STARLIFTER PERFORMANCE ESTIMATESOrbiter/Booster length: 394.4 ft, wingspan: 385 ft (extended)Orbiter empty weight: 1,100,000 lbs, Booster empty weight: 900,000 lbs

Orbiter main tank volume (H2 & O2) = 226,018 ft^3Auxiliary tank volume (O2) = 13,823 ft^3 x 2 tanks per veh. = 27,646 ft^3Booster main tank volume (H2 & O2) = 481,951 ft^3

Total tank volumes: Orbiter: 253,664 ft^3 Booster: 509,597 ft^3(Assuming a 2.7061 ratio of H2 to O2 by volume) then:

Orbiter H2 volume: 185,219 ft^3, H2 weight: 820,520 lbs (at 29.3 psi)Orbiter O2 volume: 68,445 ft^3, O2 weight: 4,859,595 lbs (at 22 psi)Total Orbiter H2 & O2 by weight: 5,680,115 lbs

Booster H2 volume: 372,095 ft^3, H2 weight: 1,648,381 lbs (at 29.3 psi)Booster O2 volume: 137,502 ft^3, O2 weight: 9,762,642 lbs (at 22 psi)Total Booster H2 & O2 by weight: 11,411,023 lbs

Orbiter weight fueled (without payload): 6,780,115 lbsBooster weight fueled: 12,311,023 lbs ( x 2 boosters)

Using Tsiolkovsky’s Ideal Rocket Equation to predict performance: DeltaV = ISP (Specific Impulse) x Gravity (9.81 m/s^2) x Natural Log (Mass before/after)

NOTE: mass is being expressed in pounds not in kilograms as it should. This is ok in this instance because we are simply determining a before/after mass ratio to plug into the Natural-Log function. The ratio will be the same whether we use pounds or convert to kilograms.

For delivery of 300,000 lb (150 ton) payload to Low Earth Orbit (L.E.O.):1 st stage : Weight before launch: 2 boosters @ 12,311,023 lbs w/ fuel + fueled orbiter @ 6,780,115 lbs + 300,000 lb payload = 31,702,161 lbsWeight after 1st stage burn: 31,702,161- (10,612,251 fuel x 2 boosters) = 10,477,659 lbs (7% 1st stage fuel held in reserve for boost-back to KTTS = 798,772 lbs per booster)Estimated average Specific Impulse during ascent: 390 seconds1st stage DeltaV = 390s x 9.81m/s^2 x Natural-Log (31,702,161 / 10,477,659)1st stage DeltaV = 4,235.8 m/s

Booster boost-back DeltaV:Weight before fuel burn = Booster (900,000 lbs) + reserve fuel (798,772 lbs) = 1,698,772 lbs: Weight after fuel burn is the empty weight of the Booster (900,000) so:Boost-Back DeltaV = 390s x 9.81 m/s^2 x Natural-Log (1,698,772 / 900,000)Boost-Back DeltaV = 2,430.5 m/s (this is about 7.15 Mach enough to cancel the Mach 6 forward velocity, and generate Mach 1.15 velocity headed back to the launch site)

2 nd Stage : Weight before 2nd stage burn: fueled orbiter @ 6,780,115 lbs + 300,000 lb payload = 7,080,115 lbsWeight after 1st stage burn: empty orbiter @ 1,100,000 lbs + 300,000 lb payload = 1,400,000 lbsEstimated 2nd stage Specific Impulse: 428 seconds2nd stage DeltaV = 428s x 9.81 m/s^2 x Natural-Log (7,080,115 / 1,400,000)2nd Stage DeltaV = 6,805.3 m/s

1st Stage DeltaV (4,187.5 m/s) + 2nd Stage DeltaV (6,805.3 m/s) = 10,992.8 m/s

Total DeltaV (10,992.8 m/s) – Estimated Gravity Losses (1,620 m/s) – Estimated Aerodynamic losses (200 m/s) + Earth Rotation Boost gained by launching towards the east near the equator (KTTS in southern Florida) (200 m/s)

FINAL DeltaV with 300,000 lb payload: 9,372.8 m/s (9,000 m/s is the minimum needed to achieve Low Earth Orbit (L.E.O.))

