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The manned reusable space system which was designed to slash the cost of space transport and replace all expendable launch vehicles. It did neither, but did keep NASA in the manned space flight business for 30 years (and counting...) Redesign of the shuttle with reliability in mind after the Challenger disaster reduced maximum payload to low earth orbit from 27,850 kg to 24,400 kg. In the mid-1960's the US Air Force conducted a series of classified studies on next-generation space transportation systems. These were to reduce the cost of launching military payloads while supporting a projected robust manned military presence in space - including large space stations and reconnaissance and strike missions. These Air Force studies finally concluded that a partially reusable vehicle was the most attractive, epitomized by Lockheed's Starlifter, which had a large drop tank but returned the engines and avionics of the vehicle for reuse. The Air Force probably spent around $ 1 billion on 'black' technology development tests at this time, including work on linear aerospike engines and high fineness lifting body shapes that would re-emerge again 30 years later in Lockheed's X-33 space shuttle successor. NASA also had ambitious plans - for large space stations, lunar bases, nuclear interplanetary rocket stages, and manned Mars expeditions. NASA went through a long iterative process in designing and selecting the space shuttle, leading ultimately to the same conclusion as the Air Force. By mid-1969, the ambitious new NASA Administrator, Tom Paine, had proposed an extensive manned space exploration program as the logical follow-on to Apollo. A new, modular, reusable space transportation system would be required to set up bases on the Moon and Mars during the 1970s and 1980s. This system would consist of a reusable space shuttle to low earth orbit space stations and interorbital and interplanetary nuclear and chemical space tugs. The first major goal was a 12-man space station by 1975. NASA awarded $2.9-million study contracts to North American Rockwell and McDonnell-Douglas in July 1969. The space station was to evolve into a 50-man space base by 1980. Additional way-stations to Mars would be deployed in geostationary, lunar and Mars orbit during the 1980s. George Mueller headed the space shuttle portion of this effort, which accelerated as the Apollo project grew to a close. NASA awarded four $0.3-million space shuttle / Integral Launch and Re-entry Vehicle ILRV Phase A study contracts to North American Rockwell, McDonnell-Douglas, Lockheed and General Dynamics in January 1969. Martin Marietta's bid was rejected, but the company continued to participate using its own funds. The ILRV requirement was for a booster/spacecraft combination with 12-crew / 2.3 - 22.7 metric ton payload capability, a 720 km re-entry cross range, and first flight by 1974. The most important mission was expected to be space station resupply payloads weighing about 11,300 kg. 120 different permutations were investigated by the contractors. The assumption of a massive cost-is-no-object future space program was that only fully reusable vertical takeoff/horizontal landing, two-stage-to-orbit concepts for the space shuttle were considered at first. NASA's Shuttle task group had already calculated the potential life-cycle costs of three classes of 22,680-kilogram payload reusable launch vehicles based on prior USAF studies:
Space Shuttle Mission Model (mid-1969)
Space Shuttle Mission Requirements (mid-1969)
In August 1969, in post-moon landing euphoria, NASA directed the Phase A contractors to concentrate only on fully reusable shuttle concepts. These were two stages, both either winged or lifting bodies, and both recovered at the launch site for reuse. Only as an afterthought, some alternate concepts were still evaluated, including Lockheed's LS200 single orbiter with drop tank, and Chrysler's SERV ballistic single-stage-to-orbit vehicle. The Phase B designs were more refined but still used the same two-stage approach. Mueller set up a NASA space shuttle task group headed by LeRoy Day to evaluate potential uses of the vehicle. The shuttle requirements had changed considerably as a result of the new post-Apollo program which required a total of 546 shuttle launches in 1975-85. In May 1970, Mueller instructed the task group to increase the payload capability to 22,680 kg to comply with US Air Force requirements, but also because there would be a need to launch vast quantities of low-density rocket propellants into Earth orbit for future space stations in geostationary and lunar orbit. The mission requirements also grew significantly more complex and diverse as the Shuttle also now had to be capable of launching unmanned satellites and planetary probes. At this point a controversy developed over the basic design approach. There were over large cross-range winged designs, medium cross-range lifting body designs, and minimal cross-range stub-wing designs. NASA's Max Faget, who had dictated the spacecraft design for the Mercury, Gemini, and Apollo programs, advocated the stub-wing design. Then the Nixon administration burst NASA's balloon. The future NASA budget would be only a fraction of Apollo-program levels. There would be no moon bases, no flights to Mars, no nuclear interplanetary stages, no space stations, no more Saturn V's, no space tugs. There wouldn't even be a space shuttle unless NASA could get the development cost down and also convince the US Air Force to use the shuttle for its launch requirements. A USAF requirement was a large cross-range to allow recovery of the shuttle orbiter at the Vandenberg AFB launch point after a single polar orbit of the earth. This was necessary for abort-once-around, quick satellite deployment, strike, or quick-look reconnaissance scenarios. This, together with wind tunnel studies indicating that Faget's straight wing was unstable at re-entry speeds, drove NASA to the delta wing. The reduction in development cost led NASA to throw away the concept of reusing anything but the engines and guidance systems. Instead the shuttle would be boosted by cheap solid fuel boosters and, taking a concept from the Air Force, the propellants would be put in a big expendable drop tank. Following the usual charade of competitive bidding, NASA picked the same prime contractor as for X-15 and Apollo, who could be trusted to build precisely the vehicle NASA had in mind. North American Rockwell was selected to build the orbiter, with its Rocketdyne Division making the main engines. Thiokol was selected on political grounds for the solid rocket boosters. Martin Marietta would build the External Tank, but at the government Saturn IC factory at Michoud. To finance the Shuttle, already-built