the interwar years of the 20th Century. Today’s earth-toorbit operations are reminiscent of the challenges and risks of Charles 

Lindbergh’s foray across the Atlantic. We have yet to build the space 

equivalent of the DC-3. Any personal agenda I may have brought to the 

debate amounts to nothing more than one day hoping to affordably buy a 

ticket to earth orbit—and one more thing. To inspire my son, and his 

generation, in the same way I was inspired by the men and women who 

built Apollo and safely delivered a dozen men to the lunar surface and 

brought them home again. I’d like to one day claim that I was part of the

next generation who built the ships that permanently bridge the gap 

between air and space—both inspiring and enabling my son’s generation 

to begin an awe inspiring journey of their own. 

This paper is dedicated to Dr. Ron Humble, an aerospace 

enthusiast with whom I had the pleasure to co-edit a text entitled Space 

Propulsion Analysis and Design, who passed away in 2002 at 44 years 

young. We’ll miss him. 

xii 
Abstract 

The principal barrier to unconstrained civil, military, and 

commercial exploitation of space is the high cost and elevated risk

associated with access to low-earth orbit and beyond. The road ahead 

remains clouded and ambiguous despite thirty years of trying. This paper 

is intended to empower senior-level decision-makers with the insight and 

objectivity needed to ask tough, probing questions, as well as provide a

course of action that illuminates a path toward truly robust, reliable, and 

inexpensive space access (R2

ISA) for the 21st Century. It will be made

clear that: 1) R2

ISA will never be achieved with foreseeable expendable 

launch vehicle technology. Reusable launch vehicles are currently the

only economically and operationally viable path toward achieving R2

ISA. 

2) Existing technology limitations make pursuit of a single-stage-to-orbit 

solution imprudent at this time, while a two-stage-to-orbit reusable launch 

vehicle is the logical next step towards R2

ISA. 3) Despite conflicting 

requirements, a common “national” solution for civil, military, and 

commercial application is both possible and desirable. 4) Fundamental 

reforms are essential to make R2

ISA a reality. 

This paper provides a concrete, quantifiable definition and a

theoretical construct for R2

ISA. It then presents the historical context, 

divergent requirements, and disparate technical perspectives that comprise 

the confusing state of affairs engulfing the space transportation debate and 

concludes with a set of lessons learned. It will then explore the current 

physical, economic, and technological challenges associated with 

achieving R2

ISA. From this, it then outlines the technical solutions to 

R2

ISA focusing on rocket and air-breathing reusable launch vehicles. A 

set of proposed “national” space launch system architecture attributes are 

introduced as well as a proposed solution to the current impasse. Finally, 

this paper will present a set of conclusions and recommendations for the 

successful pursuit of R2

ISA. Along the way, this paper will frame the 

critical issues and key questions that must be asked and answered before 

the United States commits billions of dollars towards a future space launch

architecture. A fundamental question considered is whether R2

ISA is best 

achieved through separate (but coordinated) Military, Civil, and 

Commercial endeavors, or if a single “national” solution makes sense. 

This question is vital to framing a debate that transcends technological 

xiii 
challenges, interagency rivalry, and political expediency, and may 

ultimately determine the difference between success and failure. 

 

xiv 
Decision Maker’s Guide...1 

I. Introduction 

The Air Force was born of a new technology – manned 

powered flight. Innovation will enable the Air Force to 

evolve from an air force to an air and space force on its 

path toward space.... We are now transitioning from an 

air force to an air and space force on an evolutionary path 

to a space and air force. 

—Global Reach—Global Power 

Robust/Reliable/Inexpensive space access (R2

ISA) is the holy grail 

of space transportation. It is universally accepted that the principal barrier 

to unconstrained civil, military, and commercial exploitation of space is

the high cost and elevated risk associated with access to low-earth orbit 

and beyond. It is evident that the United States has yet to introduce a 

space transportation architecture that delivers R2

ISA despite numerous

abortive attempts to do so. Equally disappointing is the clouded and

ambiguous road ahead despite thirty years of trying. It didn’t have to be 

this way.

This paper is intended to empower senior level decision-makers

with the insight and objectivity needed to ask tough, probing questions, as 

well as provide a course of action toward truly robust, reliable, and 

inexpensive space launch for the 21st Century. It will be made clear that: 

1) R2

ISA will never be achieved with foreseeable expendable launch

vehicle technology. Reusable launch vehicles are currently the only viable 

path, both economically and operationally, toward achieving R2

ISA. 2) 

Existing technology limitations make pursuit of a single-stage-to-orbit 

solution imprudent at this time while a two-stage-to-orbit reusable launch

vehicle is the logical next step towards R2

ISA. 3) Despite conflicting 

requirements, a common “national” solution for civil, military, and 

commercial application is both possible and desirable. 4) Fundamental 

reforms are essential to make R2

ISA a reality. 

This paper will begin by providing a concrete, quantifiable

definition and a theoretical construct for R2

ISA. It then presents the 

historical context, divergent requirements, and disparate technical 

perspectives that comprise the confusing state of affairs engulfing the 

space transportation debate and concludes with a set of lessons learned. It 

will then explore the current physical, economic, and technological 
2...Decision Maker’s Guide 

challenges associated with achievingR2

ISA. From this, it will then outline 

the technical solutions to R2

ISA focusing on rocket and air-breathing 

reusable launch vehicles, propose a set of “national” space launch system 

architecture attributes, and conclude with a proposed solution to the 

current impasse. Finally, this paper will present a set of conclusions and 

recommendations for the successful pursuit of R2

ISA. Along the way, this 

paper will frame the critical issues and key questions that must be asked 

and answered before the United States commits billions of dollars towards 

a future space launch architecture. A fundamental question considered is 

whether R2

ISA is best achieved through separate (but coordinated) 

Military, Civil, and Commercial endeavors, or if a single “national” 

solution makes sense. This question is vital to framing a debate that 

transcends technological challenges, interagency rivalry, and political 

expediency, and may ultimately determine the difference between success

and failure. 

R2ISA DEFINED 

Robust, Reliable, Inexpensive space access is a qualitative 

construct describing the essential characteristics of any successful civil,

military, and commercial space launch system architecture. However,

quantification is necessary to scope the problem, define usable metrics,

and permit objective decision-making. Although any quantitative 

rationale is subject to debate, a R2

ISA definition is nonetheless provided. 

Inexpensive 

The term inexpensive is used in lieu of affordable with 

premeditation. “Affordable” is whatever a customer is willing to pay 

based upon overall mission value or imperative.1

 “Inexpensive” represents 

a threshold where there is no longer a cost disincentive in considering 

space-based alternatives to specific mission requirements. The 

complication arrives when attempting to define that threshold. Typically, 

between the civil, military, and commercial arena, the lowest cost 

threshold lies within the commercial realm. Thus, the commercial 

standard is used to define the threshold for “inexpensive” space launch. 

The National Aeronautics and Space Administration (NASA)

sponsored a collaborative government/industry Commercial Space 

Transportation Study in 1994 that is recognized as the most 
Decision Maker’s Guide...3 

comprehensive and authoritative effort to date in quantifying the

commercial space market.2

 It was intended to address the global 

perception that: 1) significant untapped commercial markets existed if the 

cost of access to space could be lowered an order of magnitude or more, 2) 

a new launch system could provide such a reduction in launch costs, and 

3) such a reduction would result in a “space industrial revolution” with 

very large increases in users and traffic.3

 The analysis concluded that a 

non-linear change in demand relative to price (called market elasticity) 

begins to enter the commercial space launch market at around $1000 

(FY93) per pound to low-earth orbit and is well established at $600 per

pound.4

 For the purposes of this paper, the threshold for commercial 

viability begins at $1000 (FY02) per pound to low-earth orbit and is used

as the threshold for a second-generation launch system.

5

 Based upon the

author’s interpretation of Commercial Space Transportation Study

estimates, $100 per pound to low-earth orbit can provide a sufficient 

return-on-investment to make the subsequent development of profitable 

commercial launch system architectures self-sustaining. Thus, the price

threshold of $100 per pound is used for third generation launch systems.6

 

However, this optimism does not negate the fact that another order of 

magnitude drop in costs to $10 per pound is necessary to approach

“aircraft-like” costs. Further, the study concluded, “as a commercial 

investment measured at standard industrial return-on-investment levels, 

the investment cost for a new space launch system must be kept in the 

range of a few billions of dollars.”7

 

Reliable 

There are two distinct dimensions of reliability that have a large

impact upon overall system cost and performance. The first is the 

probability of mission failure, which is relevant to both manned and 

unmanned systems. The second is probability of catastrophic loss of 

either a mission payload or crew. The former drives critical commercial 

considerations to include customer confidence and the cost of insurance

(which can be a high percentage of launch costs) as well as directly 

impacting the probability of mission success in the military realm. The 

latter deals with crew safety and implies a “crew-rated” system demanding 

significantly higher safety margins than unmanned systems, adding 

weight, complexity, and cost. Higher margins for crewed systems can be 

addressed one of two ways. The first is to augment a non crew-rated 
4...Decision Maker’s Guide 

vehicle with a crew-escape system to make the system safer. This 

approach can provide improved safety margins, but with a significant 

penalty (typically manifested as weight) to overall system performance. 

The second is to build a vehicle with sufficiently high reliability to make 

the crew-rated distinction irrelevant, as is the case with existing aircraft

operations. 

Space launch systems have historically operated at very low 

reliability and safety thresholds relative to aircraft operations. Based upon 

Bayesian reliability theory (where the benefit of learning is factored into

statistical analyses), the first human launch of the Mercury Redstone (Alan 

Shepard) appears to have carried a 61 percent probability of vehicle 

launch failure. The first human launch of a Mercury Atlas (John Glenn) 

carried a 63 percent probability of launch failure. Both of these values are 

based upon the heritage of the their respective launch vehicle development 

to that date.8

 To deal with unacceptably low demonstrated reliability, a

crew escape system was added with a five percent probability of failure. 

With crew escape, safety now becomes the product of the two numbers

yielding a probability of both the launch vehicle and the crew escape 

system falling to three percent, or alternatively, a 97 percent probability of 

crew survival.9

 Although expendable launch vehicle reliability has

improved dramatically since the early 1960’s, demonstrated system

reliabilities to this date range between 94 and 99 percent. 

Current NASA shuttle reliability numbers vary, and there is a large

disconnect between the desired reliability of its future generation reusable 

launch vehicles and reality. One unpublished NASA source projects 

shuttle reliability against loss of vehicle as “marginally higher” than the 

originally specified 98 percent reliability (implying a catastrophic loss of 

one vehicle in every 50 launches).10 An internal NASA “bottom-up”

failure analysis predicts the probability of catastrophic failure at 1 in 

247.11 The space shuttle, despite its overwhelming complexity, is without 

question the safest and most reliable space transportation system to ever 

fly. However, relative to aircraft operations, the shuttle is a very 

dangerous machine. NASA articulated a desire to achieve a two order of 

magnitude improvement with its second-generation reusable launch 

vehicle development under the auspice of the Space Launch Initiative to 

reduce the risk of crew loss to approximately 1 in 10,000 missions.12 Note

that another order of magnitude improvement to 1/100,000 probability of 

loss of vehicle is necessary to approach “aircraft-like” reliability.13 A

more realistic and achievable failure probability goal for a second- and 
Decision Maker’s Guide...5 

third-generation reusable launch vehicle is 1/1000 and 1/10,000 

respectively. One can safely conclude that aircraft-like reliability for 

space launch systems is well beyond existing state of the art. 

Robust 

The term “robust” is a multifaceted mission and system-dependent 

component of R2

ISA that can carry significantly different meanings in the 

civil, military and commercial realms. An important goal of a national 

solution to R2

ISA is an amelioration of these differences that is addressed 

in the following chapter. This axis of the R2

ISA parameter space can be 

defined in any fashion, to include the flexibility to readily select the

mission, payload, flight path, and crew (if applicable) without significant 

configuration changes despite highly variable weather and/or operational 

conditions. For the purpose of this paper, robustness takes on the much

narrower definition of responsiveness, which is simply the launch vehicle 

preparation or recycle time, measured in hours. Figure 1 below 

graphically depicts the R2

ISA parameter space for existing expendable and 

reusable launch vehicles, as well as second-, third-, and fourth- generation 

(aircraft-like) space launch systems. 

Finally, a brief discussion of important definitions is in order. The 

FAA Office of Commercial Space Transportation defines four payload 

mass classes to low-earth orbit: small (5,000 lb or less), medium (5,001-

12,000 lb), intermediate (12,001-25,000 lb), and heavy (more than 25,000 

lbs).14 For the purposes of this paper, medium and intermediate are

merged into medium-class payloads. Low-earth orbits are used as the 

baseline for objective cost and performance assessments and to define 

some common terms. Low-earth orbits are defined as circular orbits in the 

range of 185 kilometers (100 nautical miles) to 460 kilometers (250 

nautical miles) with the lower threshold used in this paper.15 Orbital

velocity is important because it represents the minimum speed and 

direction at which immediate active propulsive thrust is no-longer required 

to counteract the earth’s gravitational acceleration and the trajectory can

be sustained indefinitely (absent losses due to drag, gravitational 

anomalies, solar activity, etc.). 
6...Decision Maker’s Guide 

Figure 1. Robust Reliable Inexpensive Space Access (R2

ISA) “Trade 

Space” 

R2ISA KEY LINKAGES 

R2ISA is directly linked to the overall vehicle design that is in turn 

influenced by an overarching space architecture. Earth-to-orbit 

operations, of which R2

ISA is directly concerned, is a part of this larger

space architecture. Since a system-of-systems perspective is essential to 

achieving an optimized R2

ISA solution, two important observations are 

now made. First, the successful oversight of these elements can only be

accomplished by very senior level decision-makers. Secondly, these

interrelationships can get very complicated very quickly. Figure 2 below 

graphically illustrates these relationships and can be considered a topical 

“roadmap” of discussion for this paper. Although the details of this figure
Decision Maker’s Guide...7 

are not discussed here, the reader is encouraged to refer to it often as more 

detailed information and actionable arguments are presented in this paper. 

Figure 2. R2

ISA Linkages to Vehicle Design and Space Architecture 
8...Decision Maker’s Guide 
Decision Maker’s Guide...9 

II. Background 

The earth is covered by two-thirds water and one-third 

space launch studies.16

—Secretary of the Air Force Sheila A. Widnall 

December 1992 

Regarding his predecessor, Daniel S. Goldin’s $1,000-perpound-to-orbit goal: I’m not an archeologist; I’m not a 

forensic pathologist. I don’t know where this stuff came 

from, and I’m really not interested in going through an 

excavation or a dig around the agency to figure out who 

came up with what number when. That is a bumper sticker, 

and I haven’t found anybody who can attest that there is 

any technology that can achieve that.17

—Sean O’Keefe, NASA Administrator 

November 2002 

HISTORICAL PERSPECTIVE: THE ROAD TO 

NOWHERE 

In the late 1960s, two very different approaches emerged as

potential means of significantly reducing the cost of access to low-earth

orbit. One approach proposed using simplified expendable boosters; the

other a winged, fully reusable, manned launch system.

18 In 1972, the U.S. 

government officially came down on the side of the winged, fully reusable

system. The space shuttle was established as America's future space 

launch vehicle and a solution to high cost. Although the ultimate shuttle 

configuration was only partially reusable, cost analysts at that time

nevertheless predicted an order of magnitude reduction in launch costs to 

$162 per pound to orbit (FY71) relative to existing expendable systems.19 

The space shuttle proved to be an engineering marvel providing a wide

range of on-orbit capabilities; however, as a launch system it was, and 

continues to be, an economic failure. In order to get the funds to build the 

shuttle during the lean 1970s, NASA was forced to trade higher 

operational costs in the future for lower developmental cost in the present.
10...Decision Maker’s Guide

The ramifications of that political trade-off eventually came home to roost, 

as shuttle operations have dominated and continue to dominate the NASA 

budget.20 The shuttle is most expensive heavy-lift launch vehicle in the 

U.S. inventory, whether based on cost per launch or on dollars per pound 

to orbit.21

In 1987, the DOD began a cooperative effort with NASA to 

develop a new, simplified rocket booster called the Advanced Launch 

System that was intended to succeed where the shuttle had failed. Once 

again, the goal was to achieve order of magnitude reduction in launch 

costs. In November 1987, Congress specified that any Advanced Launch 

System request for proposal would include the target of $370 or less per 

pound of payload to low-earth orbit.22 Budgetary and political pressures 

caused the original Advanced Launch System initiative to eventually be 

transformed into a follow-on program called the National Launch System. 

This system represented a family of simplified expendable boosters that 

depended on shuttle-derived hardware for some key components, and 

advocates claimed it held great promise for reliable, responsive space 

transportation. The National Launch System program de-emphasized the 

goal of $370 per pound to low-earth orbit. Program managers emphasized 

a simpler design approach to keep manufacturing and operating costs low,

but cost projections for the development of the new launch vehicle 

continued to rise.23 Congress cancelled the program in 1992 when nonrecurring development cost projections for the National Launch System 

exceeded $10 billion.24 The subsequent joint Space-Lifter program,

dubbed “Shapeshifter” by some due to its ability to mutate according to 

changing political demand, survived only a year and was cancelled in 

1993.25 

The post-Challenger environment was a politically turbulent and

uncertain one for space policy. Between 1987 and 1992, coinciding with 

the birth and demise of both the Advanced Launch System and National 

Launch System programs, this uncertainty manifested itself in an 

abundance of high-level space policy studies to include Pioneering the 

Space Frontier, the 1986 report of the National Commission on Space; the 

1987 Leadership and America's Future in Space report generated by a 

panel chaired by America's first woman in space, Sally Ride; the Space 

Architecture Study done by DOD in 1988; the 1990 Report of the Advisory 

Committee on the Future of the U.S. Space Program (the Augustine 

committee); the 1991 Synthesis Group Report (the Stanford committee); 

the 1992 reports from the National Space Council and the Vice-Presidents
Decision Maker’s Guide...11 

Space Policy Advisory Board, The Future of U.S. Space Launch 

Capability and a Post-Cold War Assessment of U.S. Space Policy, and the 

1992 National Research Council report From Earth to Orbit. All 

examined, and most stated as a priority, the imperative of low-cost earthto-orbit transportation.26 The National Commission on Space stated, for 

example, “the most significant contributions the U.S. government can 

make to opening the space frontier are to ensure continuity of launch 

services and to reduce drastically transportation costs.”27 One concludes 

that the need for change within the existing space transportation system 

was recognized, yet interagency/political consensus proved elusive. 

Meanwhile, an alternative to a rocket-based solution to low-earth

orbit access was being pursued. The U.S. Air Force and NASA initiated 

the X-30 National Aerospace Plane program as a single-stage-to-orbit 

vehicle using a supersonic combustion ramjet (scramjet) and slush 

hydrogen propellant. This decision was made despite the fact that a 

scramjet had never flown nor had slush hydrogen propulsion ever been 

successfully demonstrated. The program plan allocated $3.33 billion over 

eight years (soon raised to $5 billion) to build and test two vehicles (later 

cut to one) by 1990. The National Aerospace Plane was to enable aircraftlike operations with a 24-hour turnaround time and a 100-person ground 

crew. As the program progressed, the scramjet technology, thermal 

protection system, and a host of integration challenges became

problematic, causing Defense Secretary Richard Cheney to review the 

program in 1989. At this point, first flight had slipped a decade, and total

costs estimates were as much as 500 percent over initial estimates. 

Secretary Cheney terminated DOD investment in the National Aerospace

Plane, killing the program.

28 

In 1996, NASA began the X-33 competition for the nextgeneration Space Shuttle. The Lockheed-Martin Skunk Works was 

awarded the contract to build a wedge-shaped, vertical take-off horizontal 

landing lifting body powered by a linear-aerospike engine and 

incorporating a metallic thermal-protection system and composite fuel 

tanks (all unproven technologies). The sub-scale unmanned X-33 

demonstrator was designed to reach Mach 15, a velocity adequate to 

validate the technology necessary to build an orbital reusable launch 

vehicle. The X-33’s first flight was originally scheduled for March 1999;

however, numerous technical challenges slipped the launch date and added 

costs. The failure of a composite liquid hydrogen tank forced the use of a 

heavier aluminum alternative and destroyed what little design margin
12...Decision Maker’s Guide

existed. Scalability and traceability to the orbital reusable launch vehicle

was in serious doubt. A September 2000 revision of the NASALockheed-Martin agreement posited first flight in 2003, contingent upon 

the X-33 winning funding from the new Space Launch Initiative. Funding 

was not forthcoming, and the X-33 program died in late 2001 when the

USAF, after a six-month review, chose not to resuscitate it. NASA had 

spent $912 million and Lockheed-Martin an additional $365 million on the 

X-33 before it was finally cancelled.29 Also in 2001, NASA deferred 

further work on a smaller X-37 spaceplane prototype and killed the airlaunched X-34 Mach 8 hypersonic demonstrator that had gone over 

budget.30 However, The X-37 was resurrected in November 2002 when

NASA awarded a $301M contract to Boeing Phantom Works to continue 

the development of the X-37 Approach and Landing Test Vehicle as well 

as the design of the long-duration orbital vehicle.31 

NASA unveiled the $4.85 billion Space Launch Initiative, often 

referred to as the second-generation reusable launch vehicle program 

(shuttle replacement), in May 2001.32 It was a near-term (2001-2006) 

business plan for NASA and its partners, to include DOD, to investigate

new space transportation architectures and advanced technologies required 

to profitably implement them.

33 Stated objectives were: 1) “Invest in 

technical and programmatic risk reduction activities, driven by industry 

needs, to enable full-scale development of commercially competitive, 

privately owned and operated, Earth-to-orbit reusable launch vehicles by 

2005; and 2) Develop an integrated architecture with systems that build on 

commercial Earth-to-orbit launch vehicles to meet NASA-unique 

requirements that cannot be economically served by commercial 

vehicles.”34 

NASA solicited and initially received hundreds of proposals and 

selected 15 for further consideration in April 2002. Among the 15 rocketbased concepts, five were based solely on hydrogen fuel, another five used 

a mix of hydrogen and kerosene, and the last five tapped hybrid mixtures

and unconventional launch concepts.35 On 21 October 2002, NASA 

indefinitely suspended a system requirements review slated for the 

following month that would permit a down select and pursuit of three 

Space Launch Initiative designs through 2005.36 A total of $2.3 billion 

was cut from the Space Launch Initiative, with the decision on whether to 

develop the vehicle at all pushed back from 2006 to no earlier than 2009. 

The NASA justification says that they “had hoped to pay for the new 

vehicle by amortizing its estimated $10 billion development cost across 
Decision Maker’s Guide...13 

the commercial and NASA launch market. The assumptions proved too 

optimistic given the declining launch market.” Further, four independent 

cost estimates projected the total development cost of the new reusable 

launch vehicle at $30-35 billion.37 

The USAF did a better job, albeit with a significantly more modest 

goal, to successfully upgrade the Delta and Atlas expendable launch 

vehicle families (Delta IV and Atlas V) to initial operational capability in

2002. In the 1994 Space Transportation Study chaired by Air Force 

Lieutenant General Thomas Moorman, roadmaps for four space launch 

options were developed: “maintaining the status quo; undertaking a 

limited evolution of current systems or components; beginning a new, 

clean-sheet expendable launch vehicle development; and developing a 

reusable launch vehicle.”38 However, no specific recommendations were 

presented with the report. Eventually, the Evolved Expendable Launch 

Vehicle—the evolutionary approach—was initiated as a politically 

acceptable response. The primary motivation was to provide responsive, 

“assured access” to space at reasonable cost. It was never intended to be a 

cure for the transportation problem, rather only a treatment to permit a

cure to be found.39 The Air Force awarded $500M to both Boeing and 

Lockheed-Martin in October 1998 to develop their respective launch 

vehicle families and supporting launch infrastructure. The contract also

included options to purchase 19 launches worth $1.38 billion to Boeing 

and $650 million to Lockheed-Martin.40 The Evolved Expendable Launch 

Vehicle program goals included: 1) reduce launch costs by 25-50 percent, 

2) launch with 98 percent design reliability within 10 days of scheduled 

launch, and 3) capable of launching pre-integrated payloads within 45 

days of government notification.41 Current Air Force estimates are that it

will pay between $75 and $150 million for medium and heavy-class 

payloads respectively.42 Both the Delta IV and Atlas V had successful 

launch debuts in 2002. 

Yet today the space transportation situation remains virtually the 

same as it was in 1985, with only the modest improvements of the 

Evolved Expendable Launch Vehicle as the notable exception. In fact, 

within the reusable launch vehicle arena, the situation has deteriorated 

considerably since the inaugural launch of the shuttle in 1981. The shuttle 

was the first attempt beyond totally expendable rocket designs, which 

initially became operational in the 1950s. If space transportation is

assumed to drive and leverage cutting-edge technology, something is 

terribly askew. Clearly it is not another launch study that is needed. 
14...Decision Maker’s Guide

Why has so little progress been made within the realm of space

transportation for the past thirty years? First, a new space transportation

system that truly delivers R2

ISA is not really necessary to continue limping 

along and perpetuating the status quo. Between the Shuttle, Titan, Delta, 

Atlas, and other international expendable launch vehicles in existence, the 

current payload manifest can be easily met. Second, “incrementalism”

43

has permeated U.S. space policy. Elements of the government 

bureaucracy remain locked in debate over policy, requirements, 

technology, and resources that derails efforts toward rational decisionmaking. Unfortunately, the aerospace community, to include government

(DOD and NASA), academia, and industry, has done little to clarify the 

issues and help settle the debate. What is needed, more than all else is a

rational, deliberative, and enduring effort to solve the space transportation

mess. Outsourcing the problem to industry (discussed in detail in chapter 

3) won’t work. The systems are too expensive and the markets too thin to 

support this at present. Additionally, the status quo ensures that space 

launch maturation will remain stalemated, and future space exploration 

and exploitation will remain stagnant. 

Until recently, the government was assumed to bear the burden of 

all serious space transportation solutions, with NASA playing the 

historical leadership role in that endeavor. Although NASA’s agenda

was not always synchronized with commercial or military needs, (as 

repeatedly demonstrated by strife over joint programs to include 

Advanced Launch System, National Launch System, Space-Lifter, X-30, 

and Space Launch Initiative to name a few) it was intended that it be 

endowed with the resources and technical expertise to lead these efforts. 

NASA has had great success with the latter, but the resource challenges 

have worsened dramatically. The continuing operational challenges 

associated with the shuttle and the international space station will continue

to hobble NASA. Both are projected to consume more than 51 and 48 

percent of NASA’s annual $14+ billion FY03 and FY04 budgets 

respectively.

44 This results in an inability to fund its own proposed 

initiatives or demonstrate the wherewithal to see promising technology 

demonstrations to completion. 

The recent Space Launch Initiative debacle, and NASA’s decision 

to defer any decision on a second-generation reusable launch vehicle to no 

earlier than 2009, left unanswered, removes the possibility of R2

ISA for at 

least another 20 years. It is unclear whether this decision was based solely 

on current fiscal pressures, the agency’s inability to articulate a clear and 
Decision Maker’s Guide...15 

convincing case to its own administrator as to how $1000 per pound to 

low-earth orbit was possible, or the $30-35 billion price tag. In any case, 

there is no credibility to NASA’s vision for a return to the moon and 

subsequent human exploration of the solar system until it effectively 

addresses the fiscal and managerial burden of the shuttle and the 

international space station and secures R2

ISA as the key enabler for its 

space exploration initiatives. Put simply, at least for now, NASA has been 

forced to abdicate its historical role as a catalyst for U.S. space launch 

system innovation and development. 

The DOD could be allowed to pursue its own specific space launch 

systems that satisfy specific military requirements, and left to its own 

devices will likely do so. Based upon the Air Force’s commitment to the 

Evolved Expendable Launch Vehicle program, there is reason to believe it 

would stay the course on whatever future solution it chooses. However, is 

this the most prudent path for the nation? Certainly a strictly military

solution would have limited civil and commercial benefit. A successful 

national “compromise” solution to the stalemate may be the best way to

permit NASA to regain its footing and deliver valuable military and

commercial capability as well. Is a national solution to the current 

impasse possible? Does it make sense? The remainder of this paper will 

endeavor to answer these questions. 

