U.S. patent application number 14/026791 was filed with the patent office on 2014-01-16 for direct flight far space shuttle.
The applicant listed for this patent is Robert Salkeld. Invention is credited to Robert Salkeld.
Application Number | 20140014779 14/026791 |
Document ID | / |
Family ID | 40674741 |
Filed Date | 2014-01-16 |
United States Patent
Application |
20140014779 |
Kind Code |
A1 |
Salkeld; Robert |
January 16, 2014 |
Direct Flight Far Space Shuttle
Abstract
A vehicle and method for enabling propulsive flight from
suborbital altitudes and velocities directly to far space, or
beyond Low Earth Orbit (LEO), preferably without requiring
injection into or refueling in LEO. The vehicle is preferably
reusable and can withstand re-entry speeds and temperatures higher
than those that are typical for LEO vehicles. The vehicle
preferably lands horizontally, for example on a runway. The vehicle
forms the upper stage of a booster vehicle system which can be
launched either from the ground or from a subsonic air platform.
The vehicle may optionally be used for LEO missions.
Inventors: |
Salkeld; Robert; (Santa Fe,
NM) |
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Applicant: |
Name |
City |
State |
Country |
Type |
Salkeld; Robert |
Santa Fe |
NM |
US |
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|
Family ID: |
40674741 |
Appl. No.: |
14/026791 |
Filed: |
September 13, 2013 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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12115324 |
May 5, 2008 |
8534598 |
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14026791 |
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11871892 |
Oct 12, 2007 |
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12115324 |
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60829179 |
Oct 12, 2006 |
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Current U.S.
Class: |
244/159.3 |
Current CPC
Class: |
B64G 1/14 20130101; B64G
1/002 20130101; B64G 1/402 20130101; B64G 1/007 20130101; B64G
1/401 20130101; B64G 1/005 20130101 |
Class at
Publication: |
244/159.3 |
International
Class: |
B64G 1/14 20060101
B64G001/14 |
Claims
1. A reusable upper stage of a space vehicle for performing direct
flight from ignition within the atmosphere and returning to Earth
from far space; said upper stage rigidly attachable to a booster;
and said upper stage having a sufficiently high hypersonic lift to
drag ratio to enable said upper stage to remain at high altitude
long enough to dissipate thermal energy until it decelerates to Low
Earth Orbit velocity.
2. The upper stage of claim 1 having a thrust-to-mass ratio of
between approximately 1 and approximately 0.25.
3. The upper stage of claim 1 configured for horizontal Earth
landing.
4. The upper stage of claim 3 comprising wings.
5. The upper stage of claim 1 having a fineness ratio between
approximately 7 and approximately 8.
6. The upper stage of claim 1 configured to carry six personnel
with life support, environmental control and power provisions for
missions up to 10 days.
7. The upper stage of claim 1 wherein said hypersonic lift to drag
ratio is approximately 3.
8. The upper stage of claim 1 comprising separate tanks for storing
a first fuel, a second fuel, and an oxidizer, which together
comprise a dual fuel propellant.
9. The upper stage of claim 8 wherein said dual fuel propellant
comprises oxygen-hydrocarbon-hydrogen propellant.
10. The upper stage of claim 1 comprising separate tanks for
storing a fuel and an oxidizer, which together comprise a single
fuel propellant.
11. The upper stage of claim 10 wherein said single fuel propellant
comprises oxygen-hydrogen propellant.
12. The upper stage of claim 1 comprising an internal flight
deck.
13. The upper stage of claim 1 comprising end loading cargo bay
doors.
14. The upper stage of claim 1 comprising external propellant tanks
and/or external cargo pods and/or an external moon lander.
15. A space vehicle comprising a booster rigidly attached the upper
stage of claim 1.
16. The space vehicle of claim 15 consisting essentially of said
booster rigidly attached to said upper stage.
17. The space vehicle of claim 15 wherein said booster is
launchable from the ground or from an air launch platform.
18. The upper stage of claim 1 further comprising an internal bay
for transporting a separately propulsive vehicle.
19. The upper stage of claim 18 wherein said separately propulsive
vehicle comprises a characteristic selected from the group
consisting of single stage, non-aerodynamic, and reusable.
20. The upper stage of claim 18 wherein said separately propulsive
vehicle comprises a fuel supply.
21. The upper stage of claim 18 wherein said separately propulsive
vehicle comprises sufficient fuel capacity to travel to a desired
location and return to said upper stage.
22. The upper stage of claim 21 wherein said location is the moon
and said upper stage is disposed in Low Moon Orbit.
23. The upper stage of claim 22 wherein said separately propulsive
vehicle can land on the moon.
24. The upper stage of claim 18 wherein said separately propulsive
vehicle is returnable to Earth while being transported within said
bay.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a continuation of U.S. patent
application Ser. No. 12/115,324, entitled "Direct Flight Far Space
Shuttle", filed on May 5, 2008, which application is a
continuation-in-part application of U.S. patent application Ser.
No. 11/871,892, entitled "Direct Flight Far Space Shuttle", filed
on Oct. 12, 2007, which claims priority to and the benefit of the
filing of U.S. Provisional Patent Application Ser. No. 60/829,179,
entitled "Direct Flight Geolunar Personnel Shuttle", filed on Oct.
12, 2006, and the specifications and claims thereof are
incorporated herein by reference.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention (Technical Field)
[0003] The present invention relates to far space Earth launch and
shuttle transportation vehicle systems.