Now let’s calculate for a 250,000 lb payload:

For delivery of 250,000 lb (125 ton) payload to Low Earth Orbit (L.E.O.):1 st stage : Weight before launch: 2 boosters @ 12,311,023 lbs w/ fuel + fueled orbiter @ 6,780,115 lbs + 250,000 lb payload = 31,652,161 lbsWeight after 1st stage burn: 31,652,161- (10,612,251 fuel X 2 boosters) = 10,427,659 lbs (7% 1st stage fuel held in reserve for boost-back to KTTS = 798,772 lbs per booster)

Estimated average Specific Impulse during ascent: 390 seconds1st stage DeltaV = 390s x 9.81m/s^2 x Natural-Log (31,652,161 / 10,427,659)1st stage DeltaV = 4,248.1 m/s2 nd Stage : Weight before 2nd stage burn: fueled orbiter @ 6,780,115 lbs + 250,000 lb payload = 7,030,115 lbsWeight after 1st stage burn: empty orbiter @ 1,100,000 lbs + 250,000 lb payload = 1,350,000 lbsEstimated 2nd stage Specific Impulse: 428 seconds2nd stage DeltaV = 428s x 9.81 m/s^2 x Natural-Log (7,030,115 / 1,350,000)2nd Stage DeltaV = 6,928.2 m/s

1st Stage DeltaV (4,248.1 m/s) + 2nd Stage DeltaV (6,928.2 m/s) = 11,176.3 m/s

Total DeltaV (11,176.3 m/s) – Estimated Gravity Losses (1,600 m/s) – Estimated Aerodynamic losses (200 m/s) + Earth Rotation Boost gained by launching towards the east near the equator (KTTS in southern Florida) (200 m/s)

FINAL DeltaV with 250,000 lb payload: 9,576.3 m/s (9,000 m/s is the minimum needed to achieve Low Earth Orbit (L.E.O.))

Examples of different real world concepts that were studied:BAC (British Aerospace Co.) MUSTARD – late 1950’s

General Dynamics Triamese (space shuttle concept) mid 1960’s

Typical mid 1960s TSTO – Two-Stage-To-Orbit mothership/orbiter concepts:

Lockheed Star-Clipper Single-Stage-To-Orbit with discarding tank 1960’s:

Repurposed Saturn-V 1st stage with orbiter & expendable tank as 2nd stage:

Buran vs US Space Shuttle 1980’s:

Evolved shuttle with crew abort system 1990’s:

Re-usable Energia (Buran core) 1990’s:

Venturestar 1990’s:

Boeing SPS delivery HLLV (Heavy Lift Launch Vehicle) concept 1970’s

Miscellaneous proposed German, Russian, Soviet, and US boosters and orbiters:

Examples of proposed Solar Power Satellite systems:

Design criteria for the Starlifter:*Can’t use altitude compensating nozzles: – Venturestar’s aerospike engines had a lower thrust-to-weight ratio than traditional bell nozzles. Because of this, I specified that traditional altitude-specific bells would be used in place of altitude compensating hardware. (However, I think that a low altitude insert in the 2nd stage bell that can be ejected at higher altitude would be an acceptable cheap and easy substitute).

*Has to use pressure vessel fuel tanks if cryogenic fuels are used: – The need for properly shaped fuel tanks that can hold pressurized fuel without leaking were the driving factor in the final fuselage shape. Subsequently, Starlifter was designed AROUND THE FUEL TANKS with the wings retracting into the fuselage, and tucking in under the tanks. – The Venturestar team attempted to develop special lobed fuel tanks that would fit into its lifting body shape. Engineers however, were never able to develop a suitable tank that would fit properly and hold pressure without leaking. As a result of this and other technical difficulties, that project was cancelled.

*Has to use existing facilities for launch and recovery if possible: - Starlifter would require construction of new launch facilities, but is capable of using the existing runway at KTTS for orbital recovery. Also construction of a new crawler is not needed, as the launch components can be towed individually to the launch pad. - A large single stage mothership is out of the question for an HLLV of this size, as it would be too large to use existing landing facilities.