DIVERGENT REQUIREMENTS 

There are unique requirements to civil, military, and commercial 

space launch. Differences in the areas of payload size and weight, launch 

rate, payload integration, mission turn-time, and cost can be significant. A 

military vehicle is driven to high sortie rates, smaller payloads with

maximum mission flexibility, and minimum integration time. Civil 

(primarily NASA) requirements include medium to large payloads, lower 

launch rates, and deliberate/predictable payload integration. Although 

cost currently permeates all three sectors, not surprisingly, commercial 

launch vehicle service providers consider manageable costs the highest

priority. Examination of all three sectors is intended to establish some 

“common ground” from which a national solution might prove possible. 
16...Decision Maker’s Guide

Civil & Military Convergence? 

Figure 3 highlights some of the important differences in civil and 

military requirements along functional lines. Taken at face value, one 

could quickly conclude that these “requirement” sets irreparably drive 

NASA and the DOD to two completely different system solutions, where 

collaboration beyond mutually supporting technology initiatives is both 

technically unfeasible and operationally unwise. Although this thinking is 

deeply entrenched in both the civil and military space communities, there 

are “joint” solutions that merit serious consideration. It will become

evident later in this paper that the military’s need for responsive

operations, rapid turn time, and high sortie rate are the key operability 

elements that deliver order of magnitude reductions in cost. NASA’s 

experience with the shuttle will clearly illustrate this point. Further, the 

apparent differences in payload delivery can be ameliorated through other 

elements of the space transportation architecture. 

NASA’s needs fall into two general categories: 1) launch of 

satellites for environmental monitoring and planetary exploration, and 2) 

launch of astronauts and payloads in support of the manned space flight 

program to include the International Space Station. Each of these

currently require medium to heavy-lift payload capability to low-earth 

orbit. The added complexity associated with a crew-rated system may 

also need to be considered. At present, NASA is publicly committed to 

the development of an orbital space plane45 and upgrades to the shuttle to 

keep it flying until at least 2012.46 International Space Station logistical 

support may also require a heavy-lift capability that is currently being 

serviced by four Space Shuttle missions annually, at $800M per launch.47
Decision Maker’s Guide...17 

Figure 3. Civil and Military Requirements Comparison48

Military space mission areas include space force enhancement, 

space forces support, space control, and space force application resulting 

in a diverse set of specialized payloads.49 DOD payloads often use 

commercially provided medium-payload-class expendable launch vehicles

which deliver a wide variety of force enhancement systems to include 

navigation, communication, environmental, and 

information/surveillance/reconnaissance satellites. Future trends in the 

force support arena point to the proliferation of smaller satellite systems to 

include “gap filler” systems launched in time of crisis. Defining payload 

requirements for space control missions is problematic. The Air Force has 

publicly articulated its intention to develop offensive and defensive 

counterspace systems in the future.50 Virtually all of these systems will be

developed in a highly classified realm. However, it is generally accepted 

that these payloads will be relatively small (smallsat/microsat class), and 

RAND research studies indicate that there are a variety of useful smalllaunch-vehicle-class space control payloads possible.51 Force application
18...Decision Maker’s Guide

from or through space to effect the terrestrial battlespace—particularly 

lucrative, time-critical, remote, hardened, or heavily defended targets

beyond the capacity of terrestrial-based aerospace forces—is an attractive 

option for space planners. Some proposed future systems that might 

accomplish these missions could be quite large, with space-based kinetic 

and directed energy systems requiring heavy-lift low-earth-orbit payload 

delivery and other concepts demanding small to medium-class 

capability.52 

RAND concluded in 1996 that a military vehicle capable of a

1,000 to 5,000 lb payload delivery capability may be sufficient for most of 

the space control missions. Further, many of these military missions 

would require launch-on-alert within minutes to hours, lack launch 

predictability, and demand a rapid turnaround and launch reconfigurability 

to be most effective. These characteristics imply aircraft-like operations

to include alert status in times of crisis. Finally, such levels of 

responsiveness demand aircraft-like supportability and reliability achieved 

with the smallest vehicle possible.53 RAND considered the feasibility of a 

Trans-atmospheric Vehicle, or space plane, as a flexible approach to 

satisfying these requirements, where flexibility is defined as the capability 

of a space plane to deliver payloads to a variety of orbits and to operate 

from a number of different bases.54 Other considerations included launch 

infrastructure, reentry cross-range capability, and ability to perform other

missions, such as that of a long-range bomber. RAND concluded that a

first-generation space plane, once demonstrating reliable operations, could 

provide significant long-term cost savings in terms of reduced launch 

costs.55 Such a system currently appears attractive for an important

segment of the future DOD space launch requirement. However, it fails to 

address some of the larger, more recently recognized payloads concepts

such as the Space Maneuver Vehicle or Common Aero Vehicle that may

follow. 

Where might common ground between NASA and DOD exist?

The “120-Day Study” chartered by the Secretary of the Air Force and the 

NASA administrator in October 2001 was a joint NASA-USAF effort to

develop a “credible, comprehensive plan for the joint development of the

next generation of reusable launch vehicles.”56 The subsequent “Red 

Team” study observation was that the study was chartered to address a 

“point solution,” consistent only with overarching NASA secondgeneration Space Launch Initiative objectives and that a broader Air Force 

analysis of alternatives was prudent.57 Although it must be conceded that 
Decision Maker’s Guide...19 

such direction inevitably narrowed the trade space, NASA and the Air 

Force did find agreement in operational requirements to include: initial 

operational capability between 2012 and 2014, reliability between 1:750 to 

1:1000 (probability of failure), full payload deorbit mass, abort to orbit, 

rendezvous capable, and full access to a full range of orbital inclinations

(equatorial to polar and sun-synchronous). Overlap in the areas of payload 

mass and responsiveness is even more significant. USAF weapons and 

information, surveillance and reconnaissance preliminary requirements 

fell between ten and fifteen thousand pounds to low-earth orbit. NASA 

science and international space station support ranged between ten and 

fifty thousand pounds to low-earth orbit (or the equivalent of five to 

twenty thousand pounds delivered to the International Space Station).58 

The team concluded a common booster with a payload range between 

25-45 thousand pounds to low-earth orbit was possible, and a wider

payload range could permit a common orbiter.59 NASA expressed a desire

for a 48-hour call-up for crew rescue purposes in contrast to the Air 

Force’s 12-24 hour call-up for contingency operations.60 Overall, the team

concluded that “architecture options were identified that meet USAF and 

NASA needs.”61 

The Air Force believes it needs to further refine space missions,

requirements, and concepts of operations before it can commit to any joint 

DOD/NASA reusable program, although a NASA deferral of any decision 

regarding a next-generation reusable launch system appears to make this

point moot. It is noteworthy however, that the 120-Day Study Red Team 

also observed that “multiple reusable launch vehicle development 

programs was most likely unaffordable and recommended that NASA and 

DOD leadership should both commit to some flexibility on requirements 

in order to control costs.”62 Certainly the RAND and 120-Day Study 

Team conclusions bound the Air Force payload requirement between one 

and fifteen thousand pounds to low-earth orbit. The articulated NASA 

requirement exceeds forty thousand pounds to low-earth orbit. However, 

NASA’s decision to build an orbital space plane that is launched from an 

Evolved Expendable Launch Vehicle greatly reduces this requirement. 

The two remaining missions are cargo delivery to the International Space 

Station (currently Russian Progress missions delivering 4,000 pounds of 

cargo) and earth observing/interplanetary missions ranging from two 

thousand to twelve thousand pound payloads.63 Exploitable overlap now

exists between civil and military payload requirements. 
20...Decision Maker’s Guide

Commercial 

The most important attributes of any commercial launch system 

include 1) reducing costs to compete effectively in the worlds market, 2)

performing to scheduled launch manifests, 3) developing payloads sized to 

the market, and 4) providing preplanned flight profiles without major 

anticipated changes. Most commercial satellites today require medium to

heavy-lift launch vehicles. The relatively low global demand for

commercial launch services (33 actual and 32 forecast worldwide 

commercial launches in 2002 and 2003, respectively)64 and the crowded 

launch services market, which includes U.S., French, Russian, and 

Chinese, make the launch services industry very competitive. The vast 

majority of commercial missions require a geostationary transfer orbit in 

lieu of low-earth-orbit mission orbits that are the equivalent of a 20,000 to 

45,000 lb low-earth-orbit class delivery. 

The Commercial Space Transportation Study examined a plethora 

of potential commercial markets to include communications, space 

manufacturing, remote sensing, unique civil/military missions, 

transportation, entertainment, space utilities, extraterrestrial resources,

advertising, and new missions.65 Although there was significant 

variability among system architecture requirements between segments, 

some important common system attributes and requirements emerged and 

are summarized in Table 1. 

Table 1. Commercial Space Transportation Study Attributes and 

Requirements66

Category Attribute Requirement Range

Dependability High probability of launching 

on schedule 

99% within scheduled 

hour 

99.999% within 

scheduled day 

Schedule Minimum advanced booking 

time 

6 months to 24 hours 

Reliability Equal to or greater than 

existing system 

1/100000 probability of 

loss 

Cost Minimum cost per launch Less than $1000/lb to 

orbit 

minimum $10,000 per 
Decision Maker’s Guide...21 

event 

Operations Standardized and simplified 

payload interfaces 

Maximize payload 

capability with 

customer base 

Capabilities - Support multiple payload 

classes 

- Provide delivery to multiple 

destinations

- Provide on-orbit rendezvous 

and docking 

- Provide delivery and return 

capabilities 

3000 pounds suborbital package delivery 

to 7,000 pounds to 

geostationary transfer 

orbit 

Availability High probability that the 

system will be in an 

operational rather than standdown state 

90% to 99.9% 

Responsiveness Minimum response time for 

launch on need 

30 days to 24 hours 

There is a striking similarity between these commercial 

requirements and those articulated for the military case. For example, fast 

package delivery was the most operationally stressing case examined by 

the study and drove the most aggressive requirements captured in Table 1. 

It was estimated that a 3000-pound cargo capability with a 10,000 nautical 

mile capability could capture between 70,000-1,000,000 pounds of cargo 

delivery at $1000/lb and 1,000,000–100,000,000 pounds at $100/lb based 

upon 1991 prices and markets.67 One can reasonably conclude that a 

militarily suitable reusable launch vehicle that pierced the $1000/lb to 

low-earth-orbit cost threshold may have significant commercial viability. 

BROKEN PROMISES and LESSONS LEARNED 

Three experiences with reusable launch vehicles, one resulting in a 

flight vehicle (shuttle) and two that did not (X-30 and X-33), provide a 

unique opportunity to learn some important lessons. Closer examination 

of these programs will make it apparent that there is much more to high 

shuttle costs than complexity and a lack of true reusability, just as 
22...Decision Maker’s Guide

immature technology is not the only culprit behind the demise of the X-30 

and X-33. 

Space Shuttle 

As the Apollo program was approaching its zenith, NASA’s next 

goal was to first build and deploy an already mature space station design

while simultaneously beginning the development of a fully reusable twostage-to-orbit space shuttle. Shrinking budgets forced NASA to postpone 

space station development when it became apparent that it was not

economical without a low-cost supply system. Hence, shuttle was at the 

forefront of development in the 1970’s, with the space station shelved for 

at least a decade. In July 1970, NASA awarded contracts to North 

American Rockwell and McDonnell Douglas to design a shuttle with a

25,000 lb to low-earth orbit payload and 200-1500 nautical mile crossrange capability. Further in-house studies that year prompted NASA to

opt for a delta-winged orbiter with a 65,000 lb low-earth orbit payload and 

1500 nautical mile cross-range capability in 1971. This configuration 

came under intense criticism in both congress and scientific communities. 

The $10 billion developmental cost for this configuration was considered

too expensive, so alternative booster designs were studied. NASA decided 

to scrap the fully reusable shuttle for political, technical, and economic 

reasons and scrambled to complete a new plan within six months in time

for fiscal year 1973 appropriations. The final “thrust assisted orbiter 

concept” selected in 1972 consisted of a manned orbiter, expendable 

external tank, and reusable solid rocket motors characteristic of today’s 

shuttle system. 68

Cost analyses conducted in 1971 at the request of the Office of 

Management and Budget (OMB) estimated development costs of a fully 

reusable shuttle at $12.8 billion ($56.8B FY02) and stated that the 

development of a less expensive $5.15 billion ($22.9B FY02) partially 

reusable system was justified within a level of space activity between 300 

and 360 flights between 1979 and 1990. The estimated cost per flight 

was $10.5 million (FY71) ($46.6M FY02) based upon a flight rate of 50 

per year and a first launch in 1978. By 1980, development costs had 

increased 20 percent to $6.2 billion (FY71) ($27.5B FY02) and a cost per 

flight to $15.2 million (FY71) ($67.5M FY02). The spectacular firstflight of Columbia occurred on 12 April 1981. A maximum annual flight 

rate of nine had been demonstrated in 1985 before the catastrophic loss of 
Decision Maker’s Guide...23 

Challenger and her crew on 28 January 1986. Extensive redesign and 

system improvements/upgrades were incorporated into shuttle, resulting in 

maximum payload reduction from 65,000 lbs to 53,700 lbs.69 

The fundamental question is this: How could initial 1971 shuttle 

launch cost estimates of $46.6 million (FY2002) equating to $717 per 

pound to low-earth orbit have been more than an order of magnitude in 

error? At the risk of oversimplification, the answer lies with a large 

annual flight rate shortfall illustrated in the Figure 4 below. Once the 

magnitude of this impact is understood, attacking its root causes goes a

long way towards enabling R2

ISA. The original shuttle orbiter 

maintenance turnaround operations envisioned a simple to operate and 

maintain vehicle, very little infrastructure, simple payload integration, and 

low labor intensity enabling a flight rate of forty per year.70 This “vision”

of operations was rendered circa 1974, prior to a detailed systems 

definition. The shuttle flight and ground support architectures matured as 

first flight approached and grew in complexity to meet the servicing,

inspection and checkout required by the vehicle design. As a result,

spacelift performance expectations (top of the graph which is a product of

the flight rate multiplied by the single mission lift capability) versus its 

actual performance (the lowest area on the graph) in Figure 4 diverged 

radically.71 Emblematic of this performance shortfall is the current need

to remove and replace 50-100 line replaceable units between each flight 

due to failures found in flight (10 percent), on the line (55 percent), or 

while under test or inspection (35 percent), respectively.72 A stable flight

rate of about eight per year for a fleet of four vehicles had been achieved 

through 1997 with a decline to four per year by 2002. Likewise, the 

single lift capability had not met expectation, dropping from the original 

65,000 lbs concept to 50,000 lbs actually fielded for operation. Although 

the vehicle performance shortfall was small compared to the flight rate 

shortfall, the combined effect had tremendous implications on the shuttle 

flight and ground architectures, undermining total spacelift performance.73 

The flight rate shortfall can be attributed in part to the design

compromises made to lower developmental costs as well as the NASA 

approach to forgo “Y-prototype” development that permits the rigorous 

testing and experience necessary to inject operability and maintainability

improvements into the objective system design. Hence, an important 

design parameter that should be examined, estimated, and ultimately 

verified before production of the objective system is the “single vehicle” 

capability. Verification is accomplished with frequent flights of the 
24...Decision Maker’s Guide

Y-prototype in an operationally stressing environment. Arguments that a 

“stressing” high single vehicle capability rate drives unnecessarily high 

cost into prototype development fail to recognize the importance of flight 

rate capability and fall into the same trap graphically illustrated by shuttle. 

These issues emphasize the importance of the conceptual design phase as 

the first opportunity to establish first-order-of-magnitude ground 

infrastructure and cycle turn-around time requirements that are ultimately 

verified in a Y-prototype. Any launch system’s total flight rate and 

vehicle performance (payload capacity) drives total payload throughput 

and defines a system’s overall spacelift performance capability.74

Figure 4. Shuttle Spacelift Performance —Vision vs. Reality 75

X-30 and X-33 

The X-30 National Aerospace Plane program of the late 1980’s 

and early 1990’s was preceded by a program of the same name in the late 

1950’s and early 1960’s; both were cancelled before a flight vehicle was 

ever built. The first program demonstrated a number of important 

technologies such as real-time air liquefaction and hypersonic refueling.76 
Decision Maker’s Guide...25 

The second program, X-30, called for a single-stage-to-orbit, fully 

reusable system based upon a complex combined cycle propulsion concept 

with several air-breathing components. The original program goal was to 

insert a manned, air-breathing, single-stage-to-orbit vehicle into low-earth 

orbit. However, the high risks associated with the propulsion concept, as 

well as other vehicle design aspects, prevented it from proceeding beyond 

the technology development phase.77 At the time, computational fluid 

dynamics was not sufficiently advanced, nor were ground test facilities 

sufficient (upper limit was Mach 10) to preclude the need for extensive 

flight-testing. Further, the vehicle’s depressed ascent profile (to permit air

ingestion) resulted in high skin temperatures and potential heating of 

internal structure and components. This environment demanded an 

advanced thermal protection system that included active cooling of 

leading edge surfaces. The combined cycle engine was to provide smooth 

transition from a slower subsonic/transonic mode to ramjet and eventual 

scramjet mode of operation to achieve orbital velocity. The Defense 

Science Board Task Force reviewed the program in 1988 and found six 

critical technology areas: aerodynamics, supersonic mixing and fuel-air 

combustion, high temperature materials, actively cooled structures, control 

systems, and computational fluid dynamics. The Defense Science Board 

concluded that the development schedule for all of these technologies was 

unrealistic.78 The program was cancelled after $1.7 billion was spent and 

it became clear that an operational prototype would cost $10 billion or 

more.79 Remnants became an advanced technology program.

80

NASA initiated the X-33 program in 1995 with the goal of 

demonstrating key single-stage-to-orbit technologies by the year 2000, 

leading the way for an eventual operational vehicle that could replace the 

space shuttle as well as existing expendable launch vehicles.81 The X-33 

was a subscale technology demonstrator intended to show scalability and 

traceability to a full-scale single-stage-to-orbit reusable launch vehicle. 

Lockheed-Martin Skunk Works was awarded an $837 million contract on 

4 July 1996 to design a lifting body vertical-take-off/horizontal-land 

vehicle using a linear-aerospike engine. NASA had budgeted an 

additional $104 million to support its own program infrastructure, with

Lockheed-Martin investing an additional $212 million for X-33 

development. Lockheed-Martin had estimated a fleet of two to three fullsized vehicles would cost between $4.5-5.0 billion at the successful

conclusion of the X-33 program.

82 The significant technical risks outlined 

by the Lockheed-Martin program manager, Dr. David Urie, at the outset of 
26...Decision Maker’s Guide

the program included vehicle integration, structures, propulsion, and 

thermal protection. To achieve single-stage-to-orbit capability, LockheedMartin would have to successfully overcome specific design challenges in 

the X-33 that included flight stability and control (the Lockheed-Martin 

design was aerodynamically unstable); a very lightweight, structurally 

efficient vehicle; and highly efficient, performance driven propulsion.83 

Economics played a dominant role in X-33 development. 

Lockheed-Martin intended to transform the X-33 into a commercially 

viable single-stage-to-orbit reusable launch vehicle that could compete 

successfully against expendable launch systems. Corporately, LockheedMartin had to achieve revenue and profit objectives within corporate 

capital investment constraints and reasonable reusable launch vehicle 

market demand. This included recouping sub-scale X-33 and full–scale 

Venture Star development costs within a reasonable time frame at an

acceptable investment rate of return.

84 The following business objectives 

had to be satisfied to make a reusable single-stage-to-orbit vehicle 

economically viable: 1) having the first reusable launch vehicle to enter

the marketplace; 2) building a reliable launch vehicle that had a successful 

first flight; 3) meeting market-based “cost-per-pound-to-orbit” pricing 

targets; 4) designing for low operations costs; 5) establishing long-term 

cash flows and a predictable launch rate; 6) lining up customers; and 7) 

establishing good returns on a space-port type launch services.85 The X33 began development at a time where a fairly robust demand for launch 

services prevailed and was expected to continue. Based upon the 

aggregate of all three competitors’ mission models, average demand 

varied between 32 and 46 space launches per year. At that time, this 

represented nearly the combined U.S. government and commercial 

launches projected by the 1994 Moorman panel and implicitly assumed an 

anchor tenancy (guaranteed market access) for NASA and DOD 

payloads.86 There was also agreement among the three competing

contractors that DOD payloads in excess of 20,000-pound low-earth-polar 

orbits fell within the Titan-IV heavy-lift vehicle class and was outside the 

design limits of a marketable reusable single-stage-to-orbit concept. All 

three X-33 concepts were designed to capture the majority of the Delta 

and Atlas class payloads.87 

By 1999, technical problems with the vehicle’s internal composite 

fuel tanks, linear-aerospike rocket engine, and thermal protection system 

had, in turn, precipitated cost increases, revision of key performance 

objectives, and significant delays in the vehicle’s flight test schedule.88 
Decision Maker’s Guide...27 

After the failure of a composite liquid hydrogen tank in ground testing, 

NASA restructured the program as a competitor for Space Launch 

Initiative bidding. The X-33 failed to secure NASA funding after an 

agency-wide review concluded that the costs of continuing the program

outweighed the benefits they were expected to produce. Mr. Art 

Stephenson, director of NASA Marshall Space Flight Center concluded, 

“One of the lessons learned is that our technology has not yet advanced to 

the point that we can successfully develop a new reusable vehicle that 

substantially improves safety, reliability and affordability.”89 Dennis 

Bushnell, Chief Scientist of NASA Langley Research Center, concluded 

that a key lesson learned from both the X-30 and X-33 programs is that 

“revolutionary goals require revolutionary technology and a ‘going in’ 

large ‘cushion’ in terms of expectations versus metrics.”90 Both of these

perspectives place a premium on technology and claim that the technology 

isn’t ready. Technology is indeed a key enabler for R2

ISA; however, there 

are other equally important facets of space launch vehicle design directly

impacting system maintainability and operability that cannot be ignored. 

The space shuttle space-lift performance shortfall is a case in point. 

Additional Lessons Learned 

A recent paper from Booz Allen Hamilton91 cast a larger net by

examining the experience base from a broader set of past reusable launch 

vehicle projects,92 as well as other successful aerospace endeavors,93 to 

make a set of technical/programmatic and “environmental” project 

comparisons. The analysis and conclusions from this paper not only 

reinforce the shuttle, X-30, and X-33 experiences, but also add additional 

insight as well. A synopsis is highlighted in Tables 2 and 3 below. 

The most intriguing observation that can be made regarding these 

technical/programmatic lessons learned is that they could describe any 

high-risk government acquisition program, many of them successful. This 

in turn implies that the space shuttle performance shortfall and the X-30

and X-33 program failures were avoidable. Either the “system” should 

have permitted a better risk assessment and more realistic funding, or they 

should never have been funded in the first place. 
28...Decision Maker’s Guide

Table 2. RLV “Technical/Programmatic” Development Lessons 

Learned 

Characteristic Discussion

Technology In every case, including both successful and canceled

RLV programs, the technology required was more

difficult to develop than most proponents had 

originally forecast. Successful programs exceeded

original budget estimates to solve technology 

challenges and were supported to completion. 

Risk Cost, schedule, and performance are inextricably 

linked. Sufficient trade space must exist between them

to be successful. Low failure tolerance by 

management negatively impacts cost and schedule. 

Incremental versus substantial leaps in technology or 

performance is the prudent path. 

Requirements Solid mission requirement and/or clear market demand 

must exist. 

Credibility Overselling the capabilities of a vehicle results in loss

of credibility, leading to much more difficult funding. 

Shuttle, X-30, and X-33 were all oversold. 

Realistic Cost & 

Schedule 

Estimates 

Part of the overselling problem. Sometimes the result 

of honest underestimation but also from deliberate 

dishonesty. Typically, the largest cost overruns are 

associated with immature technology where precise

cost estimating is not possible. 

Organization 

Matters 

Successful X-vehicles have traditionally been

produced by organizations that were kept “lean” and

co-located as the scope of the project allowed. 

 
Decision Maker’s Guide...29 

Table 3. RLV “Environmental” Development Lessons Learned 

Lesson Learned Discussion 

Credible and 

Compelling 

Need 

A clear, credible, and compelling mission is essential. 

A solid business case to justify the substantial 

investment in lieu of technical and business risk 

associated with the development. 

National 

Commitment

Top-level, enduring commitment at the national level 

is necessary, as well as significant commercial and

public support to ensure project stability. 

Recognize 

Competing 

Interests 

A radical departure from the status quo can be viewed 

as threatening to organizations and institutions that 

benefit most directly from established approaches.

Very real political and industrial base considerations 

must be addressed. 

Realistic 

Expectations

The vehicle cannot be oversold. Although design 

variations may address additional missions, claiming at 

the outset that it will do all or most related space

missions will result in another Space Shuttle. 

Parallel 

Approaches

Parallel approaches to mitigate higher risk elements of

the supporting technology base are essential for 

program success. 

Small Steps 

versus Giant 

Leaps 

Unreasonable to expect too much from a demonstrator. 

No X-vehicle ever completed has had operational 

capability. Smaller steps associated with a clear,

sustainable evolutionary path are the proven approach. 

Focused 

Leadership 

While many organizations may contribute, leadership 

must be in the hands of a single organization, one not 

bound to legacy systems. History is clear that 

breakthrough aerospace projects have a program office 

or other organization with real authority and long-term 

commitments from the agencies and contractors

involved. 

The environmental factors highlighted above are political and 

institutional in nature. Due to the high front-loaded developmental costs 

inherent in R2

ISA, one must ask if indeed there is a compelling need—and 
30...Decision Maker’s Guide

whether an enduring national commitment can be reasonably expected. 

With the tragic loss of Columbia, the entire civil space program is under

review, with the future of NASA’s manned space flight program at the 

center of the debate. The DOD is clear in its position: “Space 

transportation represents the sine qua non of space power: unless 

sufficient lift capability becomes readily available at significantly less

cost, U.S. capabilities to place its projected systems on orbit in sufficient 

quantities to achieve mission objectives will increasingly lag behind 

demand. Major technological advances leading to improved launch 

capability will be needed to achieve the very first of USSPACECOM’s 

objectives for the future—Assured Access to Space—without which its 

other objectives may remain beyond reach.”94 American commercial 

space launch is in siege mode, with the current glut of global expendable 

launch capacity exceeding demand by a factor of 300-400 percent. The 

compelling need exists, but it must now be recognized, communicated, 

and acted upon by senior leadership. Experience has taught us that a 

successful reusable launch vehicle development program hinges as much 

upon a national commitment to a politically sensitive, realistic, focused, 

evolutionary solution as it does to making the appropriate design choices 

responsive to a compelling mission/market defined by clear and 

unambiguous requirements. 
Decision Maker’s Guide...31 

III. Defining the Problem 

The last ten percent of performance generates one-third the 

cost and two-thirds of the problems. 

—Augustine’s Law Number XV 

I think that if I’d been at Kitty Hawk in 1903 when Orville 

Wright took off, I would have been farsighted enough, and 

public spirited enough—I owed this to future capitalists—to 

shoot him down. I mean, Karl Marx couldn’t have done as 

much damage to capitalists as Orville did.95

—Warren Buffet 

This section introduces two key underlying elements of R2

ISA, 

namely: the Newtonian physics and thermodynamics that define the 

fundamental limits of space technology, and the linkages between the 

inelastic space transportation market (broadly defined as civil, military, 

and commercial) and the prohibitively expensive space technology upon 

which it is based. 

ROCKET SCIENCE 101 

The past thirty-year history of space transportation indicates that 

there has been a basic misunderstanding by some who should have known 

better that the laws of gravity and thermodynamics are non-negotiable. 

