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Small Unit Space Transport And Insertion (SUSTAIN): How to Do It

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7th Responsive Space Conference
RS7-2009-1002
Small Unit Space Transport
And Insertion (SUSTAIN): How
to Do It and Use It as a Driver
for Low-Cost Responsive
Orbital Launch
John M. Jurist
Odegard School of Aerospace Sciences
David C. Hook
Planehook Aviation Services, LLC
David Livingston
Odegard School of Aerospace Sciences
7th Responsive Space Conference
April 27–30, 2009
Los Angeles, CA
RS7-2009-1002
Small Unit Space Transport and Insertion (SUSTAIN): How to Do It and Use It as a
Driver for Low-Cost Responsive Orbital Launch
John M. Jurist
Odegard School of Aerospace Sciences, University of North Dakota
1540 Lake Elmo Drive, Suite 5, Billings, Montana 59105-1798; (406) 245-5704
jmjurist@gmail.com
David C. Hook
Planehook Aviation Services, LLC
600 Windhaven Drive, San Antonio, Texas 78239-2133; (210) 653-8842
David Livingston
Odegard School of Aerospace Sciences, University of North Dakota
P. O. Box 95, Tiburon, California 94920; (415) 435-6018
dlivings@davidlivingston.com
ABSTRACT
Cheap, rapid orbital launch (Responsive Space) has been elusive. Three potential approaches, all with different
policy and economic implications, are considered.
The first, exemplified by Virgin Galactic’s SS-2, evolves current attempts at suborbital space tourism or humantended science involving brief flights with apogees above 100 km into point to point suborbital transport and
eventually into orbital transport.
The second, exemplified by SpaceX’s Falcon series, evolves more traditional aerospace technology with improved
management. This is intended to drive down launch costs on the margin and evolve eventually to the point where
responsive and inexpensive space access is accomplished.
The third, and in our opinion most viable short term approach, uses stated national security needs to accelerate
development and to exploit currently existing or partially developed and demonstrated technology.
We consider the potential synergism between the recently released US Air Force request for information on placing
a small unmanned aerial vehicle (UAV) anywhere in the world from the continental US within 2 hours and the
Small Unit Space Transport and Insertion (SUSTAIN) requirement defined by the US Marine Corps. SUSTAIN
specifies a needed capability to place a squad of 13 marines and field supplies anywhere in the world from the
continental US within 2 hours.
Potential solutions considered for SUSTAIN and rejected include:
•
•
•
A DC-X like vertical take-off rocket-powered vehicle that decelerates and lands under rocket power and
then returns under rocket power without refueling and refurbishing cannot be developed and fielded in a 5
to 10 year period.
An aerospace plane that would most likely require development for more than a decade and would also
require a landing field near the target area.
Placing and staffing a constellation of up to 12 space stations with re-entry vehicles. This is technically
possible but economically implausible.
Each requirement (the USAF UAV anywhere and SUSTAIN) can be met with launch vehicles from the Microcosm
Scorpius launch system family, specifically the Sprite and Exodus vehicles. These vehicles can inexpensively
Jurist
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AIAA/7th Responsive SpaceВ® Conference 2009
satisfy both programmatic launch requirements within the 5-10 year time frame using current technology and are
pressure-fed, liquid-fuelled, ablatively-cooled, composite 3 stage vertical take-off rocket-powered launchers.
For SUSTAIN, a man-carrying capsule decelerates aerodynamically during re-entry, decelerates further with a
parachute or parasail, and cushions the final impact with small solid-fuelled rockets. Extraction of individual team
members could be accomplished by using Fulton Recovery Systems on them individually or by lifting the capsule
containing the team to several thousand feet AGL with the capsule abort rocket system and then snagging it in
midair with a cargo aircraft.
The basic technology for this approach has been demonstrated over the past ВЅ century. Most of the technology
elements for the Scorpius Sprite and Exodus have been demonstrated and even flown. Therefore, SUSTAIN could
be implemented rapidly and inexpensively. The major remaining developmental element is the capsule.
Major impediments to implementing SUSTAIN fall within the political, economic, and policy arenas. A side benefit
of combining our approach to SUSTAIN with the USAF Responsive Space concept is an end result of a simple,
cheap, responsive space launch vehicle series with remaining development costs spread between several programs.
KEYWORDS
of that technology. The far term solutions could
potentially exploit better materials, exotic propulsion
schemes, etc.
Responsive Space, Launch Vehicles, Scorpius,
Microcosm, SUSTAIN, USMC, USAF, Fulton
Recovery System, FRS, Vertical Takeoff and Landing,
VTOL, Delta-V, Hypersonic, Saanger, Space Shuttle,
STS, C-130, C-17, C-5, Parasail, Recovery Aircraft,
Specific Impulse, Exodus, Sprite, Wake Turbulence,
Double-Ply Parachute, Gas Bag, Political, Policy and
Economic Issues, Space Pearl Harbor, Rumsfield Space
Commission Report, Space Weapons Systems, Outer
Space Treaty, OST, SpaceX, Falcon, Virgin Galactic,
SS-2, Suborbital Space Tourism, Rapid Orbital Launch,
DC-X, Unmanned Aerial Vehicle, UAV
OVERVIEW
In a hypothetical future, the United States may lack the
current network of foreign bases. The ability to project
payloads and small force units rapidly to any part of the
world is essential to US interests.
A low earth orbit (LEO) requires approximately 90
minutes. Allowing 15 minutes for orbital insertion, and
15 minutes for de-orbit, re-entry, and landing therefore
places any location on earth within about 75 minutes of
travel time from any other location. Current proven
aviation technology is limited to several thousand miles
per hour (MPH). Yet, flying the worst case distance of
about 12,500 miles in no more than two hours requires
an average speed of at least 6,250 MPH. While
technically possible, the budgets and effort required to
develop, create, test, and field an aircraft to fulfill the
requirement almost certainly fall outside the 5 to 10
year range we assume for the near term. The use of,
and travel through, space would be required to solve
this problem in the near term.
REQUIREMENTS
The USAF requirement is to insert a UAV into any area
in the world within 2 hours. The USMC requirement
for SUSTAIN is to insert a Marine squad of 13 riflemen
and field supplies into any potentially hostile area in the
world within 2 hours and retrieve them.
SPECIFICATIONS
The UAV payload for the USAF need as described in a
Request for Information (RFI) due March 15, 2009, is
assumed to mass roughly 450 pounds delivered over the
target within two hours of flight time.
The US lacks both the basic knowledge and the
technology for developing a vehicle capable of
continuous acceleration of more than one gravity for
prolonged periods of time. That approach is therefore
relegated to the very far term if ever. However, recent
developments make near term solutions for the
specified problems possible.
The SUSTAIN specification would include life support
for up to 4 hours, masses of 220 pounds per man plus
150 pounds of supplies times 13 men in the squad. This
implies a deliverable payload of 4,810 pounds plus the
life support and retrieval systems.
Near term requirements for both programs are assumed
to be in the 5 to 10 year range with the far term being
anything thereafter. The near term solutions are
dependent on existing technology or slight extensions
Jurist
Without large research and development budgets, most
proposed approaches are not feasible. Several are
feasible but not practical with current technology, and
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AIAA/7th Responsive SpaceВ® Conference 2009
beyond any near term capability for a single vehicle and
would require enormous development effort and cost
for a technological gamble.
development timetables would exceed one decade.
Others are not practical with current technology
because of the high costs of implementation.
Technological risks also increase with highly elegant
approaches to meeting the stated requirements.
The approach to the target area and landing would be
extremely noisy for the period of the rocket landing
burn and could potentially subject the vehicle to ground
fire during a vulnerable portion of its flight. The
vehicle itself would not be capable of carrying any
meaningful armor with current materials and would be
vulnerable to combat damage during the squad insertion
process.
ALTERNATIVE (REJECTED) APPROACHES
Several of the impractical approaches to SUSTAIN are
worth further discussion. The SUSTAIN requirement is
more rigorous than the USAF UAV placement
requirement, so it will be discussed in more detail.
Meeting the UAV placement requirement can be
accomplished as part of the launch vehicle development
path for SUSTAIN.
Current practice using helicopters involves approaching
the landing zone, disgorging the troops, and flying back
to a secure area (a forward base) if needed to await the
retrieval process. During retrieval, the helicopter flies
back to the landing zone, retrieves the troops and flies
away. This process can use local tactical air support to
suppress resistance if needed. Unfortunately, this
approach is impossible with a DC-X type vehicle
because the already high delta-V requirement increases
enormously.
