Author: Hubert P. Davis, PE
Title: "5*Star Booster: Launch & Construction for
Space Solar Power"
Launch and construction of the massive space structures required for
Space Solar Power (SSP) requires careful attention to the inter-relationships
of these two functions. The size of the launch vehicle will be dictated
by the largest item which must be produced on Earth. A concept for
a reusable, winged launch vehicle intended to achieve these functions will
be presented along with a preliminary structural concept for the satellite
itself, and for its deployment and assembly in space.
Differing views of the means to launch the elements for the SSPI were presented by the 1979 NASA/DOE study and the 1996 study by SAIC and Futron. The 1979 approach was to perform a detailed "point design" of the 5,000 MWe satellite and to synthesize dedicated, series burn two stage winged fully reusable vehicles capable of 250 to 450 tonnes of payload per launch to place its elements and the supporting space infrastructure into low Earth orbit (LEO).
The 1996 NASA "Fresh Look" study depended upon use of the
prospective commercial single-stage-to-orbit vehicle, Venturestar, aided
by a ground-based electro-magnetic accelerator to provide 500 m/sec. of
the ideal velocity to reach orbit. The accelerator will increase
its payload capability from a nominal 11.3 tonnes to 25.6 tonnes per launch
to reach a 250 km circular orbit from the latitude of Kennedy Space Center,
Florida. The Venturestar payload bay will measure 4.6
x 4.6 x 13.7m.
THE SSPW MISSION MODEL
The vehicle described is tasked with providing primary transportation to deliver materials for eventual construction in the geo-stationary orbit (GEO) of 300 Gwe ( baseload) of solar power satellite generation. A preliminary "figure-of-merit" of 10 tonnes per Mwe, based loosely upon the 1979 study, was used to derive an interim estimate of the flights required to place the 300 Gwe satellite constellation. An "orbit burden factor" (OBF) of 1.26 was estimated to account for launch of the electric propulsion system, its propellants, and infrastructure needs, to derive a total of 38,000 flights of the nominal 100 tonne payload per flight space launch vehicle occurring during the 8 year placement interval, 2012 - 2020. Peak flight rate was 5,000 flights per year, or 16 flights per day, based upon a 305 day per year launch schedule.
Much of the mass of the SSP might possibly be obtained from the Moon,
perhaps reducing the number of flights by a factor approaching 10, but
with the "front end" burden of establishing the lunar infra-structure needed.
These data will be refined as the Workshop proceeds and additional information
THE SPACE LAUNCH VEHICLE
The location of the space launch facility was assumed to be near the
equator, on the premise that the gains in efficiency for the mission model
to be addressed made this investment worthwhile. The full benefits
of the Earth's rotational rate are gained by launching near the equator
and the need is eliminated to conduct a 28.5 degree plane change during
in-space transportation to GEO. The destination orbit was assumed
to be at 500 km altitude to decrease concerns about aerodynamic influences
upon construction of large tenuous structures in lower altitude orbits
and, perhaps, to encounter a less severe orbital debris environment.
Benefits were gained by deriving the Booster of this two stage launch vehicle from its predecessor, StarBooster(, a fully reusable liquid flyback booster (LFBB) by Starcraft Boosters, Inc., headed by Astronaut Buzz Aldrin, which has been defined by work accomplished over the past 2 1/2 years. StarBooster( houses a Zenit 1st stage, powered by a single RD-170 rocket engine, the world's most powerful rocket engine, delivering 1 GN-sec. of total impulse to its rocket stage(s) plus satellite payloads to a velocity of up to Mach 6.5. StarBooster is considered to be usable in the near term for economical placement of communications satellites to 6.5 tonnes or greater to the geo-transfer orbit (GTO) and to provide a demonstration article for the proposed LFBB for the Space Shuttle and perhaps for the USAF Spaceplane. It is planned that significant operational experience will be gained before onset of the Space Power program by the commercial use of StarBooster(.
5*StarBooster( uses five RD-170 rocket engines, and 1,594 tonnes of oxygen/kerosene propellants. It is equipped with four turbofan engines for subsonic cruise back to a runway landing at the launch site. Its weight at landing is 265 tonnes, compared to the Boeing 747-400F airliner's 302 tonnes. The primary construction material is aluminum, used as a "heat sink" during re-entry.
5*StarBooster( delivers a new fully reusable, winged hydrogen/oxygen Orbiter and its payload to an ideal velocity of 2.8 km/sec. This Orbiter bears a strong resemblance to its predecessor - - the Space Shuttle Orbiter, and is powered by five of the same engines - - the Space Shuttle Main Engines (SSMEs). Performance was calculated assuring that the loss of a single engine on both the Booster and Orbiter would not preclude safe return - - "engine-out" capability was provided in full
The hydrogen fuel for the Orbiter is carried in external tanks, conserving a large portion of the otherwise required body volume, reducing both weight and cost. It has a propellant capacity of 1,075 tonnes, an estimated dry mass of 108 tonnes, plus 18.4 tons of external hydrogen tanks, compared to the 160 tonnes empty mass of the Boeing 777-300 airliner. Almost 30 tonnes of orbit maneuvering propellant is carried to permit the circularization in LEO, de-orbit, and to allow up to a 0.1 degree plane change on arrival in orbit and for beginning re-entry for return to base.