[0004] 2. Description of Related Art
[0005] Note that the following discussion refers to a number of
publications by authors and year of publication, and that due to
recent publication dates certain publications are not to be
considered as prior art vis-a-vis the present invention. Discussion
of such publications herein is given for more complete background
and is not to be construed as an admission that such publications
are prior art for patentability determination purposes.
[0006] Familiar far space personnel vehicle concepts include
non-propulsive ballistic capsules such as the non-reusable American
Apollo Command Module, as well as reusable non-propulsive concepts
such as the Ballistic Crew Exploration Vehicle now under
development. Other far space concepts include reusable propulsive
aerodynamically lifting and horizontal Earth landing shuttles such
as in accordance with U.S. Pat. No. 5,090,642, Feb. 25, 1992. These
prior art shuttles start from low Earth orbit (LEO), which is
defined herein as having an altitude of less than approximately
1000 km, but do not participate in an Earth ascent launch.
Typically, one or more stages are required to reach LEO, and then
another stage is required to go to a higher orbit.
BRIEF SUMMARY OF THE INVENTION
[0007] The present invention is a reusable upper stage of a space
vehicle for performing direct flight. The upper stage preferably
comprises a thrust-to-mass ratio of between approximately 1 and
approximately 0.25. The upper stage is preferably configured for
horizontal Earth landing. The upper stage preferably comprises a
fineness ratio between approximately 7 and approximately 8. The
upper stage is optionally configured to carry six personnel with
life support, environmental control and power provisions for
missions up to 10 days. The upper stage preferably comprises a
sufficiently high hypersonic lift to drag ratio to enable efficient
re-entry from far space, preferably approximately 3. The upper
stage optionally utilizes dual fuel propellant, preferably
comprising oxygen-hydrocarbon-hydrogen propellant. The upper stage
alternatively optionally utilizes single fuel propellant,
preferably comprising oxygen-hydrogen propellant. The upper stage
preferably comprises an internal flight deck. The upper stage is
optionally configured to load and unload cargo at a back end of the
upper stage. The upper stage preferably comprises external
propellant tanks and/or external cargo pods and/or an external moon
lander.
[0008] The present invention is also a space vehicle consisting
essentially of a booster and the aforesaid upper stage. The booster
is preferably launched from the ground or from an air launch
platform.
[0009] The present invention is also a method of travelling to far
space, the method comprising the steps of launching a space vehicle
comprising a booster and a reusable propulsive upper stage;
activating upper stage propulsion at a suborbital altitude; and the
upper stage performing direct flight; wherein the upper stage does
not enter or exit Low Earth Orbit during ascent to far space. The
launching step preferably comprises launching from the earth's
surface or from an air launch platform. The method preferably
further comprises the step of decelerating the upper stage in
preparation for re-entry. The method preferably further comprises
the step of the upper stage re-entering Earth's atmosphere from far
space. The method preferably further comprises the step of the
upper stage landing horizontally on Earth's surface. The landing
step optionally comprises gliding.
[0010] An object of the present invention is to provide a vehicle
and techniques for improved personnel and/or cargo transport
between the Earth and far space.
[0011] Another object of the present invention is to provide a
vehicle system in which a reusable personnel and/or cargo
transporter preferably functions as both the upper ascent stage of
an Earth launch vehicle and as a far space shuttle performing
horizontal landing Earth return, within present-day practical
limits.
[0012] Another object of the present invention is to provide direct
personnel flight from the Earth surface into far space transfer
trajectories without any pause in low Earth orbit (LEO).
[0013] A further object of the present invention is to provide
direct far space personnel flight using a dual-functioning
apparatus enabling reduced vehicle complexity and cost, reduced
flight time and improved overall safety and reliability.
[0014] Other objects, advantages and novel features, and further
scope of applicability of the present invention will be set forth
in part in the detailed description to follow, taken in conjunction
with the accompanying drawings, and in part will become apparent to
those skilled in the art upon examination of the following, or may
be learned by practice of the invention. The objects and advantages
of the invention may be realized and attained by means of the
instrumentalities and combinations particularly pointed out in the
appended claims.
BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS
[0015] The accompanying drawings, which are incorporated into and
form a part of the specification, illustrate several embodiments of
the present invention and, together with a description, serve to
explain the principles of the invention. The drawings and the
dimensions therein are only for the purpose of illustrating one or
more particular embodiments of the invention and are not to be
construed as limiting the invention. The boosters, engines, and
other components depicted in the drawings are simply used for
example purposes and are not to be construed as limiting the
present invention. In the figures:
[0016] FIGS. 1A-1D depict embodiments of a personnel-only direct
flight far space vehicle of the present invention;
[0017] FIGS. 2A-2E depict embodiments of a personnel/cargo direct
flight far space vehicle of the present invention;
[0018] FIGS. 3A-3C depict an air launch embodiment of a
personnel-only direct flight far space shuttle utilizing the
.DELTA.IV Medium+(5,4) booster;
[0019] FIG. 3D depicts the existing AN-225 aircraft for size
reference;
[0020] FIG. 4A depicts an air launch embodiment of a personnel-only
direct flight far space shuttle utilizing the .DELTA.IV "short"
booster;
[0021] FIG. 4B depicts an alternate use of the air launch
platform;
[0022] FIG. 4C depicts the existing 777-200 aircraft for size
reference;
[0023] FIGS. 5A-5B depict single fuel CFD and IFD configurations of
the present invention;
[0024] FIGS. 6A-6B depict a dual fuel engine cycle schematic and a
dual fuel O.sub.2/MMH/H.sub.2 engine;
[0025] FIGS. 7A-7B depict dual fuel CFD and IFD Endloader
configurations of the present invention;
[0026] FIG. 8 depicts five boosted launch options usable with the
present invention;
[0027] FIG. 9 depicts an EL configuration with an enlarged cargo
bay;
[0028] FIG. 10A depicts an extended mission configuration of an
embodiment of the present invention;
[0029] FIG. 10B shows the embodiment of FIG. 10A configured with a
.DELTA.IV Heavy derivative;
[0030] FIG. 11A depicts an embodiment of the present invention
capable of LMO with a moon lander;
[0031] FIG. 11B shows the embodiment of FIG. 11A together with the
Ares V Heavy Earth launchers;
[0032] FIG. 11C shows an LEO version of the embodiment of FIG. 11A;
and
[0033] FIG. 11D shows the embodiment of FIG. 11C together with a
.DELTA.IV Heavy Earth launcher.