Apollo hardware that would have supported a second Skylab mission was sent to museums and American manned space flight went into a long hiatus in the 1970's. Budget cuts and overruns eventually reduced the number of shuttles built from five to four and delayed the first flight from 1978 to 1981 (thereby ruining the plan to save Skylab 1 on an early shuttle mission). But the development cost was indeed minimized - the shuttle ended up costing $ 6.744 billion in 1971 dollars, versus $ 5.15 billion estimated - less than a quarter of the Apollo program cost and a very modest overrun in comparison to some other programs. The pretext for the shuttle was that it would be much cheaper than expendable launch vehicles and would replace them all. Production was accordingly terminated by the US government of Delta, Atlas, and Titan vehicles. NASA staff and contractors were under incredible pressure to justify this decision by increasing the shuttle launch rate, lowering the turn-around time, and thereby reducing the cost per launch. When the shuttle Challenger exploded and the entire US space lift program was shut down for almost a year, the fallacy of this decision was exposed. The US Air Force and commercial users returned to use of expendable launch vehicles. When the shuttle began flying again, it was only for NASA programs. In the final analysis the shuttle came up short in three areas. First, the shuttle orbiter ended up almost 20% over its specified weight - resulting in it being unable to boost the US Air Force's payloads into polar orbits from Vandenberg. Lighter filament-wound casing Solid Rocket Boosters were being developed for use in flights from Vandenberg, but even this did not seem enough. After the Challenger explosion the USAF was able to extricate itself from the Shuttle program. The Vandenberg launch complex, built at the cost of billions, was mothballed. The Air Force started a new costly development program to design the Titan 4 expendable rocket for its large military payloads. The second shortcoming of the shuttle was that it failed utterly to reduce the cost of putting payloads into orbit. The shuttle program inherited from Apollo huge fixed costs - the Manned Spaceflight Center in Houston, the cadres of government and contractor workers at the Kennedy Space Center, and so on. The result was that there was a fixed base cost of around $ 2.8 billion per year, just to keep all those people and facilities in place, even if no flights were undertaken at all (as occurred after the shuttle disaster). The marginal cost of each flight added to this base was under $ 100 million. Seen this way the shuttle was almost competitive expendable boosters - but didn't come anywhere near the reductions NASA promised when development started. But if the usual number of flights per year was divided by the total annual costs, the cost per launch was $ 245 million, significantly more than a Titan or Proton launch with the same payload. The final shortcoming was that the shuttle was designed as if it had the inherent operating safety of an airliner. It was not equipped with any provision for crew rescue in case of booster failure during ascent to orbit, or being stranded in orbit, or structural failure during re-entry. The crew was not even provided with spacesuits, despite the lessons of the Soviet space program. This seemed an extraordinary act of engineering hubris, given that contemporary military aircraft were equipped with pressure suits and ejection seats. But the weight problem also meant that there was no margin for crew safety measures without (to NASA) unacceptable impact to the net payload. If the shuttle failed as a space truck, it succeeded in keeping America in the manned spaceflight business in the face of low public interest and political support. With the excuse of delivering payloads to orbit, NASA got to fly up to seven astronauts and run a host of supplementary experiments and payloads with each flight. With construction of the international space station beginning, NASA was looking forward to finally using the shuttle for its intended purpose. Due to the lower than planned flight rate, NASA's contractors were confident they could keep the existing shuttles flying through 2030. The real test came when (as was inevitable) another shuttle was lost. Following the Columbia disaster, NASA finally realized it could not make the shuttle safe. The only way to continue American manned spaceflight would be to develop a replacement manned spacecraft with an escape system, and meanwhile fly the shuttle as little as possible. NASA decided to complete the International Space Station in order to keep its international partners happy, then retire the shuttle by 2010. It was to be replaced by a modernized Apollo capsule, dubbed the Orion. The shuttle turned out to be a fifty-year detour to nowhere. When the Orion program started, NASA hoped to have the sort of lunar base by 2020 they would have had by 1980 if it had continued with Apollo rather than started the shuttle program. Yet even that dream receded further into the future, as Manufacturer: NASA. Launches: 120. Failures: 1. Success Rate: 99.17%. First Launch Date: 1981-04-12. Last Launch Date: 2007-10-23. Launch data is: continuing. LEO Payload: 24,400 kg (53,700 lb). to: 204 km Orbit. at: 28.50 degrees. Payload: 12,500 kg (27,500 lb). to a: space station orbit, 407 km, 51.6 deg inclination trajectory. Apogee: 600 km (370 mi). Associated Spacecraft: ACTS, AERCam, AFP-675, AS 4000, ASC, Atlantis, BremSat, Challenger, Chandra, Columbia, CRO, CTA, Discovery, DSCS III, DSP Block 14, Endeavour, ERBS, Eureca, Galileo, Galileo Probe, GLOMR, GRO, HS 376, HS 381, HS 601, HST, IAE, IBSS, Insat 1, IRT. Other Associated Spacecraft: ISS Unity, KH-12, Lacrosse, Lageos, LDEF, Magellan, Magnum, Mightysat 1, Mir-Shuttle Docking Module, MPEC, NUSAT, OAST-Flyer, ODERACS, ORFEUS, PAMS, PDP, SAC-A, SDS-2, Simplesat, Spacebus 100, Spacehab, Spacelab, Spartan, SPAS, SSF, Starshine, TDRS, TSS. Further Associated Spacecraft: UARS, Ulysses, WSF, NASDA Japanese Experiment Module, Transhab Module, Space Station Options 1993, Alpha Lifeboat, Spacedock, X-38, International Space Station, NASA ACRV, Space Station Fred, Industrial Space Facility. Liftoff Thrust: 25,751.600 kN (5,789,190 lbf). Total Mass: 2,029,633 kg (4,474,574 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft). Development Cost $: 10,100.000 million. in: 1977 average dollars. Launch Price $: 245.000 million. in: 1988 price dollars. Total Production Built: 5. Flyaway Unit Cost $: 63.000 million. in: 1988 unit dollars. Cost comments: Shuttle has high fixed costs and low marginal costs. Cost per mission dependent on rate. Flyaway cost is marginal cost for extra mission. Launch cost is cost per flight at 6 per year.