Vertical Take-Off and Landing
This approach would develop a vertical take-off and
vertical landing rocket vehicle similar to the old DC-X
vehicle. The vehicle would be launched from the
continental US, fly an appropriate suborbital or
fractional orbital trajectory, re-enter the atmosphere,
and then land vertically under rocket power. Retrieval
of the squad would reverse the process.
Aerospace Plane
Half way around the world is approximately 12,500
miles. Flying this worst case distance in no more than
two hours requires an average speed of at least 6,260
MPH. The overall energy requirements of a supersonic
or hypersonic aircraft flying at these speeds through the
atmosphere are enormous. Such an aircraft does not
exist, and developing it would be very expensive in
both time and money.
Examples of current alt.space development of this
technology include Blue Origin (Jeff Bezos) and
Armadillo Aerospace (John Carmack)1,2.
This approach, while interesting for the future, has
several fatal flaws for SUSTAIN.
The DC-X concept allows vertical landing, but it was a
test vehicle only. Landing on rockets burns lots of fuel
and drives up the mission delta-V and thus the required
mass ratio markedly. Such a vehicle would require up
to twice the one-way mission delta-V in order to
complete the mission without refueling at the target
zone. As will be shown later, the launch delta-V for a
suborbital ballistic trajectory half way around the world
is the same as for launch to LEO. Reducing the
percentage of the Earth’s circumference between the
launch and landing point does not proportionally reduce
the launch delta-V.
If the postulated hypersonic transport vehicle flies
outside the atmosphere for part of its mission, it is
basically a winged orbital or suborbital manned space
vehicle – a space plane.
A space plane concept similar to the old Saanger
vehicle will not allow vertical insertion into an
unprepared area. Landing in or near the target area
would require an airfield which may or may not be
available for use. Such a vehicle could be used to
transport troops to a target area for an airborne assault
as it flies past the landing zone, but retrieval would then
require additional aircraft.
The one-way delta-V would be increased considerably
(potentially as much as doubled) by the fuel
requirements for approach and landing of the vehicle
over that required for launch and insertion into the
appropriate trajectory.
Space plane concepts have been proposed multiple
times and have never worked to any extent other than
with the current Space Transport System (STS or Space
Shuttle), which would not be able to initiate extraction
and return after landing in the absence of extensive
infrastructure.
The total mission insertion and retrieval delta-V,
therefore, could be as much as four times that required
for a launch into low circular Earth orbit. This is well
Jurist
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AIAA/7th Responsive SpaceВ® Conference 2009
There is an inverse relationship between the required
number of bases and the cross-range capability of the
re-entry vehicles in order to provide world-wide
coverage. For example, if two hour polar circular orbits
are assumed, 12 bases spaced at polar-intersecting plane
inclination intervals of 30 degrees would allow at least
one base to approach any spot in the world within two
hours. Re-entry vehicle cross range capability of one
twelfth of the Earth’s circumference, or approximately
2,100 statute miles, would provide the requisite
coverage.
The payload requirements for SUSTAIN, as will be
shown later, are perhaps 15 percent of the STS payload
capacity. Developing an entirely new, smaller vehicle
along the STS model would be very expensive and
would require at least a decade to develop and test.
Evolution of current winged suborbital space vehicle
concepts into something that is satisfactory for the
SUSTAIN requirement would take a lot of time and
funding.
Commercial suborbital space planes under development
might climb vertically to a motor burnout speed of
perhaps 3,700 feet per second (ft/sec), or roughly 1,130
meters per second (m/sec) at an altitude of about
170,000 feet, or about 52 kilometers (km) before
coasting to an apogee above 100 km. Given that the
delta-V for a ballistic trajectory half way around the
world is essentially the same as for launching into LEO
(7,905 m/sec plus allowances for drag, gravity, and
other losses), currently conceived crewed suborbital
vehicles are inadequate because they fall far short of the
required delta-V capability and lack the required
thermal protection systems for atmospheric re-entry at
large fractions of orbital speeds. Examination of the
speed versus range relationship in Figure 1 shows that
current suborbital space plane concepts provide
negligible range, expressed in degrees of Earth
circumference, as point to point transportation systems.
This approach would cost many billions or even
trillions of dollars and is not politically feasible in the
current economic climate.
Rocket-Launched Capsule and Aerial Recovery
The second technically feasible approach is to use a
rocket to boost a man-carrying capsule either into an
LEO trajectory followed by fractional orbit and
subsequent re-entry or to boost the capsule into a
ballistic trajectory with re-entry over the target zone.
This approach could use cheap expendable vehicles
rather than expensive reusable vehicles that would
require extensive maintenance, would be vulnerable to
opposition near the landing zone, etc.
As in the old Mercury, Gemini, and Apollo projects, the
capsule would re-enter, slow by parachute (or modern
parasail), and then land after an approximately vertical
approach. Adding mass for thermal protection for
ballistic re-entry with aerodynamic braking is much less
massive than the mass penalty of capability for a
rocket-decelerated re-entry with minimal aerodynamic
loading.
Suborbital Ranges
Delta-V (vs circular
100%
80%
60%
40%
The former Soviet Union developed the capability to
return their capsules to land without injuring the
occupants by employing small retrorockets firing just
before impact, reducing the capsule sink rate under
parachute support down to something sustainable by the
crew. This capability could be incorporated into the
SUSTAIN capsule.
20%
0%
0
45
90
135
180
Range (deg)
Figure 1. Launch Delta-V versus Range.
TECHNICALLY FEASIBLE APPROACHES
Retrieval could be accomplished in one of two ways.
Current and near term rocket technology suggests
several approaches to meeting the SUSTAIN
requirement.
First, the squad could re-enter the capsule, fire the abort
and escape rocket on the capsule to loft it to several
thousand feet above ground level, and have it snagged
by an approaching cargo aircraft. At this altitude, the
aircraft would be relatively resistant to small arms fire
from the ground. This technology was demonstrated
operationally more than 40 years ago with air retrieval
of film canisters from reconnaissance satellites. It was
also demonstrated in retrieving downed airmen and
Space Stations
The first technically feasible approach is to launch and
maintain a fleet of orbiting bases (space stations) in
relatively low polar orbits, each of which contains the
requisite troops and a re-entry vehicle.
Jurist
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AIAA/7th Responsive SpaceВ® Conference 2009
even grounded gliders in the late 1940s and early
1950s.
FIRST APPROXIMATION
The assumption of a 450 pound UAV to be delivered
implies a total payload after re-entry of perhaps 1,000
pounds. The basic SUSTAIN payload to be delivered
by soft-landing is perhaps 220 pounds per rifleman plus
perhaps 150 pounds in supplies each. With a life
support mass budget of 100 pounds per person, and
reasonable estimates for a capsule, heat shield, abort
system, extraction system, etc., the total to be softlanded is roughly 15,000 pounds.
Second, the individual squad members could be
retrieved individually by the same approach (Fulton
Recovery System or FRS)3.
The rocket-launched capsule and recovery approach to
SUSTAIN will be explored in detail in this paper.
MULTIPLE STAGE BALLISTIC VEHICLE
The fundamental problem with a ballistic vehicle is the
mission delta-V, or specified speed change. In order to
travel half way around the earth, the delta-V is
essentially equal to the circular velocity at sea level
(7,905 m/sec, 25,936 ft/sec, or 17,684 MPH) ignoring
aerodynamic losses (drag), gravity losses, and terminal
losses during final approach and landing. All of these
components add to the delta-V requirement. Gravity
and drag losses may add 10 to 20 percent to the delta-V
needed for attaining LEO.
Taking the basic delta-V for a 180 degree range flight
and adding 15 percent (1,186 m/sec) for gravity,
aerodynamic, and terminal maneuvering losses, 9,091
m/sec (29,828 ft/sec) is needed for either launch
mission. The requirements of 1,000 pounds re-entered
over target or 15,000 pounds soft landed on target can
be met by a partly-developed launch vehicle family.
Having a short response time implies the use of LOX
and hydrocarbon (RP-1 or kerosene) propellant.
Current experience with LH2 suggests that long term
storage of a viable vehicle for responsive launch is a
difficult problem. Present technology supports using a
multistage RP-1 and LOX propelled vehicle if the
purely ballistic approach is desired.