Ignore or defy them at your own peril. This section endeavors to briefly 

introduce, in as plain English as the subject will permit, the basics that 

must be understood and mastered to permit any meaningful technical 

discourse on the subject of R2

ISA. 

The introduction of several basic equations is necessary to best 

explain fundamental concepts and commonly used measures of merit. 

English units are used throughout the text and in the tables to maximize 

familiarity and comfort with these topics. Unfortunately, English units are

ill-suited for use within the context of modern science and engineering, as

the recent loss of the Mars Climate Orbiter in September 1999 will attest. 

Hence, only metric units are used in the equations below.96 
32...Decision Maker’s Guide

Energy is the Key 

The challenge associated with delivering any payload to its final 

mission trajectory/orbit can be reduced to two fundamental activities: 

first, the acceleration of a payload from the earth’s surface to orbital 

velocity; and second, the subsequent energy added to “raise” the orbit to 

its final state. Both elements involve adding large amounts of energy to 

the system, kinetic energy with the former, potential energy with the latter. 

Equation 1 quantifies this relationship. 

 

Orbital Energy = ⎥

⎦

⎤ ⎢

⎣

⎡

+ ∫

r

r gdr v

m

2 0

2

 and 

∫ = = + r

r gdr v

m

OrbitalEnergy

oe

2 0

2

where oe = specific orbital energy 

v = velocity ( 2m s ) 

g = acceleration due to gravity ( 2m s ) 

 = Earth radius, 6370 kilometers (km) 0r

r = r + altitude (km) 0

Equation 1. Specific Orbital Energy 

Figure 5 illustrates graphically that objects in a reference 100 

nautical miles (nm) (or 185 km) altitude low earth orbit (where 0 r r is a

mere 1.029) have specific orbital energy that is overwhelmingly 

comprised of specific kinetic energy. In fact, the ratio of kinetic versus 

potential energy is about 20 to 1. Alternatively, moving from low-earth to 

geostationary-earth orbit (36,200 km from the center of the earth where 

0 r r is 5.68), specific orbital energy is comprised mostly of specific

potential energy. The ratio of kinetic to potential energy is less than 0.1, 

or 1 to 10 for this case. This has profound implications for space 

transportation architectures. The challenge of earth-to-orbit launch 

systems is a kinetic one, where enormous energy is required to generate 

high levels of propulsive thrust in a short period of time to accelerate a

payload to very high velocity (in excess of 7800 m/s). This is an 

inherently inefficient and potentially dangerous process due to the high 
Decision Maker’s Guide...33 

energy densities required. Alternatively, additional energy added after 

orbital velocity is achieved can be accomplished in a much more 

deliberate fashion. Lower thrust systems operating at much higher 

efficiencies for longer duration provide invaluable flexibility (as well as 

innovation) in space propulsion that is not typically enjoyed in earth-toorbit launch systems. Ideally, any intelligent, optimized space 

transportation architecture built from the ground up would minimize the 

burden on the launch vehicle while fully leveraging the benefits of onorbit space propulsion to the maximum extent possible. For a wide variety 

of reasons, mostly historical, today’s space “architecture” is woefully suboptimized. For example, geostationary satellite delivery typically uses the

launch vehicle upper stage in a direct ascent into a geostationary transfer 

orbit. This is done despite the fact that a wide array of space propulsion

technologies could do so much more efficiently. However, due to a suboptimized architecture, these options are either impractical or simply do 

not exist. Any future space transportation architecture must address this

issue directly. 

Figure 5. Specific Orbital Energy97
34...Decision Maker’s Guide

Measures of Merit 

The most common propulsion performance parameter is specific 

impulse (Isp), with higher values indicating greater efficiencies. It 

compares the thrust generated by a propulsion system to the propellant

mass flow rate and can be considered a measure of thrust generation 

efficiency (much like gas mileage in your car). 

mg0

F I sp & = ; g0 c I = sp

where = specific impulse (s) sp I

 F = thrust magnitude (N) 

m& = propellant mass flow rate (kg/s) 

g0 = acceleration due to gravity, 9.807 ( 2m s ) 

c = effective exhaust velocity (m/s) 

Equation 2. Specific Impulse and Effective Exhaust Velocity 

Specific impulse permits the direct comparison of propulsion 

systems using different types of propellant (such as a hydrogen versus 

kerosene fueled liquid rocket engines), or dissimilar systems (such as a 

hydrogen/kerosene fueled liquid rocket and a supersonic combustion 

ramjet). The unit of seconds is disconcerting to some, but the concept is 

not that unlike miles per gallon as an efficiency measure for automobile 

internal combustion engines or specific fuel consumption of aircraft 

turbojet engines. However, the mass flow component of specific impulse

is different for rocket systems (where all of the propellant is contained 

within and consumed by the vehicle) and airbreathing systems (where

internal fuel is mixed with external air). The specific impulse calculations

for airbreathing systems do not need to account for the external air flowing 

through the engine, thus resulting in much higher effective specific 

impulse. 

Table 4 presents a synopsis of applicable propulsion technology 

for both earth-to-orbit as well as on-orbit applications. While more 

versatile chemical systems operate over a wide range of thrust levels, they 

do so at low to moderate specific impulse. Alternatively, high specific

impulse electric systems deliver very low thrust, while nuclear systems 

carry sufficient political and environmental concerns to make approval for 
Decision Maker’s Guide...35 

their use in a near-earth environment unlikely. Solar thermal propulsion is 

representative of a developing class of on-orbit space propulsion 

technologies where thrust levels are high enough to make them attractive

as orbital transfer vehicles that operate at specific impulse levels two to

four times higher than traditional chemical systems. 
36...Decision Maker’s Guide

Table 4. Propulsion Technology Performance Summary98

Technology

sp I (s) Thrust 

(lbs) 

Advantages Disadvantages

Cold Gas 

- N2

- H2

60 

250 

0.02-10 • simple 

• safe 

• low 

contamination 

• low specific 

impulse 

Chemical 

- Liquid 

 - Monopropellant

 - Bipropellant

- Solid

- Hybrid

140-235

320-460

260-300

290-350

0.02 to

2,700,000 

• high thrust

• heritage 

• moderate

performance 

• combustion

complications 

• safety concerns 

Nuclear 

- Solid Core 

- Liquid Core 

- Gas Core 

800-1100

3000

6000

Up to 

2,700,000 

• high specific 

impulse 

• unproven

• politically 

unattractive 

• expensive 

• low thrust/weight

Electric 

- Electrothermal 

- Electromagnetic 

- Electrostatic 

500-1000

1000-7000 

2000-10000

0.00002 

to 5

• very high

specific 

impulse 

• high system mass 

• low thrust levels 

• limited heritage 

Solar Thermal 700-1000 1-10 • high specific 

impulse 

• limited heritage 

Another useful performance parameter is effective exhaust velocity 

“c” where the force of thrust “F” in the specific impulse equation is

replaced with “ m& c ” and appropriate substitutions yield g0 c I = sp . Thrust, 

whether aircraft, rocket, or from a deflating toy balloon is generated 

through the acceleration of the exhausted propellant. Thrust is maximized 

for a given mass of propellant by accelerating it to the highest velocity 

possible, hence higher exhaust velocities typically indicate more highly 

efficient propulsion and provides another method of direct comparison 

between propulsion systems. Engine thrust-to-weight is an intuitively 

straightforward parameter that captures the “engineering” efficiency of an 

engine. It is the total thrust generated by the engine divided by the engine 

weight, where higher ratios are desirable. 

Table 5 highlights the range of performance both within and 

between chemical rocket and airbreathing propulsion systems suitable for 

terrestrial and earth-to-orbit applications. The very large increases in 
Decision Maker’s Guide...37 

specific impulse in airbreathing versus pure rocket-based systems make 

their use in space launch systems highly desirable. However, hypersonic 

airbreathers remain technologically immature, and the benefits of higher 

specific impulse are partially offset by lower thrust, thrust-to-weight, and 

limits to its operating range. 

Table 5. Key Propulsion Performance Characteristics 

Demonstrated 

Technology *Isp (s) 

Sea Level

Thrust 

(lbs) 

Thrust 

to

Weight

Range of

Operation 

Chemical Rockets

- Liq Oxy/Hydrogen (SSME)99

- Liq Oxy/Hydrogen (RS-68)

100

- Liq Oxy/Kerosene (RS-27A)101

- Solid

- Hybrid

454 

410 

302 

260-300

290-350

418,000 

608,000 

200,000 

59.8 

41.8 

79.1 

No

Restrictions 

(Sea Level 

to Vacuum) 

**Airbreathers 

- CF6-80C2B6 (B-767)

102

- F100-PW220(F-15C)103

- Notional Hydrogen Scramjet104

- Notional Kerosene Scramjet105

10770 

1713

1-3000

1000

60,800

23,770

6.3 

7.3 

0-Mach .9

0-Mach 2+ 

Mach 5-20 

Mach 5-8 

*Note that rocket Isp values are ideal vacuum performance. Seal level (launch) 

performance is typically only 80% of the ideal value. 

** Includes the mass of fuel only 

Other important vehicle “technology” metrics are the inert and 

propellant mass fractions as well as the structural ratio. The vehicle mass 

relationships are highlighted both graphically and mathematically in 

Figure 6. 
38...Decision Maker’s Guide

Figure 6. Two Stage Vehicle Mass Contributors 

prop inert

inert

inert

m m

m

f

+ = ; 

prop inert

prop

prop m m

m

f

+ = ; prop inert f = 1 ; − f 

payload

inert

m

m

R =

where = inert (structural) mass fraction inert f

prop f = propellant mass fraction 

R = structural ratio 

= inert (structural) mass (kg) minert

= propellant mass (kg) mprop

= payload mass (kg) mpayload

Equation 3. Launch Vehicle Mass Fractions 

The inert mass fraction ( ) is an excellent indicator of the 

overall structural efficiency of a launch vehicle, or any aerospace vehicle. 

Lower values indicate higher efficiency. The structural ratio (

inert f

R ) is the 

ratio between inert vehicle mass and payload mass, with lower numbers

approaching unity most desirable. Structural ratio is a direct measure of 

how efficiently overall vehicle dry weight is allocated to useful payload. 
Decision Maker’s Guide...39 

Table 6 below shows representative mass fractions and vehicle structural 

ratios for both expendable and reusable launch vehicles as well as aircraft. 

Table 6. Mass Characteristics of Aerospace Systems 

Vehicle 1st Stage 2nd Stage 

inirt f prop f * R inirt f prop f inirt f prop f

Delta II 0.06 .91 2.4 .055 .94 .137 .86 

Atlas II 0.06 .91 1.8 .062 .94 .11 .89 

Titan II 23G 0.04 .96 3.6 .033 .96 .092 .91 

Pegasus (L-1011 stage 0) 0.10 .90 5.0 .084 .92 .096 .90 

Space Shuttle System .134 .85 11.2 .15 .85 .12 .87

Space Shuttle Orbiter .73 .16 2.8 

X-15106 .42 .54 24

X-33 0.09 .88 3.3 

F-15107 .46 .28 1.7

B-777-300108 .52 .21 1.9

*Launch Vehicle payloads normalized to 100 nm 28.5 deg incl except Titan II (polar)

**Sum of vehicle inert and propellant mass fractions do not add to unity due to

inability to fully extract payload and wet vs dry weights from data sets 

 

There are some important trends revealed in Table 5 that warrant

comment. First, all of these are or were operational vehicles, with the 

exception of the X-33, which did not fly. Note how expendable, single 

use, space launch vehicles tend to have inert vehicle mass fractions

between 0.04 and 0.10, indicating very highly structurally efficient 

systems. At the other end of the spectrum you find highly reusable 

aircraft with inert mass fractions between 0.42-0.52. The higher mass 

fractions are primarily the result of heavier (less expensive) structural 

materials, higher margins of safety, and the operability and maintainability

inherent with high reusability. When one considers the amount of design 

work and resources that are allocated to keeping an aircraft as light as 

possible, one can begin to appreciate the implications of inert mass

fractions below 0.10. Now consider the required inert mass fraction

required for the performance driven X-33. The technologies envisioned 

for this vehicle were beyond existing state-of-the-art, with virtually no 

design margin available for future weight growth. The wide disparity 

between demonstrated mass fractions of expendable and reusable systems

imply that there are some formidable technological challenges in 
40...Decision Maker’s Guide

transitioning from expendable to reusable launch vehicles as well as 

making reusable launch systems “aircraft-like.” 

The ideal rocket equation is an elegantly simple equation that 

provides powerful insight into the first order concerns surrounding launch 

vehicle design and performance: 

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

∆ =

f

i

sp m

m

V g0 I ln = ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

+

+ +

inirt payload

inirt prop payload

sp m m

m m m

g0 I ln

where ∆V = change in velocity (m/s) 

= initial mass (kg) minitial

= final mass (kg) m final

Equation 4. Ideal Rocket Equation 

Here, represents the change in velocity (without accounting 

for losses due to gravity, steering, atmospheric drag, and earth’s rotation)

that a vehicle can attain based solely upon specific impulse and 

initial/final vehicle mass. This equation, coupled with a velocity budget 

associated with a given mission, defines the boundary conditions for 

launch vehicle design. An example of such a budget for ascent into 

selected low-earth orbits for rocket based launch systems is summarized 

below in Table 7. 

∆V

Table 7. Velocity Budgets to Low-Earth Orbits for Selected Launch 

Vehicles109

Vehicle Orbit: p a h xh

Inclination (deg) 

∆VLEO ∆Vgrav ∆Vsteering ∆Vdrag ∆Vrot Σ∆V

Delta 

7925

175x319

33.9 7842 1150 33 136 -347 8814

Atlas I 149x607

27.4 7946 1395 167 110 -375 9243

Titan IV/ 

Centaur 

157x436

28.6 7896 1442 65 156 -352 9207

Space 

Shuttle 

196-278

28.5 7794 1222 358 107 -395 9086

Units of ∆V in m/s. 
Decision Maker’s Guide...41 

The first term (∆VLEO) is calculated using the ideal rocket equation. 

The additional values are added velocity required to overcome gravity, 

steering, and drag. Note relative magnitudes of these values with both the 

ideal velocity required and the relative uniformity in the gravity and drag

losses among pure rocket trajectories. These losses increase for 

airbreathers that spend more time in the atmosphere before attaining

orbital velocity. Also note that ∆Vrot is negative because of the benefit 

these trajectory flyouts receive from the earth’s west to east rotation. 

Hence velocity budgets can be built and used to determine basic vehicle 

requirements (initial and final mass, propulsion efficiency, etc.) that 

deliver a desired capability (payload to a desired orbit). 

Figure 7 shows inert mass fraction as a function of specific 

impulse with two curves derived from the ideal rocket equation, 

representing ∆Vs of 8800 and 9300 meters per second (28,900-30,500 ft

per second), that bound the velocity change required to achieve low-earth 

orbit insertion. Real first stages (those on existing or historical launch 

vehicles) are plotted against these curves. The shaded region represents 

the trade space that is not feasible for single-stage-to-orbit vehicles. 

Figure 7 illustrates the impact of higher inert mass fractions

demanding higher specific impulse to attain orbital velocity. For example, 

a single-stage-to-orbit with an average specific impulse of 430 seconds 

(typical liquid oxygen/hydrogen rocket) must have an inert mass fraction 

of at least 0.125 (no more than 12.5% structure) to achieve orbit with no 

excess margin to deliver any useful payload. An added payload 

requirement would drive this number even lower. However, reusable 

launch vehicles must be much more structurally robust than their 

expendable cousins, demanding inert mass fraction of at least 0.20-0.30.

This would imply a requirement for a propulsion system with an average 

specific impulse in excess of 550 seconds, a technology not yet available 

for practical chemical rockets. Unfortunately, a reusable launch vehicle

with an inert mass fractions below 0.10 has yet to be demonstrated either. 

This type of basic analysis graphically illustrates the challenges inherent 

with any single-stage-to-orbit program and puts the technical hurdles 

confronted by the X-30 and X-33 into proper perspective. Thus, in 

proposing these two programs, either someone wasn’t asking the right 

questions, or someone else wasn’t telling. 

 
42...Decision Maker’s Guide

Figure 7. Feasible Regions for Launch Systems110 

Staging 

The ideal rocket equation makes it clear that certain specific 

impulse and vehicle inert mass fraction combinations make earth-to-orbit 

missions impossible. One way to overcome this limitation is through 

staging. For example, from Figure 7, one could surmise that a singlestage-to-orbit vehicle with a specific impulse of 290 seconds (a liquid

oxygen/kerosene combination) and aggressive inert mass fraction of 0.05 

cannot achieve orbit. However, staging permits the “discarding” of inert 

vehicle mass at some point in the trajectory. So, what is the right number 

of stages for a specific combination of technology and mission 

constraints? Figure 8 is based upon the ideal rocket equation and 

illustrates the benefit of staging for an orbital mission requiring a ∆V of 

9000 meters per second (29,527 ft/sec), inert (structural) mass fraction of 

0.08, and specific impulse of 420 seconds. 

Notice that in this example overall system mass is decreased by 54 

percent when adding a second stage and the marginal benefit of three or 

more stages. This is typical for a wide range of specific impulse/inert 

mass fraction combinations for earth-to-orbit launch vehicles and is why 
Decision Maker’s Guide...43 

most launch vehicles use two stages to achieve orbital velocities. 

Additional parallel first stages (such as solid strap-on boosters) are used to 

increase the initial vehicle thrust to weight ratio as well as payload

capacity to low-earth orbit. 

Figure 8. Staging Impact Upon Vehicle Mass111

Launch Vehicle Technology 

Technology and system integration are the “brick and mortar” 

upon which R2

ISA is built. It is universally acknowledged that the most 

challenging element to the development of any complex system is system 

integration—the creation of new system that is greater than the sum of its

parts. Surprisingly, there is also consensus within the space launch 

community where the bulk of the enabling technology development 

remains. As previously discussed, advances in four areas to include 

propulsion, advanced materials and structures, thermal protection (not 

relevant to expendable launch systems), in addition to vehicle integration 

are critical to achieving R2

ISA. 

Table 8 is a top-level structural, propulsive, and mechanical 

comparison of existing aircraft and reusable/expendable launch vehicles. 

This table lists some of the inherent engineering design characteristics of 

aircraft and space launch vehicles and highlights a few of the challenges in 

developing more “aircraft-like” space launch vehicles. The wide disparity 
44...Decision Maker’s Guide

in performance is in part an artifact of the inherent demands of space 

launch, but much of it is a direct result of the performance maximization 

mindset permeating U.S. launch vehicle development since the early days

of intercontinental ballistic missiles in the 1950’s. Performance 

maximized systems typically lack design margin, or the difference 

between designed and demanded performance, that in turn has a very large 

impact on vehicle reliability, robustness, and cost. 

Table 8. Aircraft to Launch Vehicle Comparison112

Characteristics Aircraft STS

(Orbiter) 

ELVs 

Structures: 

- Factors of Safety 1.5 1.4 1.25 

- Gross Liftoff Weight (Klbs) 618 4,426 1,888 

- Design Life (Missions) 8,560 100 1 

Propulsion: 

- Thrust (Vacuum, Klbs) 30-60 470 200 to 

17,500 

- Thrust/Weight Ratio 4.5 74 60 to 140 

- Operating Temperature (°F) 2,550 6,000 500 to 5,000 

- Operating Pressure (PSI) 140 2,970 500 to 1,200 

- Cruise Level Power 25% 109% 100% 

Mechanical: 

- Specific Horsepower (hp/lb) 2 108 3 to 18 

- RPM 13,450 35,014 5,000 to 

34,000 

The reasons for the wide range of structural mass fractions

between expendable and reusable systems can be seen here. Lower factors 

of safety,113 volumetric efficiency, and single use, are characteristics that

result in highly specialized, highly efficient expendable launch vehicles.

Much smaller aircraft represent the opposite end of this spectrum, 

demanding higher factors of safety and structural mass fractions (refer to 

Table 5) up to an order of magnitude greater. The inherent nature of 

earth-to-orbit operations demands very powerful (thrust) and lightweight 

(high thrust-to-weight) propulsion, which in turn requires demanding 

performance maximization that includes much higher operating 
Decision Maker’s Guide...45 

temperatures and pressures as well as operation at or on excess of design 

limits (cruise level power) to accomplish the mission. The impacts of 

cruise power are well documented. For example, the operational life the 

Space Shuttle Main Engine is reduced by nearly a factor of ten when 

operated at 109 percent of rated thrust. Reusable launch vehicle 

propulsion will need to operate well beneath its design limits to permit

aircraft-like operations in the future. The disparity in the mechanical 

demands placed upon rocket-based turbomachinery (pumps) versus

aircraft (turbines) is truly remarkable. The high rotational velocities 

coupled with the power that drives rocket turbo-pumps makes them the 

single most failure prone and potentially dangerous component on a space

launch system. The development of space launch vehicles that approach 

aircraft-like operations will need to have large performance margin built 

into its systems to minimize “redline” operations. 

Propulsion. The most important performance related propulsion 

characteristics are high thrust at high total system performance (specific 

impulse), high structural efficiency (thrust to weight), accomplished with 

the highest density propellants possible (smaller vehicle size), and robust 

performance margin. The operability requirements for rapid turnaround 

and low cost operation demand durability, damage tolerance, ease of

inspection, and capability for rapid and safe shut-down. Performance and 

operability inevitably come into direct conflict with each other. 

The highest specific impulse and thrust-to-weight rocket 

propulsion systems commonly used in space launch are liquid 

hydrogen/oxygen (LH2/LO2) fuel/oxidizer systems. There are several 

drawbacks to this propellant combination. First, the cryogenic properties 

of liquid hydrogen create unique fuel handling and storage problems. It is 

difficult to contain, prone to leaks, and always an explosive hazard. 

Further, its very low boiling point (-423° F) introduces unique material 

compatibility issues. Although all of these problems have been 

successfully addressed since the earliest days of modern rocketry, the 

added complexity makes a flexible, responsive launch vehicle and ground 

support system problematic. The second drawback to liquid hydrogen is 

its very low density, demanding much larger and often complex propellant 

tanks manifesting themselves into larger, heavier vehicles. Hence, despite

their lower performance, alternative higher density oxidizer/fuel 

combinations such as liquid oxygen combined with kerosene (RP-1), 

methane, as well as hydrogen peroxide (oxidizer) and kerosene are

attractive alternatives. Other more exotic and even higher density 
46...Decision Maker’s Guide

materials (HEDM) are under investigation. A synopsis of these propellant 

combinations is summarized below. 

Table 9. Properties of Candidate RLV Propellants114 

Propellant Tfp

(°K) 

Tbp

(°K) 

Pvap

(Mpa) 

Density

(kg/m3

) 

Stability / 

Handling/ 

Storage 

 Characteristics 

(O/F) 

Isp (sec) 

Oxygen O2 54 90 5.07 @

154°K 

1142 Good/

Good/ 

Cryogenic 

N/A 

Hydrogen H2 13.8 20.3 1.29 @

32.8°K 

71 Flammable/

Flammable/

Cryogenic 

(3.8) 

435 

w/O2

Kerosene 

(RP-1) 

CH1.97 229-

291 

445-

537 

2275 @ 

344°K 

810 Flammable/

Flammable/

Good 

(2.2) 

321 

w/O2

Hydrogen

Peroxide 

H2O2 267.4 419 345 @

298°K 

1414 Unstable >414°K 

Flammable/

1%/yr decomp 

(7.5) 

298 

(w/RP-1) 

Tfp = freezing point; =boiling point; =vaporization pressure; Tbp Pvap

(O/F)= oxidizer to fuel ratio; 

sp I

= Vacuum Specific Impulse; 0°K=-460°F// 72°F=295°K 

Hence propellant selection is a very important design parameter for 

launch vehicle design consideration. When different propellant

combinations are used in a staged vehicle, the maximum ∆V is derived 

from higher specific impulse upper stages. Higher density propellants can 

have a very large impact upon vehicle size (note that kerosene is 11.4 

times denser than liquid hydrogen) making the denser propellants most 

attractive in the larger first stage. 

One of the best first-order indicators of conventional rocket engine 

design complexity is chamber pressure. Higher chamber pressures equate 

to higher overall performance, but unfortunately, also demand heavy,

complex, and often failure prone turbo-pumps as well. Further, higher 

pressures equate to higher operating temperatures, often driving an engine 

from passive cooling to much more complex active cooling. Rocket 

engines operate at maximum performance when the exit pressure equals 

ambient pressure, a condition present during a very small segment of the
Decision Maker’s Guide...47 

engine’s operation. Conventional rocket engines typically have fixed bell 

nozzles resulting in sub-optimal performance. Variable expansion nozzles

are typically heavy, complicated and cumbersome, rarely making their 

implementation worthwhile. However, the linear-aerospike engine, like 

that pursued on the X-33, uses a fixed-center “spike” constraining the flow 

between nozzles but permitting the outside flow to freely expand to 

ambient pressure. Hence, the nozzle can be optimized to operate at a high 

altitude or vacuum condition without experiencing significantly degraded 

performance at low altitude.115 Despite some minor performance 

drawbacks at certain flight transients, these engines can operate at much 

lower chamber pressures and can be designed to be modular, greatly 

enhancing propulsion integration with the remaining vehicle structure.116

In the final analysis, rocket propulsion is on the performance edge 

of demonstrating single-stage-to-orbit capability at current specific 

impulse and thrust-to-weight ratios. They also currently lack the

operability and maintainability necessary for R2

ISA. A combination of 

higher specific impulse and thrust to weight with proven reliability and 

durability is necessary for rockets to become an attractive alternative for 

either single or two-stage-to-orbit vehicles. 

Advanced Materials and Structures. Minimizing inert mass

fractions is key to maximizing the payload capability of space launch 

vehicles. Development of advanced materials, particularly reinforced

composites and metal matrix composites for both structural and thermal

performance are central to this success. These advanced materials exhibit 

strength-to-weight and stiffness-to-weight ratios two to five times higher 

than aluminum or titanium, the current staple for aerospace vehicles.117 

These materials include graphite epoxy (five times stronger per unit 

weight than aluminum, which is the primary shuttle structural material), 

aluminum/lithium alloys, and others. According to some analyses, 

advanced composite materials and lightweight metal alloys may permit 

launch vehicle structure weight to be reduced by up to 35 percent.118 

Despite some unique flight vehicle environmental challenges,

(temperature, material compatibility, etc.), intelligent application of these

advanced materials is a prerequisite to building design (in this case 

structural) margin and robustness onto future reusable launch vehicles. 

Thermal Protection. Load bearing structures such as aluminum 

and titanium cannot survive the severe thermal loads associated with 

reusable launch vehicle atmospheric reentry. None of the classic solutions 

of the past, to include missile reentry vehicle ablative ceramics, or space 
48...Decision Maker’s Guide

shuttle silica glass tiles are suitable. Light, inexpensive, low maintenance, 

launch systems with rapid mission turnaround and high payload mass 

fractions require a durable, robust, and lightweight thermal protection 

system. Reusable launch vehicle aerodynamic design can be optimized to 

reduce peak heating and overall thermal loads on the vehicle, thereby

decreasing the degree of thermal protection required. RAND concluded 

that the technology is sufficiently mature to fabricate a thermal protection 

system that is more reliable and less expensive to maintain than the current 

shuttle ceramic tile system that requires 17,000 man-hours for

refurbishment after every flight.119 Provided reentry temperatures less 

severe than those encountered on the shuttle, existing advanced thermal 

protection materials, have the potential to provide the robustness and 

durability required for a next generation reusable launch vehicle.120 Peak

temperatures generating the highest thermal loads would still likely

require reinforced carbon-carbon, but combinations of metallic panels are 

suitable for most other vehicle locations. The thermal regime referenced

is summarized in Table 10. Temperatures associated with the Boeing

Reusable Aerodynamic Space Vehicle are consistent with the RAND 

conclusions discussed above. RAND further observed that “although 

metallic panels have higher density than ceramic tiles, a metallic TPS may 

be lighter and simpler by eliminating the need for a complex adhesive 

system such as that used on the space shuttle. The panels may also serve 

as aerodynamic load bearing structures, eliminating the necessity for an 

underlying airframe.”121 

Table 10. Selected Temperature Distributions122

National 

Aerospace 

Plane

Boeing RASV Space Shuttle 

Nose, leading edges 3000-4000 °F 1800-2770°F 2300-2800°F 

Lower fuselage/ 

wings 

N/A 1200-1800°F 1800-2300°F 

Upper fuselage/ 

wings 

N/A 900-1200°F 600-1800°F 

Note that the two most common Shuttle thermal protection

materials, namely reinforced carbon-carbon found on the shuttle nose and 

leading edges as well as the silica glass tiles on the lower fuselage and 
Decision Maker’s Guide...49 

wings, have single-mission and 100-mission temperature ratings of 

3300/2700°F and 2600/2300°F respectively.123 Even the moderated

Boeing Reusable Aerospace Vehicle temperature regime described above 

exceeds the 100-mission life temperature of reinforced carbon-carbon.