Shuttle ASRM.
Shuttle using Advanced Solid Rocket Motors (development cancelled 1993).
Liftoff Thrust: 28,193.000 kN (6,338,038 lbf). Total Mass: 2,100,293 kg (4,630,353 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft). Flyaway Unit Cost $: 829.000 million. in: 1985 unit dollars.
Shuttle ISS.
Redesign of the shuttle with reliability in mind after the Challenger disaster reduced maximum payload to low earth orbit from 27,850 kg to 24,400 kg. When the decision was made to move the International Space Station to a high-inclination 51.6 degree orbit, net payload to the more challenging orbit dropped to unacceptable limits. The situation was improved by introduction of the Super Lightweight External Tank, which used 2195 Aluminium-Lithium alloy as the main structural material in place of the 2219 aluminium alloy of the original design. This saved 3,500 kg in empty mass, increasing shuttle payload by the same amount. The tank was first used on STS-91 in June 1998.
LEO Payload: 27,500 kg (60,600 lb). to: 204 km Orbit. at: 28.50 degrees. Payload: 16,050 kg (35,380 lb). to a: space station orbit, 407 km, 51.6 deg inclination trajectory. Liftoff Thrust: 28,190.000 kN (6,337,360 lbf). Total Mass: 2,040,000 kg (4,490,000 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft).
Shuttle FR-3.
General Dynamics shuttle proposal phase A of October 1969. Unwinged flat-bottom configuration booster and orbiter with V butterfly-tails. General Dynamics received a $0.15-million Phase-A extension from NASA/Marshall to further study its shuttle concepts. The Triamese design was abandoned in September 1969 after more detailed analysis indicated that the development cost showed no considerable advantages vs. traditional two-stage systems. It had proven difficult to have one aerodynamic shape serve both as booster and orbiter; too many compromises had to be made. The designers then tried to "stretch" the Triamese orbiter but this, in turn, reduced the design commonality and hence cost savings. The company finally settled for a similar "FR-3A" two-stage design with V-tails to provide hypersonic roll control and aerodynamic stability. The booster/orbiter staging point was at 56.7km altitude and 193 seconds after launch when the vehicles were flying at 3325 meters/second. The straight-sided bodies were designed to accommodate cylindrical propellant tanks efficiently. Variable geometry switch-blade wings would still be used for subsonic flight and landing. This would have reduced the thermal protection system requirement (GD was the only company that still proposed a "hot structure" metallic TPS at the end of Phase A) . But the General Dynamics concept also had the same problem as MSC’s/North American’s DC-3 design since it had comparatively poor re-entry cross range capabilities Manufacturer: Convair. Liftoff Thrust: 30,243.000 kN (6,798,896 lbf). Total Mass: 2,557,880 kg (5,639,160 lb). Core Diameter: 11.19 m (36.71 ft). Total Length: 86.00 m (282.00 ft).
Shuttle LS A.
Lockheed shuttle proposal phase A of December 1969. X-24B lifting body orbiter with delta-wing booster. Lockheed’s $0.15-million shuttle Phase-A contract with NASA’s Marshall Space Flight Center concentrated on fully reusable versions of the Starclipper after NASA rejected all partially reusable concepts in August 1969. The final design consisted of a delta-planform lifting-body orbiter and a body-wing first stage booster. Lockheed also considered a "triamese" configuration of its lifting body design, but ultimately rejected the approach since the booster/orbiter propellant crossfeed system only produced a marginal performance advantage while introducing additional complexity and operational risk into the design. The orbiter’s large wingtip fins were expect to increase the vehicle’s subsonic lift-to-drag ratio and provide increased stability over the entire speed range. An additional plus was the lifting-body delta-wing concept would meet Air Force cross range requirements. Lockheed recommended a 22,680kg payload capability although a smaller 11,340-kg version also was investigated. Another important Lockheed contribution was a new thermal protection system made of silica fiber "tiles" which provided better insulation than metallic "shingles" and Lockheed therefore proposed to build the basic structure of aluminium rather than titanium. The final Space Shuttle design would use this approach and by the end of 1969, all Phase-A contractors except General Dynamics were proposing a silica-tile based TPS. The Lockheed Phase-A shuttle would have cost $5.51 billion (=$25B at 1999 rates) to develop. This cost included five booster/orbiter pairs for 175 horizontal and 25 vertical flights. Two boosters+orbiters would have been transformed into operational vehicles and Lockheed would have built five additional orbiters & two boosters for the fleet. The cost per flight would have been $1.255 million ($5.7M in 1999 $’s), and the inflation-adjusted transportation cost per kilogram was to be $251/kg over 1000 flights including R&D amortisation. Manufacturer: Lockheed. Liftoff Thrust: 28,463.700 kN (6,398,894 lbf). Total Mass: 1,632,012 kg (3,597,970 lb). Core Diameter: 10.00 m (32.00 ft). Total Length: 80.00 m (262.00 ft).
Shuttle MDC.