Considering a reduced range does not relax the delta-V
requirement much, as was shown in Figure 1. An
impulsive launch half way around the earth (180
degrees) requires essentially 100 percent of circular
velocity. An impulsive launch one quarter of the way
around the earth (90 degrees) reduces this to only 90
percent of circular velocity. A range of 45 degrees
brings it down to about three fourths of circular
velocity. Figure 1 shows the impulsive launch delta-V
as a fraction of circular velocity for various ranges
expressed in degrees of Earth’s circumference.
Reliability for a responsive launch also requires getting
the parts count down to an absolute minimum. The
desired characteristics of such a vehicle include a low
aspect ratio to reduce wind shear sensitivity during
launch, low cost if it is not recoverable and reusable,
and simple to reduce storage and maintenance costs.
Modularity with commonality of parts should be
attempted to further reduce costs and complexity. As
will be shown later, much of the preliminary
development of a usable vehicle has been accomplished
by Microcosm, Inc.
The conclusion to be gained here is that reducing the
range does not yield a proportional benefit in reduction
of the delta-V requirement. Therefore, equipping two
bases in the United States for SUSTAIN operations –
one on each coast – might reduce the maximum range
for world-wide coverage by about 1,500 miles, or from
180 degrees to about 158 degrees, but has negligible
effect on the mission delta-V requirement.
THE SUSTAIN CONCEPT
The technically feasible approach is to use a three stage
expendable rocket to boost a man-carrying capsule
either into an LEO trajectory followed by a fractional
orbit and subsequent re-entry (preferred) or to boost the
capsule into a ballistic trajectory with re-entry over the
target zone.
Although a lower mission delta-V might be needed for
a shorter range mission, meeting the requirement for
insertion anywhere in the world requires that the launch
system be capable of launching the payload into low
earth orbit. A lower range ballistic trajectory for
shorter range missions has an additional disadvantage –
that of re-entry decelerations.
Reducing these
decelerations to tolerable levels for both UAV insertion
and for SUSTAIN can be accomplished if a fractional
orbital trajectory is used to reduce re-entry angles and
thus re-entry decelerations. This approach could still
meet the 2 hour flight time requirement.
Jurist
The capsule would re-enter, slow aerodynamically to
several hundred ft/sec, decelerate further with a
parachute or modern parasail to approximately 30
ft/sec, and then land after an approximately vertical
approach. The capsule could employ small retrorockets
firing just before impact to reduce the capsule sink rate
down to effectively zero at “impact.”
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AIAA/7th Responsive SpaceВ® Conference 2009
Retrieval can be accomplished by the squad re-entering
the capsule, firing the abort and escape rockets on the
capsule to loft it to several thousand feet above ground
level (AGL), and snagging it with an approaching
aircraft – perhaps a C-130 – as were the old
reconnaissance satellite film canisters. At this altitude,
the aircraft would be relatively immune to small arms
fire from the ground. Alternative retrieval could snag
individual team members with FRS by a cargo aircraft.
of 484 ft/sec and an apogee height of 2,970 feet. This
abort rocket system, designed for a thrust of 85,860
pounds and a 3.21 second burn at a propellant
consumption rate of 373 pounds per second (specific
impulse of 230 seconds) would result in about 6
gravities of perceived axial acceleration upwards
(eyeballs in) for the capsule occupants. Immediately
after the abort system rocket burnout, the perceived
axial acceleration would be about -1.5 gravities
(eyeballs out) as a result of air drag.
SUSTAIN CAPSULE
Large Diameter Capsule
One of several feasible scenarios would use a capsule
configured as a truncated cone with a 60 degree apex
angle and a base outside diameter of 16 feet. The low
subsonic drag coefficient would be approximately 0.4.
Capsule contents would include 13 Marine riflemen at
an average of 220 pounds each with 75 pounds of
consumable supplies (ammunition and ordinance) and
75 pounds of returnable supplies (weapons,
communications gear, body armor, etc.) each. The
capsule life support system, sustainable for four hours,
would weigh 100 pounds per person or 1,300 pounds
total and would provide mixed oxygen – nitrogen
atmosphere at the target zone ambient pressure.
The capsule occupants, assumed to average at the 90th
percentile US male size, would be placed supine in
three stacked rows of 5, 5, and 3 or 6, 5, and 2 abreast
from bottom to top.
The capsule weight budget is shown in Table 1.
Table 1. Capsule Weight Budget (Pounds).
The parachute or parasail system, designed for a sea
level sink rate of about 33 ft/sec, would be deployed at
apogee. With a drag coefficient of 1.4, a parachute
canopy would be about 100 feet in diameter to yield the
design sink rate of about 33 ft/sec at sea level. Such a
parachute might weigh 150 pounds or so. It is
potentially possible to include a reserve parachute,
given the budgeted weight of the parachute system in
Table 1. The landing retrorockets would be fired about
8 feet off the ground to reduce the sink speed by about
33 ft/sec with a one-half second burn of 196 pounds per
second of propellant, for a total thrust of 45,180
pounds. Again, the design specific impulse is 230
seconds. The maximum perceived axial acceleration
upwards would be three gravities.
After a nominal flight and atmospheric re-entry, the
capsule heat shield could be optionally jettisoned over
the landing zone and the parachute or parasail would be
deployed. Landing would be as in the abort scenario
described above.
After landing, the parachute or parasail would be
jettisoned, the hatch opened, and the squad deployed.
Two subsidiary scenarios can be considered for squad
retrieval.
Liftoff
Basic capsule
Heatshield assembly
Abort structure
Abort propellant
Retrorocket structure
Retrorocket propellant
Parachute assembly
13 Marines
Consumable supplies
Returnable supplies
Life support system
Totals
Liftoff
abort
5,000
Landing
retro
Retrieve
liftoff
5,000
5,000
750
1,500
1,200
100
100
300
2,860
975
975
1,300
1,500
1,200
100
100
300
2,860
975
975
1,300
5,000
750
1,500
1,200
100
100
300
2,860
975
975
1,300
15,060
14,310
15,060
1,500
1,200
100
2,860
975
1,300
12,935
In the event of a launch abort, the capsule heat shield
would be jettisoned to reduce weight and the abort
rocket system fired. For a pad abort, this system would
be capable of lofting the capsule to a maximum speed
Jurist
The first is to destroy the capsule and retrieve the
individual squad members using the FRS and an
approaching aircraft such as the C-130.
The second subsidiary scenario is, upon ground mission
completion, to have the squad members dump their
consumable supplies, re-enter the capsule, close the
hatch, and fire the launch abort system. Since the
capsule is considerably lighter than at launch, the abort
system accelerates the capsule to 534 ft/sec and a
nominal apogee of 3,240 feet. The perceived upward
acceleration is initially 6.6 gravities and a maximum of
about 6.7 gravities (eyeballs in).
The upward
acceleration does not increase further as the abort
system propellant burns off because of progressively
increasing air drag. Immediately after rocket burnout,
air drag induces a perceived downward acceleration of
about -2.3 gravities (eyeballs out).
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AIAA/7th Responsive SpaceВ® Conference 2009
Candidate aircraft for either squad or capsule recovery
include the C-130, C-17, or C-5.
Recovery would be accomplished by snagging the
capsule with an appropriately sized cargo plane (C-17
or C-5) as it approaches apogee. This process is
provisionally described in a later section of this report.
The cargo bay dimensions for the three candidate
aircraft are (length by height by width):
The latter secondary scenario provides several
advantages.
First, the FRS-type concept uses a
parachute snagging system that has been field tested
and also would allow for a second full-size parachute to
be incorporated into the capsule if capsule-based
recovery is used. This concept would allow for a
reasonable survival option for the capsule-riding
Marines should all attempts at snagging them fail.
Second, having a second full-size parachute allows for
a reserve chute upon re-entry or after an on-pad abort
,should the first one fail. This increases the probability
of squad survival and recovery from a plausible failure
mode.
•
•
•
The larger capsule would not fit inside a C-130, but
would fit the larger aircraft. The smaller diameter
capsule would fit in all three aircraft. All three aircraft
have sufficient lift capacity to handle the capsule.
Aircraft characteristics as they affect recovery, safety,
and mission security issues are discussed later.
Cruise speeds for the three aircraft are 318 knots for the
C-130 and 450 knots for the C-17 and C-5. The
estimated downrange distances are roughly 1,520,
2,180, and 2,140 nautical miles 5 hours after takeoff for
the C-130, C-17, and C-5, respectively.
Small Diameter Capsule
A second scenario is to configure the capsule as a
cylinder approximately 8 feet in outside diameter by
about 18 feet long and topped by a truncated cone with
an apex angle of 60 degrees. The squad would be
stacked in four rows of three abreast with the remaining
squad member in the fifth (top) row. The capsule
weight budget would be essentially the same as for the
previously described 16 foot diameter conical capsule.