This punctuates the challenges and importance of reentry energy 

management and an appropriate thermal protection schema. 

Vehicle Integration. Structurally efficient and maintainable 

vehicles demand the effective integration of lightweight, high-strength 

composites and metal alloys into vehicle structures to include the dual role

of tanks and metallic thermal protection systems as structural load bearing 

elements. The efficient integration of propulsion elements into the

vehicle structure, particularly airbreathing hypersonic propulsion solutions

is also a major design challenge. The success of these integration efforts 

is essential for minimizing inert vehicle mass and maximizing design 

margin or useful payload delivery to low-earth orbit. Further, vehicle 

design integration must ensure simple subsystem interfaces and 

maintenance accessibility to minimize labor intensity, ground handling,

and mission-turn time. Careful design trades between operationally 

desirable but structurally parasitic wings in lieu of structurally efficient but 

control limited lifting bodies must also be carefully considered.

124 

EXIT STAGE RIGHT 

This section addresses the present technology limitations that drive 

any existing space launch vehicle to a two-stage-to-orbit-solution. 

Challenges of Single-Stage-to-Orbit Vehicles 

Historically, single-stage-to-orbit designs of the 1960’s and 1970’s 

had been considered more technically challenging than two-stage-to-orbit 

designs due to technological limits on achievable vehicle inert mass 

fractions and propulsion specific impulse. These limitations resulted in 

extremely performance sensitive single-stage-to-orbit designs with large 

gross lift-off weights and hypersensitivity to any performance shortfall. 

The National Aerospace Plane unsuccessfully attempted the development

and integration of very high specific impulse, airbreathing scramjet 

technology into a reasonably sized single-stage to orbit vehicle until it was 

abandoned in 1992. The development of modern composite materials, 
50...Decision Maker’s Guide

advanced lightweight alloys, and thermal protection systems in the 1980’s 

and early 1990’s indicated that vehicle structures might be reduced by as 

much as 35 percent.125 This reassessment did not go unnoticed by NASA, 

prompting pursuit of the X-33 program in 1996. 

A single-stage-to-orbit space launch system holds the promise of 

providing the simplest reusable space launch architecture possible, yet the 

enabling technology to achieve it remains beyond the state-of-the-art. As 

previously discussed, a viable singe-stage-to-orbit vehicle is challenging 

because of the high ∆V and aggressively low inert mass fraction required. 

Efforts such as X-33, shown conceptually in Figure 9 and outlined in 

Table 11, heavily leveraged technology to push the performance envelope

with high performance rocket propulsion, advanced lightweight highstrength materials, and innovative thermal protection systems. The X-33 

program was intended to demonstrate technology scaleable and traceable 

from a subscale vehicle to a full-scale single-stage-to-orbit rocket.126 

Since the X-33 design competition was so keen (a successful X-33 design 

was considered an inside track to the build the successor to the space 

shuttle) and the design approaches so varied, it is instructive to examine 

the vehicle “vital statistics” for the all three proposed objective reusable 

launch vehicle systems. 

Figure 9. X-33-Based SSTO RLV Design Concepts 
Decision Maker’s Guide...51 

Table 11. X-33 Based Single-Stage-to-Orbit Design Concepts 

Description

Characteristic LockheedMartin127McDonnell 

Douglas

128

Rockwell129

Type Lifting Body

VTHL

Ballistic VTVL Winged VTHL

Length (ft) 127 185 213

Width (ft) 128 48.5 103

GLOW (lb) 2,186,000 2,400,000 2,200,000

Empty Weight (lb) 197,000 219,000 296,000

Propellant Weight

(lb)

1,929,000 2,136,000 1,861,000

Payload Weight (lb) 59,000 45,000 43,000

Propulsion 7 RS2200 Linear 

Aerospike Engines

8 Rocketdyne 

RS-2100 Engines 

6 Rocketdyne 

RS-2100 Engines 

Propellant LH2/LOX LH2/LOX LH2/LOX

Engine Vac. Isp (sec) 445 450 450

Engine T/W 83 to 1 83 to 1

Chamber Pres. (psia) 2250 3250 3250

Payload Size (ft) 15x45 16.5x35 15x45

Inert Mass Fraction 0.090 0.091 0.135

Propellant Mass 

Fraction

0.882 0.890 0.846

Payload Mass

Fraction

0.027 0.019 0.020

RDT&E COST 2-3 vehicles@ 

$5B

1 vehicles@ $4-7B 1 vehicles@ $5-

8B

VTHL=Vertical Take-off Horizontal Land; VTVL=Vertical Take-off Vertical Land;

GLOW=Gross Lift-off Weight; T/W=Thrust-to-Weight; RDT&E=Research

Development Test & Evaluation; LH2/LOX=Liquid Hydrogen/Oxygen

Although the vehicles depicted above are very different in outward 

appearance, they share many common underlying characteristics. The 

most notable but predictable characteristic common to all three vehicles is 

the aggressive application of “cutting-edge” technology to expand the

performance envelope sufficiently to make a single-stage-to-orbit vehicle 

practical. As previously discussed, existing technology resulted in 

vehicles cursed with very small payload fractions and razor thin 

performance margins that, in turn, demanded very aggressive inert mass 

fractions and high performance propulsion. The inert mass fractions
52...Decision Maker’s Guide

ranging between 0.09-0.135 for the proposed X-33 is reasonable for an 

expendable launch vehicle, but unrealistic for reusable systems. 

Unreasonable demands on the propulsion side of the equation were no

different. The thrust-to-weight ratios for the X-33 design proposals are 

consistent with the National Research Council conclusion that a singlestage-to-orbit reusable launch vehicle would require a minimum thrust-toweight value between 75 and 80.130 The Space Shuttle Main Engine

Block II+ with a sea-level thrust to weight ratio of 58 remains the most 

advanced liquid hydrogen/oxygen rocket engine ever built. NASA is

planning to improve to thrust-to-weight to near 70 in the Block III version. 

Hence, all proposed X-33 propulsion solutions reside beyond the existing 

state of the art. Any performance shortfall of the already aggressively 

specified RS-2100 or revolutionary linear-aerospike engine would 

drastically reduce payload capacity for any of these designs. Not 

surprisingly, high performance liquid hydrogen/oxygen is incorporated

into all three designs despite the low propellant density (and increased 

vehicle mass) associated with it. Lockheed-Martin chose to accept 

additional risk by incorporating the then unproven linear-aerospike engine 

to gain a five to eight percent performance improvement.131 

Aggressive inert mass fractions also have significant design 

implications. The vertical take-off mode proposed by all three designs 

effectively constrains vehicle launch operations to a well-equipped 

spaceport. This design choice is made to eliminate the need for a robust 

landing gear (it supports an order of magnitude less vehicle mass at 

landing) and structurally inefficient but beneficial wings (also useful for 

significant cross-range), both necessary for a horizontal take-off. Neither 

a launch-certified horizontal take-off and landing gear nor wings appear 

realistic for single-stage-to-orbit application until much improved inert 

mass fractions and structural margin is demonstrated. Also, despite

successful incorporation of current state-of-the-art materials technology, 

low inert mass fractions will make it difficult, in the foreseeable future, for

the vehicle to compete with single stage expendable rockets for the 

launching of heavy payloads (i.e. 25,000 pounds or more).132

All three designs were emblematic of a performance-driven 

solution. A slight miscalculation or performance shortfall results in the 

elimination of useful payload capacity. Squeezing out the last drop of 

specific impulse or shaving off the last pound also costs money and 

reduces system reliability and robustness. It cannot be overemphasized 

that performance-driven designs are ill suited to deal with the inevitable 
Decision Maker’s Guide...53 

weight growth that comes with design maturation, but more importantly, it 

makes R2

ISA highly unlikely. Typically, resource and system

performance constraints force sub-optimization along one or more of the

R2

ISA axes. Experience repeatedly teaches that this is done at great peril.

A balanced design approach that addresses the demands inherent with all 

three axes is necessary to prevent the cost and performance shortfalls

graphically illustrated by shuttle and the existing expendable launch 

vehicle fleet. 

Figure 10 quantitatively illustrates the single-stage-to-orbit 

performance trap discussed above. Specific impulse (propulsion 

performance limit) is plotted against vehicle initial mass with an inert 

mass fraction (structural performance limit) overlay. For this specific

example, it becomes readily apparent that any propulsion or structural 

performance shortfall will have very large consequences for the overall 

vehicle. Despite the obvious technical challenges associated with singlestage-to-orbit vehicles, the concept still remains the optimum R2

ISA

“endgame” solution. However, another aggressive, near term,

performance driven, single-stage-to-orbit R2

ISA solution is doomed to 

failure. Yet it must be emphasized that the same technologies that enable 

single-stage-to-orbit are very useful in two-stage-to-orbit systems, 

suggesting that their continued pursuit remains valuable. 

Figure 10. Parametric Analysis of a Single-Stage-to-Orbit Vehicle 
54...Decision Maker’s Guide

 Single Stage vs. Two-Stage-to-Orbit Solution 

The single-stage-to-orbit example previously discussed demands

beyond current state-of-the-art technology to deliver useful payloads to 

low-earth orbit. As mentioned earlier, a rocket-based solution continues to 

require liquid hydrogen/oxygen propellant resulting in volumetrically

larger vehicles and a minimum thrust to weight ratio between 75 and 80,

very light structures delivering a vehicle inert mass fraction below 0.10, as 

well as a very robust thermal protection system. All of these demands 

become tradable design margin with two-stage-to-orbit solutions. 

The idea of a fully reusable two-stage-to-orbit system is not new. 

The original World War II German Sanger concept proposed a supersonic 

Mach 6 first stage, with a rocket-based orbital vehicle as the second

stage.133 The Sanger had both the first and second stages optimized and 

fully integrated into a combined vehicle configuration. Today, the Sanger 

II upper stage designed to achieve orbital velocity is rocket powered, 

while the first stage can be either rocket powered, airbreathing, or both. A 

two-stage-to-orbit vehicle could be launched either horizontally or 

vertically, with both first and second stages typically landing horizontally.

First and second stage engines can operate concurrently at launch/take-off 

to provide additional thrust, but these designs require fuel and/or oxidizer

transferred from the first to the second stage as propellant is consumed in

the latter. Subsonic vehicle separation has been demonstrated between a 

highly modified B-747 and the space shuttle orbiter, and supersonic 

separation up to Mach 3 speeds with an SR-71 air-launched ramjet 

powered drone.134 The reusable two-stage-to-orbit second stage will still 

require a robust thermal protection system to withstand the full rigors of 

reentry with minimal maintenance, but permits a much smaller orbital 

vehicle as well as eliminating the need for a complex thermal protection

system for the first stage. Given similar technologies (i.e., the same 

propulsion system and inert mass fraction), a two-stage-to-orbit vehicle 

will always have a better payload mass fraction than a single-stage-to-orbit 

vehicle designed for the same mission. Only when the single-stage-toorbit inert mass fraction approaches zero does its payload mass fraction 

approach that of a two-stage-to-orbit vehicle.135 
Decision Maker’s Guide...55 

SPACE ECONOMICS 101 

The Space Launch Market 

The U.S. space transportation market is a $7 billion industry

comprised of 39 percent civil (manned) low-earth orbit missions; 28 

percent commercial geostationary-earth-orbit; 16 percent military

geostationary-earth-orbit; six percent military low-earth orbit; six percent 

civil low-earth orbit, and the remaining five percent to civil earth escape

and commercial low-earth orbit missions.136 Estimates in forward demand 

are historically both highly variable and optimistic. Much of the

geostationary spacecraft launch demand is being offset by dual launch 

capability, and commercial low-earth-orbit demand remains uncertain. 

The 1994 Commercial Space Transportation Study indicates that present 

demand is linear (inelastic) to price down to approximately $1000/lb to 

low-earth orbit. Interestingly, tourism is the one new market that appears

to dominate all others, indicating a highly elastic demand curve. There are 

undoubtedly others that are yet unidentified and will remain so until price 

barriers are lowered. Historical market conditions (high price, low 

volume) maximize industry revenue when demand is linear (or nearly so) 

with price. This is emblematic of a classical market failure, which 

ventures into the realm of a private cost, public benefit problem. If the 

price is halved, volume doubles, and total revenue remains constant, hence

there is no market incentive to invest. Reality to date supports this 

assessment, as no investment in a new space launch system has returned 

that investment in real terms.137

Transportation industries have bedeviling contradictions in the past

and remaining today. On the one hand, they are a vital part of a robust

economy. On the other, they are historically government subsidized, 

money losing propositions for investors. Any government policy maker,

corporate CEO, or entrepreneur who believes that the current economic 

state of affairs in space transportation is amenable to profitable 

commercial enterprise (outside of very limited niche markets) is sorely 

mistaken. However, that said, there is nothing unusual about this, and 

certainly nothing inherently “wrong” with the future potential space 

transportation market. The trick will be finding the right chemistry to

unlock that potential. 
56...Decision Maker’s Guide

Warren Buffet cites the automobile and the airplane as two 

illuminating examples of 20th Century transportation market failures with 

his commentary on the latter below:138

The other truly transforming business invention of the first 

quarter of a century, besides the car, was the airplane—

another industry whose plainly brilliant future would have 

caused investors to salivate. So I went back to check out 

aircraft manufacturers and found that between the 1919-39 

period, there were about 300 companies, only a handful 

still breathing today. Move on to airline failures. Here is a 

list of 129 airlines that in the past 20 years filed for 

bankruptcy. Continental was smart enough to make that 

list twice. As of 1992, ...the money that had been made 

since the dawn of aviation by all of this country’s airline

companies was zero. Absolutely zero.139

One conclusion that can be drawn from Mr. Buffet’s observations 

are that access to public capital markets for the billions of dollars required

for space launch system development will be very difficult. The potential 

for an attractive return on investment simply does not yet exist. More

than $1 billion has been privately invested in various space launch startups 

since 1980, excluding the Evolved-Expendable-launch Vehicle. The 

cumulative “return” on this investment of social resources is less than 

$600 million from 35+ revenue launches of Pegasus and Athena vehicles. 

Even ignoring the time value of money, it is evident that private 

investment in new space transportation assets has not been justified by the 

return on investment. This is exactly what we would expect if the 1994 

Commercial Space Transportation Study data were broadly correct.140 

Note that the only “successful” new entrant into the industry (Orbital) 

undertook private development of the Pegasus vehicle following the 

implied government endorsement of an “anchor tenancy,” a form of 

government incentive.141 All of this implies that access to U.S.

government demand may be required for private or partly private 

financing of new launch assets.142

This situation is not at all unusual. There has been significant 

subsidization of railroad, automobile, commercial aviation, and space 

launch infrastructure. Further, within the aviation and space industry, 

national security imperatives were responsible for absorbing the cost 
Decision Maker’s Guide...57 

burden of technological advance as well as flight vehicle development. 

This was the case from 1910 through 1960 in the aviation industry to 

include the enabling technology for the development of the DC-3 and B707. The same was true for the space launch market. Once President

Eisenhower made ICBM development a national priority in 1954, it would 

ultimately dwarf that of the Manhattan Project in both size and funding. 

The Thor, Atlas, Titan, (forerunners of the Delta, Atlas, and Titan launch 

vehicles) and Minuteman missiles would be successfully developed in just 

eight years.143 The latest Evolved Expendable Launch Vehicle systems, 

namely the Delta IV and Atlas V, are direct descendants and beneficiaries 

of the immense capital investment made nearly fifty years ago. The only 

other major U.S. launch systems since that time, namely Saturn and 

Shuttle, were also government-financed systems. This will not change

until there is a robust commercial space transportation industry to support. 

Consequently, any illusion that the “smart money” from

either the largest aerospace corporations or venture capitalists will finance

future-generation launch systems before the government paves the way for 

several orders of magnitude reduction in launch costs is an exercise in 

wishful and unrealistic thinking. This problem is compounded by the fact 

that the space transportation market has no incentive to grow until the cost 

savings from an advanced space transportation architecture is 

demonstrated. The current state of affairs is a classic “chicken or the egg” 

scenario emblematic of a market failure. 

Launch Vehicle Economics 

Figure 11 provides a macro-view of existing expendable launch 

vehicle space launch cost trendlines of geosationary-transfer-orbits for 

U.S., French, Russian, and Chinese launch services. It also captures the 

operational partially reusable 1st Generation RLV launch systems (namely 

Pegasus and Shuttle). This section will provide the economic detail that

drives these cost trendlines for Western systems. 
58...Decision Maker’s Guide

 

Figure 11. ELV Space Launch Cost Trendlines144

One approach to space launch system cost accounting is to

describe total system costs as the sum of direct and investment costs: 

Csystem ≡ [Chardware + Cpropellant + Coperations]DIRECT 

+ [Cinfrastructure + Cdevelopment + Cmoney]INVEST

Equation 5. Total System Cost Elements 

where investment costs include infrastructure (spaceports and support 

equipment), vehicle development costs, and cost of money (all costs to be

recouped via vehicle operations). Direct costs include vehicle hardware,

propellant, and vehicle operations costs. Propellant costs are generally 

insignificant and will not be discussed further.145 Cost comparisons

between systems is difficult due to the proprietary nature of data 

associated with commercial systems. However, in spite of these 

limitations, some important observations regarding launch system costs 

can be made. 

Investment costs are typically the subject of intense scrutiny due to 

the upfront commitment of capital and the seemingly unavoidably large 

price tag that comes with it. Although this paper only briefly addresses 

infrastructure costs and the cost of money, it is important to recognize that 

these factors weigh heavily on any commercial decision to develop a 
Decision Maker’s Guide...59 

space launch system. Developmental costs often dominate investment

costs and are critically important because they are usually constrained by 

an upper threshold that, if breached, results in elevated managerial and 

political scrutiny, often putting the entire development effort in jeopardy. 

Subsequent design compromises, often made during the developmental 

phase as a consequence of constrained funding, come home to roost later 

as much higher production and operations and maintenance costs. 

Direct costs typically dominate the overall system cost. Not

surprisingly, hardware costs comprise the lion’s share of expendable 

launch vehicle direct costs while operations and maintenance costs may 

comprise as much as 70 percent of the total direct costs for reusable 

launch vehicles.146 Much of the reusable launch vehicle direct costs can 

be traced back to underlying technology, design, manufacturing, 

operability, and maintenance choices that are made during system design 

and development. Hence, insight into direct costs can guide decisions

regarding the entire launch system architecture. Table 12 provides 

specific data highlighting some relevant launch vehicle cost data.

Table 12. Launch Vehicle Costs 

System “Dry”

 Mass 

(lb)147

Payload to 

Low-Earth 

Orbit 

 (lb) (ref)148

Price to 

Low-Earth

Orbit 

($/lb)149

Specific 

Developmental Cost

($ per lb)150

Specific 

Hardwar

e Cost

($/lb)151

Specific 

Operational 

Cost ($/lbm)

Labor 

Intensity 

(hrs/lb)152

SSTS 615,679 63,433

(127) 

15,000 114,000 18,900 1

Pegasus 5395 976 (115) 13,500 Not Avail 

X-15 13,873 N/A N/A 102,000 8960 0.5

X-33 76,534 N/A N/A 20,000 3340 N/A

X-34 16,500 N/A N/A 14,000 1190 N/A

Delta II 36,388 11,330 

(115) 

4854 9,500 10

Atlas II/ 

 Centaur 

22,100 18,982

(115) 

5136 31,600 40

Titan IV/ 

 Centaur 

132,800 41,000

(110) 

4951 Not avail 40

Notes 

(alt) nm 

28.5 deg

incl 

FY2000 FY2002 FY2002

See Figure

11 

 

Direct Launch Costs. Launch vehicles are, pound for pound,

more expensive than almost any other manufactured product—

significantly higher than the cost of aircraft. Table 13 depicts a relative 
60...Decision Maker’s Guide

cost comparison between several different transportation systems. These 

high specific costs are directly tied to technical complexity, design

decisions, and manufacturing processes. 

Table 13. Transportation Vehicle Hardware Costs per Pound153

Vehicle Relative Cost Cost (FY02 $/lb) 

Space Shuttle Orbiter 11.1 16,473 

Atlas IIA 2.5 3,708 

Delta II 7925 0.97 1,438 

F-15 Fighter Aircraft 1.0 1,488 

Commercial Jet Airliner 0.33 496 

Automobile 0.005 7.27

Table 13 is a graphic illustration of how the demands of increasing 

performance result in increased cost. Two root causes are implied, 

namely, system complexity and technology. Expendable launch vehicle 

single-use permits inert mass fractions as low as five percent of total dry 

vehicle mass contrasted with 13 percent for shuttle and a “stageequivalent” inert mass fraction above 50 percent for a front-line fighter.154 

Reusable systems are typically more expensive than expendable systems 

because they simply weigh more, typically by a factor of five to ten.

However, note the radical departure of the space shuttle orbiter from this 

trendline. This departure speaks directly to the technical complexity 

argument. The requirement to minimize inert mass fractions and 

maximize useful payload capability drove U.S. launch vehicle 

development to maximize system performance. This approach minimized 

performance margin and added complexity that ultimately manifested 

itself into expensive expendable launch vehicles (Delta, Atlas, Titan) and 

reusable launch vehicles (Pegasus and Shuttle). 

Operational launch costs can represent a very large fraction (up to 

70 percent) of overall direct costs of low-earth-orbit payload insertion.

Claybaugh makes the case that labor intensity L* (man-hours/lb) is a 

useful surrogate for relative comparison of operational costs between 

systems. Figure 12 identifies two distinct trendlines between military and 

commercial aircraft (“sonic trendline”)155 and space launch vehicles

(“orbital trendline”) that are independent of vehicle mass. 
Decision Maker’s Guide...61 

Figure 12. Labor Intensity of Aerospace Systems156 

Note that both existing expendable and reusable space launch 

systems exhibit prohibitively high labor intensity, implying a lucrative 

area to pursue in any effort to reduce overall launch costs. The sonic 

trendline indicates a labor intensity (L*) of 0.0010, while the orbital 

trendline suggests 10 man-hours/lb. This approximate five order of 

magnitude difference in labor intensity between aircraft and space launch

systems implies that operation and maintenance costs must be a paramount 

consideration in any future designs. Contemporary Ariane IV and space 

shuttle157 launch systems are two notable departures from the orbital

trendline showing a one order of magnitude improvement in labor

intensity over more venerable launch systems to near one man-hour/lb.

Two other systems of note falling between these two trendlines lines are 

the SR-71 and X-15 with an L* of 0.013 and 0.5 man-hours/lb 

respectively.

158 Arguably, both vehicles should deviate upward from the 

sonic trendline as they begin to more closely resemble spacecraft rather 

than classical aircraft. This data begins to quantify the very large disparity

that exists between existing space launch and aircraft system labor 

intensity and provides a technique to bound operational costs as a function 

of vehicle mass. However, this approach begs the question: is labor 

intensity merely a surrogate for the increased vehicle complexity that 

inevitably migrates into larger aerospace systems? If so, it may suggest 

that other metrics can measure this phenomenon more directly. 
62...Decision Maker’s Guide

Investment Costs. Investment cost estimates can “make or break” 

a launch system development decision. Although little data exists to 

support a strong statistical correlation between vehicle dry mass and 

facility costs, one does exist for vehicle dry mass and development costs.

Sixty years of aerospace company proprietary data shows the impact of 

technological progress. Cost per unit mass tends to remain about the same

as vehicle performance increases over time, or alternatively, similar 

performance vehicles cost less per unit mass over time.159 Historical U.S.

expendable launch vehicle development, all with a governmental 

development legacy, demonstrate developmental costs ranging from $10-

40 thousand (FY02) per pound. Although developmental cost data for 

reusable launch vehicles is sparse, Claybaugh calculated costs for the

government-developed X-15 and space shuttle systems at $102 and $114

thousand (FY02) per pound respectively. He reports costs of $25, $20,

and $14 thousand per pound for the Boeing-777 (modern reusable

benchmark), X-33, and X-34 respectively. The fact that neither X series

vehicle actually matured to flight status likely biases these values

significantly downward. Claybaugh concludes from this data that 

government-developed reusable launch systems are much more expensive 

than commercially developed vehicles.160 However, when you consider 

the fact that governmental subsidization was involved in the development 

of virtually all U.S. expendable and reusable launch systems, this 

conclusion may be incomplete. The cost of money comes into 

consideration when infrastructure and developmental costs must be

amortized over some meaningful period of time. Within the realm of the 

commercial world, investment decisions are dominated by maximizing 

return on investment. Hence, the need to recoup initial outlays as quickly 

as possible, often in as few as five but no more than ten years, must occur

to permit any space launch development venture to be deemed

commercially viable. The bottom line is that the calculations embedded 

in investment cost estimates are based upon assumptions—assumptions 

that often do not withstand the test of time—as the space shuttle clearly 

illustrates. 

Expendable Launch Vehicles. How cheap can expendable 

launch vehicles get? Empirical cost estimating data previously presented 

can be applied to a very simple linear cost model to provide an answer to

this question.161 The following examples all assume investment costs 

(infrastructure, development, and cost of money) are zero, and that 

propellant is free to leave the following: 
Decision Maker’s Guide...63 

Csystem ≡ Chardware + Coperations = RfCh + RL * CL = R(fCh + L*CL) 

where Csystem = specific system cost to deliver low-earth-orbit payload

($/lb) 

R = structural ratio (inert structural mass/payload mass) 

f = expended hardware mass fraction (1.0 for an expendable 

launch vehicle) 

Ch = specific hardware cost ($/lb) 

CL = specific labor cost ($/hr) 

L* = labor intensity (man-hours/lb); 

 ($75 U.S. burdened rate)162 

Equation 6. Simplified Expendable Launch Vehicle Cost Expression 

Example A in Table 14 is representative of a typical present day

commercial booster delivering payloads to low-earth orbit for around 

$5000 per pound. For further simplification, ignore all operations costs 

for Examples B and C. Example B shows that a $2000/lb cost using 

typical present day structural ratios and hardware costs implies that the 

system must be one third reusable. Alternatively, Example C illustrates 

the reduction of the structural ratio to two, implying the highly unlikely 

scenario of single-stage-to-orbit technologies at present day hardware 

costs.163 Example D considers the “Big Dumb Booster” as some

proponents affectionately call it, where a very simple pressure-fed, all 

composite structure yields a “heavy” design with a structural ratio of six.