DETAILED DESCRIPTION OF THE INVENTION
[0034] The present invention encompasses a vehicle for transporting
personnel and/or cargo that preferably substantially participates
in an Earth launch as the transatmospheric upper stage for Earth
launchers, as disclosed in IAC 05-D2.3.08, "Geolunar Shuttle as
Upper Stage for Heavy Earth Launchers," 56.sup.th IAC, Fukuoka,
Japan, October 2005. This invention optionally comprises a
personnel shuttle with a suitable booster that can be air launched
from a horizontal takeoff subsonic aircraft such as introduced in
papers: 1) "Single-Stage Shuttles for Ground Launch and Air
Launch", Astronautics and Aeronautics, Vol. 12, No. 3, March 1974;
2) "Space Rescue and Other Space Missions from Existing Airstrips",
IAF-A-76-06, XXVI IAC, Anaheim, Calif., October 1976. The entire
contents of the aforesaid publications are incorporated herein by
reference.
[0035] The present invention preferably allows for direct flight
from upper stage suborbital ignition (i.e. during ascent to orbit)
to far space, and a return to Earth preferably utilizing a
horizontal landing, while bypassing the steps of injecting into and
ejecting from LEO. The invention also replaces the conventional
expendable upper stage booster. The start thrust-to-weight ratio
for the vehicle of the present invention is preferably less than or
equal to approximately 1, can be as low as about 0.25 and can be
the same for both an Earth launcher upper stage and a reusable far
space transfer shuttle. Thus, the same vehicle can perform both
functions, that is be the upper stage of an Earth launcher and a
far space transfer shuttle. The low thrust-to-weight ratio can be
achieved while still providing sufficient propulsion to reach far
space.
[0036] This novel self-launch, integration step combining the
functions of an expendable Earth ascent upper stage with those of a
far space shuttle, in the same vehicle, confers at least five new
benefits: 1) elimination of the costs incurred by an expendable
upper stage; 2) elimination of the performance penalties incurred
by injection into and ejection from LEO; 3) elimination of time
spent in LEO during which the far space shuttle is vulnerable to
simple inexpensive ground fire; 4) reduction of far space trip time
and associated life support and power weight requirements; and 5)
simplification, reliability and safety improvement for the overall
transportation mission.
[0037] As used throughout the specification and claims, "far space"
means any orbit, altitude, or earth escape beyond Low Earth Orbit
(LEO), including but not limited to Medium Earth Orbit (MEO), Low
Moon Orbit (LMO), Geosynchronous Orbit (GEO), or any Earth-Moon
libration or Lagrange point. The orbit may have any inclination,
eccentricity, or direction.
[0038] As used throughout the specification and claims, "direct
flight" means a propulsive flight that provides thrust and velocity
increase which starts from a suborbital condition (but above the
earth's surface) and proceeds directly to a desired far space
location.
[0039] As used throughout the specification and claims, "booster"
means a detachable portion of a launch vehicle, at least a part of
which portion is ignited either at the Earth's surface or from an
air launch platform such as an aircraft. A booster may comprise
more than one booster stage, where "stage" means a propulsive unit
that comprises both its own propellant and at least one engine. An
example of a booster is the booster portion of a heavy Earth
launcher.
[0040] The present invention is directed to a system including a
process and apparatus for the delivery and return of personnel and
cargo, including but not limited to high value, special mission,
ascent, or Earth return cargo, between the Earth and far space, and
preferably comprises a reusable propulsive vehicle which functions
as both the transatmospheric upper stage in an Earth liftoff and
ascent trajectory and as a far space shuttle. The vehicle
preferably comprises a reusable propulsive dart-shaped vehicle,
preferably comprising a high hypersonic lift-to-drag ratio of
approximately 3, enabling aerodynamic maneuverability which can be
used for timely deployment and manageable re-entry from far space
preferably to horizontal landing, e.g. on a runway, at any location
on Earth. Such a landing is typically, though not limited to, a
gliding landing. It is preferably compatible with any suitably
capable expendable or reusable Earth liftoff first stage booster.
Lifted to high altitude within the atmosphere preferably by the
first ignition booster stage(s), the contemplated vehicle can
continue directly to far space. The vehicle preferably comprises a
single upper stage preferably comprising an initial thrust-to-mass
ratio as low as about 0.25. The present invention may optionally be
air launched from a horizontal takeoff aircraft. In this
embodiment, smaller first ignition boosters may be employed. The
far space vehicle of the present invention may optionally be used
to transport personnel and cargo to and from LEO.