The McDonnell Douglas Space Shuttle Phase A studies were conducted under contract NAS9-9204. Their baseline Class III vehicle design was completed in November 1969 after 13 alternate configurations had been considered. The two-stage-to-orbit vehicle had a gross mass of 1,550,000 kg and a 11,300 kg payload was accommodated in a 4.6 m x 9.2 m payload bay. Both the booster and orbiter would be towed horizontally to the pad, then tilted vertically and mated on the pad. The entire vehicle would lift off vertically, with the orbiter separating at 64 km altitude and 9800 kph. This staging velocity was slightly lower than that of the North American Phase A concept since the McDonnell Douglas orbiter had a slightly higher propellant mass fraction due to its lightweight design. The booster stage was a delta-winged vehicle, 59.5 m long, powered by ten Pratt & Whitney 188,000 kgf engines. The 15% thick wing included fuel tanks for the JP-4 fuel for the turbojet fly back engines. The liquid oxygen tank was positioned forward to minimise engine gimbal requirements during boost. The heat protection system consisted of titanium structural panels, with Rene-41 shingles on the hottest areas. The orbiter was a lifting body using the HL-10 shape, 32.6 m long, powered by two 188,000 kgf engines, with the payload bay at the centre of gravity. A two-man crew would fly the orbiter. The orbiter's main engines would shut down at an altitude of 80 km, putting the vehicle into an itinital 80 x 185 km orbit. They would be restarted at 10% thrust at apogee half a world away, circulising the orbit at 185 km. The liquid oxygen tanks were forward, and three liquid hydrogen tanks aft. All the tanks were integral with the orbiter structure. Re-entry would be at a -1.5 deg flight path angle, at 50 deg angle of attack held until the lower fuselage reached 1200 deg C. Then the lifting body's bank angle controller would turn the vehicle to ensure that the 1200 deg C temperature would not be exceeded. In order to achieve the 725 km maximum cross range specified, the lifting body would bank after re-entry but while still travelling between 7400 and 4500 kph. The all-metallic thermal protection system used materials as appropriate to the heating rates on each portion of the external surface - titantium, Rene-41, nickel-chromium, with columbium-752 shingles on the hottest areas. For ferrying of the orbiter on earth, a special wing with 2 x 10,000 kgf TF39 turbofans would be bolted into the payload bay. McDonnell Douglas estimated the total development program would cost $ 5.946 billion and be completed in 21 months (!) This would include production of two orbiters and two boosters. For operations a fleet of three orbiters and two boosters would be used. Assuming 12 flights were made per year, delivery cost to orbit would be $54 per kg. At a 100 flight per year rate, this would drop to $30 per kg. The preferred launch site was McConnell AFB, Kansas. Manufacturer: Douglas. Liftoff Thrust: 19,213.400 kN (4,319,344 lbf). Total Mass: 1,578,231 kg (3,479,403 lb). Core Diameter: 8.00 m (26.20 ft). Total Length: 71.00 m (232.00 ft). Flyaway Unit Cost $: 48.000 million. in: 1985 unit dollars.
Shuttle MDC A Alternate.
McDonnell-Douglas shuttle proposal phase A of November 1969. Delta wing first stage and HL-10 lifting body second stage. McDonnell-Douglas (McDAC) decided not to pursue the Langley HL-10 design any further and instead proposed a number of comparatively low-cost straight wing orbiter designs. Typical was the company’s "drawbridge wing" shuttle orbiter from late 1969. It would have re-entered with folded wings for high-crossrange (2800km) military missions. The cross range with extended wings would have been only 400km but the re-entry heat loads also would be less severe. The gross lift-off weight of this system was 1,587.5 t and McDonnell-Douglas estimated it would cost $6.5 billion to develop in 1969 dollars (=$29.5B at 1999 rates). Marginal cost per mission would be: $3.5-4.5M (at 1969 rates; $1000-$1250/kg in 1999) assuming 100 reuses of each vehicle. Besides the primary drawbridge shuttle concept, McDonnell-Douglas proposed a smaller alternative orbiter design derived from the 1968 ILRV concept. Its 4.57 * 18.3-meter cargo bay would contain a small propellant tank on some missions. It could then be capable of deploying 9,072-kilogram 7.3 meter long payloads in a fully reusable mode. 22,680-kilogram payloads occupying the entire cargo bay could still be launched if expendable external drop tanks were used. The empty tanks could be returned to Earth inside the cargo bay on some missions, or alternatively be discarded. This system actually would have been only marginally more expensive to operate ($4-5.5M/flight vs. $3.5-4.5 million for the drawbridge orbiter concept). The specific launch cost would be less, since the drop-tank orbiter would carry a significantly larger payload than the drawbridge version. The drop tank orbiter would have used a scaled-down version of the drawbridge configuration booster. This smaller system would have cost $5.5 billion ($25B at 1999 rates) to develop. McDAC also proposed a simpler and less expensive unmanned tow-back winged booster to further reduce the weight and cost the booster as well as the entire system, down to $4 billion (=$18 billion in 1999 $'s). The booster would have been snagged by a C-5A Galaxy following re-entry and then towed back to base. McDAC also investigated even simpler and cheaper interim recoverable and expendable ballistic boosters. Manufacturer: Douglas. LEO Payload: 19,958 kg (43,999 lb). to: 555 km Orbit. at: 55.00 degrees. Liftoff Thrust: 19,213.400 kN (4,319,344 lbf). Total Mass: 1,600,727 kg (3,528,998 lb). Core Diameter: 8.00 m (26.20 ft). Total Length: 71.00 m (232.00 ft). Flyaway Unit Cost $: 48.000 million. in: 1985 unit dollars.
Version: Shuttle NAR A.