The 12 hydrogen tanks form a circular pattern mounted on the nose of
the Orbiter to form the side walls of the payload fairing (PLF), using
a large nose cap to closeout the front end. The enclosed cylindrical
volume becomes the primary payload compartment: 25 m in length
and 9.1 m inside diameter.
Net payload of 93.5 tonnes is delivered to the 500 km circular orbit
from the near-equator launch site, plus the hydrogen tanks and PLF nose
cap: a total of 121 tonnes of potentially useful material, and some 4.65%
of liftoff mass, without use of accelerators. These sizable
structural elements can be used to construct the Space Solar Power satellites.
PAYLOAD STOWAGE FOR LAUNCH
5*StarBooster( stores the 154 tons of hydrogen fuel for the Orbiter
stage of this two stage to orbit (TSTO) vehicle in external tanks.
Additionally, rather than discarding these tanks on each mission, these
essential elements of the launch vehicle, after they have served their
primary purpose of containing the fuel for ascent flight, become integral
parts of the large Space Solar Power Satellite structure to be constructed
in orbit, providing the numerous compression struts needed.
These tanks were designed to be slender, with a diameter of 3.05 m (10
ft.), and 0.54 mm wall thickness, in order to use the existing tooling
and successful manufacturing techniques now in use for production of the
efficient thin-wall stainless steel propellant tanks of the Atlas and Centaur
These tanks each contain 12.8 tonnes of liquid hydrogen for use by the five Space Shuttle Main Engines (SSMEs) of the Orbiter. They have 189 m3 internal volume and are 27.1 m long, including the 45 degree ellipsoid end domes. Each of these tanks has dry mass of about 1.4 tonnes. To enable their use to support payload elements during launch and for effective subsequent use in the Space Solar Power satellite structure, a number of modifications and add-ons will be incorporated within the 15.7 m PLF envelope.
The principal add-ons are Structural Nodes at either end of the propellant tanks, which also serve multiple purposes: to securely mount the tanks to the Orbiter for ascent flight; to transfer the hydrogen fuel into the interior of the Orbiter for use by the SSMEs; to provide launch mounting for the payload elements; to house a tank stabilization system for use after arrival in orbit to assure the structural integrity of these tanks as long-lived structural elements; to house deployment mechanisms for the Photo Voltaic arrays (PVAs) carried as payload; to house the androgynous docking system which will be used to couple the tanks together in orbit; to provide multiple Remote Manipulator System (RMS) grapple fittings; to provide end support and deployment for stabilizing booms; to provide a minimal attitude control capability; and to mount a number of fittings for support of stabilizing cables and other secondary structure.
Internal and external modifications to the tank side walls will also be necessary; to provide end support for the PVA to be deployed in orbit; to externally support items of payload equipment including canister racks of Astromast( or similar extensible booms, attach closeout panels between the tanks; and to provide center supports for launching the stabilization booms stowed against the tank side walls during ascent flight. Doublers will also be required in these thin-walled tanks to distribute the loads produced by these external structures; it is possible that internal beams will also be required.
Each of the 24 PVA segments, constrained to fit in pairs within triangularly
shaped envelopes, is carried attached to a tank side wall in two rolls,
each 1.1 m diameter. Following separation of each tank in the 500 km altitude
assembly orbit, these rolls are rotated outboard and extended to
a length of 46 meters, 20% of their full 230 meter length (assuming
cell roll packing thickness of 3.5 mm - - the cells themselves are much
thinner). Deployment is by means of a set of four Astromast
(or similar actuators mounted on Structural Nodes. (Which also support
the ends of the tank for launch and are equipped with androgynous docking
LEO ASSEMBLY & DEPLOYMENT
Once the Orbiter has arrived in the 500 km assembly orbit with its payload, deployment and low Earth orbit (LEO) assembly operations will begin. Tank stability will first be assured, then the tanks with their attached payload segments will detach one by one from the PLF using the Orbiter Remote Manipulator System (RMS), docking them together to form a single gravity-gradient stabilized assembly of the 12 tanks. The twelve tanks of the launch, docked end-to-end in LEO, form an assembly 337 meters long, which when fully deployed in GEO extends to an array 464 m wide.