[0041] In one embodiment, the vehicle is designed to carry six
personnel with life support, environmental control and power
provisions for ten-day missions, and optionally comprises no
dedicated internal cargo bay such as in the first generation NASA
shuttle. Internal volume designed for personnel and their
provisions can be exchanged for other limited high value cargo. The
contemplated reusable vehicle incorporates internally its own
ascent propellants and may optionally carry external propellants in
expendable drop tanks to provide adequate overall vehicle
performance, as well as external pods to accommodate ascent-only
cargos.
[0042] In another embodiment, the vehicle, with reduced ascent
propellants of its own and a smaller launch booster, may be
employed as a shuttle for LEO missions, in which ascent-only cargos
may, for example, include propellants for high-velocity LEO and
other special missions.
[0043] The vehicle preferably incorporates and consolidates
performance, economic and operational advantages of two individual
system concepts: 1) combining propulsive elements (i.e. propellant
and engine systems) with mission elements (i.e. personnel, cargo
and control systems) into an integrated vehicle in place of
segmented vehicles, which confers performance gains and shared
structural and subsystem benefits (see "Comparison of Separate and
Integral Spacecraft," AIAA Journal of Spacecraft and Rockets," Vol.
6, No. 11, November 1969, incorporated herein by reference); and 2)
employing a slim, high-fineness-ratio (for example, between
approximately 7 and approximately 8) aerodynamic shape facilitates
direct superorbital re-entry from far space distances, using its
inherent high hypersonic lift-to-drag ratio to remain at high
altitude long enough to dissipate thermal energy until decelerated
to LEO velocity, from which final decent can be made within
established thermostructural technologies, and with long cross
range maneuver and horizontal landing (see "New Concept for Far
Orbit Transportation," IAF-80-F-243, XXXI IAC, Tokyo, Japan,
September 1980, incorporated herein by reference). Configuration
and aerothermodynamic features of such a propulsive high-fineness
vehicle are discussed in U.S. Pat. No. 5,090,692 and in "Geolunar
Shuttle: Earth Launch Options; Growth Using Lunar Propellants,"
IAF-01-V.3.09, 52.sup.nd IAC, Toulouse, France, October 2001, which
references are both incorporated herein by reference.
[0044] Because the vehicle of the present invention would typically
re-enter earth's atmosphere from far space, its re-entry velocity
will be closer to approximately 7 miles per second, rather than the
approximately 5 miles per second for vehicles re-entering from LEO.
Thus the present vehicle preferably comprises a slim dart
configuration, preferably comprising highly swept wings, as shown
in the figures. This provides a sufficiently high lift to drag to
optionally enable deceleration to LEO velocities if desired. Also,
this, combined with use of advanced structural materials, provides
resistant to the resulting higher re-entry temperatures. In
addition, a new re-entry strategy is required. (See, for example,
R. Salkeld, R. Beichel, and R. Skulsky, "A Reusable Space Vehicle
for Direct Descent from High Orbits," Astronautics and Aeronautics
19, Apr. 19, 1981, pp. 46, 47, and 53, incorporated herein by
reference.)
[0045] An upper stage personnel-only embodiment of the present
invention is shown in FIGS. 1A and 1B. Suggested locations for
H.sub.2 tanks 10 and O.sub.2 tanks 20 are shown. The vehicle
preferably comprises internal flight deck (IFD) 30 and airlock 40.
Table 1 below lists example cargo capacities, weight limits, engine
designs and re-entry data. FIG. 10 depicts an embodiment preferably
comprising one RL-60 or MB-60 engine 50 suitable for reaching Low
Moon Orbit and other far space shuttle capability beyond LEO; the
shuttle is carried by, for example, a .DELTA.IV Heavy Baseline
Booster. FIG. 1D depicts an embodiment suitable for reaching LEO
wherein the shuttle is carried by, for example, a .DELTA.IV Medium
Booster. In this embodiment, a RL10B-2 engine may be used in place
of the RL-60 or MB-60 engine, but the vehicle would then be able to
only transport less than four tons of cargo. Any other booster or
Earth launch concept, including but not limited to the Atlas V or
Ariane V ECB, may alternatively be employed.
TABLE-US-00001 TABLE 1 Parameter Earth-LMO-Earth Earth-LEO-Earth
Cargo, round trip Personnel (days) 6(10) 6(10) Other, (margin) lbm
1,600 9,300 Gross start weight, lbm 162,300 47,300 Dry weight, lbm
21,500 16,000 Engine 1xRL or MB-60 1xRL or MB-60 Re-entry planform
loading 18.4 18.4 (w/6 people), lbm/ft.sup.2 Re-entry cross-range,
n. mi. 4,500 4,500
[0046] While the previous embodiment is directed toward personnel
transportation with no dedicated cargo bay, dedicated cargo
carrying capacity, optionally combined with personnel
transportation, may alternatively be employed within the vehicle's
capability and design. Such a vehicle preferably comprises internal
dedicated cargo bay 60, preferably 12.times.30 feet in size, in
addition to the volume allocated for personnel transport, as shown
in FIGS. 2A-2E. FIGS. 2A and 2B show a configuration comprising
conventional flight deck (CFD) 70. Table 2 illustrates example
cargo capacities and other parameters of the vehicle of this
embodiment. FIG. 2C shows an alternate configuration comprising
internal flight deck 30, with the crew seated near the center of
the vehicle instead of at the front. FIG. 2D depicts an embodiment
suitable for reaching Low Moon Orbit (LMO) and other far space
distances, preferably comprising three RL10B-2 engines 80, wherein
the shuttle is carried by, for example, .DELTA.IV Heavy Derivative
Booster. FIG. 2E depicts an embodiment suitable for reaching LEO,
preferably comprising two RL10B-2 engines 80, wherein the shuttle
is carried by, for example, a .DELTA.IV Heavy Baseline booster in
either a tandem or parallel configuration. Any other booster or
Earth launch concept, including but not limited to the Atlas V or
Ariane ECB, may alternatively be employed.