Table 2. Capsule Configuration.
Configuration
The abort, re-entry and landing, and squad recovery
scenarios for the cylindrical capsule would be the same,
except the drag coefficient would be a bit lower than for
the conical capsule (assumed to be 0.35 instead of 0.4).
The overall aerodynamic drag at any airspeed would be
reduced because of the lower cross sectional area of the
capsule and the lower coefficient of drag. As a result,
the abort burnout speed and apogee would be increased.
The re-entry profile would be changed, and the terminal
descent speed at parachute deployment for landing
would be increased. If the capsule abort rocket is used
for recovery of the capsule, the burnout speed and
apogee would be increased, but the smaller diameter
would allow the snag to be performed with a smaller
and more suitable cargo aircraft (the C-130).
Truncated
Cone
Cylinder plus
Truncated Cone
Outside diameter
16
Cone apex angle
60
Diameter at truncation
3
Height
11.3
External volume
922.5
Assumed drag coefficient 0.40
8
60
3
22.3
1,014.7
0.35
feet
degrees
feet
feet
cubic feet
Abort
apogee
burnout speed
initial acceleration
peak acceleration
initial deceleration
4,493
530
6.0
6.2
-0.4
feet
feet/second
gravities
gravities
gravities
989
915
feet/second
feet/second
2,970
484
6.0
6.0
-1.5
Landing chute deployment speed
at10,000 feet
462
at 5,000 feet
428
Recovery
apogee
3,240
burnout speed
534
initial acceleration
6.6
peak acceleration
6.7
initial deceleration
-2.3
As an alternative, the smaller diameter capsule could be
equipped with wings to allow some cross-range
capability. Detailed discussion of this alternative is
beyond the scope of this paper.
5,303
feet
596
feet/second
6.6 gravities
6.8 gravities
-0.6 gravities
LAUNCH VEHICLE
A conceptual vehicle capable of lofting roughly 15,000
pounds to LEO is within reach technologically since
similar vehicles have been demonstrated for more than
four decades. In order to hold down costs and
maintenance difficulties as well as increase reliability,
pressure-fed rocket motors should be used rather than
motors supplied by turbopumps.
This reduces
Table 2 summarizes the aerodynamic implications of
the smaller diameter capsule. Equal weight budgets
and equal performance of the abort and landing system
rockets are assumed for both capsule configurations.
RECOVERY AIRCRAFT
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C-130: 40 ft by 9 ft by 9ВЅ ft,
C-17: 88 ft by 121/3 ft by 18 ft, and
C-5: 121 ft by 13ВЅ ft by 19 ft.
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AIAA/7th Responsive SpaceВ® Conference 2009
system cost equivalent is 1,132 man-years or $285
million (2007).
complexity and thereby increases reliability. With
standard aerospace engineering techniques, the requisite
pressurization of the propellant tanks drove up inert
weight as well as development costs.
Modern
composite materials bring light weight, very robust
propellant tanks into the realm of possibility.
Launch vehicle development cost calculations assume
the team experience adjustment factor of 1.15 used for
the abort motor system development. Table 4 shows
the calculated development costs in equivalent manyears and in billions of 2007 dollars.
The basic conceptual vehicle for SUSTAIN is a three
stage pressure-fed RP-1 and LOX propelled, vertical
take-off, expendable rocket with an effective payload of
15,060 pounds plus a 15 percent margin. For the clean
sheet cost estimation, a three stage vehicle is assumed
with the characteristics summarized in Table 3. In this
table, the initial and final accelerations are the axial
accelerations experienced by the vehicle at the
beginning and the end of each stage burn. The cost
estimates are based on standard aerospace design and
construction practices without the extensive use of
modern composite materials.
Table 5 summarizes the clean sheet project research and
development costs. Developing the requisite vehicle
from the bottom up would cost more than an estimated
$14 billion using the methods of Koelle and standard
aerospace methodology, with total project development
costs of about $19 billion.
Table 4. Launch Vehicle Development Cost.
Stage 1 motor
Stage 2 motor
Stage 3 motor
Table 3. Three Stage Launch Vehicle.
Stage 3
Stage 2
Stage 1
23
6,860
42
12,526
35 percent
10,440 ft/sec
Total motor
Delta-V allocation
Stage delta-V
Motor Isp
Stage mass ratio
323
1.935
319
3.390
Stage 1 vehicle
Stage 2 vehicle
Stage 3 vehicle
286 seconds
3.110
Total vehicle
Payload plus margin
Stage mass fraction
Motor T/W ratio
17,319
18.5
85.0
Stage structure
Propellant
Stage GTOW
4,668
20,563
42,550
Nominal thrust
Burn time
Initial acceleration
Final acceleration
123,394
53.8
2.9
5.6
Motor mass
1,452
Stage net mass fraction
15.6
42,550
11.0
85.0
17,836
144,308
204,693
204,693 pounds
10.0 percent
85.0
62,693 pounds
564,235 pounds
831,621 pounds
45,082
8.0
5,132
3,202
2,142
$ 1.293 billion
0.807
0.540
10,476
2.640
22,343
14,301
7,664
5.630
3.604
1.931
44,307
$ 11.165
Man-years
pounds
percent
2007 cost
Capsule
19,494
Abort motor
1,132
3 stage vehicle motors 10,476
3 stage vehicle
44,307
$ 4.912 billion
0.285
2.640
11.165
Total
$ 19.003
75,410
This quantity of development funding is almost
certainly beyond current US political and economic
feasibility. An alternative approach is to exploit that
which already exists and which was developed under
other funding arrangements.
CLEAN SHEET DEVELOPMENT COSTS
Clean sheet development costs for SUSTAIN are
developed using the methods devised by Dietrich
Koelle with 2007 US engineering costs in the aerospace
industry4.
For example, as part of the Responsive Space
initiatives, Microcosm developed a conceptual low cost
rapidly deployed launch system (Scorpius). Much of
the preliminary development cost for this series of
vehicles has been spent and hardware has been
developed.
Capsule development assumes a technical status
adjustment factor of 1.15 and a team experience
adjustment factor of 1.35. This results in a cost
equivalent of 19,494 man-years or $4.9 billion (2007)
for capsule development.
THE MICROCOSM APPROACH
The analysis given above suggests a concept close to
the Scorpius launch system studied, partially developed,
The abort motor development assumes a lower team
experience adjustment factor of 1.15. The abort motor
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2007 cost
Table 5: SUSTAIN Development Cost.
388,917 1,496,918 pounds
118.4
107.8 seconds
1.9
1.8 gravities
6.4
5.6 gravities
4,575
9.2
Man-years
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AIAA/7th Responsive SpaceВ® Conference 2009
and demonstrated by Microcosm5. The Sprite vehicle
payload of 1,060 pounds to LEO and the Exodus
vehicle payload of 19,700 pounds to LEO are sufficient
to meet the launch requirements of the USAF UAV and
the SUSTAIN capsule, respectively, with comfortable
margins.
The Scorpius system uses a series of seven essentially
identical modules – 6 for the first stage in a hexagonal
array and one for the second stage nested in the center
of the first stage array. All are pressure-fed with
ablatively-cooled motors, thus minimizing the
propulsion system parts count. Each individual module
is extremely simple and easily storable with minimal
difficulty. By running the 6 first stage modules in
parallel with the second stage located within the
hexagonal array, the overall vehicle aspect ratio is low
and sensitivity to wind shear is minimal. That allows
launch in hurricane level winds.
The overall characteristics of the Scorpius launch
vehicle family are shown in Table 6.
Table 6. Scorpius Vehicle Family.
Sprite
LEO Payload
Launch Price
Overall Height
Pod Diameter
Vehicle Diameter
GLOW
Engine Configuration
Stage 1
Stage 2
Stage 3
Max Axial Accel.