*All parts must be re-used: – Boosters retain swing-wing, and landing gear of the orbiter and save some of their onboard fuel for a “boost-back” to the launch site.

*Has to have limited development costs: The main advantage of the Triamese over a Siamese or TSTO with mothership 1st stage, is that the Triamese is predicted to have lower development costs due to the fact that you are essentially designing one airframe, and then adapting it for use as a booster, and as an orbiter. The orbiter and boosters share many components, and would theoretically require less testing, development, and engineering versus the dissimilar Siamese.

*Must have limited environmental impact: Starlifter burns H2 and O2, of which the primary emission is H2O (water vapor) 2*H2 + O2 = 2*H20. The shuttle also burned hydrogen; however the solid rocket boosters created combusted aluminum byproducts which rained down on the Florida Everglades with every launch.

*Must have increased safety over previous systems: The crew module was designed to separate from the orbiter in the event of either a launch or orbital abort scenario. The crew module has an emergency use ablative heat shield for re-entry, and ballistic parachutes for descent. Solid rocket boosters have been eliminated adding another layer of safety.

Design compromises:*Starlifter cannot match a pure multistage rocket with throw away components in terms of liftoff weight, payload mass fraction, and overall launch efficiency. This is because Starlifter must drag its recovery hardware along for the ride. All of the heat-shields, wings, and landing gear needed to recover the components become dead weight during the ascent. This along with the added development costs, are the price that must be paid for system re-usability. However, the cost savings associated with recycling the hardware do begin to pay off when you are guaranteed to perform hundreds of launches as would be the case for SPS construction.

*Cost savings due to hardware recycling, and low turnaround time between launches is partially offset by the high amount of maintenance that has been historically required for re-usable launch vehicles. Some of these issues could be mitigated through the use of advanced low maintenance, and long lasting materials, such as the advanced metallic thermal protection system that was developed for Venturestar, along with carbon fiber, lithium-aluminum alloys, carbon nano-tubes, and even some of the advanced materials being developed for Orion.

*Retracting the wing into the fuselage allows for a smaller heat-shield, improved glide ratio with the wings extended, and lower drag / tighter packaging of the Triamese stack. Problems could arise if the wing mechanism jams or malfunctions post re-entry. The crew could be saved by the escape module, but the orbiter would be lost. Also some method of sealing the wing retraction slot would be necessary to prevent superheated plasma from entering the fuselage during re-entry. Fuselage rigidity is also compromised by the need to for reliefs to be cut into the fuselage bulkheads in order to make room for the wing to retract into the fuselage (see image).

*Another design compromise is that the payload bay is separated from the crew module by about 100’. Steve Wilson came up with the idea of installing a horizontal elevator trolley system to move astronauts back and forth between the crew module and the payload bay.

So how close is this to a realistic design? I’d like to think that it’s pretty close. The orbiter and boosters would need to meet the maximum weight requirements in order to meet performance predictions. Also the engines would need to meet minimum thrust requirements which are somewhat high for their size. A more realistic redesign might be to put 7 engines on the boosters, and 5 on the orbiter.

Special thanks to:Steve Wilson for scripting and flight testing.AI for his water tower models which were used in the launch facility.Lance Lobo for his sky diver, which I transformed into the MMU, Russian SPK, and Zero-G-ForkliftJohnny Forehead for his high-resolution parachutes.Mr. Scott for his superb liftoff sounds.Kovarik from turbosquid.com for his Caterpillar 797F heavy tow-truckMichael Carbagal and NASA for the space shuttle launch gantry which was downloaded and modified from nasa.govRobert Gaskell, “The Planetary Institute”, and the Japanese Space Agency JAXA for the 3D model of Itokawa AsteroidAustin Meyer and Laminar Research for providing us all with such a well stocked sand-box.

Possible future addon situations:*Delivery and assembly of mars bound habitat modules.*Assembly of habitat modules, and refueling depot modules onto the surface of Itokawa.*Assembly of deep space telescopes and sensors in the shadow of one of the SPS platforms