TABLE-US-00002 TABLE 2 Parameter Earth-LMO-Earth Earth-LEO-Earth
Cargo, round trip Personnel (days) 2(10) 2(10) Other, lbm 20,700
30,200 Gross start mass, lbm 271,600 122,800 Dry mass, lbm 37,500
32,800 Engines (RL10B-2) 3 2 Re-entry planform loading 22.8 26.8
(RT cargo), lbm/ft.sup.2 Cargo bay, ft 12 .times. 30 12 .times. 30
Cargo density (RT cargo), 6.3 9.4 lbm/ft.sup.2 Re-entry
cross-range, n. mi. .+-.4,500 .+-.4,500
[0047] The present invention may optionally utilize subsonic air
launch to space using existing airstrips, which may have
performance, operational and economic benefits (see as introduced
in "Multiple-based Air and Ground Launch for Inspection, Rescue and
Other Space Missions," AIAA Journal of Spacecraft and Rockets, Vol.
6, December 1969, incorporated herein by reference). Current
commercial jet engines in the 100,000 lbf thrust class enable
development of subsonic aircraft weighing at least two million
pounds at takeoff (see "Super Duper Jumbo," AIR and SPACE
Smithsonian, Vol. 21, No. 2, July/August 2006, incorporated herein
by reference). Scaled configurations of the present invention in
conjunction with subsonic aircraft and launch platforms are shown
in FIGS. 3A-3D and 4A-4C.
[0048] FIGS. 3A and 3B depict personnel-only direct flight far
space shuttle 82 attached to .DELTA.IV Medium+(5,4) booster 85
mounted on a large subsonic air launch platform, which comprises,
for example, four GE90-115B engines 90 and optionally 2 or more
RD-180 engines 100, and is comparable in scale to the AN-225,
currently the world's largest aircraft, shown in FIG. 3D. Use of
this booster enables the shuttle to achieve LMO or other far space
orbits beyond LEO, for example via rocket assisted pullup with a
launch at about 60,000 ft. altitude and a 45.degree. flight path
angle. Table 3 contains non-limiting parameters for this
embodiment. FIG. 3C shows a subsonic platform carrying an alternate
mission pod.
TABLE-US-00003 TABLE 3 .DELTA.IV M + (5,4) Far space Vehicle Far
space Vehicle Parameter Aircraft Booster Earth-LEO-Earth
Earth-LMO-Earth Cargo, round trip Personnel (days) 6(10) 6(10)
Other, lbm 850,000 162,300 28,500 4,800 Gross start mass, lbm
1,896,300 672,200 162,300 162,300 Dry mass, lbm 601,300 71,900
21,500 21,500 Engine(s) 4xGE-115B 1xRS68 1xMB-60 1xMB-60 2xRD-180
4xGEM-60 RL-60 RL-60
[0049] FIG. 4A depicts personnel-only direct flight far space
shuttle 102 attached to .DELTA.IV "short" booster 105, which
comprises one RS-68 engine 110, mounted on a subsonic air launch
platform which comprises, for example, four JT9D-7J engines 120 and
one RD-180 100 engine, and is comparable in scale to the 777-200
aircraft, shown in FIG. 4C. Launch is preferably via rocket
assisted pullup with a launch altitude of about 60,000 ft. and a
45.degree. flight path angle. The .DELTA.IV "short" booster is a
.DELTA.IV Medium booster modified only by shortening propellant
tankage. Table 4 contains non-limiting parameters for this
embodiment. FIG. 4B shows a subsonic platform carrying an alternate
mission pod.
TABLE-US-00004 TABLE 4 ".DELTA.IV Short" Far space Vehicle
Parameter Aircraft Booster Earth-LEO-Earth Cargo, round trip
Personnel (days) -- -- 6(10) Other, lbm 280,000 47,300 2,300 Gross
start mass, 628,900 233,000 47,300 lbm Dry mass, lbm 229,600 27,900
16,000 Engine(s) 4xJT9D-7J 1xRS-68 1xMB-60 1xRD-180 or RL-60
Oxygen-Hydrogen Propulsion
[0050] The present invention can incorporate single-fuel
(oxygen-hydrogen or O.sub.2--H.sub.2) propulsion or dual fuel
(oxygen-hydrocarbon-hydrogen, or O.sub.2-MMH--H.sub.2) propulsion.
In the previous example embodiments, although oxygen-hydrogen is
depicted because the appropriate engines (RL10B-2, RL-60, and
MB-60) are in service or under development, any fuel may be
employed. In one embodiment, the dimensions of the shuttle vehicle
closely match those of the Delta IV Heavy upper-stage/payload
assembly, and both use the same RL10B-2 engine.
[0051] If the propellant loading of the contemplated vehicle is
controlled to match the start mass of the Delta IV Heavy
upper-stage/payload start mass, which may not be an optimal
condition, then estimating the subsequent vehicle ascent
performance can be performed for three modes of operation: 1)
ascent to LEO as a next generation LEO vehicle; 2) refueling in LEO
and ascent to and return from far space, as a far space vehicle;
and 3) direct ascent to and return from far space, from liftoff and
boost by an earth launcher, for example a 7-segment Delta IV Heavy
Derivative Earth launcher. Two possible vehicle configurations are
a Conventional Flight Deck (CFD) vehicle, shown in FIG. 5A, or an
aft-shifted control station and cargo bay or Internal Flight Deck
(IFD) vehicle, shown in FIG. 5B. For essentially the same vehicle
mold lines, the IFD option offers improved vehicle balance, more
efficient internal volume utilization, and increased propellant
capacity resulting in 30-40 percent increases in far space round
trip cargo to dry mass and cargo to gross mass ratios when fully
refueled in LEO. However, round trip LEO and direct flight far
space cargos are reduced 5-10% because of the assumed constraint on
the ascent start mass. Table 5 shows various parameter values for
these configurations.