North American's Phase A shuttle design was completed under contract NAS9-9205 in December 1969. North American had learned that the way to win a NASA design competition was to adhere to the design favoured by Max Faget, so they proposed a two-stage-to-orbit vehicle, with both booster and orbiter being of Faget's straight-wing, low cross-range configuration. Faget disliked the Lockheed Starclipper and other lifting-body designs since they had poor low-speed handling characteristics and would be difficult to develop since the structure was tightly coupled with the aerodynamics. Faget preferred a simple winged design. His solution was dubbed the 'DC-3' and sought to alleviate the problem of re-entering at a 60 deg angle of attack, essentially accomplishing a ballistic re-entry like an Apollo capsule. This would only expose the flat underside of the vehicle to high heating rates, as most of the thermal energy would go into the shock wave forming in front of the vehicle. The high drag also shortened the duration of the heat pulse, yet did not exceed acceptable crew deceleration load factors beyond 2 g's. The DC-3 wing would only be optimised for subsonic flight and landing. But the low lift-to-drag ratio re-entry profile advocated by Faget would limit the DC-3's cross range to 430 km, far less than the USAF requirement. The gross lift-off mass of North America's elaboration of Faget's concept was 2,030,000 kg. Both booster and orbiter would be towed horizontally to the launch complex, and be lifted vertically and mated on the pad. The booster was 85.4 m long, with a 2219-Al aluminium structure, with a thermal protection system consisting of Lockheed LI-1500 tiles with a density of 63 kg/sq m. The wing had a span of 74.4 m and was built of 6A1-4V titanium. 11 x 230,000 kgf Rocketdyne Lox/LH2 engine provided boost at lift-off. 4 x JT9D-15 turbojet engines, fed by 26,200 kg of fuel, provided a 500 km cruise back range. The booster was unmanned during space missions, but there were provisions for two crew to fly the aircraft during ferry flights. The orbiter was 61.6 m long, and had a 263 sq m wing with a span of 44.5 m. Two crew plus ten passengers could be accommodated in the orbiter, plus 5,580 kg cargo in a 4.6 m x 18.3 m payload bay. The orbiter was equipped with 2 x 231,000 kgf Rocketdyne Lox/LH2 engines. It was anticipated that improved engines would increase the rated payload to orbit by 4500 kg. For atmospheric cruise, 4 x JT3D-7 jet engines, rated at 8600 kgf each, were mounted in nacelles on the upper wing surface. Internally, the payload bay was placed at the centre of gravity, with the two liquid hydrogen tanks aft, and the liquid oxygen tank forward. The propellant tanks were suspended within the orbiter structure and made of 2219-Al aluminium. The structure was built of 6A1-5V titanium, and the thermal protection system consisted of LI-1500 tiles (with a density of 63 kg/sq m), developed by Lockheed. The orbiter would separate from the booster at 61 km altitude and 3,300 m/sec velocity. The orbiter's main engines would fire until the spacecraft was placed in a minimum 92 x 185 km orbit. It was expected that orbital operations - circularisation, rendezvous and docking with a space station in a 480 km circular orbit, deorbit - would take 370 m/sec delta-V. The orbiter was designed to accomplish a total 600 m/sec delta-v - a 60% margin. Expected re-entry heating rates on the orbiter were 1650 deg C on the leading edge, and 790 deg C over 80% of the lower surface. North American estimated their space launcher would take 4.5 years to develop and productionise, at a cost under the $5.912 billion ceiling set by Faget. In order to fly 50 missions per year, it was estimated that six vehicles would be required, together with 200 maintenance staff. Manufacturer: North American. Liftoff Thrust: 24,948.000 kN (5,608,533 lbf). Total Mass: 2,036,734 kg (4,490,229 lb). Core Diameter: 9.88 m (32.41 ft). Total Length: 102.00 m (334.00 ft).
Shuttle DC-3.
Marshall Spaceflight Center shuttle concept of April 1970 using Faget low cross range stub-winged booster and orbiter. Payload for the Faget vehicle was to be only 5,700 to 6,800 kg to low earth orbit, and the system was to be operational by the end of 1975, after the last Apollo flight. MSC-001 was the first in a long series of Faget straight-wing designs. It featured an 2.4 m x 9.1 m payload bay, and a cross-range of just 300 km. The orbiter was equipped with two booster engines (XLR-129 modifications with 134,700 kgf), 2 orbital manoeuvring engines (RL10's with 6800 kgf), and six air-breathing engines (RB162-86's of 2,400 kgf burning JP-4 jet fuel). The booster would be equipped with booster engines, and Pratt and Whitney TF-B turbofan engines of 8,100 kgf for flyback. The launch scenario was for the booster to take the orbiter to altitude, release it, and then land at a down-range airfield. It would be refuelled there with jet fuel and fly back on its turbofan engines to the launch site. The orbiter had a 27.7 m wingspan with a 14 deg wing leading-edge sweep. The aluminium structure was protected by a silica-based thermal protection system. The booster would be 61.9 m long, have a 43 m wingspan with the same 14 degree sweep, and a total wing area of 264 square m. The leading edge would be protected by a pyrolised carbon laminate, and the lower surface by a silica-based thermal protection system. The payload would be delivered into a 500 km orbit at a 51 deg inclination. It was expected a fleet of six orbiters and four boosters would undertake 30 flights per year, each spacecraft having a life of 100 flights. It was expected a 48-hour reaction time between order for launch and launch would be possible. Total development cost of the orbiter was estimated as $2.77 billion, with the first article costing $171.2 million. The booster would cost $3.142 billion to develop, with a first article cost of $236 million. LEO Payload: 5,700 kg (12,500 lb). to: 500 km Orbit. at: 28.00 degrees. Liftoff Thrust: 11,433.200 kN (2,570,286 lbf). Total Mass: 998,775 kg (2,201,921 lb). Core Diameter: 5.08 m (16.66 ft). Total Length: 74.00 m (242.00 ft). Flyaway Unit Cost $: 38.000 million. in: 1985 unit dollars.
SERV.