(Note that liquid pump-fed systems operate with tank pressures in the tens 

to hundreds of pounds per square inch, while pressure-fed systems

demand tank pressures in the thousands of pounds per square inch. You 

are trading the complexity of the pump for a savings in weight versus the

simplicity of design for increased structural weight to accommodate the

high tank pressure). Using a specific hardware cost of $300/lb (where 

$1000/lb is typical) that has been demonstrated by some solid rocket 

motors, yields $2250/lb to low-earth orbit. Claybaugh concludes that 

when including propellant, a best case, minimum cost (not price) 

expendable launch vehicle might demonstrate $2000/lb to low-earth 

orbit.164 To emphatically make the case that expendable systems will 

never achieve R2

ISA, Example E illustrates the unachievable low structural 

ratio necessary for the high pressure “big dumb booster” to achieve

$1000/lb. Example F shows the lowest possible cost that might be 
64...Decision Maker’s Guide

achieved in the distant future with as yet unknown ultra-high strength, 

inexpensive structures, free labor, free (or subsidized) propellant, and no 

investment costs (amateur rocket builders?). Note that the NASA Highly 

Reusable Space Transportation Study evaluated a Highly Evolved 

Expendable Launch Vehicle in 1995 using aggressive, highly innovative

designs and concluded that the lowest cost attainable was about $800 per

pound ($944 FY02) to low-earth orbit.165 This reasonably approximates 

the Example F “Low Ball Booster.” 

Table 14. Simplified Expendable Launch Vehicle Cost Excursions 

Description Tag Csystem = R ( f Ch + L * CL ) 

Current A 5250 3 1 1000 10 75

$2000/lb goal B 2000 3 .67 1000 0 0

$2000/lb goal C 2000 2 1 1000 0 0

Big Dumb

Booster 

D 2250 6 1 300 1 75

Super Cheap 

& Sleek 

E 1000 2.7 1 300 1 75

Low Ball 

Booster 

F 1050 3.5 1 300 0 0

Space Shuttle. The space shuttle merits detailed discussion as it 

represents not only a first-generation reusable launch vehicle but has been 

subjected to very detailed “root-cause” analysis, particularly in the areas 

of cost and flight regeneration or cycle time. Figure 13 highlights the 

$3.27 billion resource allocation that generated four shuttle missions in 

Fiscal Year 2002. Nearly two-thirds of the shuttle budget is allocated 

toward fabrication, refurbishment, or flight preparation of the physical 

flight vehicle. Surprisingly, only about one quarter of the budget supports 

the highly visible ground and flight operations at Kennedy and Johnson 

Space Centers respectively. It also suggests that calling the shuttle a 

reusable launch vehicle is a misnomer, as clearly, more than one-third of 

the budget is directed at the refurbishment and reloading of the solid 

rocket motors and external tank/main engine production. 

Considered alone, Figure 13 might suggest that hardware 

dominates the shuttle budget. This is not the case. A functional 

breakdown of shuttle costs using NASA definitions of direct 
Decision Maker’s Guide...65 

(landing/recovery, vehicle assembly/integration, launch, payload/crew 

preparation, and turnaround), indirect (vehicle depot maintenance, 

traffic/flight control, and operations support infrastructure), and support 

(shuttle logistics and operations planning management) comprise 10, 20, 

and 70 percent of the total shuttle budget respectively. The latter is 

comprised of very labor-intensive activities across all of the NASA 

Centers to get the shuttle prepared for the next launch. The nature of these 

activities is highlighted in Figure 14.

Figure 13. Fiscal Year 2002 Space Shuttle Budget Summary166
66...Decision Maker’s Guide

Figure 14. Time Allocation at the Space Shuttle Orbital Processing 

Facility167

Figure 14 suggests that the unplanned troubleshooting and repair 

of hardware drives a lack of confidence in the system that is not restored

until required inspections and checkouts for flight certification are 

completed. These activities combined with complicated, manpower

intensive, vehicle servicing are responsible for over 75 percent of shuttle 

cycle time. The cycle times for shuttle mission STS-81 represents the 

lowest post-Challenger annual recurring cost per pound achieved by the 

shuttle program.

168 Another insightful perspective on STS-81 task 

durations is to examine time allocated to subsystems that include 

structural/mechanical, propulsion, power management (mechanical and 

electrical), thermal management (active and passive) comprising 26, 18,

16, and 16 percent of orbital processing facility time respectively.169 Not 

surprisingly, these areas correlate with the key reusable launch vehicle 

technology enablers previously discussed in chapter 3 to include materials 

and structures, propulsion, thermal protection. 

Other Launch Cost Considerations. The expense of launching 

payloads into space today is very high. Launch vehicles and their 

operation -- whether expendable or reusable, small or large -- cost tens of
Decision Maker’s Guide...67 

millions to hundreds of millions of dollars per flight. These costs are in 

addition to the usually very expensive payload the launch vehicle is

carrying. A payload budget planner must allocate such a significant 

portion of the budget to launch services that these considerations can have 

a powerful ripple effect on all aspects of the space mission. The cost of 

space vehicles has become almost inextricably linked to the cost of launch, 

and reducing the cost of space systems and missions is largely dependent 

on achieving lower space transportation prices.170

The nature of space launch has wide-ranging, pervasive impacts on 

the design and operation of spacecraft. Launch costs often are a large 

portion of space system lifecycle cost and hence heavily influence satellite 

capability, weight, volume, and complexity, as well as mean mission 

duration, deployment options, constellation quantities, and cost. A retired 

TRW executive stated that because launch systems cost so much, satellite 

designers always take the smallest and least expensive launch system 

possible and spend large amounts of effort and money trying to get their 

space vehicles to meet booster weight and volume constraints. He cited 

instances where designers spent up to $230,000 per pound (FY02) in 

taking the last few pounds out of a satellite so they could meet the selected 

launch vehicles orbital lift capability.171 172 There are also more subtle 

interactions between payload and launch costs. As launch costs increase, 

so do payload costs. To reduce risk of on-orbit failure and the probability

of re-launch, some payload subsystems are made triple redundant, 

increasing the cost and weight of the satellite. If launch costs can be

reduced significantly, it may no longer be necessary to design to such high 

levels of redundancy. In addition, a capability such as an orbital space 

plane or orbital transfer vehicle could recover payloads in orbit, and if

payloads were designed modularly, they could be quickly repaired onorbit. Such payloads could cost considerably less than existing satellites.

An orbital space plane and/or orbital transfer vehicle linked to R2

ISA

might enable a new era of low cost access to space.173 
68...Decision Maker’s Guide
Decision Maker’s Guide...69 

IV. The Search for Solutions 

Everything should be made as simple as possible, but not 

simpler. 

—Albert Einstein 

This chapter explores key design attributes and enabling 

technology necessary to deliver R2

ISA. Consider the reinforcing

relationship between vehicle design and R2

ISA illustrated in Figure 15 

below. Robustness, as defined in chapter 1, is determined by the vehicle 

performance (maximum payload mass per flight) and sustainable flight 

rate (flights per year). These two factors multiplicatively determine the 

maximum vehicle throughput (measured in payload mass/year) that a 

launch vehicle can deliver to low-earth orbit. Recalling the flight rate 

capability shortfall experienced by the shuttle in chapter 2 speaks directly

to this relationship. Reliability also has direct, traceable linkages to 

vehicle design. Space launch vehicle reliabilities today typically gravitate 

towards 98 percent (1/50 probability of failure) for both expendable and 

reusable launch systems, with a goal of 99.9 percent (1/1000 probability of 

failure) reliability for a second-generation reusable launch system. Costs, 

both direct and investment, were addressed at length in chapter 3. Design 

decisions directly impact developmental, infrastructure, hardware, and 

operations costs to such an extent, that arguably, the upper limits of 

vehicle affordability are determined (although impossible to precisely 

quantify) before the first piece of system hardware is ever fabricated. 

These annual costs ($/year) are balanced against payload throughput 

(payload mass/year) to determine overall cost effectiveness ($/lb) of

payload delivery to low-earth orbit. This construct emphasizes the 

internally reinforcing nature of robust, reliable, and inexpensive systems 

and operations. As the shuttle clearly illustrated, robust operations 

characterized by high throughput are a prerequisite to affordability. 

Pursuit of one at the expense of the other is sheer folly. This chapter will 

address the key design attributes and enabling technology that delivers 

R2ISA. 
70...Decision Maker’s Guide

Figure 15. Space Launch Design Linkages to R2ISA 

THREE CAMPS 

There is no overwhelming consensus within the aerospace 

community regarding the most promising and lowest risk approach to 

R2ISA. In fact, there is no universal agreement on a definition for robust, 

reliable, and inexpensive operations. This is not surprising when 

considering the political, economic, and social dimensions coupled with 

the technical, programmatic, and engineering complexities and resource 

constraints shrouding space launch. At the risk of oversimplification, the

debate falls within three general camps: 1) the expendable “big dumb 

booster” contingent; 2) rocket-centric reusable launch vehicle proponents 

(who presently appear to be in the majority); and the airbreathing, 

hypersonic reusable launch vehicle advocates. Overlaid upon this debate 

are numerous agendas and divisions to include manned versus unmanned; 

single-stage versus two-stage-to-orbit; civil versus military, etc. Despite 

these complexities, it is helpful to briefly examine the arguments, both pro 

and con, promulgated by each camp. 
Decision Maker’s Guide...71 

Expendable Launch Vehicles 

The venerable expendable launch vehicle has been and remains the 

dominant method for earth-to-orbit operations and has consistently 

demonstrated the lowest costs. Medium to heavy lift Western systems are 

in the $4-5,000/lb cost range, while Russian/Chinese systems are in the 

$2,000/lb cost range for payload delivery to low-earth orbit. Economic 

dislocations and inexpensive manpower rather than differences in 

technology or vehicle complexity largely explains this price discrepancy. 

The qualitative argument suggests that applying simple design 

practices, leveraging “off-the-shelf” technology, and using readily 

available inexpensive materials can produce the “big dumb booster” as an 

affordable low cost alternative. A simple design means fewer parts and 

interfaces, translating into lower first unit and recurring manufacturing 

costs as well as lower operating costs. Inexpensive boosters could create a 

greater market demand for launchers, permitting increased production runs

that translate into even greater economies of scale. Further, economies of 

scale are one area in which expendable launch vehicles always have an 

advantage over reusable ones.174 The longer-term evolution of expendable 

vehicles is driven by development of ultra high strength-to-weight 

materials and more powerful and efficient propulsion. Advances in 

materials will result in lighter and stronger vehicles and improve structural

efficiency. Stronger structures permit higher operating pressures that in 

turn permit the transition from pump-fed liquid rocket systems to much 

simpler pressure-fed systems. This transition would permit a dramatic 

reduction in propulsion system complexity and a marked improvement in 

reliability and manufacturability. Further, advances in new propellant 

technology, to include High Energy Density Matter,175 could be leveraged 

to significantly improve rocket performance (both specific impulse and 

thrust-to-weight) and significantly reduce vehicle size and gross lift-off 

weight. Advocates claim order-of-magnitude cost reductions from current 

prices to low-earth orbit are possible at much lower developmental risk 

and cost than any reusable system. 

Expendable launch advocates also provide quantitative cost 

arguments supporting its advantage over reusable systems. Basically,

expendable launch vehicle proponents concede an advantage to reusable 

systems in direct launch vehicle hardware costs, but argue advantages in 

nearly all other cost factors to include: 1) becoming cheaper with time 

(learning curve); 2) free of recovery and refurbishment costs making 
72...Decision Maker’s Guide

vehicle operations less expensive; 3) lower required vehicle reliability 

permitting less demanding performance and reliability margins; 4) easy

accommodation of technological/performance upgrades; 5) simpler flight

operations; 6) much lower up-front development (due to lower 

complexity) and production costs. James Wertz recently developed a

purely analytic economic model permitting useful direct comparisons 

between expendable and reusable launch vehicles. He concludes that 

expendables remain somewhat lower-cost than the reusables over flight 

rates ranging from 10 to 10,000 flights per year and that the result is 

relatively insensitive to the input assumptions.176 He indicated that a 

factor of five to ten near term reduction in expendable launch vehicle costs

and a far-term reusable cost reduction of two to three from current costs 

are achievable.177 The former implies expendable launch costs between 

$500 and $1000 per pound to low-earth orbit. There is no known 

exploitable technology known to date to enable this. Wertz’ analysis is 

flawed in that it identified development costs as the principle reusable cost 

driver but over-weighted the increased manufacturing efficiencies 

associated with the learning curve, skewing the expendable launch vehicle 

results.178 This type of well-intended but flawed economic argument is 

not uncommon and illustrates the pitfalls of purely econometric analyses 

that lack sound technical grounding and fail to leverage known empirical 

cost estimating relationships. 

Reusable Rocket and Airbreathing Launch Vehicles 

The latter two camps both advocate the reusable launch alternative,

but paths quickly diverge from there. Rocket proponents are quick to 

point out that a reusable earth-to-orbit system is in operation, and what is 

required to achieve R2

ISA is the integration of a set of technology 

advances in materials, propellants, propulsion cycles, “designer-aero,” and 

launch assist.179 These technologies can significantly reduce vehicle inert 

mass fractions, provide higher specific impulse and thrust-to-weight 

propulsion systems, potentially via High Energy Density Materials 

propellants, as well as lower the aggregate vehicle ∆V requirement to orbit 

via reduced drag or fixed ground launch assist. A look at the aborted 

Space Launch Initiative second-generation shuttle replacement is 

instructive in that the rocket-based solutions proposed were large (millions

of pounds at launch) and expensive ($30-35 billion in development costs). 
Decision Maker’s Guide...73 

Hence, recent experience implies that a pure rocket-based reusable launch

vehicle as NASA envisioned is currently unaffordable. 

A very attractive alternative to rockets is air-breathing propulsion.

The two first-order potential performance benefits of this approach are 

much smaller vehicles operating at extremely high specific impulse. 

Figure 16 clearly illustrates that nearly two-thirds of shuttle’s 4.5 million 

pound gross lift-off weight is oxidizer. A mere four and one-half percent

mass reduction in liquid oxygen carried within its external tanks doubles 

the payload capacity to low-earth orbit! 

Oxidizer (liq)

31%

Oxidizer (solid)

34%

Fuel

20%

Orbiter

4%

Inert

10%

Payload

1%

Figure 16. Space Shuttle Mass Distribution 

Air-breathing propulsion gets its oxidizer from the surrounding 

atmosphere in lieu of carrying it internally, resulting in much smaller

vehicles. The effective specific impulse of hydrogen and hydrocarbonbased airbreathing systems as a function of Mach number is captured in 

Figure 17. 
74...Decision Maker’s Guide

 Figure 17. Airbreathing versus Rocket Performance180

Turbojets compress the incoming air to high pressure with 

turbomachinery prior to fuel mixing and combustion. Ramjets slow 

incoming air to subsonic velocities (with no moving parts) before the 

mixing of fuel and subsequent combustion, while supersonic combustion 

ramjets (scramjet) do this while the flow remains supersonic. The 

potential performance benefit over pure rockets is enormous. However, 

three remaining challenges associated with airbreathing technology 

include lower thrust, low engine thrust-to-weight, and vehicle integration

challenges relative to rockets. 

Airbreathing propulsion successfully applied through most of the 

Mach range to orbit (0-25) can have a very substantial improvement over

rockets in vehicle robustness and reliability. Steve Cook, Deputy Program 

manager for NASA Next Generation Launch Technology cites a 10 to 25 

percent improvement in structural ratio, increased inert mass fraction to 

0.15 from 0.05, a factor of four to five reduction in total thrust required, 

less than half of the total gross weight (permitting horizontal take-off), and 

a factor of three decrease in performance sensitivity to weight growth.181 

Reliability and safety are improved through increased abort options, 

powered landing capability, and more benign failure modes due to lower

power densities. Further, mission flexibility is enhanced through a factor 

of two to three increase in launch windows, a factor of 2.5 improvement in 

cross-range, self transport/ferry, orbit rendezvous two to three times faster, 

and increased basing options.182 Unfortunately, hypersonic airbreathing 

propulsion technology is not sufficiently mature today for use on a
Decision Maker’s Guide...75 

second-generation launch vehicle. Further, only an aggressive research 

and development effort can make it available by 2010. 

Despite the promise of air-breathing hypersonics, there are many 

skeptics. Among them is Mr. Dennis Bushnell, Chief Scientist at NASA 

Langley Research Center, who was deeply involved in the X-30 National 

Aerospace Plane program. His views include: 

Regarding rocket alternatives: The usual commentary 

regarding such is that Rockets are (forever) evolutionary, 

"not much left there,” and therefore, Air-breathing systems

of various persuasions, even with all of its warts, is the 

"only way ahead." (This may be true for the military

flexibility metrics, but not the NASA cost metrics.) There 

are, in fact, a plethora of advanced concepts which could

seriously revolutionize rockets – IF – we spent the 

time/treasure to go there. Overall, these technologies offer

much more performance and a greater return on investment 

than air-breathing systems over similar time frames. These 

open up the design space and offer the opportunity to truly 

revolutionize space access (as opposed to the known 

agonies/disastrous sensitivities of air-breathing systems).183 

Airbreathers will, in all probability, increase the cost(s) and 

reduce overall safety for routine space access.

Maintenance requirements for airbreathers is wholly 

unknown. Particularly worrisome are thermal/acoustic 

fatigue problems.

184 A higher Isp does not necessarily

imply reduced costs. There are no system studies which

indicate airbreathers will/are expected to come anywhere 

near even an idealized factor of two, much less orders of 

magnitude improvement. Available estimates on idealized 

systems predict an approximate 35 percent cost reduction;

this will disappear as we “get real.”185 

Once again, at the risk of oversimplification, consider the R2

ISA 

trade space. Expendable launch vehicle advocates are comfortable with 

the “disposable” nature of their systems and believe that technology can 

deliver an order of magnitude improvement in cost at the same levels of 

robust/reliable operations experienced to date. The rocket-centric reusable 

launch vehicle proponents believe that the successful maturation of a small 
76...Decision Maker’s Guide

subset of promising technologies will enable R2

ISA. Meanwhile, the 

hypersonic advocates make a compelling case for robust operations, a

potentially strong but as yet unproven cost benefit, and a wholly unproven 

reliability argument. The truth is out there. 

PROPULSION: THE KEY ENABLER 

The previous chapter introduced four key enabling technology 

areas to include propulsion, advanced materials and structures, thermal 

protection, and vehicle integration. However, propulsion technology has 

traditionally been the most important pacing factor in flight system

development since the Wright brothers, and it is especially true for space 

systems. This is not surprising when one considers the enormous amount 

of power generated by these systems to accelerate payloads to orbital 

velocity. Fundamentally, the means by which this is accomplished and the 

efficiency of the energy conversion process in turn drives the vehicle

design and supporting system architecture more than any other single 

factor. A brief assessment of both rocket and airbreathing propulsion and 

the types of vehicles they enable is presented to permit an objective 

assessment of current capability and establish what currently lives within

the realm of the possible, the improbable, and the impossible. 

Rocket –Based Reusable Launch Vehicles 

Modern rockets have dominated space transportation for nearly

half a century, but their performance still fails to enable R2

ISA. The most

important elements of rocket performance that directly impact vehicle 

performance are total thrust, specific impulse, and engine thrust-to-weight. 

Improved performance for both kerosene and hydrogen-based engine 

technology is valuable, particularly the latter’s use as an enabler for a

rocket-based single-stage-to-orbit vehicle. However, the more immediate 

need is for demonstrated operability, maintainability, and reliability that 

permits dozens of relights with little or no human intervention. The liquid 

oxygen/hydrogen Space Shuttle Main Engine186 is the most advanced 

“reusable” engine of its kind. While it has seen substantial use in an

operational environment, it has yet to fly on successive shuttle missions 

without removal after each flight.187 The latest 7,480 pound block IIA

version can produce a maximum of 418,660 pounds of thrust at a specific 

impulse of 363 seconds at sea level (512,950 pounds, 452 seconds in 
Decision Maker’s Guide...77 

vacuum) resulting in a 56 to 1 thrust to-weight ratio. Performance is

sufficient for a two-stage-to-orbit vehicle, but it is unlikely its underlying 

technology will ever permit the level of reusability necessary for R2

ISA. 

Proposed systems by major rocket engine developers are pursuing 

development of 100/50 (total mission life/missions to overhaul) systems

with thrust to weight ratios in the mid-seventy range.188 More subtle

design parameters that directly influence engine reliability and reusability

include operating pressures, combustion cycles, combustion efficiencies,

and active versus passive cooling techniques. Rockets with these 

characteristics would be an invaluable R2

ISA enabler, but will not come

cheap. The Apollo-Saturn expendable F-1 and J-2 engines and the

reusable Space Shuttle Main Engine cost between $1.7-2.5 billion to 

develop and flight qualify, with one quarter of these costs attributable to 

the basic development program and the remainder in the “fail-fix” test 

mode. Boeing Rocketdyne recently demonstrated a much lower overall 

cost and more favorable development-to-test cost ratio with the RS-68 

expendable rocket engine that may indicate a sustainable trend for a new 

reusable design.189 However, there can be no mistake that such an engine

will cost multiple billions of dollars to develop and qualify for flight.

Continuing advances in performance and operability and maintainability 

are necessary to successfully demonstrate highly reliable and reusable 

rocket propulsion as a “threshold” enabler for R2

ISA. 

One system concept that bridges the gap between rockets and

airbreathing hypersonics is the air-augmented rocket. In this concept, the 

rocket engine is placed in the flow path of an airbreathing engine. At 

take-off the rocket engine, operating fuel rich, ejects this fuel-air mixture

into the airbreathing engine duct. Analysis indicates a 15 percent increase 

in thrust at zero velocity and a 50 percent thrust enhancement at Mach 

2.190 This approach is one of many that might be used in a rocket-based

combined cycle engine that is discussed in the next section. 

In light of aborted X-33 single-stage-to-orbit and Space Launch 

Initiative two-stage-to-orbit programs, is there a rocket-based solution 

leveraging existing technology that can deliver R2

ISA? A 1996 RAND 

study investigated the utility, feasibility, and cost of procuring a transatmospheric vehicle, or orbital space plane, capable of accomplishing 

military missions.191 A wide range of potential candidates were studied, to 

include three X-33 candidate designs, the X-34, and numerous commercial 

and government air launched space planes (to include the Air Force

Phillips Laboratory Black Horse). The study considered the Northrop 
78...Decision Maker’s Guide

Grumman two-stage-to-orbit space plane particularly promising. 

Estimated developmental costs of $677 million (FY96) would be required 

to build one sub-scale prototype X-vehicle and one full-scale operational 

prototype.192 The vehicle concept included an air-launched “mini-shuttle” 

second stage with a Boeing-747 first stage. The mini-shuttle was an 

180,000 lb horizontal take-off horizontal land system with liquid 

oxygen/hydrogen engines delivering one to six thousand pounds to lowearth orbit (significantly lower payloads to polar orbits).193 More

importantly, however, total system life cycle costs, when compared to an

existing Pegasus expendable system, were very favorable. Despite a large 

disparity in developmental non-recurring engineering costs ($149 million 

to $677 million), a total life cycle cost of $7.9 billion for six transatmospheric vehicles delivered 600 launches (100 launches per vehicle) at 

a cost of $2,900 per pound to low-earth orbit, while an equal sum applied 

to Pegasus expendable launches would buy 290 launches at a cost of 

$34,000 per pound to low-earth orbit for the same overall life cycle cost. 

The level of detail of this analysis strongly suggests that a rocket-based 

reusable launch vehicle architecture delivering routine operation and a 

significant cost savings can be achieved if a high vehicle flight rate can be

sustained.194 However, it also further reinforces the conclusion that

breaking the $1,000 per pound to low-earth-orbit threshold remains highly 

unlikely within existing rocket performance constraints. 

RAND came to the following conclusion in its 1996 report on 

Military Trans-atmospheric vehicles (space plane): 

The X-33 competition, much like the NASP program, has 

focused attention on SSTO vehicles. RAND believes

TSTO TAV concepts deserve equal attention if delivering 

small to medium sized payloads to LEO is viewed as a 

primary mission need. Air-launched TSTO TAV concepts

appear particularly promising from a cost standpoint 

because the first stage aircraft could be based upon a 

commercial civil transport. In addition, they may provide 

an evolutionary development path to full reusability and 

aircraft-like levels of responsiveness for orbital vehicles. 

In contrast, SSTO systems may be more challenging 

technically, much more costly to build, and would be so 

large they could not meet military responsiveness needs.195 
Decision Maker’s Guide...79 

These conclusions are not only fundamentally sound, but perceptive in the 

sense that it anticipated the elevated risks associated with the single-stageto-orbit X-33 program that was ultimately cancelled in 2001. NASA 

reached a similar conclusion when it pursued a two-stage-to-orbit solution

to its short-lived Space Launch Initiative second-generation reusable

launch vehicle. Finally, it suggests that if a flexible and affordable vehicle 

might be possible with a subsonic first stage, then a two-stage-to-orbit 

vehicle with a higher staging Mach number (6 to 10 for example) might 

deliver heavier payloads to low-earth orbit with similar benefits. Yet an 

air launched “mini-shuttle” that delivers one to six thousand pounds to low 

earth-orbit satisfies a smaller niche focused on military payloads. Since 

the preliminary Space Launch Initiative designs intended to deliver 50,000 

pounds to low-earth orbit had a $30-35 billion developmental price tag, 

the question remains: can a rocket-based earth-to-orbit system that 

satisfies military, civil and commercial requirements deliver R2

ISA? 

Answer: Not with existing technology, and certainly not based upon 

NASA’s acquisition philosophy. 

Airbreathing Reusable Launch Vehicles 

The chequered history of hypersonics is one of underlying promise 

pursued in a series of “fits and starts” that has never permitted the 

formation of a stable and coherent long-term technology development 

process. Current interest in hypersonics falls into three general categories: 

hypersonic missiles (engaging time-critical targets) operating between 

Mach 4-8; a hypersonic cruiser (global reconnaissance/strike and 

commercial transport) operating between Mach 4-10; and reusable launch 

vehicles (Mach 8-15 and higher) for space access.196

Hypersonic propulsion relies on a steady air within a suitable band 

of dynamic pressure that defines an atmospheric flight corridor that is 

illustrated in Figure 18. There are three primary regions of hypersonic 

operations. Region 1, which includes operations up to about Mach 8, 

requires the least complex hypersonic vehicles from a technical 

perspective, and appears reasonably achievable and tolerable for sensor 

operation and weapon delivery.197 Region 2, between Mach 20-25 and 

outside the sensible atmosphere, is the most attractive area for hypersonic

missions because of launch flexibility, short flight time, ease of 

maneuverability, and the relatively benign space environment.198 Region 

3 involves sustained cruising flight in the atmosphere between roughly 
80...Decision Maker’s Guide

Mach 8-20. Operation within this region is plagued with high skin 

temperatures, control and maneuverability difficulties, ionized boundary 

layers that make sensor operations difficult, and high infrared 

signatures.199 Hence, it is advisable to avoid Region 3 as a steady-state 

operating or “cruise” environment to minimize exposure to this harsh 

environment. 

Figure 18. Notional Hypersonic Corridor 

Leveraging the benefits of the atmosphere at hypersonic velocities 

comes at a price. The stagnation temperature of air rises dramatically with 

increasing Mach number.200 For example, a vehicle traveling at Mach 4,

6, and 8 encounters corresponding stagnation temperatures of 1,100°F, 

2,500°F, and 4,200°F respectively. With heat addition due to propulsive 

combustion, even higher temperatures of about 4,000°F, 4,400°F, and

5,100°F must be endured.201 No known or projected materials with a 

practical scramjet application could survive these temperatures without 

active cooling at or near Mach 8.202 Supersonic combustion also 

introduces some propulsion/vehicle integration challenges. High airflow 

velocities leave very little time for the introduction, mixing, and 

combustion of the fuel. Hydrogen is an ideal scramjet fuel choice because

it provides higher specific impulse (relative to hydrocarbon-based fuels), 

enables very rapid combustion, can be used for active cooling where high 

temperatures are encountered, and indeed, is the only viable propellant
Decision Maker’s Guide...81 

choice above Mach 8. However, even with hydrogen, scramjet operations 

approaching orbital velocities (above Mach 15) typically demand that the 

entire vehicle act as an integrated inlet/mixer/combustor. The much 

denser hydrocarbon alternative (more appealing for lower stages and/or 

Mach numbers) reaches a stoichiometric limit

203 at Mach 8 and becomes 

unusable above these velocities. 