Stage 1
Thrust, vacuum
Thrust, sea level
Gross Mass
Stage 2
Thrust
Gross Mass
Stage 3
Thrust
Gross Mass
Liberty
Exodus
1,060
5.2
54.2
42.0
11.2
80,500
4,200
11.6
83.0
66.38
174.8
322,300
19,700
26.7
143.5
111.2
29.7
1,497,000
6x20K
1x20K
1x2.6K
5.9
6x80K
1x80K
1x9K
5.9
120,000
101,000
65,600
480,000
405,000
262,000
2,310,000 lbs
1,944,000 lbs
1,220,000 lbs
22,300
10,900
89,200
43,600
429,000 lbs
202,700 lbs
2,300
3,005
9,200
12,020
44,300 lbs
55,900 lbs
lbs
million
ft
in
ft
lbs
6x385K
1x385K
1x39K
5.9
gravities
Scorpius has a proprietary pressurization system that
end-runs many of the disadvantages of conventional
pressurization systems and provides roughly a 50
percent mass savings over stored Helium gas and
associated tankage for propellant pressurization. This
system uses Tridyne (Helium with small,
noncombustible quantities of Hydrogen and Oxygen
added)6. The Tridyne is run over a catalyst bed and the
Hydrogen and Oxygen react to form water vapor and
release heat. Some of the heat is transferred back to the
stored Tridyne tank to prevent cooling by expansion as
the gas is depleted. The remainder heats the Helium
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9
(and the small amount of water vapor) used for
propellant pressurization to a typical temperature of 200
to 250 degrees F.
The most important aspect of Scorpius is that all major
systems have been tested in flight. A basic module was
flown with a partial propellant load (to avoid overflying
the range limits) at White Sands some years ago. The
flight took place within hours of pulling up to the
concrete pad with vehicle, fuel truck, etc. This
demonstrated responsiveness.
A scalable and
inexpensive growth path has been defined.
This history of Scorpius launch system development
establishes a significant portion of the overall
technology.
The estimated remaining development cost for the
complete Scorpius series up to and including the
Exodus launch vehicle is less than $256 million. The
unit launch costs for Exodus are estimated to be $26.7
million. Therefore, incorporation of the Scorpius
Exodus into SUSTAIN would reduce estimated
development costs by about 71 percent relative to the
clean sheet approach. The remaining high cost item
would be the capsule.
Using a cheap, modular, very simple expendable launch
vehicle may seem to be a backwards approach, but an
expendable Scorpius can be envisioned with unit costs
much less than the amortized research and development
on a reusable launch vehicle unless many thousands of
flights are envisioned per vehicle. SUSTAIN would be
anticipated to be a low frequency use system that would
not allow for amortizing the development costs incurred
in a dedicated reusable launch vehicle.
Extensive reusability is not easily attained from current
technology, which has demonstrated partial reusability
of on the order of only 100 flights. One can argue that
airplane-like operations are not going to be attained
with reusable space launch vehicles in the near term
regardless of what one hears from the alt.space
community. This has been extensively discussed in
other documents for commercial suborbital sounding
rockets and for orbital launchers7,8,9.
SQUAD RECOVERY
Recovery would be accomplished by a variant of the
technology used to recover film canisters from
reconnaissance satellites in the early 1960s.
Two possible scenarios are envisioned. The first is to
recover individual team members by FRS and a cargo
aircraft (C-130, C-17, or C-5). If the FRS is used, the
recovery aircraft would slow from cruising speed to an
AIAA/7th Responsive SpaceВ® Conference 2009
indicated airspeed of between 130 and 140 knots (220
to 236 ft/sec) at a designated initial point. Reducing
speed reduces the acceleration experienced by the
individual Marines during recovery.
The second scenario is to recover the capsule by the
same method with the team contained inside the
capsule.
Large Diameter Capsule
The larger diameter capsule could be recovered with a
C-17 or C-5 at 180 knots.
Redundancy can be provided by using three aircraft in
order to have three attempts at successfully snagging
the capsule. The three aircraft would approach the
vertical line defining the team capsule trajectory after
the abort system rocket is fired for extraction.
Figures 2 and 3 show the altitude and vertical speed
histories of the capsule after the abort system rocket is
fired to initiate extraction.
Altitude History
3,500
just before burnout as air drag becomes significant.
Right after burnout, air drag provides a perceived axial
deceleration of about 2.3 gravities (eyeballs out) which
tails off as the capsule approaches apogee. This is
shown in Figure 4. These acceleration levels are easily
tolerated by supine capsule occupants if they are
secured properly to resist the negative accelerations
experienced after rocket burnout.
All aircraft fly to an initial point and then decelerate
and descend to an assumed true airspeed of 180 knots
(304 ft/sec) for the C-17s or C-5s. The altitudes and
timing are approximate since they are based on the
assumption that the aircraft can maintain constant speed
using autothrottles while climbing or diving on a
circular trajectory with a radius of 700 feet.
The provisional assumption is that a trailing net
arrangement is used to capture the capsule. Momentum
transfer between the capsule, moving vertically with
essentially no horizontal momentum, and the recovery
aircraft, with all momentum oriented essentially
horizontally, could occur over a tolerable time frame by
using an aircraft winch arrangement that reels out at a
constant force and then reels the captured capsule back
toward the aircraft for docking and stowage.
2,500
Acceleration History
2,000
1,500
1,000
500
0
0
5
10
15
20
25
30
35
Time (sec)
Figure 2. Large Capsule Altitude History.
Axial Acceleration
Altitude (ft)
3,000
7
6
5
4
3
2
1
0
-1
-2
-3
0
5
10
15
20
25
30
35
Time (sec)
Vertical Speed History
Figure 4: Large Capsule Acceleration History.
Speed (ft/sec)
600
500
400
300
200
100
0
0
5
10
15
20
25
30
35
Time (sec)
Figure 3. Large Capsule Vertical Speed History.
The first aircraft approaches the capsule line of flight
by flying level at lowest altitude, the second approaches
three seconds (912 feet) behind the first in level flight at
a higher altitude (about 184 feet higher than the first),
and the third approaches three seconds behind the
second in level flight at an intermediate altitude (about
108 feet below the second aircraft). If wake turbulence
is a problem, the headings could be varied during the
approach by perhaps 10 degrees with all approaching
the same vertical line. Alternatively, the aircraft could
fly the same heading while separated laterally by one or
two wingspans.
During squad extraction, the perceived axial
acceleration (eyeballs in) during the abort system rocket
The first aircraft attempts to snag the capsule 3 seconds
burn starts out at about 6.6 gravities, peaks at about 6.7
gravities, and then drops down to about 5.1 gravities
before apogee, when it is ascending at 99 ft/sec and is at
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AIAA/7th Responsive SpaceВ® Conference 2009
As in the case of the large diameter capsule recovery,
redundancy could be provided by using three aircraft in
order to have three attempts at successful snagging of
the capsule. In this instance, a C-130 could be used for
recovery and that scenario will be discussed. As with
the large capsule, the three aircraft would approach the
vertical line defining the trajectory of the team capsule
after the abort system rocket is fired for extraction.
an altitude of 3,098 feet AGL. To meet at the capsule
line of flight with a climb rate of 99 ft/sec, the aircraft
pitches up 228 feet before the capsule line from an
altitude of 3,060 feet. After the attempt, it immediately
breaks away and resumes level flight.
The second aircraft attempts to snag the capsule at
apogee, when it has a climb rate of approximately zero
and an altitude of 3,244 feet AGL. In order to attempt
recovery, it arrives at the capsule line of flight 3
seconds after the first aircraft. After the attempt, it
immediately breaks away and resumes level flight.
Figure 5 and Figure 6 show the altitude and vertical
speed histories of the capsule after the abort system
rocket is fired to initiate extraction.
The third aircraft attempts to snag the capsule 3 seconds
after apogee, when it is descending at 95 ft/sec and is at
an altitude of 3,101 feet AGL. To meet the capsule, it
pitches down 219 feet before the capsule line from an
altitude of 3,136 feet. After the attempt, it also
immediately breaks away and resumes level flight.
Altitude History
6,000
Altitude (ft)
5,000
The aircraft making the successful capture reels in the
capsule and stows it for proper center of gravity
maintenance of the aircraft as rapidly as possible. All
three aircraft then join up into whatever formation is
appropriate and head for home. The breakaway
maneuvers could assume different headings for each
aircraft if it does or does not make the successful snag
prior to the join up for departure from the target area.
4,000
3,000
2,000
1,000
0
0
5
10
15
20
25
30
35
40
Time (sec)
Figure 5. Small Capsule Altitude History.
Vertical Speed History
Whether the team recovery is successful or not, the
entire capsule and team recovery would be determined
in less than 20 seconds after the capsule abort rocket is
ignited. This hopefully minimizes the time during
which a potential enemy can react.
700
Speed (ft/sec)
600
In the scenario outlined above, air supremacy is
assumed with fixed antiaircraft installations suppressed
or destroyed. Belly armor for the aircraft might be
considered if .50 cal ground fire is anticipated or if it is
not assumed to be suppressed.