Chrysler ballistic single stage to orbit alternate shuttle proposal of June 1971. This was the most detailed design study ever performed on a VTOVL SSTO launch vehicle. The 2,040 tonne SERV was designed to deliver a 53 tonne payload to orbit in a capacious 7 m x 18 m payload bay. The Chrysler SERV single-stage-to-orbit ballistic vehicle was the subject of a six-volume report produced under the $ 1.9 million NASA contract NAS8-26341. The booster could be launched from the existing LC39 built for the Saturn V. SERV would be built at NASA's Michoud facility and transported by a 'Bay'-class vessel modified to carry the wide load through the existing inland waterway system between Michoud and Cape Canaveral. SERV was a squat 27.4 m in diameter and 20.3 m tall. A payload of 52,800 kg, housed in a 7 m x 18.3 m cargo bay, could be transported to a 185 km/28.5 deg orbit. The vehicle was powered by a 12-module aerospike engine, 26.6 m in diameter and 2.5 m tall, producing 2.45 million kilograms of thrust at a specific impulse of 347 seconds at lift-off. The engine could be throttled to 80%, and the turbopumps were interlinked, so that the failure of any one pump could be compensated for by bringing the others up to 120% of their rated capacity. Protective doors covered the engine during the base-first re-entry, which would be accurate enough to bring the booster to within 6500 m of the aim point.. After slowing to subsonic speed,. 28 x 11,400 kgf turbojet engines powered by JP-4 fuel braked the spacecraft to a hover and soft touchdown on landing pads 2.8 km from the Vertical Assembly Building at the Kennedy Space Center. For manned missions, a MURP spaceplane would be used for separate return of the crew to earth. Total development costs was estimated as $3.565 billion, with each SERV costing $350 million in FY1971 dollars, and being rated for 100 flights over a 10 year service life. As had Philip Bonob at Douglas before them, the Chrysler team, led by Charles Tharratt, fervently believed that they had the best solution to providing America with routine access to space. But NASA was wedded to the concept of a winged shuttle and never gave SERV any serious consideration. Manufacturer: Chrysler. LEO Payload: 52,800 kg (116,400 lb). to: 185 km Orbit. at: 28.50 degrees. Liftoff Thrust: 25,795.300 kN (5,799,014 lbf). Total Mass: 2,040,816 kg (4,499,229 lb). Core Diameter: 27.40 m (89.80 ft). Total Length: 20.30 m (66.60 ft). Development Cost $: 3,565.000 million. in: 1971 average dollars. Flyaway Unit Cost $: 350.000 million. in: 1971 unit dollars.
Shuttle H33. Status: Study 1971. Grumman/Boeing alternate shuttle proposal of July 1971. Shuttle orbiter with drop tanks, delta booster. On 29 December 1970 Grumman and Boeing received contract NAS9-11160 to study two-stage-to-orbit shuttle configurations using both internal and external liquid hydrogen tanks. Reviews with NASA in January and March 1971 showed there could be significant weight, risk, and cost reductions through use of a booster with a heat-sink airframe and an orbiter equipped with an external liquid hydrogen tank. In April 1971 NASA authorised the contractors to make a detailed study of the most promising configuration - a three-engine orbiter with an external liquid hydrogen tank and heat-sink booster. Staging would be at 7660 kph instead of the 11,000 kph optimum for the all-recoverable two-stage-to-orbit designs. This lower separation velocity meant a smaller booster requiring less thermal protection and using less JP-4 fuel for the return to base. The team quickly established that using 2 x 250,000 kgf engines on the 405 tonne orbiter resulted in unacceptable abort constraint. Therefore 3 x 188,000 kgf baseline Phase B engines would have to be used. Configurations used in the trade study were the H-33 three-engine orbiter with an external tank and the G-3 three-engine fully-reusable orbiter. The H-33 orbiter used 3 x 188,000 kgf main engines and 4 x JTF22A-4 turbofans. The orbiter's 55 deg sweep delta wing provided a cross range of 2040 km and a landing speed of 333 kph. Ferry range would be 556 km using the jet engines. The orbiter weighed 89,300 kg empty and had two internal liquid oxygen tanks with a capacity of 320,044 kg below the payload bay. Two external liquid hydrogen tanks, each 31 m long and 4.5 m in diameter, held 2 x 27,000 kg of the low-density fuel. The orbiter used titanium as its basic structural material. The H-33 booster weighed 99.835 kg empty. It was equipped with 12 SSME and 12 x GE F101 or P&W JTF22A-4 turbofan engines, which gave it a 715 km ferry range. The 10 m diameter fuselage used Saturn S-IC propellant tanks, mounted in a hot structure that required no thermal protection system. R&D was estimated at $ 2.674 billion for the orbiter, $ 32.6 million for the external tanks, $ 2.181 billion for the booster, and $ 893 million for flight test and project management, a total of $ 5.878 billion. This did not include $1.0165 billion of government furnished equipment (primarily engines). Grumman estimated that production orbiters would cost $615 million each; boosters $ 274 million; and each pair of expendable external tanks, $740,000. The marginal costs for each H-33 flight were calculated as $ 4.2 million. Given a March 1972 go-ahead, first orbital flight could be accomplished by April 1978. The team estimated 3,000 people would be needed to support a launch rate of 75 flights per year, working out to $540,000 per flight costs for manpower. Manufacturer: Grumman, Boeing. Liftoff Thrust: 22,148.000 kN (4,979,068 lbf). Total Mass: 1,963,916 kg (4,329,693 lb). Core Diameter: 10.00 m (32.00 ft). Total Length: 90.00 m (295.00 ft).
Shuttle HCR.