Once assembly of the twelve tanks is complete, each of the attached payload segments will be deployed. The two 1.1 m diameter, 25 m long reels of PVA attached to each tank and the four deployment mechanisms of each tank are rotated to their final deployment position. Next the racks of canisters for the extensible masts are positioned and communication paths to the deployment mechanisms established. Finally, the four stabilization booms per tank deploy and unfold to their full 25 m length. The array is thus launched in 24 segments per flight, two each attached to the interior of each tank. I have named it the Self-Erecting Solar Array, SESA(
Deployment of the PVAs will now take place, to a length of 46 m, about 20% of their full extension. As the canisters of extensible mast elements carried in the four deployment mechanisms are exhausted during deployment, new canisters are delivered from the racks and the mast elements attached to the now partially extended masts. This step will be repeated, using five of the 20 canisters provided for each mast. Cable stabilization will provide rigidity to the partially deployed PVAs. The resultant assembly, in gravity gradient stabilization mode in LEO, will be 97 m wide and 337 m long, and will have 8.7 MWe power output capability when facing the sun in LEO. This power will subsequently be used by a separately delivered electric propulsion system for transfer of the assembly to GEO. With the 93.5 tonnes payload capability of the 5*StarBooster and use of the 18.4 tonnes of hydrogen tanks not classified as payload, this assembly will have a mass of 112 tonnes.
If the packing density of the PVAs on the reels requires 3.5 mm per
layer, fully deployed length of each array is 230 m. (Full array
extension is deferred until arrival in GEO and subsequent further assembly.)
Once fully extended, each of these assemblies, now about 465 m wide and
337 m long, provides 138,000 m2 of area, which, if 24% efficient solar
cells are used, produce 43.7 MWe. If 75% of the payload mass
(70 tonnes) is allocated to the PVAs, the required cell figure-of-merit
is 624 W/ kg. A total of 23.5 tonnes plus the 18.4 tonnes of
tanks will be available for structure and other subsystems.
The number of these assemblies, each produced by a single flight of the 5*StarBooster HLLV, which are to be assembled in LEO for each transfer flight is yet to be determined, but is expected to be between two and ten. Definition of the electric propulsion system and the transfer mission will occur at a future date.
Ten of these assemblies are delivered to GEO and docked end-to-end to the growing GEO Space Solar Power Satellite, stabilized by cables interconnecting the 480 25 m long stabilization booms. They are mounted in pairs on opposing sides of a central 2.1 km diameter ring which will support the microwave power transmitters. This ring, itself, is composed of hydrogen tanks and stabilized by a system of cables. Each of these PVA "petals", fully deployed, is capable of producing 434 MWe output, at an array efficiency of 24%.
Ten launches of 5*StarBooster will provide a 3400 m x 465 m, 1120 tonnes, SSP module or "petal" in GEO, not including the launches required for the electric propulsion system, its propellants, other required subsystems, nor the microwave power transmission system.
At an end-to-end efficiency of 48% (DC output of the PVAs to AC
power on the terrestrial grid), 24 such modules, or "petals" are
assembled in GEO to complete a 5,000 MWe output SSP (on Earth to the terrestrial
grid). The completed SSP array will thus have an outside diameter
of 8.9 km and mass of 27,000 tons not including the transmitter(s),
yoke(s), attitude control systems, and other required features. Two
hundred forty launches of 5*StarBooster provide these major assemblies
of the 5GWe SSP in LEO.
TECHNICAL ISSUES & CONCERNS
Much work lies ahead to refine the design of this integrated space launch vehicle/solar power construction system , to assure its long term structural integrity in the presence of the several loading conditions and micro-meteorite hazards to be experienced, to thoroughly plan its complex operations, to derive highly credible mass properties, and to provide credible cost estimates.
An area of concern is the structural adequacy of the 3.4 km long "petals" using only a number of "docked" 3.05 m diameter hydrogen tanks as the central compression member. The NASA/DOE "Reference System" of 1977 provided a single planar solar array 5.3 km wide and 10.5 km long supported on a truss structure built in space 470 m deep - - much deeper than the design proposed here. This truss was made up of triangular cross section built-up beams with 7.5 m x 7.5 m "bays".
Elements of each beam were roll-formed in space from coils of aluminum strip stock 0.96 mm (.038 inches) thick and 60 cm (12 inches) wide, forming a triangular cross section, open on one side, structural element. 31.4 cm (12.4 inches) long per side.
Although the 3.05 m diameter stainless steel hydrogen tanks, when stabilized, clearly provide much more compressive strength than can the 0.31 m individual beam elements of the "Reference System", comparison of the hydrogen tank compressive strength with the 7.5 m box "bays" is less clear, as is comparison with the overall 470 m deep completed truss of the NASA/DOE design.
Supplemental stabilization means for these slender columns may be required, to be designed by analysis and confirmed by tests. Due regard must be paid in this refinement of the present design to the absolute necessity of conserving mass.
Grateful acknowledgment is made of the computer aided design support
provided by Mr. Scott Lowther and the helpful suggestions, encouragement,
and comments offered by Mr. Gordon Woodcock.