TABLE-US-00005 TABLE 5 Conventional Flight Deck Internal Flight
Deck Delta IV Heavy Launcher Baseline Deriv. (7 Seg.) Baseline
Deriv. (7 seg.) Destination LMO LMO LMO LMO (LEO (Direct (LEO
(Direct LEO Refuel) flight) LEO Refuel) flight) Parameter Cargo,
round 13,700 7,600 9,400 13,200 10,700 8,400 trip, kgm Gross start
55,700 73,000 123,200 55,700 84,100 119,200 weight, kgm Dry Weight,
14,900 14,900 17,000 15,300 15,300 17,800 kgm Engines 2 2 3 2 2 3
(RL10B-2) Re-entry 119 94 110 108 108 119 planform loading,
kgm/m.sup.2 Cargo bay, m 3.7 .times. 9.1 3.7 .times. 9.1 3.7
.times. 9.1 3.7 .times. 9.1 3.7 .times. 9.1 3.7 .times. 9.1 Cargo
143 79 98 138 111 86 density, kgm/m.sup.3 Re-entry .+-.8,300
.+-.8,300 .+-.8,300 .+-.8,300 .+-.8,300 .+-.8,300 cross-range, km
Crew, if 2 2 2 2 2 2 required
Mixed Density-Impulse Propulsion
[0052] A mixed density-impulse propulsion principle may optionally
used by the vehicle, which enables the use of smaller engines. This
principle is also called "mixed-mode," "dual-fuel", or
"tripropellant." An oxygen-monomethylhydrazine-hydrogen
(O.sub.2/MMH/H.sub.2) orbit-to-orbit "tug," can attain round trip
geosynchronous cargo to gross mass and cargo to dry mass ratios,
21% and 47% respectively, higher than an O.sub.2/H.sub.2 tug of
equal LEO gross start mass. (See R. Salkeld and R. Beichel,
"Mixed-Mode Propulsion for Full Capability Space Tugs", AAS-75-162,
21.sup.st AAS Annual Meeting, Denver, Colo., August 1975,
incorporated herein by reference.) These benefits are realized even
when the tug is shortened by 40-60 percent. The O.sub.2/MMH/H.sub.2
engine design is described in FIG. 6A, which shows an example dual
fuel engine cycle schematic comprising heat exchanger 200, MMH
supply 210, LOX supply 220, LH.sub.2 supply 230, regenerative
cooling region 240, and radiation cooling region 250. Preferred
materials are indicated. FIG. 6B is a schematic of a dual fuel
O.sub.2/MMH/H.sub.2 engine producing a vacuum thrust of
approximately 20,000 pounds (LBF). Table 6 below compares engine
characteristics for different engine versions. A gas-gas
staged-combustion cycle can be adopted to allow stable dual-fuel
operation with a common injector and combustion chamber. The
O.sub.2/MMH/H.sub.2 propellant combination and engine design
described above can be incorporated into the present invention to
take advantage of the superior cargo to gross mass and cargo to dry
mass ratios.
TABLE-US-00006 TABLE 6 Dual Fuel Hydrogen Engine Basic (O.sub.2
Version Version Characteristics MMH) (O.sub.2/MMH/H.sub.2)
(O.sub.2/H.sub.2) Vacuum Thrust (lb) 20,000 20,000/13,500 13,500
Specific Impulse (sec) 393 393/469 469 Expansion Ratio 400 400 400
(A.sub.e/A.sub.t) Chamber Pressure 2,700 2,700/1,800 1,800 (psia)
Mixture Ratio (O/F) 1.70 1.70/7.0 7.0 Oxidizer Flow (lb/sec) 12.0
12.0/25.2 25.2 Fuel Flow (lb/sec) 18.9 18.9/3.6 3.6 Engine Weight
(lb) Fixed Nozzle 270 310 270 Rolling Nozzle 300 340 300
[0053] FIG. 7A depicts an O.sub.2/MMH/H.sub.2 CFD configuration.
The preferred location of MMH tank 260 is indicated. Because the
small size of the O.sub.2/MMH/H.sub.2 engine permits shifting the
cargo bay fully aft, the weight and complexity of side-loading
cargo bay doors may be eliminated and the overall vehicle size and
dry mass is reduced, resulting in an O.sub.2/MMH/H.sub.2 IFD
"Endloader" (EL) configuration, shown in FIG. 7B. Table 7 below
compares the CFD and EL configurations.
TABLE-US-00007 TABLE 7 "Conventional" Flight Deck "Endloader" Delta
IV Heavy Launcher Baseline Deriv. (7 Seg.) Baseline Deriv. (7 seg.)