McDonnell-Douglas/Martin Marrietta shuttle high cross-range proposal phase B of December 1970. Swept wing booster, delta wing orbiter. The McDonnell Douglas/Martin Marrietta Phase B shuttle proposal was designed under Contract NAS9-100959. The team proposed that the originally requested 6,800 kg payload would be achieved in the prototype, with the revised 11,500 kg payload to be obtained by making changes to lighten the orbiter as a result of development findings. The high cross range orbiter had a 92,300 pound empty mass and the same engine configuration as the low cross range orbiter (2 x 188,000 kgf engines and 4 x 8,200 kgf turbofan air-breathing engines). But the aerodynamics gave it a 2784 km cross range and a 300 kph landing speed. It also had a titanium structure enclosing aluminium propellant tanks, but the thermal protection system was made of a metallic cobalt super alloy, with columbium leading edge panels. The booster for both versions was the low cross range configuration design used by the same contractors in Phase A. This had a high wing to the rear, and canards forward. The empty weight of the low cross range version was 205,353 kg. The10-m body diameter was intended to allow use of existing Saturn V tooling to fabricate the propellant tanks. The high cross range booster was slightly larger and heavier to accommodate the heavier high cross-range orbiter. A number of expendable upper stages were considered that could replace the manned orbiter for cargo missions. These included:
The team also examined the costs to establish launch facilities. A Kennedy Space Center shuttle launch facility, taking advantage of existing Saturn V and Apollo infrastructure, would cost $ 87 million and take five years to build. One at Vandenberg AFB would cost $ 285 million, and one at White Sands would cost $ 317 million and take seven years to build. Manufacturer: Douglas, Martin. Liftoff Thrust: 25,839.700 kN (5,808,996 lbf). Total Mass: 1,977,415 kg (4,359,453 lb). Core Diameter: 10.00 m (32.00 ft). Total Length: 85.00 m (278.00 ft).
Shuttle LCR. Status: Study 1969. McDonnell-Douglas/Martin Marrietta shuttle low cross-range proposal phase B of December 1970. Swept-wing booster, Faget straight wing orbiter. The McDonnell Douglas/Martin Marrietta Phase B shuttle proposal was designed under Contract NAS9-100959. The team proposed that the originally requested 6,800 kg payload would be achieved in the prototype, with the revised 11,500 kg payload to be obtained by making changes to lighten the orbiter as a result of development findings. The low cross range orbiter was 113.8' WS x 147.6' x 21', a Faget configuration with a dry empty mass of 85,500 kg, equipped with 2 x 188,000 kgf engines and 4 x 8,200 kgf turbofan air-breathing engines. It had a cross range of only 370 km (less than an Apollo capsule) and a landing speed of 330 kph. The structure was built of titanium, with a columbium thermal protection system and carbon-carbon leading edge panels. The booster for both versions was the low cross range configuration design used by the same contractors in Phase A. This had a high wing to the rear, and canards forward. The empty weight of the low cross range version was 205,353 kg. The10-m body diameter was intended to allow use of existing Saturn V tooling to fabricate the propellant tanks. The high cross range booster was slightly larger and heavier to accommodate the heavier high cross-range orbiter. A number of expendable upper stages were considered that could replace the manned orbiter for cargo missions. These included:
The team also examined the costs to establish launch facilities. A Kennedy Space Center shuttle launch facility, taking advantage of existing Saturn V and Apollo infrastructure, would cost $ 87 million and take five years to build. One at Vandenberg AFB would cost $ 285 million, and one at White Sands would cost $ 317 million and take seven years to build. Manufacturer: Douglas, Martin. Liftoff Thrust: 23,994.000 kN (5,394,065 lbf). Total Mass: 1,834,830 kg (4,045,100 lb). Core Diameter: 10.00 m (32.00 ft). Total Length: 90.00 m (295.00 ft).
Shuttle LS200.
Lockheed Skunk Works alternate shuttle proposal of June 1971. X-24B lifting body orbiter with wrap-around external tank. Lockheed designed a version of their earlier 1.5 stage-to-orbit 'Starlifter' concepts meeting NASA requirements. Two versions were proposed: an LS-200-10, which would be a 1.5 stage to orbit vehicle, and the LS-200-8 and -5, which could later be uprated to the LS-400-7A two-stage-to-orbit. The Lockheed designs were all high-fineness lifting bodies derived from their 1968 LSC-8MX design for the USAF ILRV requirement. The 47.7 m long orbiter was surrounded at launch by a 58.5 m high x 8.23 m diameter V-shaped drop tank. The advantages of the design were that only a single complex vehicle needed to be designed, built, flight tested and maintained. At the same time, if at a later time an all-recoverable two-stage-to-orbit launcher was needed, the already developed and paid for orbiter could serve as the second stage of such a system. Lockheed also maintained that their delta-body design would be 9,000 kg lighter than an equivalent delta wing, while still having the same 2783 km cross-range. The Lockheed design featured an aluminium airframe, a titanium thrust structure, an LI-1500 tile thermal protection system, and a tantalum alloy nose-cap able to withstand 1650 deg C. The V-tank would be dropped at 19,800 kph, at a zero aerodynamic pressure. The booster for the two-stage variant was a scaled-up version of the McDonnell Douglas/Martin Marietta Phase B design. The orbiter in the two-stage version would delete the drop tank, seven of the nine engines, and two turbojets. The positions of the internal tanks would be reversed to compensate for the centre of gravity change created by deletion of the seven engines. Lockheed estimated there would be an 5200 kg weight penalty for an orbiter created for 1.5 stage to orbit and later used for two stage to orbit, as opposed to one designed only as the second stage. Lockheed planned to built two ground and three flight test vehicles in an $ 8 billion development program which would lead to a first flight by April 1975. The expendable drop tank would consume 24% of this amount. Cost per flight was estimated as $ 7.1 million, declining to $6.3 million by the 416th flight. Manufacturer: Lockheed. Liftoff Thrust: 21,214.400 kN (4,769,187 lbf). Total Mass: 1,730,803 kg (3,815,767 lb). Core Diameter: 4.57 m (14.99 ft). Total Length: 57.00 m (187.00 ft).