Two classes of rocket propulsion alternatives, namely turbojets and 

ramjets/scramjets and their respective performance envelopes are 

illustrated in Figure 19. Combined cycle engines couple low speed 

turbines with ramjets, scramjets, or rockets to provide thrust throughout 

the entire vehicle flight regime. Airbreathing propulsion concepts 

currently under development and their associated flight regimes to 

include: Air Turbo-Ramjets, Mach 0-6; Hydrocarbon Scramjets, Mach 5-

8; Hydrogen Scramjets, Mach 5-20+; Hydrocarbon/Hydrogen TurbineBased Combined Cycle Engines, Mach 5-10; Rocket or Turbine-Based 

Combined Cycle Engines (with hydrogen scramjet), Mach 8-15.204 The

wide variety of concepts is emblematic of immature technologies where 

no clear-cut “winners” have yet emerged. However, the current 

scattershot of development activity is at least partially due to either the 

lack of a direct linkage to a developmental flight vehicle or the meager 

funding the hypersonics community has historically endured. The

efficient combination of propulsive technologies into a single reliable 

system is critical to hypersonic vehicle development. 
82...Decision Maker’s Guide

Figure 19. Rocket/Air-Breathing Propulsion Options/Characteristics 

The most successful combined cycle engine to date is the Pratt &

Whitney J-58 Air-Turboramjet developed in the late 1950’s to power the 

SR-71 Blackbird until the 1990’s. This 6,000-pound powerplant was rated 

at 32,500 lbs maximum sea-level thrust and flew in excess of Mach 3 at 

80,000 feet. Any subsequent military air-turboramjet development is 

classified, but theoretical performance in excess of Mach 6 at altitudes of 

120,000 feet is possible.205 NASA is pursuing either solely or in

partnership with DOD and industry the development and demonstration of 

high-speed turbine, hydrogen/hydrocarbon scramjets, as well as turbine 

and rocket-based combined cycle engines. Recent combined NASA/DOD 

funding for all of these efforts has been at a modest $65-85 million per 

year and is slated to drop sharply.206 Limited testing is scheduled through

2005, with the bulk of testing occurring in the 2006-2010 timeframe.207 

Available funding overlaid upon program schedules suggests most of 

these planned tests are in jeopardy. 

There are other ongoing hypersonic technology development

activities targeted at both missiles and flight vehicles. The NASA Langley 

experimental hypersonic ground and flight test program called Hyper-X 

demonstrated hydrogen-fueled scramjet operation on the X-43A test
Decision Maker’s Guide...83 

vehicle on 27 March 2004. The 12-foot long X-43A was dropped from a

B-52B and launched to its operational altitude of 95,000 feet by a Pegasus 

booster rocket. The scramjet ignited as planned and operated for ten 

seconds, the duration of its hydrogen fuel, reaching the targeted top speed 

of Mach 7 before gliding back and impacting in the Pacific Ocean.208 The

ongoing X-43 Program represents the largest and most advanced 

hypersonic test program to date and is part of a plan to flight test many of 

the ongoing engine systems under development. The current status of the 

FY03 X-43 test series is summarized below in Table 15. 

Table 15. X-43 Series Hypersonic Tests209 

Sponsor Funding

($M) 

Goal 

X-43A NASA 185

(FY01-06) 

Flight test of Hyper-X engine to 

demonstrate hydrogen powered 

scramjet to at Mach 7 in 2003 and to 

Mach 10 in 2005210

X-43B NASA

+RBC3 

Flight test ISTAR strutjet engine to

demonstrate a hydrocarbon rocketbased combined-cycle engine in 

2006-2008.211

X-43C NASA+

USAF 

Flight test of HyTech engine and 

demonstrate hydrocarbon scramjet

performance with TBD flight

dates.212

X-43D NASA Flight test hydrogen powered 

scramjet to Mach 15 

The NASA Marshall integrated systems test of an air-breathing

rocket (ISTAR) has prompted a consortium between Aerojet, Rocketdyne, 

and Pratt & Whitney called RBC3. It is a rocket-based combined-cycle 

engine that employs a hydrocarbon-powered liquid rocket system for 

initial acceleration, with a ramjet that ignites at about Mach 2.5, followed 

by a conversion to scramjet mode at Mach 5 to accelerate to Mach 7. The 

engine can then revert back to rocket mode to continue acceleration to

higher velocities as required.213

In the wake of the cancelled X-30 National Aerospace Plane 

program, the Air Force Hypersonic Technology (HyTECH) program was
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established in 1995 with modest funding.214 Pratt & Whitney and the U.S. 

Air Force successfully demonstrated a hydrocarbon-fueled scramjet 

engine under this program operating between Mach 4.5 to 6.5 in January 

2001. The next milestone is a full-scale system designated the Ground 

Demonstrator Engine that will weigh less than 200 lbs and result in a 

flight-worthy engine as early as 2004.215 NASA selected General Electric

in 2002 to begin a turbine-based combined cycle ground demonstration of 

a Mach 4+ turbine engine suitable for a future turbine-based combinedcycle engine with testing planned in the 2006-2008 timeframe.

216 Finally,

the Defense Advanced Research Project Agency and the U.S. Navy plan 

to air launch a powered prototype hypersonic missile in late 2004. The 

four-year Hypersonic Flight (HyFly) program established in January 2002 

plans to demonstrate a Mach 6 dual combustion ramjet217 integrated into a

missile with a 400-600 nautical mile range.218 The U.S. Air Force 

Scientific Advisory Board in 2001 was highly critical of the lack 

continuity and requirements-driven systems engineering driving U.S. Air 

Force hypersonic research. Since that time, an ongoing effort to 

harmonize hypersonic research across all of DOD and NASA appears to 

be coalescing into a “national hypersonic strategy.”219

The ultimate success of a hypersonic airbreathing propulsion 

system on a missile, hypersonic vehicle, or earth-to-orbit launch system 

depends upon its seamless integration with the overall vehicle structure. 

There are three major vehicle design classes to include the most common 

winged-body (space shuttle), the familiar lifting body (X-33), and the less

familiar waverider (X-30). The latter deserves special attention since any 

hypersonic aerospace vehicle will likely be a waverider.220 

In simplest terms, a waverider is any vehicle that uses its own 

shock wave to improve its overall performance, and a correctly designed 

vehicle can ride this wave to produce greater lift (compression lift) and 

less drag, resulting in higher lift to drag ratios and improved performance. 

For example, the delta-winged XB-70 Valkyrie experimental aircraft used 

compression lift to increase overall lift by 35 percent at Mach 3 and 

70,000 feet altitude. Range was so dependent upon this benefit that in the 

event of a loss of one of its engines, it was more economical to go to 

afterburner than to fly in off-design conditions.221 Waveriders only begin 

to become practical at Mach 4 or higher where the shock wave remains

close to the surface of the vehicle.222 The maximum lift to drag ratio 

naturally decreases with Mach number, but actual flight test experience 

has shown that actual measured values fall substantially below even what 
Decision Maker’s Guide...85 

theory predicts. This phenomenon has been dubbed the Lift over Drag 

“L/D” barrier. The waverider design regains performance much closer to 

theoretical limits through compression lift that results from designing a 

vehicle to take advantage of the vehicle’s own shock wave as illustrated in 

Figure 20 below. The precise vehicle shape is primarily dependent upon 

the Mach number and the shock angle that provides the best performance 

(maximum lift over drag ratio). Figure 21 illustrates waveriders optimized 

for Mach 6, 14, and 25 flight. 

In addition to improved performance through a higher lift to drag 

ratio, waveriders also present a shape that can accommodate airbreathing

propulsion integration. As previously discussed, the high flow velocities 

result in very short residence times for compression, mixing, and 

combustion. This problem is addressed most effectively by leveraging the 

entire integrated propulsion/vehicle system to efficiently perform this task. 

Figure 20. Construction of a Body “Waverider” Producing a Conical 

Shock223 
86...Decision Maker’s Guide

Figure 21. Waveriders Optimized for Different Mach Numbers and 

Conditions224

Dr. Kevin Bowcutt and others at the Boeing Phantom Works in

Huntington Beach, California, are collaborating with Air Force Space 

Command and NASA on a horizontal takeoff, two-stage-to-orbit

waverider concept known as FASST (Flexible Aerospace System Solution 

for Transformation).225 The concept would operate from a standard 

runway with a waverider first stage using a hydrocarbon-fueled, turbinebased combined-cycle propulsion system (based upon the NASA Glenn 

Research Center Revolutionary Turbine Accelerator (RTA) program) and 

operate to at least Mach 4 and 70,000 feet altitude.226, 227 The second 

stage orbital vehicle is also a waverider design using a hydrogen-fueled 

rocket-based combined cycle engine transitioning to a pure rocket mode at 

Mach 12.228 Upon separation, the orbiter stage proceeds to conduct onorbit operations. Both stages are recovered with an autonomous 

horizontal runway landing.229 The FASST concept can deliver significant

military capability to include rapid flexible global strike to any global 

position within 90 minutes, replenishment, augmentation, and integration 

of surface/air/space Intelligence, Surveillance, and Reconnaissance assets, 

as well as offensive and defensive space control capability. The concept 

would deliver 25,000 lbs to low-earth orbit, suitable for civil and 

commercial payloads. The first stage boost vehicle used in a standalone 

configuration as a Mach 4 strike/reconnaissance penetrator has significant 
Decision Maker’s Guide...87 

commercial potential as well.230 This specific FASST design concept 

(smaller payload delivery capable vehicles were studied) has a 1.2 million 

pound take-off gross weight (less than either the Antonov-225 or planned 

Airbus-380) with 550,000 and 650,000 pounds allocated to the first and 

second stages respectively. It would be approximately 200 feet in length, 

45 feet in height, with a wingspan of 90 feet.231 Takeoff speed is 250 knots 

based upon an initial thrust-to-weight of 0.4 with a first stage fly-back

range of 130 nautical miles. Design characteristics such as these begin to 

approach the “aircraft-like” operations necessary for R2

ISA. The vehicle 

concept is presented in Figure 22. 

Figure 22. Boeing FASST Concept232

There are some notable advantages as well as concerns regarding

this concept. Nine advantages of this approach include 1) a design 

supporting a national (civil/military/commercial) solution to R2

ISA; 2) an

inherently robust system architecture that provides significant 

performance and design margin; 3) the ability of the first stage to 

accommodate alternative orbital stage solutions (such as a pure rocket); 4)

the availability of the first stage for a host of adjunct missions; 5) a 

horizontal takeoff and landing capability that provides important 

flexibility and robustness over vertically launched systems; 6) a single 

propulsion system in each stage that significantly reduces the complexity 

of engine integration and access; 7) elimination of first stage thermal 

protection requirements precipitating greatly reducing operations and 

maintenance costs; 8) utilizing technology of direct benefit to a potential 

future single-stage-to-orbit successor, 9) inert mass fractions of 45 and 22 

percent for the first and second stages, respectively, with an additional 15 

percent design margin, represent very reasonable mass assumptions for
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existing state-of-the-art. However, there are also some deficiencies with 

this concept that deserve noting. First, the use of densified “slush” liquid 

hydrogen for the second stage increases risk and degrades operability and 

maintainability. Second, the higher dry vehicle mass relative to pure

rocket-based solutions implies a higher developmental cost that can only 

be recovered through significantly lower total system life cycle costs. 

Finally, the advantages of a low staging Mach number are partially offset 

by an additional performance penalty assumed if the second stage

propulsion system reverted to a pure rocket-based solution with a higher 

optimum staging Mach number. However, despite these concerns, the 

FASST concept is fundamentally sound. More importantly, such a design 

successfully “closes” while simultaneously delivering more than 20,000 

lbs to low-earth orbit. Its first stage can accommodate military, civil, and 

commercial upper stages with the potential to evolve into a thirdgeneration space transportation system. Such an approach is one that 

should be seriously considered as a second-generation reusable launch 

system in the pursuit of a national solution to R2

ISA. 

The generation-after-next reusable launch system may no longer 

need to be a two-stage-to-orbit system. The NASA Highly Reusable 

Space Transportation Study completed in 1997 was intended to seek 

innovative concepts and identify advanced technologies to further reduce 

earth-to-orbit launch costs to $100-$200/lb—numbers that included 

development, production, operations, and amortization costs.233 

Approximately twenty concepts were examined in varying levels of detail 

that mapped various operational approaches (to include vertical takeoff/ 

horizontal & vertical landing and horizontal takeoff/horizontal landing) to 

propulsion approaches (to include all-rocket, rocket-based combined 

cycle, and scramjet). The “Argus” concept was the most promising “nearterm” (within10 years) approach that emerged from the study and was a 

horizontal takeoff/horizontal landing vehicle consisting of a largely 

axisymmetric, cylindrical-winged body using a hydrogen fueled rocketbased combined cycle engine and a ground–based catapult launch assist.234 

The concept vehicle body is 171 feet in length and 17.1 feet in diameter 

with a total wingspan of 53.1 feet.235

The baseline Argus vehicle assumes an 800 feet per second (545 

mph) subsonic launch assist with a 0.6 g acceleration and 1.2 g launch 

abort requiring a five mile track. The liquid oxygen/hydrogen 

supercharged ejector ramjet engine accelerates the vehicle to Mach 2 

under supercharged ejector ramjet power. Argus transitions to a fan-
Decision Maker’s Guide...89 

ramjet mode between Mach 2-3 and intercepts a constant 1,500-pound per 

square foot dynamic pressure corridor to Mach 6. The engine operates in 

ramjet mode from Mach 3 to 6 and transitions to pure rocket mode at 

Mach 6 to accelerate a 20,000-pound payload into a circular 100 nautical 

mile, 28.5 degree inclined mission orbit.236 Argus assumed the

availability of very aggressive materials and subsystems technologies to 

include Titanium-Aluminide used as structure and embedded in metal 

matrix composites reinforced with Silicon-Carbide fibers. These materials 

permitted a vehicle dry weight (with 15 percent margin) of 75,500 pounds 

and a takeoff gross weight of approximately 597,000 pounds yielding a 

very aggressive inert mass fraction of 12.6 percent.237 

Advantages to this approach include 1) the underlying technology 

that would support development of an operable supersonic ejector ramjet

engine is well understood; 2) the launch assist has numerous synergistic 

effects in addition to added ∆V, to include a landing gear sized for landing 

weight in lieu of much higher takeoff weight, a much smaller wing 

platform, and a 25-30 percent increase in engine performance at vehicle 

rotation; 3) very simple design: and 4) very low vehicle dry weight and 

small vehicle footprint. Serious drawbacks include a very aggressive inert 

mass fraction, very tight design and performance margins (inherent with 

all single-stage-to-orbit vehicles), and the loss of operational flexibility 

assumed through the use of ground launch assist. It must be noted, 

however, that unrealistically low inert mass fractions plague all launch

vehicle designs. Hence, the best hedge against low inert mass fractions is 

the design and performance margin inherent with the design itself. These 

two examples illustrate how single-stage-to-orbit vehicles such as Argus 

typically lack margin, demanding a very high degree of confidence in 

subsystem performance gained through demonstration before a vehicle is 

built. Alternatively, a two-stage-to-orbit system such as FASST has 

sufficient inherent margin to permit some maneuvering room within the 

design space and provide higher confidence that the objective system can 

meet its overall system performance specification. 

“NATIONAL” SYSTEM ATTRIBUTES 

There is a two-part question that can now be answered: How can 

R2ISA be achieved, and is a national solution possible? A brief synopsis of 
90...Decision Maker’s Guide

the findings and intermediate conclusions presented thus far will frame the

answer. 

First, a quantifiable definition of R2

ISA is necessary. The next 

(second) generation launch system should demonstrate a sustainable 

launch recycle time of 100 hours or less, a mission reliability rate of 99.9 

percent or greater, and a cost per pound to low-earth orbit of $1000 or less. 

The rationale behind these numbers is grounded in supporting analysis as 

well as moderated by current technological limitations and practical 

considerations. Next, the linkages between R2

ISA and the overall vehicle 

design as well as its relationship with an overarching space transportation

architecture make it clear that R2

ISA must be considered in a very broad 

contextual framework. Emphasis on a system-of-systems perspective that 

optimizes the performance trade-offs between the earth-to-orbit and space 

segments keeps all elements of the architecture in proper balance and the 

senior decision-maker focused. 

The second set of issues surround the apparent conflict between 

civil, military, and commercial mission requirements. A 20,000 lb

payload mass to low-earth orbit is a reasonable compromise reached 

between the civil and military communities and valuable to the

commercial sector as well. The military’s preference for a launch-ondemand system with short launch and recycle times is not a priority for 

civil missions. It is, however, a necessary precursor to lower costs. Hence 

amelioration through aggressive movement toward the military 

requirement is appropriate. Reliability of 99.9 percent or greater also 

exhibits an acceptable degree of convergence between the civil and

military communities for a second-generation vehicle. Although the civil 

requirement includes man-rated systems, this objective exceeds

demonstrated shuttle reliability by more than a factor of ten. Higher 

reliability coupled with lower launch costs and sharply reduced cycle 

times is sufficient to induce substantial growth within the commercial 

sector. Clearly, however, the commercial market is the most sensitive to

any demonstrated reliability less than those currently enjoyed by today’s 

airline/package delivery operations. 

Lessons learned are abundant and provide invaluable insight into 

what won’t work. First, there is the realization that R2

ISA is not really

necessary to continue the very expensive, moderate to high risk, low 

launch rate (about thirty annually) environment that perpetuates the status 

quo. Between the Shuttle, Titan, Delta, Atlas, and other foreign 

expendable launch vehicles, the current payload manifest can be easily 
Decision Maker’s Guide...91 

met. Second, “incrementalism” has permeated U.S. space policy for more 

than three decades. The bureaucratic debate over government policy, 

requirements, technology, and resources has derailed any coherent longterm strategy. A close look at the shuttle program illustrates how up-front 

design compromises intended to lower developmental costs later

manifested themselves in astronomically high operations and maintenance

overhead. Third, the X-30 and X-33 programs clearly illustrated that a

neither an airbreathing or rocket-based single-stage-to-orbit system is

currently feasible. The recent Space Launch Initiative restructure 

confirms that a rocket-based two-stage-to-orbit solution conforming to 

rigid NASA requirements and programmatic constraints is unaffordable. 

NASA subsequently chose to defer a second-generation reusable launch 

vehicle development decision to no earlier than 2009. This decision 

removes any possibility of R2

ISA until 2020 or beyond without external 

intervention.

Valuable technical/programmatic lessons learned, to include

technology, risk, requirements, cost/schedule estimates, and organization, 

punctuate the need for major change. Environmental lessons learned to 

include credible and compelling need, recognition of competing interests, 

realistic expectations, national commitment, parallel approaches, small 

steps versus giant leaps, and focused leadership cannot be ignored. These 

experiences in aggregate strongly suggest that a successful reusable launch 

vehicle development program hinges on an evolutionary, capabilitiesbased approach. It also suggests that success rests as much upon a 

national commitment to a politically sensitive, realistic, focused, 

evolutionary solution as it does to making the appropriate design choices 

that are responsive to a clear requirement/market. 

Chapter 3 introduced the technical and economic nature of the 

barriers to R2

ISA. An understanding of the fundamental metrics 

commonly used within the space launch arena provide a powerful tool to

compare expendable versus reusable launch vehicles as well as 

quantifying the impact of technological progress. Four main technology 

areas—propulsion, advanced materials, vehicle integration, and thermal

protection—represent the major requisite technologies for R2

ISA. These

topics, in turn, explain the challenges with single-stage to-orbit and why it 

cannot currently deliver R2

ISA as well as explaining why two-stage-toorbit vehicles can. This difference can be summed up on a single 

concept—design margin—two-stage-to orbit designs have it, single-stageto-orbit designs currently do not. The economic side of the equation is
92...Decision Maker’s Guide

dominated by a market failure in space transportation and makes it clear 

that it is unreasonable for the government to expect or demand a large 

commercial cash commitment until R2

ISA is demonstrated. The swiftest 

and most certain solution demands a disciplined and well-conceived 

government policy backed up by decisive action. 

The economic discussion continues with the arguments explaining 

why expendable launch vehicles can never achieve R2

ISA. A further 

examination of direct launch costs shows a four to five order of magnitude 

difference in labor intensity between aircraft and launch vehicles 

(expendable and reusable alike). This is a chasm that must be bridged if 

space launch operations ever hope to approach routine space flight 

operations akin to what we enjoy today in the terrestrial domain. A close 

examination of shuttle operations not only reinforces this observation, but

also explains how $3.2 billion is spent to support four shuttle launches in 

2002. A functional breakdown of shuttle costs using NASA definitions of 

direct, indirect, and support comprise 10, 20, and 70 percent of the total 

shuttle budget respectively. The latter is comprised of very laborintensive activities across all of the NASA Centers to get the shuttle

prepared for its next launch. 

Chapter 4 centers around technical solutions to the space launch 

impasse surrounding R2

ISA. A close examination of reusable rocket 

technology indicates that additional maturation is required before a pure 

rocket-based solution to R2

ISA is viable. A much closer examination of 

the air-breathing hypersonic status shows that despite the great promise

these technologies afford, the uncoordinated allocation of far too few 

resources has brought inevitably slow progress to date. However, recent 

advances in supersonic and hypersonic air-breathing propulsion indicate 

that near-term solutions to R2

ISA may be possible. 

A summary of major conclusions are that 1) R2

ISA will never be 

achieved with expendable launch vehicle technology while reusable 

launch vehicles are the only economically and operationally viable path 

toward achieving R2

ISA; 2) current technology limitations make pursuit of 

a single-stage-to-orbit solution imprudent at this time while a two-stageto-orbit reusable launch vehicle is the logical next step towards R2

ISA; 3)

despite conflicting requirements, a common “national” solution for civil, 

military, and commercial application is both possible and desirable. One 

possible national solution to R2

ISA is now presented. There are 

undoubtedly others. 
Decision Maker’s Guide...93 

ONE VISION OF SUCCESS

There are a number of vital prerequisites to R2ISA to secure the

road to success. First is the commitment to an evolutionary, capabilitiesbased approach. The space shuttle’s first flight as a prototype was also the 

first flight of the objective system. The design compromises during 

development created shuttle’s unforeseen yet overwhelming ground 

support infrastructure and high labor intensity that could not be reasonably 

exorcised from the system. A robust X-vehicle and Y-prototype 

development approach is essential for both cost and risk management. Dr.

Terry Bahill surveyed 20 projects of widely varying complexity and 

reported his results in the book Metrics and Case Studies for Evaluating 

Engineering Designs. He observes that breakthrough design approaches 

may cost three times what a continuous improvement model does for the 

same performance. He further concludes that projects with “high political 

risk” should be designed with “stable intermediate forms.”238 A second 

vital characteristic of R2

ISA is the utilization of X-vehicles that are part of 

a technology maturation program focused on risk reduction. Y-prototypes 

are built to demonstrate performance, operability, and maintainability as 

well as scalability and traceability to an objective system design. Yprototypes should also have sufficient “residual” capability to capture 

enough missions to justify a reasonably high sortie rate over the short term 

(3-5 years). Once a Y-prototype has served its purpose, it is scavenged for 

parts and sent to the museum. Third, ad-hoc oversight as well as

cooperative agreements between agencies for the purpose of technology 

maturation is a recipe for mediocrity and inefficiency. A stable, wellfunded, joint system program office pursuing the design and development

of an objective system that serves federal agency stakeholders (such as

DOD and NASA) that is at the same time accountable to a higher 

authority is key. These three elements should be considered a prerequisite 

for any serious effort toward R2

ISA. A scenario follows. 

A joint DOD/NASA/DARPA program office should be 

immediately established to develop an overarching space transportation

architecture designed to satisfy military, civil, and commercial space 

requirements. A reusable two-stage-to-orbit earth-to-orbit launch vehicle 

capable of delivering a minimum of 20,000 lbs to low-earth orbit and 

satisfying second-generation R2

ISA criteria should be its top priority.

Currently, about two-thirds of all commercial payloads are satellites
94...Decision Maker’s Guide

between 1-10,000 pounds launched directly into geostationary orbits at 

22,500-mile altitudes.239 Direct insertion of these payloads via a space 

launch vehicle is a very inefficient and expensive practice. A much more 

economical approach is to boost the satellite into low-earth orbit and then 

use a much higher efficiency and flexibly responsive reusable orbital

transfer to finish the job. This permits much smaller launch vehicles due 

to the significantly reduced payload requirements, particularly for 

geostationary transfer orbit (the upper stage(s) are no longer required). 

The objective system first stage is Mach 6-10 horizontal take-off 

horizontal land waverider powered by either a turbine-based or rocketbased combined-cycle engine. A precursor Y-prototype with a 10,000 lb 

to low-earth-orbit capability uses separate hydrocarbon turbine or airturboramjet propulsion system to accelerate to Mach 4-6 before engaging 

a rocket-based second stage for sub-orbital or orbital missions. The

airbreathing first stage can accomplish a powered return to the launch site. 

Separate DOD and NASA upper-stages are tailored to their specific

unique mission requirements. The DOD can develop any combination of 

a separate Space Maneuver Vehicles for space control and earth/space 

reconnaissance, a Modular Insertion Stage for affordable space access, or 

a Common Aero Vehicle to enable prompt global strike. Additional 

military first-stage utility can be explored through tests as a strategic 

reconnaissance and global strike platform capable of reaching any point on 

the earth within two hours and returning to its launch location. A

commercial variant will be the most inexpensive means for low-earth-orbit 

satellite insertion. The commercial variant could also demonstrate the 

ability for global two to three hour package delivery and act as a testbed 

for a potential passenger vehicle. NASA requirements will focus on a 

second stage that maximizes payload delivery to the international space

station. 

The benefits of such a system are numerous. First, it provides a

flexible space transportation solution for DOD, NASA, and commercial 

customers. Second, the core first stage has significant military 

(reconnaissance/strike) and commercial (package delivery/passenger) 

potential. Third, the core first stage vehicle can experience the high flight

rates necessary for R2

ISA. Fourth, a horizontal take off and land system 

with fly-back capability provides a wide array of basing/landing options. 

The use of hydrocarbon fuels keeps vehicle size and dry-weight low 

(relative to rockets and hydrogen-fueled reusables) enabling a vehicle to 

approach aircraft-like operations. Finally, a sub-scale Y-prototype 
Decision Maker’s Guide...95 

powered by turbine/air-turboramjet engines and rockets could be built 

using existing technology. This vehicle would provide the system 

experience necessary to build a highly maintainable and operationally 

efficient objective system, possess residual space launch capability, and 

act as a testbed for sorely needed hypersonic research. The supporting Xvehicle development program would validate the combined cycle engine 

concepts that merit use on the objective system. As rocket and hypersonic

propulsion technology continued to mature, one would likely prove 

superior for incorporation into third-generation single-stage-to-orbit

launch systems. The knowledge and experience gained through the 

approach outlined above would prove integral to its success. 

The linkages between vehicle design and R2

ISA are now well 

established. However, discussion of the relationship between R2

ISA and 

the role it plays within the larger space transportation architecture must 

also be addressed. Figure 23 illustrates this architecture and the role of 

R2ISA within it. 

Figure 23. R2ISA and the Space Transportation Architecture: Vision

versus Reality 

The U.S. space transportation architecture is built to support the 

varied civil, military and commercial space missions and their payloads. 

In a mature architecture, one would expect a top-down flow of 

requirements and few limiting constraints preventing self-optimization of 
96...Decision Maker’s Guide

the elements of the architecture. This creates an environment for the 

proliferation of diverse and inexpensive (likely smaller) payloads 

responsive to changing mission requirements through easy 

upgrade/replacement as the situation dictates. The space infrastructure 

would grow to accommodate easy access and replacement of these 

payloads as well as other scientific and commercial endeavors. Orbital 

transfer vehicles could economically ferry systems to and from their 

mission orbits, and sufficient space-based infrastructure would be built to 

support this activity. R2

ISA permits the fluid movement of personnel, 

payloads and equipment from earth to orbit. Dedicated ground support

infrastructure is minimized, leveraging heavily upon the existing 

military/commercial air transportation infrastructure and existing 

spaceports. A robust space industrial and technical base is free to focus its

resources on new products and services rather than the care and feeding of 

the space transportation architecture itself. However, space, the “dormant 

frontier” as some have called it, today endures a very different reality.