Destination LMO LMO LMO LMO (LEO (Direct (LEO (Direct LEO Refuel)
flight) LEO Refuel) flight) Parameter Cargo, round 13,100 17,000
7,800 15,700 14,100 10,900 trip, kgm Gross start 55,700 110,300
121,200 55,700 91,500 123,300 weight, kgm Dry Weight, 15,500 15,500
16,500 12,700 12,700 14,500 kgm Engines 2 2 3 2 2 4
(O.sub.2/MMH/H.sub.2) Re-entry 119 135 101 119 112 119 planform
loading, kgm/m.sup.2 Cargo bay, m 3.7 .times. 9.1 3.7 .times. 9.1
3.7 .times. 9.1 3.7 .times. 9.1 3.7 .times. 9.1 3.7 .times. 9.1
Cargo 135 176 82 163 147 86 density, kgm/m.sup.3 Re-entry .+-.8,300
.+-.8,300 .+-.8,300 .+-.8,300 .+-.8,300 .+-.8,300 cross-range, km
Crew, if 2 2 2 2 2 2 required
[0054] Results indicate that for essentially the same vehicle mold
lines as the CFD configuration, the dual-fuel option discussed
above offers about 100% and 50% increases in far space round trip
cargo to dry mass and cargo to gross mass respectively, when fully
refueled in LEO, but 10-25% decreases for LEO and direct far space
flight because of the assumed limitations on ascent start mass and
staging velocity. However, because of the smaller vehicle size and
dry mass, compared to the O.sub.2/H.sub.2 CFD reference vehicle,
the EL dual-fuel design provides not only about 100% and 50%
increases in far space round trip cargo to dry mass and cargo to
gross mass respectively when fully refueled in LEO, but about 35%
and 10% increases in these ratios for round-trip LEO and direct
flight far space missions.
Shuttles on Heavy Launchers
[0055] In addition to the above possibility of using the
contemplated vehicle to replace high energy upper stage (HEUS) and
payload of Delta IV Heavy launchers, the present invention can
similarly be incorporated in other heavy launchers, such as a
hypothetical Atlas V Heavy, where it could replace the Centaur
upper stage and payload, with about the same performance as with
the Delta IV Heavy. Alternatively an embodiment of the present
invention could substitute for the B5-A H28 upper stage of the
flight-demonstrated European Ariane 5 ECB heavy launcher. Other
possibilities include using heavier launchers consisting of first
generation shuttle boost and other existing boost elements.
[0056] FIG. 8 shows liftoff weights and cargo capacities in tons
for five such launch options, all pictured with the single fuel CFD
configuration described above. The "Shuttle Launcher Derivative" is
similar to various NASA shuttle element combinations, which can
place 100 metric tons in LEO. The "SSTO" utilizes a development of
the 8.7 m diameter shuttle external tank with existing RD-180 and
Space Shuttle Main Engines (SSME) to enable an
oxygen-kerosene-hydrogen single-stage-to-orbit heavy launcher also
sized to place 100 metric tons in LEO. Once in LEO, any or all SSTO
engines can be returned for use, for example in a shuttle cargo
bay.
Far Space and Cargo Considerations
[0057] Transport operations to and from Earth-Moon libration points
L1 and L2 are potentially of interest for scientific and
operational reasons. Round-trip velocity requirement between LEO
and either L1 or L2 are about 1 km/sec less than between LEO and
LMO. Therefore, shuttle cargo capabilities are increased. For
example, for single fuel CFD and dual fuel EL shuttles as upper
stages for the Delta IV Heavy launcher, L1 and L2 round trip cargos
are about 15 tons and 23 tons respectively, compared with 8T and
17T for the LMO round trip. These gains, as well as the dual-fuel
advantage, are similar for the other four Earth launchers
considered.
[0058] The present invention preferably comprises reusable vehicles
that can serve as upper stages of existing or foreseeable Earth
launchers, can perform shuttle missions to low Earth orbits (LEO),
and with LEO refueling, or direct flight (with or without external
tanks) from heavier Earth launchers, to far space orbits such as
LMO or to Earth-Moon libration points (L1; L2). The present
invention may be configured to be compatible with known aerodynamic
and thermostructural technologies, existing rocket engines and
known engine technologies.
[0059] Use of internal control stations rather than conventional
cockpits, and dual-fuel (O.sub.2/MMH/H.sub.2) rather than
single-fuel (O.sub.2/H.sub.2) propulsion, lead to more efficient
vehicle volume utilization and increased system performance.
Specifically, dual-fuel propulsion can increase cargo to gross mass
and cargo to dry mass ratios by up to 50% and 100% respectively,
depending on operational mode, and make possible "Endloader" (EL)
configurations, thereby shrinking the vehicle and eliminating heavy
sideloading cargo doors.
[0060] The above disclosed embodiments are normalized to the same
cargo bay size, 3.7.times.9.1 m (12.times.30 ft.), or about 96
m.sup.3. An estimated range of LEO cargo masses of 10-15 tons
indicates average cargo densities of 104-156 kg/m.sup.3 (6.5-9.8
lb/ft.sup.3), which is in the upper range of what is commonly
considered realistic. However, for the estimated range of far space
cargo masses of 10-30 tons (perhaps larger, if lunar propellants
refueling is contemplated), cargo densities increase
unrealistically for the 3.7.times.9.1 m bay size. This suggests
that larger cargo bays may merit consideration, and that the EL
design may be advantageous. An example of an EL configuration with
an enlarged cargo bay of approximately 4.6.times.12.2 meters
(15.times.40 ft), depicted on a Delta IV Heavy Launcher, is shown
in FIG. 9. The round trip cargo for this configuration for LEO is
approximately 8.9 tons and is approximately 12.7 tons for LMO with
LEO refueling. The gross start mass is 106.2 tons, and the dry mass
is 18.7 tons. Thus the cargo volume is more than doubled, and the
cargo mass is decreased by the 4-5 ton increase in shuttle dry
mass. Alternatively, for one way up cargo, external cargo bays may
be employed.