500
400
300
200
100
0
0
5
10
15
20
25
30
35
40
Time (sec)
Figure 6. Small Capsule Vertical Speed History.
Other scenarios could be developed that use multiple
aircraft approaching either tighter or looser windows
around capsule apogee. Compensation for varying
target zone elevations, variations in air temperatures,
etc. need to be analyzed in more detail.
Figure 7 shows the small capsule acceleration history.
Acceleration History
Axial Acceleration
A relatively small parachute could be deployed just
after apogee to slow capsule descent in case the three
recovery attempts fail. This might allow one aircraft a
chance to go around and make a fourth attempt at
recovery. Alternatively, a second full-sized parachute
could be used as a reserve during re-entry for insertion
and as a survival option during capsule recovery if
aircraft recovery fails.
7
6
5
4
3
2
1
0
-1
0
5
10
15
20
25
30
35
40
Time (sec)
Small Diameter Capsule
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AIAA/7th Responsive SpaceВ® Conference 2009
Figure 7. Small Capsule Acceleration History.
During squad extraction, the perceived axial
acceleration (eyeballs in) during the abort system rocket
burn starts out at about 6.6 gravities, peaks at about 6.8
gravities, and then drops down to about 6.7 gravities
just before burnout because of air drag. Right after
burnout, air drag provides a perceived axial
deceleration of about 0.6 gravities (eyeballs out) which
tails off as the capsule approaches apogee.
These
acceleration levels are easily tolerated by supine
capsule occupants if they are secured properly to resist
the negative accelerations experienced after rocket
burnout.
All three C-130s fly to an initial point and then
decelerate and descend to an assumed true airspeed of
180 knots (304 ft/sec). As with the large diameter
capsule, the altitudes and timing are based on the
assumption that the aircraft can maintain constant speed
using autothrottles while climbing or diving on a
circular trajectory with a radius of 700 feet.
A trailing net arrangement is used to capture the
capsule. Momentum transfer between the capsule,
moving vertically with essentially no horizontal
momentum, and the recovery aircraft, with all
momentum oriented essentially horizontally, could
occur over a tolerable time frame by using an aircraft
winch arrangement that reels out at a constant force and
then reels the captured capsule back toward the aircraft
for docking and stowage. By using a slower C-130
instead of a larger aircraft, the momentum transfer
problem is reduced significantly.
The first aircraft approaches the capsule line of flight
by flying level at lowest altitude, the second approaches
three seconds (912 feet) behind the first in level flight at
a higher altitude (about 183 feet higher than the first),
and the third approaches three seconds behind the
second in level flight at an intermediate altitude (about
108 feet below the second aircraft). As with the larger
capsule scenario, the smaller C-130 headings could be
varied during the approach by perhaps 10 degrees with
all approaching the same vertical line, or they could fly
the same heading while separated laterally by one or
two wingspans in order to reduce wake turbulence.
The first aircraft attempts to snag the capsule 3 seconds
before apogee, when it is ascending at 97 ft/sec and is at
an altitude of 5,157 feet AGL. To meet at the capsule
line of flight with a climb rate of 97 ft/sec, the aircraft
pitches up 224 feet before the capsule line from an
altitude of 5,120 feet. After the attempt, it immediately
breaks away and resumes level flight.
Jurist
The second aircraft attempts to snag the capsule at
apogee, when it has a climb rate of approximately zero
and an altitude of 5,303 feet AGL. In order to attempt
recovery, it arrives at the capsule line of flight 3
seconds after the first aircraft. After the attempt, it
immediately breaks away and resumes level flight.
The third aircraft attempts to snag the capsule 3 seconds
after apogee, when it is descending at 96 ft/sec and is at
an altitude of 5,159 feet AGL. To meet the capsule, it
pitches down 221 feet before the capsule line from an
altitude of 5,195 feet. After the attempt, it also
immediately breaks away and resumes level flight.
The aircraft making the successful capture reels in the
capsule and stows it for proper center of gravity
maintenance of the aircraft as rapidly as possible. All
three aircraft then join up into whatever formation is
appropriate and head for home. The entire capsule and
team recovery would be determined in less than 25
seconds after the capsule abort rocket is ignited.
As with the larger capsule, parachutes could be used to
slow capsule descent to enable an alternative if the
primary aircraft recovery attempts fail.
Variant Recovery
An additional recovery scheme has operational security
implications. That is to use the abort system rocket to
escape the landing zone area. Therefore, the capture
aircraft could snag the capsule some distance from the
landing zone and at an apogee when the capsule has a
horizontal vector. That approach would add to the
uncertainty of the escape direction from the enemy
point of view and possibly provide additional
operational security for the team recovery process. In
addition, the potential momentum transfer problem
would also be partially mitigated.
Since the Scorpius Exodus system provides a
significant payload mass margin, the margin could
potentially be sacrificed to provide a larger, higher
performance escape rocket. If the smaller diameter
capsule system is considered with the abort rocket
system mass increased from 1,500 pounds to 4,075
pounds and a propellant mass increase from 1,200
pounds to 3,260 pounds with all other parameters held
constant, having the capsule pitch 55 degrees at 30
degrees per second starting one second after ignition
increases burnout altitude from 898 feet to 3,350 feet.
Burnout speed is increased from 560 ft/sec to 1,065
ft/sec with range at burnout increased from 197 feet to
3,151 feet. Apogee is shifted from 4,031 feet AGL and
3,131 foot range to 7,724 feet AGL and 12,498 foot
range. Horizontal speed at apogee is increased from
200 ft/sec to 541 ft/sec.
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AIAA/7th Responsive SpaceВ® Conference 2009
No attempt at optimizing any parameters was made in
this analysis.
If the small diameter capsule were equipped with
deployable wings, additional horizontal distance from
the landing zone could be obtained during the initial
phase of the extraction process. Detailed consideration
of this option is beyond the scope of this paper.
Figure 8 shows the two trajectories for the small
diameter capsule when it is pitched right after lifting off
with the standard abort system motor and with the
larger motor that sacrifices the payload margin of the
Exodus launch system.
Small Diameter Capsule Trajectories
system? Will two or three Marines need to stay behind
to guarding the capsule? Is there some form of
camouflage to be employed to allow for the full squad
to execute their mission with a reasonable belief that
there will be no extraction-ending compromise to the
vehicle during their absence?
As a security consideration, warning of aircraft
approaching the target zone may mislead targeted
personnel to assume that the planes carry troops for a
ground operation when they are actually for extraction
of troops either in the SUSTAIN capsule or
individually. This provides a potential miscue to an
enemy, exploiting a classical element of Clausewitz’s
“fog of war.” Aircraft operations could take place
during the night with night vision goggles used by the
flight crews to provide additional security.
10,000
For the purpose of this report, air supremacy is assumed
in terms of suppression of fixed antiaircraft
installations. If the environment is denied as described
above, air cover of some type is assumed unless
operational covertness can be reasonably assured.
Altitude (ft)
8,000
6,000
4,000
2,000
0
0
5,000
10,000 15,000 20,000 25,000
Range (ft)
Standard abort motor
Use of the capsule for recovery versus individual
recovery and a one-way capsule will influence
operational considerations.
Large abort motor
Figure 8. Pitched Capsule Trajectories.
SECURITY, SAFETY, AND LOGISTICS
Detailed description of the SUSTAIN system requires
consideration of the anticipated insertion environment
for the Marines:
•
•
•
Permissible: Everyone on the ground is happy
to see them and is helpful.
Semi-permissible: Locals are neutral to the
presence of US troops on the ground or it is
uncertain as to the ultimate intentions of the
locals.
Denied: The Marines should anticipate hostile
reactions up to and including lethal force upon
disembarking.
The environment is important as the Marines will
probably need to be able to leave the capsule location to
perform their mission. If they return to a compromised
or damaged vehicle, then their extraction is more
complicated even if the individual recovery scenario is
planned.
There is an implied issue of the method of maintaining
the security and viability of the capsule during the
mission. Will the capsule have an active self-protection
Jurist
For example, in a one-way capsule insertion in response
to a non-combatant emergency evacuation operation,
the capsule would be going into at least a semipermissible environment, if not totally denied. This
would reduce the mission requirement for capsule
protection. The National Command Authority could
authorize a mission using such a vehicle under InExtremis conditions.