Earth-to-orbit operations are unpredictable, risky, and expensive. Beyond 

the heavily subsidized civil space program, only the most critical defense 

payloads or economically compelling commercial missions are pursued.

To reduce the risk of on-orbit failure and the probability of re-launch, 

many payload systems are made doubly or triply redundant, increasing the 

cost and weight of the satellites. If launch costs can be reduced 

significantly, it may no longer be necessary to design to such high levels 

of redundancy. Further, a large, inefficient, labor-intensive ground 

support infrastructure is required for both the orbital payloads as well as 

the launch vehicles themselves. This situation limits the opportunities 

afforded the industrial/technical base, stunting its growth, limiting its 

economic viability as a commercial profit center, and even threatening its 

existence. 

Furthermore, there are subtle interactions between payload and 

launch costs. As launch costs increase, so do payload costs. It may be 

possible to recover payloads in orbit, and if payloads were designed 

modularly, they could be quickly repaired on-orbit. Such payloads could 

cost considerably less than existing satellites. Such a capability may 

enable a new era of low-cost access to space.”

240

R2ISA is the enabling infrastructure element that can overcome the 

current impasse, but maximum benefit can only be realized if the other 

architecture elements can evolve with it. For example, the current annual

commercial space launch market demand is about thirty launches. Most of 
Decision Maker’s Guide...97 

these involve large systems requiring insertion into a geostationary 

transfer orbit, unserviceable by a 20,000 lb to low-earth-orbit payload 

capability. However, these same satellites could be delivered to low-earth 

orbit where an on station orbital transfer vehicle could economically insert 

these satellites into their mission orbits. Although the incentive to build 

large, redundant satellites will quickly erode once R2

ISA is demonstrated,

it makes better economic sense to have these smaller payloads ready from 

the outset. The space infrastructure would quickly grow to accommodate

this change, which could include the servicing of orbital transfer vehicles 

as well as certain high value payloads once it made good economic sense. 

The complexity and scope of the issues discussed suggest that a 

new centralized management authority is appropriate. A National 

Aerospace Transportation Agency responsible for developing an 

overarching space transportation architecture, accountable to the 

President, and responsive to military, civil, and commercial needs could

provide that authority. This agency would not be designed to compete 

with NASA, but rather take responsibility for the development and early 

operation of a complete space transportation architecture focused on

securing long term economic viability, compatibility and interoperability 

between architecture elements, and the optimum allocation of limited 

resources. Ultimately, it would spin off all major elements of the space 

architecture back to NASA, DOD, other government agencies (i.e. FAA) 

and commercial industry before dissolving itself. Or alternatively,

become a National Aerospace Transportation Authority responsible for the 

efficient operation of the space transportation infrastructure it creates.

During the build-out of this architecture, the Agency must have the 

authority to “coercively” integrate and ameliorate disparate Civil, DOD, 

and commercial requirements into a single joint space transportation 

architecture. The Joint Program Office responsible for second generation 

space launch system development would be comprised of DOD, NASA,

and DARPA personnel and take these adjudicated requirements to design, 

develop, and test the joint space launch system designed to deliver R2

ISA. 

The agency would require its own budget, but initial budget

authority as well as programs would be extracted from the relevant

developmental activities ongoing within DOD, NASA, and DARPA. This

includes the existing remnants of NASA’s Space Launch Initiative, the 

proposed National Aerospace Initiative effort, the coalescing National 

Hypersonics Strategy, X-Vehicle programs to include X-37 and X-43-like 

programs. The total budget authority as well as the oversight and 
98...Decision Maker’s Guide

management of the space shuttle and International Space Station would 

also fall under the new agency within two years. This should free NASA 

to pursue its original charter in lieu of managing burdensome cash 

strapped programs. Current plans to develop an Orbital Space Plane 

should be accelerated in conjunction with either man-rating the EvolvedExpendable Launch System or a developing companion crew escape 

system for the space plane. Once both are in place and crew transfer to 

and from the international space station is validated, shuttle operations 

should cease immediately. The fleet should then be “mothballed” or 

partially dismantled, and all planned shuttle life-extension upgrades 

cancelled. All unused budget authority reverts to the Agency. Ideally,

Agency funding should come under the oversight of only one 

authorization and appropriation subcommittee in each house of Congress. 

Some may consider the methods described above draconian or politically 

unwieldy. While one may concede either, thirty years of failure highlight 

the need for systemic change. 
Decision Maker’s Guide...99 

V. Conclusions and Recommendations 

If a sufficient number of management layers are 

superimposed on each other, it can be assured that disaster 

is not left to chance. 

—Augustine’s Law Number XXVI 

By the time the people asking the questions are ready for 

the answers, the people doing the work have lost track of 

the questions. 

—Augustine’s Law Number XXX 

What is R2ISA and why is it important? 

Robust, Reliable, and Inexpensive access to space (R2ISA) is the 

cornerstone of continued U.S. civil, military, and commercial leadership in

space. Existing expendable and reusable launch systems fail to 

demonstrate R2

ISA by a wide quantifiable margin and artificially constrain

U.S. space exploitation. A second generation reusable launch vehicle 

capable of a 100 hour launch re-cycle time, no higher than a 1/1000 

probability of loss of vehicle, at a cost of less than $1000 per pound to 

low-earth orbit has the potential to unleash an increasing non-linear 

demand for civil, military, and commercial space activity. 

R2ISA must be developed as an element of a larger space 

transportation architecture 

R2ISA is a key component of a larger space transportation 

architecture that includes missions and payloads, space infrastructure, 

earth-to-orbit operations, ground infrastructure, and a technical/industrial 

base. Virtually all existing satellites are 20,000 pounds or less, defining a

quantitative upper limit to low-earth-orbit payload delivery capability. 

Space launch vehicles are a very inefficient way to drive payloads to

mission orbits/velocities once a minimum threshold orbital velocity is

achieved. By leveraging a robust set of highly efficient orbital 

maneuvering vehicles and a 20,000 pound upper payload limit, much 
100...Decision Maker’s Guide 

smaller and operationally efficient reusable space launch capability that 

delivers R2

ISA is enabled. Fully matured R2ISA will make smaller and 

lighter satellites more attractive. Smaller payloads, coupled with suitable 

technology, can result in even smaller and supportable third-generation 

launch systems, drastically reducing the size and scope of a costly, laborintensive ground support infrastructure. This in turn permits the 

reallocation of a limited technical/industrial base toward the development 

of new products and services in lieu of servicing and adding cost to an 

already bloated space transportation architecture. Opening up the entire 

space transportation “trade-space” is an essential prerequisite to achieving 

R2ISA. 

Economic Reality Discourages Investment in R2ISA 

Economic factors play a pivotal role in enabling R2ISA. The 

commercial space launch market is characterized by a market failure 

similar to that experienced by past commercial rail, automobile, and air

transport systems. Market elasticity (a one percent drop in price results in

a greater than one percent increase in demand) enters the commercial

launch market at around $1000 per pound to low-earth orbit. Current 

prices between $2000 per pound for Russian/Chinese systems and $5000 

per pound for Western systems that result in an inelastic market (a one

percent drop in price produces no more than a one percent increase in 

demand), combined with relatively poor reliability and robustness,

undermine any economic incentive for commercial enterprise to invest and 

improve launch vehicle performance. Hence, it is unreasonable for the 

government to expect any voluntary large-scale commercial participation 

in a space launch system until R2

ISA is demonstrated. A closer

examination of space launch vehicle direct launch costs shows a four to 

five order of magnitude difference in labor intensity between aircraft

systems and launch vehicles for both expendable and reusable systems.

The underlying root causes for this gap must be eliminated to permit

R2

ISA. Finally, economic constraints combine with fundamental

technology limitations to prevent expendable launch vehicles from 

achieving R2

ISA. 
Decision Maker’s Guide...101 

Technology Investment is a Key Enabler for R2ISA 

Demonstrated ability to harness technology to overcome the 

fundamental physical challenges of accelerating payloads to orbital 

velocity remains an R2

ISA limiting factor. Two-stage-to-orbit solutions 

are necessary for second-generation reusable space launch, while more

operationally and economically efficient single-stage-to-orbit vehicles are 

possible for third and fourth-generation reusable launch systems. Existing 

technology demands single-stage-to-orbit inert (structural) mass fractions

below 10 percent of overall vehicle mass. Although this has been 

demonstrated for expendable launch systems, it is currently unrealistic to 

build single-stage-to-orbit reusable vehicles with structural mass fractions

less than 20 to 30 percent of total vehicle mass. Current technology 

provides insufficient design and performance margin to support the 

development of a single-stage-to-orbit vehicle capable of R2

ISA. 

However, sufficient margin exists to support a two-stage-to-orbit R2

ISA

that simultaneously demonstrates future key enabling single-stage-to-orbit 

technologies. Four main enabling technology areas include propulsion, 

advanced materials, thermal protection, and vehicle integration. 

Additional technology maturation is required before a purely reusable 

rocket-based R2

ISA solution is viable. However, government investment 

in both rocket and combined-cycle hypersonic engine technologies is 

orders of magnitude below what is necessary to enable R2

ISA. Recent 

advances in supersonic and hypersonic air-breathing propulsion indicate 

that a near-term two-stage-to-orbit solution is possible, specifically for

lower Mach number first stage applications. 

Mission Requirement Adjudication

Seemingly divergent civil/military/commercial requirements are

reconcilable, implying that a “national” space launch solution is possible. 

A 20,000-pound payload mass to low-earth orbit, supplemented with a

robust space infrastructure to include orbital transfer vehicles, can satisfy 

all civil and military requirements and is valuable to the commercial sector 

as well. The military’s need for a launch-on-demand system with short 

flight-recycle times is not a fundamental priority for civil missions, but it 

is a necessary precursor to lower costs. Reliability of 99.9 percent (1/1000 

probability of loss of vehicle) or greater exhibits an acceptable degree of

convergence between the civil and military reliability requirements in the
102...Decision Maker’s Guide 

near term. The combined impact of higher reliability coupled with lower 

launch costs and sharply reduced recycle times of second-generation 

R2ISA is sufficient to induce substantial growth within the commercial 

sector. 

Lessons Learned 

Today’s space launch environment is laden with an inertia making 

progress difficult. R2

ISA is not necessary to meet today’s low global space 

launch demand that perpetuates the status quo characterized by an inelastic 

market providing expensive, risk laden and inflexible launch services.

Existing capacity for launch services far outstrips demand, removing the

incentive for even modest incremental improvements in existing systems 

until absolutely necessary, as well as completely undermining new

commercial ventures intending to enter the market. The space shuttle 

program was originally intended to deliver revolutionary capability at 

revolutionary prices ($385 per pound to low earth orbit in 1971) that that 

today exceeds $10,000 per pound. The shuttle’s first flight as a prototype 

was also the first flight of the objective system. Hence, design 

compromises in development that created shuttle’s unforeseen yet

overwhelming ground support infrastructure and high labor intensity could 

not be reasonably exorcised from the system. Further, an order of 

magnitude shortfall in annual flight rate (from forty to four) artificially 

ballooned the program overhead, dwarfing incurred marginal individual 

mission costs. A functional breakdown of shuttle costs using NASA 

definitions of direct, indirect, and support comprise 10, 20, and 70 percent 

of the total $3.2 billion FY02 shuttle budget respectively. The latter is 

comprised of very labor-intensive activities across all of the NASA 

Centers to get the shuttle prepared for the next launch. A solution to the

shuttle conundrum has to be found that does not compromise

commitments to the international space station and manned space flight,

while freeing sufficient resources to develop a space launch system that 

delivers R2

ISA. 

The aborted X-30 and X-33 programs make it clear that neither an

airbreathing nor rocket based single-stage-to-orbit system is ready for 

development and production. The recent Space Launch Initiative 

restructure indicates that a rocket based two-stage-to-orbit solution

conforming to NASA requirements and programmatic constraints is 
Decision Maker’s Guide...103 

clearly unaffordable. However, these lessons do not imply that pursuit of 

R2ISA is unachievable or ill advised. On the contrary, it merely demands a 

national commitment to an evolutionary, capabilities based spiral

development approach in lieu of a technology leveraged, high risk,

performance driven point design solution. R2

ISA demands leveraging 

technology to infuse system reliability, robustness, and margin rather than 

an over-fixation on maximizing performance. 

Breakthrough design approaches may cost three times what a 

continuous improvement model does for the same performance, and 

projects with “high political risk” should be designed with stable 

intermediate forms. Continuity provides a symbiotic maturation of the 

vehicle design with the supporting ground infrastructure that can only be 

gained through operational experience. Prudent use of X-vehicles for

technology maturation and Y-prototype development for risk reduction 

can fill the void that lies between a computer aided design and the 

production system. Many X-vehicles designed and flight-tested to 

demonstrate high-risk technology solutions have been very successful. 

However, no X-vehicle intended to fly as a prototype for a follow-on 

production system has ever flown. The X-30 and X-33 clearly illustrate 

this point. A Y-prototype is most useful in demonstrating successful 

system integration and is invaluable tool for controlling cost and risk for

moderate to high-risk system development. Y-prototypes must clearly 

demonstrate scalability and traceability to an objective system design as 

well as possess sufficient “residual” capability to capture enough missions

to justify a reasonably high sortie rate over its relatively short 

“operational” life (3-5 years). Residual capability ensures a sufficient 

sortie rate that in turn provides the operational experience necessary to 

learn the tough lessons that come only with flying and operating the 

vehicle. This is the best way to gain sufficient insight to infuse operability 

and maintainability into the objective system. Thirty years of failure

should make it clear that taking shortcuts will only lead to more failure. 

R2ISA is enabled by continuity in system development, not through a 

discontinuous, revolutionary vehicle design approach. 

Vital Reform 

Ad-hoc oversight or unenforceable cooperative agreements

between agencies as an enabler for R2

ISA is also highly suspect. There is 
104...Decision Maker’s Guide 

an array of promising space technology programs funded across the

federal government that are part of “technology roadmaps” that lead to 

nowhere. Nowhere is there a link to an objective launch system that might 

enable some rational prioritization of scarce resources. Today it is 

currently done at best by cooperative, well-intentioned “enlightened” 

intuition, and at worst by political or personal agendas. Further, the lack

of an objective system removes the imperative for any sorely needed 

increases in budgets or accelerations to programs. The solution demands a

stable, well-funded joint system program office serving all federal agency 

stakeholders (including DOD and NASA) chartered to deliver R2

ISA and 

held accountable to a higher authority. A joint DOD/NASA/DARPA

program office should be immediately established to develop an 

overarching space transportation architecture designed to satisfy military, 

civil, and commercial space requirements. A reusable two-stage-to-orbit 

earth-to-orbit launch vehicle capable of delivering a minimum of 20,000 

lbs to low-earth orbit and satisfying second-generation R2

ISA criteria 

should be its top priority. 

R2ISA is the key element and current limiting factor to the 

development of a robust space transportation architecture. The complexity 

and scope of circumstances that have resulted in a 30-year impasse with

no foreseeable solution demands the establishment of a National 

Aerospace Transportation Agency (NATA) responsible for the 

development of an overarching space transportation architecture 

responsive to civil/military/commercial mission needs. The agency would 

be led by a Presidential appointee guided by an Executive Steering 

Committee comprised of NASA administrator; Deputy Secretary of

Defense; Director, Defense Developmental Research and Engineering;

Undersecretary of the Air Force for Space; and designated representatives 

from industry to serve in an advisory capacity. This agency would 

complement rather than compete with NASA by taking direct control of

the development, early operation, and successful maturation of a complete 

space transportation architecture focused on securing long term economic 

viability, compatibility and interoperability between architecture elements 

as well as determining the optimum allocation of limited resources.

Eventually, it would spin off all major elements of the space architecture 

back to NASA, DOD, other government agencies (i.e. FAA) and 

commercial industry before ultimately dissolving itself. Alternatively, if 

deemed appropriate, the agency could divest itself of all developmental 

activities and become the National Aerospace Transportation Authority.
Decision Maker’s Guide...105 

The Agency must have the authority to fully integrate and ameliorate 

disparate Civil, DOD, and Commercial requirements into a single joint

space transportation architecture. The Joint Program Office would take

responsibility for the research, development, test and evaluation of any Xvehicle and Y-prototype systems and be comprised of DOD, NASA, and 

DARPA personnel. The using organizations (i.e. NASA, USAF, NRO,

etc.) would fund operations and maintenance of the prototype systems that 

were supporting their missions. Where “voluntary” utilization may not be 

forthcoming, it may need to be mandated. All ongoing DOD, NASA, and 

DARPA programs deemed relevant to R2

ISA architecture development 

would be transferred to the Agency. Recent historical precedent for such

action can be found with the Presidential actions surrounding the 

transitioning of the Ballistic Missile Defense Office into the Missile 

Defense Agency. 

NATA would require its own budget with initial budget authority 

as well as programs transferred to the agency from DOD, NASA, and

DARPA. This includes the remnants of NASA’s Space Launch Initiative, 

DOD’s proposed National Aerospace Initiative, any coalescing programs

within the National Hypersonics Strategy, and the X-37 and X-43 

programs. The oversight, management, and budget authority of the space 

shuttle and International Space Station would also fall under the new

agency within two years. This action would free NASA to pursue its

original charter in lieu of managing burdensome cash strapped programs.

Current plans to develop an Orbital Space Plane should be accelerated in 

conjunction with man-rating the Evolved-Expendable Launch System.

Once both are in place and crew transfer to and from the International 

Space Station is validated, shuttle operations should cease soon thereafter. 

Ideally, agency funding should come under the oversight of only one

authorization and appropriation subcommittee in each house of Congress. 

A Proposed Solution 

A two-stage-to-orbit space-launch system using a first stage 

waverider design leveraging combined cycle engine technology is the best 

and fastest route to R2

ISA. First, it provides a flexible space transportation

solution for DOD, NASA, and commercial customers. Second, the core

first stage has significant military (reconnaissance/strike) and commercial 

(package delivery/passenger) potential. Third, the core first stage vehicle 
106...Decision Maker’s Guide 

can experience the high flight rates necessary for R2

ISA. Fourth, a

horizontal takeoff and land system with fly-back capability provides a 

wide array of basing/landing options. The use of hydrocarbon fuels keeps 

vehicle size and dry-weight low (relative to rockets and hydrogen-fueled 

reusables) enabling a vehicle to approach aircraft-like operations. Finally, 

supporting X-vehicle program can validate combined cycle engine 

concepts that merit use on the objective system while a sub-scale Yprototype first stage could be built today using existing propulsion

technology (either rockets, advanced turbines, or pulse detonation) until a 

turbine combined cycle engine was ready. This vehicle would provide the 

system experience necessary to build a highly maintainable and 

operationally efficient objective system, possess residual space launch 

capability, and act as a test bed for sorely needed hypersonic research. 

The two-stage-to-orbit objective system would deliver up to 

20,000 pounds to low earth-orbit (100 nm, 28.5 deg inclination) using a

Mach 4-6 horizontal take-off / horizontal land waverider powered by 

hydrocarbon-fueled Revolutionary Turbine Accelerator derived turbinecombined-cycle-engine with the capacity to return to its launch site. The 

second stage would ideally be an airbreathing waverider using a rocketbased combined cycle engine, but could implement a simpler 

wing/fuselage rocket design as an alternative if sufficient second stage 

performance can be demonstrated. Separate DOD and NASA upperstages could be developed and tailored to accommodate their unique

mission requirements. 

The DOD could develop any combination of a separate Space

Maneuver Vehicle for space control and earth/space reconnaissance, a

Modular Insertion Stage for affordable space access, or a Common Aero

Vehicle to enable prompt global strike. Additional military first stage 

utility can be explored through tests as a strategic reconnaissance and 

global strike platform capable of reaching any point on the earth within 

three hours and returning to its launch location. A commercial first stage 

variant will enable inexpensive low-earth-orbit satellite insertion,

demonstrate global two to three hour package delivery, and serve as a test 

bed for a potential future passenger vehicle. NASA requirements will 

likely focus on a second stage maximizing payload delivery to the 

International Space Station. As rocket and hypersonic propulsion

technology continues to mature, one will likely prove superior for

incorporation into third-generation single-stage-to-orbit launch systems. 
Decision Maker’s Guide...107 

The knowledge and experience gained through the approach outlined 

above would prove integral to its future success. 

In the Final Analysis... 

R2ISA as part of an integrated space transportation infrastructure is

the key to unlocking the limitless civil, military, and commercial potential 

of space. This paper has introduced one potential path to achieving it that 

if fully implemented, can guarantee ultimate success. The United States 

with the close cooperation of NASA, DOD, industry, and academia must 

lead the way to muster and organize the means to build a sufficiently 

robust, reliable, and inexpensive reusable launch system to evolve space 

exploration and exploitation beyond the realm of limited government and 

commercial enterprise into the public domain. Mark Twain wrote more

than a century ago, "Twenty years from now you will be more 

disappointed by the things that you didn't do than by the ones you did do. 

So throw off the bowlines. Sail away from the safe harbor. Catch the 

trade winds in your sails. Explore. Dream. Discover." These words were 

meant to inspire the individual. Today they hold new meaning for the

Nation. It is time to roll up our sleeves and get on with building R2

ISA. 

To do so will free countless individuals of future generations to “Explore, 

Dream, and Discover.” America and the world will be better for it. 
108...Decision Maker’s Guide 
Decision Maker’s Guide...109 

Notes 

1

 This was the U.S. Air Force motivation behind “assured access” that 

ended USAF reliance on a single launch system after the near 

simultaneous loss of Challenger and other expendable launch vehicle 

failures in the 1986-1987 timeframe. Assured access included maintaining 

the viability of two separate and distinct launch vehicle vendors within the

Evolved Expendable Launch Vehicle program. Hence high costs do not 

necessarily impede established high-value missions such as the insertion 

of a lucrative commercial communication or sensitive military payloads

into geostationary orbit. 

2

 “Commercial Space Transportation Study,” NASA Langley 

Research Center, Hampton, VA, 1994. 

3

 Commercial Space Transportation Study, Section 1.1. 

4

 Commercial Space Transportation Study, Section 4.1.3. 

5

 Inflating $600 and $1000 from FY94 dollars, based upon the U.S. 

Consumer Price Index (CPI) to 2002 yields $723 and $1211 respectively.

The $800 average between the two FY94 values inflates to $969 in CY 

2002 dollars. This value is conservatively rounded to $1000. “What is a

Dollar Worth?” Federal Reserve Bank of Minneapolis, on line, Internet, 

30 November 2002, available from

http://minneapolisfed.org/research/data/us/calc. 

6

 The National Aeronautics and Space Administration uses the first,

second, and third-generation nomenclature in its Integrated Space 

Transportation Plan describing the space shuttle as a first-generation 

reusable launch vehicle. Subsequent systems are described as second and 

third-generation systems respectively. 7

 Commercial Space Transportation Study, Section 4.1.3. 

8

 William R. Claybaugh, Economics of Space Transportation, AIAA 

Professional Development Short Course, (Washington, DC: American 

Institute of Aeronautics and Astronautics, 2002), 148. Percent

probabilities quoted are at 95% confidence. 

9

 Claybaugh, 155-159. 

10 “A Review of Space Shuttle Data: Replaced Items – LRUs per 

Flight,” National Aeronautics and Space Administration, on-line, Internet, 

30 November 2002, available from
110...Decision Maker’s Guide 

Notes 

http://science.ksc.nasa.gov/shuttle/nexgen/Nexgen_Downloads/A_Review

_of_Space_Shuttle_LRU_Data.doc. 

11 Claybaugh, 161-162. 

12 National Aeronautics and Space Administration, Introduction to

NASA’s Integrated Space Transportation Plan and Space Launch 

Initiative, (May 17,2001), 3. 

13 Robert Wall, “Pentagon Accidents, Fatalities Rise,” Aviation Week

and Space Technology, October 7, 2002, 31. DOD Class A tri-service 

accidents per 100,000 flight hours typically range between 1 and 1.5. 

14 “Space Transportation Costs: Trends in Price Per Pound to Orbit 

1990-2000,” Futron Corporation, Bethesda, MD, 6 September 2002, 2-3. 

15 Orbit inclination ranges from zero to 180 degrees. A zero inclined 

orbit travels over the equator in a west to east (prograde) direction.

Rotating this orbit ninety degrees results in a polar (90 degree inclined) 

orbit, while rotating it a full 180 degrees results in an equatorial orbit 

traveling in a westerly (retrograde) direction. A due east launch from

Cape Canaveral (at 28.5 latitude) defaults into a 28.5 degree inclined orbit 

and maximizes the benefit of the earth’s rotational velocity to achieve

orbit. 

16 As quoted by AIA President Don Fuqua, Military Space (12 

December 1994), 1. 

17 Frank Morring, Jr., “NASA to Leave ISS Crew Size Issue Open,” 

Aviation Week and Space Technology, 18 November 2002, 45. 

18 Paul Dergarabedian, Comments on National Space Transportation, 

position paper, The Aerospace Corporation, El Segundo, CA, 1988, 2-3. 

19 Steven J. Isakowitz, Joseph P. Hopkins, and Joshua B. Hopkins, 

International Reference Guide to Space Launch Systems, Third Edition,

(Washington D.C.: American Institute of Aeronautics and Astronautics, 

1999), 415. The shuttle was estimated to cost $10.5 million and designed 

to deliver up to 65,000 pounds to low earth orbit. 

20 Joan Johnson-Freese and Roger Handberg, Space the Dormant 

Frontier: Changing the Paradigm for the 21st Century (Westport, CT: 

Greenwood Publishing Group, 1997), 89. 

21 “Legislative Update,” Aerospace America, November 2002, 12. 

FY03 NASA appropriations bill provided $3.2B for four space shuttle 
Decision Maker’s Guide...111 

Notes 

launches supporting the International Space Station. This translates into 

$800M per launch, or approximately $16,000 per pound to LEO. 

22 Scott Pace, “U.S. Space Transportation Policy,” Space Policy, 

November 1988, 321. 

23 John R. London, LEO in the Cheap: Methods for Achieving 

Drastic Reductions in Space Launch Costs, Air University Research 

Report No. AU-ARI-93-8 (Maxwell AFB, AL: Air University Press, 

1994), 10-11. 

24 Jeffery M. Lenorovitz, “Congress Cancels NLS Launcher Family,” 

Aviation Week and Space Technology, 12 October 1992, 20. 

25 Johnson-Freese, 24. 

26 Johnson-Freese, 23. 

27 The National Commission of Space, Pioneering the Space Frontier

(New York: Bantam Books, 1985), 12. 

28 Matt Bille, Tony Williams, Vic Villhard, and Lisa Reed, “Dreams 

to Reality: The Path to the RLV,” AIAA, IEEE, and AAS Core 

Technologies for Space Systems Conference Proceedings, November 

2002, 4. 

29 Paul Hoversten, “U.S. Air Force Won’t Fund X-33, X-37 Space 

Planes,” AviationNow.com, 6 September 2001, on-line, Internet, 18 

December, 2002, available from 

http://www.aviationnow.com/avnow/news/channel_space.jsp?view=story

&id=news/smil090601.html

30 Bille, 5-6. 

31 “Boeing’s Phantom Works gets $301M X-37 contract,” St Louis 

Business Journal, 25 November 2002, n.p., on-line, Internet, 18 April

2003, available from http://stlouisbizjournals 

.com/stlouis/stories/2002/11/25/daily23.html. 32 The Space Launch Initiative is a direct descendant of the NASA 

Space Transportation Architecture Study that concluded that $5B of 

government investment would be needed to enable commercial RLV 

development. Refer to http://www.hq.nasa.gov/office/codea/codeae/ 

stas_results.html. 