[0061] FIG. 10 depicts an extended mission version of the present
invention, preferably comprising additional life support, internal
operations volume, and storable propulsion for extended missions
(e.g. up to about fifty days or more for a crew of four). Such
missions could include exploration visits to Earth-approaching
asteroids, Earth-Sun Lagrangian points, or servicing and repair of
distant astrophysical systems. Example parameters are shown in
Table 8. The vehicle of this embodiment preferably comprises
1.times.RL or MB-60 engine 400, operations bay 410 (preferably
about 18 feet in length), utility bay 420, and flight deck 430. In
this embodiment, post-escape maneuvers and Earth-return require a
total velocity gain of approximately 5900 ft/sec (1.2 km/sec),
which is preferably provided by two expendable propulsion pods
preferably using oxygen monomethylhydrazine propellants and the
O.sub.2/MMH engine 440 described herein. These propellants,
adequately insulated, are considered storable for this mission time
frame. Alternate storable propellant combinations such as
oxygen/methane may be used, but these may have different trade offs
of performance, cost, and operational benefits. As shown in FIG.
10B, this embodiment can be the upper stage of a .DELTA.IV Heavy
derivative using 6 GEM-60 solid rocket motors 320, as disclosed in
IAC 05-D2.3.08, "Geolunar Shuttle as Upper Stage for Heavy Earth
Launchers," 56.sup.th IAC, Fukuoka, Japan, October 2005,
incorporated herein by reference. This vehicle combination has a
gross liftoff weight of approximately 978 tons.
TABLE-US-00008 TABLE 8 EXPENDABLES CORE O.sub.2/H.sub.2 O.sub.2/MMH
Parameter VEHICLE TANKS (2) PODS (2) Cargo Personnel (4) 1000 -- --
Env. contr./life 9000 supp. (50 days) Mission equipment 2000 Gross
start mass, lbm 58,700 115,200 17,100 Dry mass, lbm 20,700* 5200
800 Engines 1xRL or -- 1xO.sub.2/MMH MB-60 Re-entry planform 20.3
-- -- loading, lbm/m.sup.2 Re-entry cross-range, km .+-.4,500 -- --
*Including 15% margin
[0062] FIG. 11A presents a direct space far space shuttle
embodiment which can serve as the upper stage for the projected
Ares V heavy Earth launcher. This embodiment preferably comprises a
crew of two 500 and external tanks 510. The vehicle can attain LMO
preferably carrying two separable propulsive elements: (1) crewed
moon lander/ascent vehicle 520 carrying the equivalent of four
people and one ton of mission equipment to and from the Moon
surface from LMO, which is preferably carried within the
12.times.30-foot internal cargo bay; and (2) unmanned expendable
moon lander 530 which can deliver 7.1 tons of cargo to the moon
surface from LMO, which is preferably carried externally on the
vehicle itself. The vehicle is preferably propelled by three RL or
MB-60 rocket engines, and each of the separable vehicles is
preferably propelled by one RL-10 engine (which comprises a nozzle
expansion ration .epsilon.=77). Specifications for this embodiment
are in Table 9. The vehicle is shown in FIG. 11B on an Ares V heavy
Earth launcher, with a gross liftoff weight of approximately 3700
tons.
TABLE-US-00009 TABLE 9 MOON ORBIT SHUTTLE WITH ARES V RESUABLE
EXPENDABLE LEO MOON MOON CARGO SHUTTLE CORE CREW EXTERNAL LANDER
WITH Parameter VEHICLE LANDER.sup.c TANKS (ONE WAY UP) .DELTA.IV
HEAVY Cargo 14,300 19,400.sup.c Personnel 6(12) 4(4) -- -- 2(10)
Env. contr./life supp. 7500 1000 -- -- 2000 Mission equipment --
2000 -- -- Gross start mass, lbm 220,800 36,100 463,100 28,500
157,400 Dry mass, lbm 48,600.sup.a 12,100.sup.b 23,200 4100.sup.b
41,800.sup.a Engines 3xRL or 1xRL-10 -- 1xRL-10 1xRL-10B-2 MB-60
Cargo bay, ft 12 .times. 30 11 .times. 11 -- (0-9) .times. 45.sup.d
12 .times. 30 Cargo density, lbm/ft.sup.3 10.6/3.6 4.8 -- 7.9 7.9
Re-entry planform 23 -- -- -- 25 loading, lbm/m.sup.2 Re-entry
cross-range, .+-.4,500 -- -- -- .+-.4,500 n. mi. .sup.aIncluding
15% margin .sup.bIncluding 10% margin .sup.cRound trip
.sup.dTapered
[0063] FIG. 11C shows a LEO version of this embodiment, which is
preferably propelled by 2.times.RL10B-2 engines and preferably
comprises cargo bay 330 (preferably 12.times.30 ft.). This version
can attain LEO (220 nautical miles, 28.7.degree.) as an upper stage
of the .DELTA.IV Heavy Earth launcher with a round-trip cargo of
the equivalent of two people and ten tons, and having a gross
liftoff weight of approximately 835 tons, as shown in FIG. 11D. The
last column of Table 9 includes parameters for this
configuration.
[0064] Although the invention has been described in detail with
particular reference to these preferred embodiments, other
embodiments can achieve the same results. Variations and
modifications of the present invention will be obvious to those
skilled in the art and it is intended to cover all such
modifications and equivalents. The entire disclosures of all
patents, references, and publications cited above are hereby
incorporated by reference.
* * * * *