What if each of the 13 Marines had an ejection seat in a
one-way capsule? It would add to both weight and
complexity, but it would also offer the unique
possibility of using a double-ply parachute which could
be both parachute (aerodynamic load from beneath) and
gas bag when inflated. Then, the only thing that would
absolutely have to be protected in the capsule for
extraction is a lighter-than-air gas generator or tank.
All the recovery aircraft has to do is chase down each
of its floating Marines and snag them using FRS. It
would also be possible to transport additional units into
the landing zone to help extract the trapped members at
a US Embassy, for example. This would reduce the
mission requirement for three cargo aircraft to
maneuver in close proximity to one another. Two could
be used, but the mission could get away with only one.
A gas bag sufficient to lift a Marine may intuitively
appear to be excessively large. An approximation of
the size of the gas envelope for this concept can be
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AIAA/7th Responsive SpaceВ® Conference 2009
augment, or rescue the team if it were not able to use
aerial recovery for some reason.
estimated. At one atmosphere and freezing point of
water (760 mm Hg and zero degrees C), the net
buoyancy of Helium is 0.06942 pounds per cubic foot.
A 220 pound marine carrying 75 pounds of returnable
supplies and 50 pounds for the harness and balloon
envelope will require a minimum of 345 pounds of lift
to be neutrally buoyant. That requires 4,970 cubic feet
of Helium or the equivalent of a sphere about 21.2 feet
in diameter. If the Helium is supplied from a Tridyne
gas generator rather than simply from compressed gas,
it will be heated above ambient temperature and can
provide additional buoyancy.
Once the capsule is brought aboard the aircraft and
stowed, alternate team extraction scenarios must be
considered as a function of recovery aircraft type.
Those considerations include dealing with jammed
capsule hatches, and stabilization and treatment of
injured Marines by a medical team aboard the aircraft.
Potentially hazardous materials aboard the capsule as it
is docked and stowed could include returning
munitions, residual abort/escape system rocket
propellant (burning slivers), unused landing rocket
propellants, attitude control system propellants, unfired
explosive bolts, etc. If the attitude control system were
to avoid hydrazine in favor of less toxic propellants
such as gaseous oxygen and methane or alcohol,
material hazards within the aircraft cargo bay could be
minimized.
If the capsule is designed for one-way use, then the
vehicle equipment and instrumentation could be more
spartan, reducing the opportunity for unintended
capture and technology transfer. It also reduces the
operational need for an Emergency Seek and Destroy
mission to prevent the capsule from getting into
unfriendly hands if the capsule destruct is not used for
some reason.
The flight duration for the Marines is specified as two
hours. In a true In-Extremis situation, the Marines
would probably have a team ready for immediate
deployment at each launch location within the
continental US. One could assume one hour to brief
and pickup the combat kits, and one hour to launch. It
is then two hours to target once the decision is made to
launch. So the time to target is in reality about four
hours from "Go."
If the capsule is intended for single use even if the
squad is recovered within the capsule, the squad
members could exit the capsule into the larger cargo
aircraft and the capsule could be dumped. Under this
scenario, the capsule could be torn apart within the
cargo aircraft after the mission if necessary to extricate
wounded or injured Marines and begin their treatment.
If the C-130 is used to recover the smaller diameter
capsule, there is a question about aircrew safety since
the capsule dimensions effectively prevent anyone from
passing from the forward end of the cargo compartment
past the capsule to the rear of the aircraft. This limits
both the ability to extricate wounded Marines while
keeping the capsule intact, and also the ability to deal
with a fire or hazardous material leak. Otherwise, the
recourse is to jettison the capsule or deal with the
consequences until the aircraft lands.
An important consideration involves the transport
aircraft flying to the extraction location. Assuming the
use of recovery C-130s with the smaller diameter
capsule, and assuming that they are ready for
immediate launch, by the time they climb to cruise
altitude and are moving toward the extraction point, the
C-130s could only get about 1,200 nautical miles
downrange in that four hours – optimistically with no
wind. If the Marines turn the aircraft into flying gas
tanks, they could get another two hours downrange
without needing to refuel, assuming that they are going
back to where they came from. So with minimum fuel
reserves upon return, the combat radius for this mission
would be about 1,500 miles at best. If the KC-130s in
this scenario can take on fuel, then they could certainly
go much further.
Another safety issue relates to the use of pressure suits
on the Marines. Current military regulations require
use of such suits above altitudes of 50,000 feet. If the
Marines must take action immediately after exiting the
capsule, they would rapidly overheat if still in pressure
suits while expending metabolic energy at a rapid rate.
Yet, removal of the suits takes significant and possibly
very valuable time. The existence of an injured Marine
either in field conditions or aboard the recovery aircraft
would require at least partial removal of the suit to gain
access for treatment.
This is why the operational environment is so
important. Based upon the consumables assumed with
no food, water or medical kits noted, the 13 Marines
probably have enough ammunition for a 20 to 30
minute running fire fight. A team going into an InExtremis situation is going to need timely extraction,
re-supply, or cached supplies.
The short rough field landing capability of the C-130 is
an attractive feature because it could potentially land
near the target zone in an emergency to resupply,
Jurist
14
AIAA/7th Responsive SpaceВ® Conference 2009
be able to respond or address the issues in a way that is
supportive to the SUSTAIN project.
The discussion above demonstrates the criticality of
defining the environment the USMC envisions. One
must conclude that the air-breathing extraction asset
will be the long pole in this operational tent.
Because SUSTAIN utilizes space for its operating
window, it may be depicted as a space weapon system.
Not only is weaponizing space not permitted by
existing United Nations agreements, it is an
overwhelming concern to many people interested in
space development14. They genuinely express their fear
and concern about the US moving in this direction.
Advocates of SUSTAIN will need to carefully
characterize and define the concept. If SUSTAIN is
identified as a space weapons system, advocates may
well have an even more difficult time selling it. As we
are about to embark upon a new Administration and
Congress in very challenging economic, political, and
national security times, SUSTAIN will undergo its
planning in times of great uncertainty. This must not be
overlooked or discounted.
POLITICAL, POLICY, AND ECONOMIC ISSUES
The political, policy, and economic issues are also
formidable challenges to realizing SUSTAIN. These
issues will require decision making that may prove to
be more difficult than any technical and engineering
challenges. Some of these challenges are discussed
below.
At present, many different US space policy
recommendations come from a variety of organizations,
panels, and committees. Some go directly to Congress.
Others are the subject of important writings,
conferences, books, and various other venues. US
space policy is political and in play. Space policy
recommendations include those from, among others:
•
•
•
•
•
•
A positive aspect of using aircraft for atmospheric
return instead of returning the team through suborbital
space relates to the Outer Space Treaty. In a scenario
where the team of Marines captures or recovers and
extracts a nuclear weapon from foreign interests,
transporting the weapon back to the continental
U.S. through the atmosphere avoids any possible
complications associated with the Outer Space Treaty
that might entail by transporting the weapon through
space.
Institute For Defense Analysis 2008 report 10;
The National Space Forum 2008, sponsored by
the Eisenhower Center for Space and Defense
Studies at the US Air Force Academy and by
the Center for Strategic and International
Studies (CSIS) in Washington, D.C.;
Various other CSIS space policy statements;
The Center for Defense Information (CDI);
and
Private organizations such as the Secure World
Foundation11.
Relevant books on the subject have been
published in the popular press, such as
Twilight War12.
If SUSTAIN is not characterized as a space weapons
system, it is still going to need to be compatible with
other space policy recommendations and programs.
There will be a finite amount of money made available
for space development programs, even on the defense
side of the budget. With national security space a
priority, growing recognition of the need for
operationally responsive space to protect our military
and commercial space assets, the need for replacing
aging legacy hardware, and the deed to establish space
infrastructure, SUSTAIN will encounter a very
competitive environment for project approval and
spending.
These and other communication routes have the access
needed to reach members of Congress, policy makers,
staff, and those that count when lobbying for project
funding and for authority with their policy
recommendations, theories, ideas, research, and
conclusions.
With the political diversity and the variation in policy
recommendations represented in today’s US society, it
is likely that space policy will reflect and mirror back to
us that which is happening in other sectors of our
nation. We may see policy recommendations for a
military presence in space to defend against or ward off
a potential “Space Pearl Harbor” as cited in the 2001
Rumsfeld Space Commission Report13, or the opposite
extreme of having no space military presence, or
something in-between. In order to sell SUSTAIN to
Congress and to the American people, we must be
aware of the differences in policy recommendations and
Jurist
The various space policy entities need to be heard and
their arguments need to be understood so that
SUSTAIN advocates can effectively and appropriately
respond. An approach that is inclusive rather than
exclusive to a particular point of view stands a better
chance of ultimate approval. In any event, policy
organizations will at the very least carry out oversight
and watchdog efforts on SUSTAIN, so working with
them from the outset can only facilitate support for the
project and help bring SUSTAIN to operational status.