33 National Aeronautics and Space Administration, Introduction to

NASA’s Integrated Space Transportation Plan and Space Launch 

Initiative, 17 May 2001, 4. 
112...Decision Maker’s Guide 

Notes 

34 “The Administration/NASA Integrated Space Transportation Plan,” 

National Aeronautics and Space Administration, n.p., on-line, Internet, 25 

October, 2002, available from http://www. 

aerospace.nasa.gov/home&home/glenn/invasp/tsld004.htm. 35 Stephen Trimble, “NASA Whittles SLI Concepts to 15,”

AviationNow.com, 30 April 2002, on-line, Internet, 18 December 2002, 

available from http://www.avaiationnow.com/

avnow/news/channel_space.jsp?view=story&id=news/ssli043002.html. 36 “NASA Postpones Review of SLI Program,” AviationNow.com, 22 

Oct 2002, on-line, Internet, 18 Dec 2002, available from 

http://www.aviationnow.com/avnow/news/channel_ 

space.jsp?id=news/csli102202.html. 37 Jefferson Morris, “NASA Budget Amendment Delays Next-Gen 

RLV, Cuts SLI By $2.323 Billion,” AviationNow.com, 13 November 

2002, on-line, Internet, 18 December, 2002, available from 

http://www.avaiationnow.com/avnow/news/channel_space.jsp?view=story

&id =news/snextgen121302.html. 38 Robert K. Steele and Peter L. Portanova, “EELV: New Acquisition 

Strategy for a New Launcher,” Aerospace America (July 1996), 38. 

39 Johnson-Freese, 90. 

40 Glenn W. Goodman, Jr., “Assured Access to Space,” Armed Forces 

Journal International, July 2002. As of 26 August, 2002, the Air Force 

has awarded 21 launch contracts to Boeing worth $1.5B and seven to 

Lockheed-Martin worth $500M. 

41 Sidney Kimhan, Janet Grondin, et.al., “EELV Program – An

Acquisition Reform Success Story,” Program Manager, vol. 28 no. 3 

(May-June, 1999): 87-88. 

42 “Military EELV Launches Could Save $10B,” Aviation Week and 

Space Technology, 26 August 2002, 23. 

43 Charles E. Lindbloom, “The Science of Muddling Through,” Public 

Administration Review, no. 19 (1959), 79-88. 

44 “NASA Readies 2nd Generation RLV Project for the Chop,” 

SpaceDaily, 22 Oct 2002, n.p., on-line, Internet, 25 October 2002, 

available from http://www.spacedaily.com/news/ rocketscience-02zh.html. 

45 The Crew Return Vehicle (CRV) will initially fly to the 

International Space Station (ISS) unmanned using a heavy-lift configured 
Decision Maker’s Guide...113 

Notes 

EELV to serve as a lifeboat to return astronauts in the event of an

emergency aboard ISS. This configuration may eventually be crew rated, 

eliminating the requirement to use the Space Shuttle for ISS mission crew 

transport. 

46 “Reassessing Space Shuttle Upgrades Strategy,” SpaceRef.com, 3

April 2002, n.p., on-line, Internet, 28 February 2003, available from 

http://www.spaceref.com/news/ viewers.html?pid=5124. 

47 “Legislative Update,” Aerospace America, November 2002, 12. 

48 Jess Sponable, “Military Spaceplane S&T Demonstrator: Two

Stage Hypersonic Testbed,” Air Vehicles Directorate, Air Force Research

Laboratory, Wright Patterson Air Force Base, OH, 1 July 2002, 3. 

49 “Air Force Space Command Strategic Master Plan,” Headquarters

USAF Space Command, 9 January 2000, on-line, Internet, 18 April 2003, 

available from

http://www.spacecom.af.mil/hqafspc/library/AFSPCPAOffice/chapterone.

htm. Definitions include Space Force Enhancement: operations conducted 

from space to support forces; e.g., navigation, communications, 

reconnaissance, surveillance, ballistic missile early warning, and 

environmental sensing. Space Forces Support: sustain, surge, or

reconstitute elements of military space systems or capabilities (includes 

spacelift and satellite command and control). Space Control: ensuring 

friendly use of space while denying its use to the enemy. To accomplish 

this, space forces must survey the space environment, protect our ability to 

use space, prevent adversaries from interfering with that use, and negate 

the ability of adversaries to exploit their own forces. Space Force 

Application: attacks against terrestrial, airborne, or space targets carried

out by weapons operating in or through space. 

50 Strategic Master Plan FY04 and Beyond, United States Air Force 

Space Command, 1-21, on-line, Internet, 5 November 2002, available 

from http://www.spacecom.af.mil/hqafspc. 

51 Daniel Gonzales, Mel Eisman, et al., Proceedings of the Rand 

Project AIR FORCE Workshop on Transatmospheric Vehicles, RAND, 

MR-890-AF, 1997, xiv. 

52 Gonzales, xiv. 

53 Gonzales, xv. 54 Gonzales, 16. 
114...Decision Maker’s Guide 

Notes 

55 Gonzales, 17. 56 “NASA-USAF Reusable Space Launch Development: 120 Day 

Study Overview,” presented to the National Aerospace Initiative Steering

Group, 5-7 March 2002, 4. 

57 John Hawley, “Red Team Observations and Recommendations on 

120-Day NASA/USAF RLV Study,” 2002, 4. 

58 Paul Branch, “30 Day Launch Forecast,” HQ USAF/XOO, the 

Pentagon, 16 January 2003. The Soyuz-U launch vehicle delivers the 

Progress spacecraft (unmanned Russian space station resupply ferry) to

low-earth orbit. Progress weighs 16,095 pounds, carries 3,970 pounds of 

consumables in the cargo compartment and 2,072 pounds of fuel (for orbit 

insertion and de-orbit) in the service module. 

59 “120 Day Study Overview,” 25. 

60 “120 Day Study Overview,” 16. 

61 “120 Day Study Overview,” 25. 

62 Hawley, 5. 

63 “2003 Aerospace Source Book,” Aviation Week and Space 

Technology, 13 January 2003, 160-164. 

64 “2003 Aerospace Source Book,” 151-152. 

65 Commercial Space Transportation Study, Section 2.0. 

66 Commercial Space Transportation Study, Appendix A, Figure 1-1. 

67 Consolidated Space Transportation Study, Section 3.5.1. 

68 Stephen J. Isakowitz, International Reference Guide to Space 

Launch Systems (Washington D.C.: American Institute of Aeronautics 

and Astronautics, 1991), 246-246. 

69 Isakowitz, 1991, 247-248. 

70 Carey M. McCleskey and Edgar Zapata, “Designing for Annual 

Spacelift Performance,” International Astronautical Federation, IAF-98-

U.2.05, 1998, 6. 

71 McCleskey, 7. 

72 McCleskey, 9. 

73 McCleskey, 7. 

74 McCleskey, 7. 

75 McCleskey, 6. 

76 Gonzales, xii. 

77 Gonzales, xii. 
Decision Maker’s Guide...115 

Notes 

78 Gonxales, 37. 

79 Gonzales, xii. 

80 Gonzales, 37. 81 Gonzales, 33. 82 S. Dornheim, “Follow-on Plan Key to X-33 Win,” Aviation Week 

and Space Technology, 8 July 1996, p.20. 

83 Gonzales, 42-44. 

84 Gonzales, 25. 85 Gonzales, 25. 86 Gonzales, 26. 87 Gonzales, 26. 88 Status of the X-33 Reusable Launch Vehicle Program, United

States Government Accounting Office (GAO/NSIAD-99-176), August 

1999, 2. 

89 Frank Mooring, “NASA Kills X-33, X-34, Trims Space Station,”

Aviation Week and Space Technology, 17 June 2001, n.p., on-line, 

Internet, 18 December 2002, available from http:// 

Aviationnow.com/avnow/content/publication/awst/20010305/nasa.htm. 90 Dennis Bushnell, Chief Scientist, Langley Research Center, email 

memo with Dr. Ron Sega, Director USD/DDR&E, “NASA Gen 

3/Airbreathing to Orbit Evaluation,” 20 March 2002. 

91 Matt Bille, Tony Williams, Vic Villhard, and Lisa Reed, “Dreams 

to Reality: The Path to the RLV,” AIAA, IEEE, and AAS Core 

Technologies for Space Systems Conference Proceedings, November 

2002. 

92 Including the X-20, X-30, X-33, X-37, Space Shuttle, Soviet, and

commercial RLVs. 

93 Including the X-15, U-2, SR-71, F-117, B-2, and Apollo programs. 94 Department of Defense Space Technology Guide FY2000-01, 

United States Department of Defense, Ch 4-11. 

95 Loomis, 220. 

96 The reader is encouraged to visit http://members.aol.com/luthben/ 

unit.html for a very helpful discussion on English to Metric conversion, 

particularly the relationship between pound-force, pound-mass, slug, and 

kilogram. 
116...Decision Maker’s Guide 

Notes 

97 William H Heiser and David T. Pratt, Hypersonic Airbreathing 

Propulsion (Washington D.C.: American Institute of Aeronautics and 

Astronautics, 1994), 128. 

98 Humble, Table 1.2, 11. 

99 “SSME Specifications,” Andrews Space and Technology, n.p., online, Internet, 20 January 2003, available from http://spacetech.com/

spacedata/engines/ssme_specs.shtml. 100 “RS-68 Engine,” Boeing, n.p., on-line, Internet, 20 January 2003, 

available from http://www.boeing.con/defense-space/space/propul/ 

RS68.html. 101 “RS-27A Propulsion System,” Boeing, n.p., on-line, Internet, 20

January 2003, available from http://www.boeing.com/defensespace/space/propul/delta.html. 

102 “2003 Aerospace Source Book,” 122. 260,800 lbs maximum 

thrust at seal level, 9,670 lbs dry weight; specific fuel consumption 0.334 

lb/hr/lb; Avweek 2003 Source Book, 122. 

103 “2003 Aerospace Source Book,” 132. 23,770 lbs maximum thrust 

at sea level; 3,265 lbs dry weight; specific fuel consumption of 2.1 

lb/hr/lb. 

104 Uwe Hueter and Charles McClinton, “NASA’s Advanced Space 

Transportation Hypersonic Program,” American Institute of Aeronautics 

and Astronautics, 11th AIAA/AAAF International Conference on Space 

Planes and Hypersonic Systems, (AIAA-2002-5175), September 2002, 2. 

105 Hueter, 2. 

106 “X-15 First Flight: X-15 Concept Evolution,” NASA, Figure 3, online, Internet, 20 January 2003, available from 

http://www.hq.nasa.gov/office/pao/History/x15conf/concept.html. The X15 had a launch weight of 33,000 pounds and a re-entry weight of 14,600 

pounds. A total of 308 pounds was allocated to avionics with another 300 

pounds allocated to the pilot and associated life support equipment. Hence 

an approximated payload mass of 600 pounds was used for this 

calculation. 

107 “2003 Aerospace Source Book,” 40. 

108 “2003 Aerospace Source Book,” 70. 

109 Humble, 66. 

110 Humble, figure C.1, 716. 
Decision Maker’s Guide...117 

Notes 

Ti Al 3

111 Humble, figure 2.11, 64. 

112 Whitehair, et al, “Operational Design Factors for Advanced Space 

Transportation Vehicles,” International Aeronautical Federation, IAF92-0879, 1992. 

113 A structure designed to successfully withstand the maximum

anticipated stresses experienced within its performance envelope possess a 

safety factor of 1.0. FAA Regulations Part 31 requires a minimum safety

factor 1.5, and higher safety factors by exception. 

114 Humble, Appendix B. 

115 Gonzales, 63. 116 Gonzales, 64. 117 Gonzales, 67. 118 Gonzales, xviii. 

119 Gonzales, xx. 120 “Titanium aluminides such as TiAl and developed for NASP 

offer the advantages of a high maximum operating temperature (800°C) 

compared to other titanium compounds (500-700°C), improved oxidation 

and creep resistance, and relatively low density to the aluminum content. 

They are however, presently more brittle than common titanium

compounds at room temperature. Nickel based alloys such as Inconel 617 

have significantly higher maximum operating temperatures (e.g., 1100°C 

for Inconel 617), allowing use on lower fuselage surfaces and other hightemperature areas. They are creep and oxidation resistant, but have 

significantly higher density than aluminum alloys.” Gonzales, 72-73. 

121 Gonzales, xx. 122 Gonzales, 67. 123 Gonzales, 69. 124 Low lift to drag (L/D) ratios at subsonic speeds and moderate to 

high L/D at supersonic speeds have stability concerns during landing. 

125 Jay P. Penn, SSTO vs. TSTO Design Considerations – An 

Assessment of the Overall Performance, Design Considerations,

Technologies, Costs, and Sensitivities of SSTO and TSTO Designs Using 

Modern Technologies, The Aerospace Corporation, Space Technology & 

Applications International Forum (STAIF-96), 7-11 January 1996, 

Albuquerque, NM. 
118...Decision Maker’s Guide 

Notes 

126 Gonzales, 42. 127 S. Dornheim, “Follow-on Plan Key to X-33 Win,” Aviation Week

& Space Technology, July 8, 1996, 20. 

128 “NASA Nears X-33 Pick,” Aviation Week and Space Technology, 

17 June 1996, 29. 

129 Gonzales, 48-49. 

130 Gonzales, xix. 

131 Rocket engine thrust is at a maximum when exhaust exit pressure 

is equal to ambient pressure, hence performance is maximized by a

variable exit nozzle configuration that permits exit pressure modulation 

during vehicle ascent. The linear-aerospike engine not only permits such 

modulation, it operates at much lower chamber pressures (the most 

important design parameter impacting rocket engine complexity). 

132 London, 100. 

133 Gonzales, 34. 134 Gonzales, 34. 135 London, 99. 

136 Claybaugh, 87. 

137 Claybaugh, 92-94. Two “arguably” private investments in space 

transportation have occurred: Pegasus and Athena. On the basis of 

return-on-investment, no purely private investment in new space

transportation assets has been successful. Based upon FY99$, the 

estimated investment and estimated ten year ROI for Pegasus and Arthena 

was $73.8M/(-39.1%) and $145.2M/(-61.9%). A ten-year treasury rate is 

used for discounting purposes, assumes 30% of launch revenue is returned 

(highly unlikely), and does not consider costs associated with correcting 

launch failures. 

138 Carol Loomis, “Mr Buffet on the Stock Market,” Fortune, 22 

November 1999, 218-220. Regarding the automobile: “All told there 

appear to be 2,000 car makes, in an industry that had an incredible impact 

upon people’s lives. If you had foreseen in the early days of cars how the 

industry would develop, you would have said, “here is the road to riches.” 

So what did we progress to by the 1990’s? After corporate carnage that 

never let up, we came down to three corporate car companies—themselves 

no lollapaloozas for investors. So here is an industry that had an 
Decision Maker’s Guide...119 

Notes 

enormous impact on America—and also an enormous impact, though not 

the anticipated one, on investors.” 

139 Loomis, 220. 

140 Claybaugh, 98. 

141 Claybaugh, 99. 

142 Claybaugh, 90. 

143 Walter J Boyne, “The Man Who Built the Missiles,” Air Force 

Magazine, October 2000, 84-86. 

144 Under Secretary of Defense for Acquisition and Technology, 

Space Launch Modernization Plan, Office of Science and Technology 

Policy, 1995. 

145 Data for delivery of propellants to NASA Kennedy Space Center 

in January 2000 included Kerosene (RP-1), Liquid Hydrogen, and Liquid 

Oxygen at $0.278, $1.300, and $0.064 per pound respectively. 

Claybaugh, 24. 

146 E. Zapata and Dr. A. Torres, “Space Transportation Cost Modeling 

and the Architectural Assessment Tool – Enhanced,” International 

Astronautical Federation, IAF-99-IAA1.1.01, 1999, 3. 

147 Claybaugh, 28 except as noted. Pegasus, Atlas IIAS, Delta II, and 

Titan IV: Isakowitz, 1991. 

148 “Space Transportation Costs: Trends in Price Per Pound to Orbit 

1990-2000,” Futron Corporation, Bethesda, MD, 6 September 2002, 2-3 

except as noted. Titan IV, Isakowitz, 1991, p. 268. 

149 Futron, 2-3 except as noted. Titan IV, Isakowitz, 1991, p. 268 

($154 Million in FY1990 inflated to $203 million in $FY00). 

150 Ted Nicholas and Rita Rossi, “U.S. Space Data Book,” Data 

Search Associates, Fountain Valley, CA, February 1996, page 11-11 

except as noted. All values inflated from FY 95 values. Delta II: 

$3440M; Atlas II: $6991M. Claybaugh, page 27 for Space Shuttle, X-15, 

X-33, and X-34. Also noted B-777 at $25,000/lb FY2002. 

151 Claybaugh, 29. 

152 Claybaugh, 18. 

153 London, 49. 

154 An F-16CJ Block 50 has the following approximate 

characteristics: 21,000 lb dry weight, 7,200 lb internal fuel, and a 37,500 

lb gross takeoff weight, implying a maximum useful “payload” of 9300 lb. 
120...Decision Maker’s Guide 

Notes 

155 The “sonic” trendline is independently verifiable in Air University

Air Force 2025 Study, Vol. 2 Ch 5, “Spacelift 2025: The Supporting 

Pillar,” 1996. 156 Claybaugh, 18. 

157 Based upon previously discussed space shuttle data, the lower 

labor intensity may be surprising. However, when one considers that the 

shuttle is a 4.5 million pound vehicle at launch, its lower labor intensity 

per pound can be mostly attributed to its enormous size (weight). 

158 Claybaugh, 18. 

159 Claybaugh, 26. 

160 Claybaugh, 27. 

161 Claybaugh, 6-31. More sophisticated cost estimating relationships 

applying linear regression techniques on data sets are commonly used in 

both government and industry. 

162 Claybaugh, 23. 

163 Claybaugh, 32-33. 

164 Claybaugh, 35. 

165 John C. Mankins, “Lower Costs for Highly Reusable Launch 

Vehicles,” Aeropsace America, March 1998, 36-37. 

166 C.M. McCleskey, “Identifying STS Cost and Cycle Time Design

Root Causes,” SLI Architecture Working Group Meeting, NASA John F.

Kennedy Space Center, Florida, 21 August 2002, Package 1, 24. 

167 McCleskey, Package 1, 23. 

168 McCleskey, Package 1, 17. A 7-8 flight per year launch rate

between 1996 and 1997 was demonstrated. The orbiter landed at Kennedy 

Space Center on 26 September 1996. Orbital Processor Facility 

turnaround occurred between 26 September and 5 December 1996. 

Vehicle Assembly and Integration began on 5 December 1996 culminating 

in a 12 January 1997 launch. 

169 McCleskey, Package 1, 12. 

170 London, 1. 

171 G.W. Elverum, Jr., “Boosters” (Transcript of a talk presented at

the Aerospace Productivity Conference), The Aerospace Corporation, El

Segundo, CA, 1987. 

172 London, 14. 

173 Gonzales, xxii. 
Decision Maker’s Guide...121 

Notes 

174 London, 100-101. 

175 HEDM: DARPA has funded HEDM research at AFRL for more 

than 20 years. AFRL has successfully synthesized the octanitrocubane

and the N5+ cation. Octanitrocubane is “hammer” stable and 30% more 

energetic than HMX. Availability of N5+ cation implies potential 

synthesis of N8. N8 can potentially be designed into a cubane structure 

that may offer a very high Isp and high energy density propellant. 

176 James R. Wertz, “Economic Model of Reusable vs. Expendable 

Launch Vehicles,” International Astronautics Federation Congress, Rio 

de Janeiro, Brazil, 2-6 October 2000, 11. 

177 Wertz, “Economic Model of Reusable vs. Expendable Launch 

Vehicles,” 14. 178 Wertz, “Economic Model of Reusable vs. Expendable Launch 

Vehicles,” 11. 179 Dennis Bushnell, Chief Scientist, NASA Langley Research Center,

private email communication, “Questions on RLVs,” 21 November 2002.

180 Hueter, 2. 

181 Steve Cook, “Next Generation Launch Technology Program,”

Presentation to the USAF Research Laboratory, Wright-Patterson AFB, 

Ohio, 18 December 2002, 28. 

182 Cook, 28. 

183 Dennis Bushnell, Chief Scientist, NASA Langley Research Center,

email to Dr. Paul Dimotakis, California Institute of Technology, 

“Revolutionary Rocket Technology,” 16 July 2002. 

184 Dennis Bushnell, Chief Scientist, Langley Research Center, email 

memo with Dr. Ron Sega, Director, OUSD/DDR&E, “NASA Gen 

3/Airbreathing to Orbit Evaluation,” 20 March 2002. 

185 Bushnell, “NASA Gen 3/Airbreathing to Orbit Evaluation.” 

186 “Space Shuttle Main Engine,” Boeing, n.p., on-line, Internet, 20

January 2003, available from http://www.boeing.com/defensespace/space/propul/SSME.html. The maximum equivalent 

horsepower developed by the Space Shuttle Main Engine is just over 

37 million horsepower, and the energy release from a single engine is

equivalent to the output of nearly eight Hoover Dams. Although not 

much larger than an automobile engine, the SSME high-pressure fuel 

turbopump generates 100 horsepower for each pound of its weight,
122...Decision Maker’s Guide 

Notes 

while an automobile engine generates about one-half horsepower for

each pound of its weight. 

187 “RLV Propulsion,” Boeing Rocketdyne presentation to Col 

Gardner, U.S. Air Force Space and Missile Center, 5 April 2002, 41. 

188 “RLV Propulsion,” 22-23. 

189 “RLV Propulsion,” 13. 

190 Ramon Chase and Ming Tang, “The Quest for Single-Stage-toOrbit: TAV, NASP, DC-X and X-33 Accomplishments, Deficiencies, and 

Why The Did Not Fly,” 11th AIAA/AAAF International Conference on 

Space Planes and Hypersonic Systems Technologies, (AIAA-2002-5143),

September 2002, 3. 

191 Gonzales, xi. 192 Gonzales, xxi. 

193 Gonzales, 51. 194 Gonzales, 22. 195 Gonzales, xviii. 

196 Cook, 26. 

197 Jack L. Kerrebrock, Hypersonic Technology for Military

Application, (Washington D.C.: National Academy Press, 1989), 11. 

198 Kerrebrock, 11. 

199 Kerrebrock, 11-12. Note that molecular oxygen and nitrogen 

dissociate at 2000 and 4000 degrees Fahrenheit respectively. Nitric Oxide 

(NO) forms at 4000 degrees Fahrenheit, and both oxygen and nitrogen 

ionize at 9000 degrees Fahrenheit. 

200 Stagnation temperature is the total temperature reached by high

velocity air that is slowed to zero velocity (stagnation). The kinetic

energy of the flow is being converted into heat, manifesting itself as

higher temperatures. 

201 A. Richard Seebass, et al., Review and Evaluation of the Air Force 

Hypersonic Technology Program, (Washington D.C.: National Academy 

Press, 1998), 6. 

202 Seebass, 1.

203 The temperature of chemical combustion no longer exceeds the 

temperature of the combustion medium. 204 Cook, 29. 
Decision Maker’s Guide...123 

Notes 

205 “The Heart of the SR-71 Blackbird, the J-58 Engine,” n.p., on-line, 

Internet, 20 January 2003, available from 

http://aerostories.free.fr/technique/J58/J58_01/page8.html. 206 Craig Covault, “Hypersonics Strategy Sets Stage for ‘Next Great 

Step’,” Aviation Week and Space Technology, 26 March 2001. 

207 Cook, 31-36. 

208 Gray Creech, “X-43 Soars on Scramjet Power,” n.p., on-line, 

Internet, 27 March 2004, available from 

http://www.nasa.gov/missions/research/x43_soars_feature.html. 209 Couvalt, “Hypersonics Strategy Sets Stage for ‘Next Great Step.’” 

210 Cook, 35. 

211 Cook, 33. 

212 Cook, 33. 

213 Covault, “Hypersonics Strategy Sets Stage for ‘Next Great Step.’” 

214 HyTech funding between FY 96 to FY 01 has never exceeded $20 

million annually. A. Richard Seebass, et al., Review and Evaluation of the

Air Force Hypersonic Technology Program, (Washington D.C.: National 

Academy Press, 1998), 49. 

215 Stanley W. Kandebo, “Landmark Tests Boost Scramjet’s Future,” 

Aviation Week and Space Technology, 26 March 2001, 58-59. Editor’s 

note: As of January 2004, Pratt & Whitney successfully completed 

ground testing of GDE-1 had been selected for a follow-on effort that is 

scheduled to result in flight tests in the 2007-2008 timeframe. See, “Air 

Force Selects Pratt & Whitney and Boeing Team For the Scramjet Flight 

Demonstrator Program,” on-line, Internet, 14 May 2004, available from 

http://www.pratt-whitney.com/pr_011304.asp. 

216 Cook, 33. 

217 Within a dual combustion ramjet, supersonic air is ingested into 

one inlet and slowed to subsonic speeds, mixed with a conventional 

hydrocarbon fuel in a fuel-rich environment, and ignited as in a ramjet. To 

break through the ramjet’s operating speed limitations, the expanding 

combustion products are then mixed with supersonic air entering through a 

second inlet and are more completely burned in a supersonic combustor. 

The dual combustion ramjet has an operating threshold of Mach 3 and a

maximum operating speed of Mach 6. 
124...Decision Maker’s Guide 

Notes 

218 Stanley W. Kandebo, “New Powerplant Key to Missile

Demonstrator,” Aviation Week and Space Technology, 2 September 2002, 

56-58. 

219 Couvalt, “Hypersonics Strategy Sets Stage for ‘Next Great Step.’” 

220 “Hypersonic Waveriders,” Summary, 1, n.p., on-line, Internet, 18 

Jan 2003, available from http://www.aerospaceweb.org/design/waverider/ 

main.shtml. 221 “Hypersonic Waveriders,” Examples, 1. 

222 “Hypersonic Waveriders,” Summary, 1-2. 

223 “Hypersonic Waveriders,” Optimized Waveriders, 2. 

224 “Hypersonic Waveriders,” Optimized Waveriders, 5-6. 

225 “Flexible Aerospace System Solution for Transformation,” Boeing, 

n.p., on-line, Internet, 8 February 2002, available from 

http://www.boeing.com/phantom/fasst.html. 

226 Marty K. Bradley, Kevin G. Bowcutt, et.al., “Revolutionary 

Turbine Accelerator (RTA) Two-Stage-To-Orbit (TSTO) Vehicle Study,” 

American Institute of Aeronautics and Astronautics, 38th Joint Propulsion

Conference (AIAA-2002-3902), July 2002, 2. 

227 “Flexible Aerospace System Solution for Transformation,” 1. 

228 Kevin G. Bowcutt, Mark Gonda, et al., “Performance, Operational 

and Economic Drivers of Reusable Launch Vehicles,” American Institute 

of Aeronautics and Astronautics, 38th Joint Propulsion Conference (AIAA2002-3901), July 2002, 6. 

229 “Flexible Aerospace System Solution for Transformation,” 1. 

230 The 1994 Commercial Space Transportation Study established

Mach 4 capability as the threshold for viable worldwide package delivery 

operation. 

231 Bradley, 9. 

232 “Flexible Aerospace System Solution for Transformation,” 1. 

233 Mankins, 37. 

234 Mankins, 39-41. 

235 J. R. Olds, P.X. Bellini, “Argus, a Highly Reusable SSTO RocketBased Combined Cycle Launch Vehicle with Maglifter Launch Assist,” 

American Institute of Aeronautics and Astronautics, 8th International 

Space Planes and Hypersonic Systems and Technologies Conference 

(AIAA-98-1557), April 1998, 3. 
Decision Maker’s Guide...125 

Notes 

236 Olds, 4-6. 

237 Olds, 5. 

238 J.A. Moody, A.T. Bahill, et al., Metrics and Case Studies for 

Evaluating Engineering Designs, (New York: Prentice Hall Publishers,

1997). 

239 “2003 Aerospace Source Book,” 155-160. 

240 Gonzales, 78-79.