15
AIAA/7th Responsive SpaceВ® Conference 2009
The skill to which these policy and economic issues are
addressed may very well hold the key to SUSTAIN’s
success. One would hope that SUSTAIN will draw
upon the same caliber of people to make its case in
these categories as it will in its technical, engineering,
and operational aspects.
The cost of SUSTAIN can only be approximated at this
point. An accurate economic analysis of the project is
required. However, even in the early stages of
discussing
SUSTAIN
with
the
appropriate
commissions, representatives, staffers, and policy
makers, the project will have to demonstrate a cost
benefit analysis showing that it is worth doing.
CONCLUSIONS
Although development and operational costs are
derived from different portions of the US budget, they
both need to be paid if the need for the project is to
proceed. The present budgeting and acquisition
processes tend to ignore this truism and therefore drive
up the total costs to the taxpayer.
Microcosm’s Scorpius launch vehicle family offers
inexpensive solutions to expressed needs by the USAF
for placing a UAV anywhere in the world from the
continental United States within 2 hours and by the
USMC for placing 13 Marines anywhere in the world
within the same time frame. Most of the systems for
their Sprite vehicle, the potential UAV launcher, have
been tested. A scalable development path exists up to
the Exodus system that can serve as a launcher for a
SUSTAIN vehicle carrying 13 men and supplies.
With today’s technology, a fully equipped Marine
squad could be deployed to just about any hot spot on
Earth in about 24 hours. If forward basing were
strategically located near the hot spot, this time interval
might be even less. To justify SUSTAIN, promoters
must show a clear benefit for deploying the Marine
squad within two hours to the hot spot versus the time it
would take using a combination of today’s military
airlift and deployment tools. The bottom line question
is: How important is meeting that two hour time limit
versus up to perhaps a 24 hour time limit that we should
spend billions of dollars on research and development
and expose the Marine squad to the added risks of using
rocket propulsion as compared to proven aviation
technology? Would the opportunity costs be worth it?
That is, could the time and money spent on working the
SUSTAIN problem be better spent on something else?
For SUSTAIN to win approval, it cannot be valuable
only on the margins. It must show a clear benefit on all
accounts.
The Sprite vehicle is estimated to cost $5.2 million per
launch at an annual flight rate of 5. Most of the first
orbital flight test vehicle has been fabricated. A
scalable path to development of Exodus shows an
estimated launch cost for this latter vehicle of $26.7
million and a total remaining development cost for the
entire vehicle family up to and including Exodus of less
than $256 million. This contrasts to an estimated
development cost of $13.8 billion for an equivalent
clean sheet vehicle, including motors, using
conventional aerospace design and fabrication
technology. The lower costs come from use of
composite materials for very robust structural
components, including tankage. The strength and
weight considerations of the vehicle propellant tanks
allow pressurization to an extent that motor turbopumps
are not required. Use of ablative motor chambers and
avoidance of turbopumps reduces the parts count and
costs significantly. Costs are so low relative to existing
launch vehicles that expendable vehicles become
reasonable if amortization of development over the
program life is considered.
Depending on the sources one deems credible, we
appear to be in uncharted economic times for our
nation. The economic condition of our nation now and
for the foreseeable future will largely dictate how we
deal with legacy infrastructure, new programs,
innovation, and much more. The political views of the
current Congress and Administration will also be
influential. Innovative projects like SUSTAIN will face
stiff competition not only with competing defense
projects, but with other federal projects in general, such
as healthcare, Social Security, or who knows what.
SUSTAIN’s case will have to be very strong with
tangible, credible, and identifiable national benefits.
USAF has been seeking responsive and lower cost
launch capability for years. The Scorpius series offers
exactly that. Since the USMC has a requirement that
can be met by the Scorpius Exodus, continuing
development of this series and the associated
manufacturing technology provides a synergism that
not only reduces launch costs. It should be noted that
USAF Security Forces and US Special Operations
Command have expressed interest in SUSTAIN. This
represents
an
opportunity
for
development
commonality to meet multiple missions on an unusual
scale.
The political, policy, and economic hurdles for
SUSTAIN may very well prove to be as challenging, if
not more so, than the technical and engineering
challenges of designing, building, testing, deploying,
and operating SUSTAIN. These broader challenges
must not be overlooked, glossed over, or treated lightly.
Jurist
16
ACKNOWLEDGEMENTS
AIAA/7th Responsive SpaceВ® Conference 2009
This work was supported in part by a grant from CRM,
Inc., Billings, Montana. We thank Drs. James R. Wertz,
Robert Conger, and Thomas Bauer of Microcosm, Inc.
for provision of data on the Scorpius program.
Portions of this paper were presented at the Project
SUSTAIN closed briefing to USMC Commandant’s
Office staff, US Security Space Office (NSSO) staff,
USAF Security Forces Center Headquarters (HQ
AFSFC) staff, and US Special Operations Command
(USSOCOM) staff on 26 February 26 2009, Lackland
Air Force Base, Texas.
REFERENCES
1.
10.
11.
12.
See Blue Origin web site
13.
http://public.blueorigin.com/index.html (accessed 5
February 2009).
2. See Armadillo Aerospace web site
http://www.armadilloaerospace.com/n.x/Armadillo
/Home (accessed 5 February 2009).
14.
3. William M. Leary, “Robert Fulton’s Skyhook and
Operation Coldfeet,” US Central Intelligence
Agency Library, Vol. 38 #5, 1995. See the CIA
web site https://www.cia.gov/library/center-forthe-study-of-intelligence/csi-publications/csistudies/studies/95unclass/Leary.html (accessed 15
January 2009).
4. Dietrich E. Koelle, Handbook of Cost Engineering
for Space Transportation Systems, Rev. 2 (D-85521
Ottobrun, Liebigweg 10, Germany, TransCost
Systems, 2007).
5. John M. Jurist, site visit, Microcosm, Inc.,
Hawthorne, CA, 7 June 2007. Additional
information was verified by James R. Wertz,
personal communication (7 June 2007). Also, see
Microcosm, Inc. web site
http://www.smad.com/ie/ieframessr2.html
(accessed 29 September 2008).
6. Shyama Chakroborty, Mark Wollen, and Lee
Malany, “Development and Optimization of a
Tridyne Pressureization System for Pressure Fed
Launch Vehicles,” American Institute of
Aeronautics and Astronautics, AIAA 2006-4716,
(July 2006).
7. John M. Jurist, Sam Dinkin, and David M.
Livingston, “When Physics, Economics and
Reality Collide: The Challenge of Cheap Orbital
Access,” American Institute of Aeronautics and
Astronautics, AIAA 2005-6620, (September 2005).
8. James R. Wertz, “Economic Model of Reusable vs.
Expendable Launch Vehicles,” International
Astronautics Federation Congress, (Rio de Janeiro,
Brazil, 2-6 October 2000).
9. John M. Jurist, “Commercial Suborbital Sounding
Rocket Market: A Role for Reusable Launch
Jurist
17
Vehicles,” Astropolitics, (Submitted 20 December
2008, accepted for publication 28 January 2009).
A. Thomas Young, Edward Anderson, Lyle Bien,
Ronald R. Fogleman, Keith Hall, Lester Lyles, and
Hans Mark, Leadership, Management, and
Organization for National Security Space: Report
to Congress of the Independent Assessment Panel
on the Organization and Management of National
Security Space (Alexandria, VA, Institute for
Defense Analysis, July 2008).
See Secure World Foundation website
www.secureworldfounation.org (accessed 5
February 2009).
Mike Moore, Twilight War: The Folly of U.S.
Space Dominance (Oakland, CA, The Independent
Institute, 3 March, 2008).
Donald Rumsfeld et al., “Space Pearl Harbor” as
cited on Page 13 in the Report of the Commission
to Assess United States National Security Space
Management and Organization (Washington, D.C.,
11 January 2001).
David Livingston et al., multiple discussions of
space weaponization on The Space Show with
guests and listeners, 26 January 2009, 13 January
2009, 12 January 2009, 1 September 2008, 5
February 2008, 31 January 2008. See The Space
Show web site http://www.thespaceshow.com .
AIAA/7th Responsive SpaceВ® Conference 2009
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