Shuttle R134C.
Rockwell/General Dynamics shuttle proposal phase B, November 1970. Delta wing high-cross range orbiter and booster. The North American Phase B design was prepared under contract NAS9-10960 and reflected the final revised shuttle specification of 11 November 1970. NAR-134-B was North American's original high cross range configuration. The delta wing had turned-up wingtips, as on the X-20 Dynasoar spaceplane of the 1960's. The orbiter was 43.4 m long, with a 38.4 m wingspan and stood 15 m high. The air-breathing cruise engines were mounted in swing-down nacelles that deployed from the belly of the orbiter. The spacecraft weighed 98,745 kg empty, and had a 9,100 kg payload. The orbiter had a 2784 km cross-range, using titanium in the wing structure, and Hynes-188 and Inconel-718 panels for the thermal protection system. The re-entry involved an initial nose-up angle of 55 deg, followed by a pitch over to 35 deg after peak heating, followed by bank manoeuvres as required to achieve the cross-range for the mission. The orbiter had a hypersonic L/D of 0.7 during the peak heating/braking manoeuvre; and 2.2 at 20 deg angle of attack for the bank manoeuvres; and 6.9 subsonic. Landing speed would be 213 kph. The B8G booster was designed by Convair and had a conventional vertical tail and horizontal stabiliser, as opposed to the 'V' tail used for the low-cross-range orbiter's booster. In three minutes it would take the orbiter to 70,000 m and 11,000 kph, hitting a peak of 4G's. After separation of the orbiter, it would spend 10 minutes in a hypersonic braking glide, finally starting its air breathing engines at subsonic speed when 500 km downrange and 6700 m altitude. The subsonic cruise back to base would take 90 minutes. The boostter had a 0.5 hypersonic L/D, and 6.7 ?/D subsonic. It was equipped with 12 x 188,000 kgf SSME engines. In the Phase B Final Report these shuttle designs was modified. The North American Rockwell/Convair team's booster was now the B9U configuration, 82 m long with a 43.6 m wingspan. Propulsion had grown to 12 x 249,000 kgf main engines, and 12 x JF22A-4 air breathing engines powered by JP-4 fuel. The NAR-161-C delta wing orbiter had replaced the upturned wingtips with a conventional vertical stabiliser. In late 1969 the USAF had indicated a preference for all-aluminium structures in the shuttle due to a titanium shortage. This requirement forced a move to non-metallic thermal protection systems, which at the time it was thought would weigh 15% less but cost 300% more. Thermal protection shingles for a titanium structure would weigh 2300 to 4500 kg less, but an aluminium structure would weight about 1800 kg more - meaning there was no essential weight difference between the two approaches. Therefore at the aluminium structure was accepted as a specification requirement. In retrospect it could hardly have been necessary to apply this requirement on a project where only a few flight vehicles were be built. It made the shuttle much more vulnerable to any breach of heat shield integrity and would lead to the death of the Columbia crew 35 years later. The resulting need for a non-metallic thermal protection system would also have enormous cost and schedule consequences for the actual program. Manufacturer: North American, Convair. Liftoff Thrust: 25,564.100 kN (5,747,038 lbf). Total Mass: 2,188,488 kg (4,824,790 lb). Core Diameter: 10.37 m (34.02 ft). Total Length: 98.00 m (321.00 ft).
Shuttle R134G.
Rockwell/General Dynamics shuttle proposal phase B, November 1970. Straight wing low-cross range orbiter. The North American Phase B design was prepared under contract NAS9-10960 and reflected the final revised shuttle specification of 11 November 1970. The low cross-range orbiter used configuration NAR-130-G, and had a payload of 20,500 kg and a cross range of 516 km. The structure was of Al-4V titanium alloy, and protected by a radiative thermal protection system, and a columbium nose cap. Haynes-188 tiles were used on the forward fuselage, and Inconel panels on the aft fuselage. The shuttle could accommodate four crew and ten passengers, plus payload in a 4.57 m x 18.3 m bay just behind the crew cabin. Two aluminium liquid oxygen tanks were located below the forward payload bay. 4 JTF22B-2 turbofan engines (versions of the F100 fighter engine without the afterburner), used liquid hydrogen as fuel. The orbiter would reenter at a 60-degree nose-up attitude, and had a hypersonic L/D of 0.56. North American had concern about the shock interaction on the forward part of the wing for the Faget configuration, but NASA insisted that their wind tunnel tests at Mach 10 showed the problem was only a bit worse than a delta configuration. In subsonic flight the orbiter would have an 8.2:1 lift to drag and a 250 kph landing speed. Study showed that for every 1% shortfall in main engine specific impulse, there would be an 11% reduction in payload. Expected temperatures were 1650 deg C on the leading edge, and 760 deg C on the belly. The B8G booster was designed by Convair and had a 'V' tail. In three minutes it would take the orbiter to 70,000 m and 11,000 kph, hitting a peak of 4G's. After separation of the orbiter, it would spend 10 minutes in a hypersonic braking glide, finally starting its air breathing engines at subsonic speed when 500 km downrange and 6700 m altitude. The subsonic cruise back to base would take 90 minutes. The booster had a 0.5 hypersonic L/D, and 6.7 L/D subsonic. It was equipped with 12 x 188,000 kgf SSME engines. In the Phase B Final Report these shuttle designs was modified. The North American Rockwell/Convair team's booster was now the B9U configuration, 82 m long with a 43.6 m wingspan. Propulsion had grown to 12 x 249,000 kgf main engines, and 12 x JF22A-4 air breathing engines powered by JP-4 fuel. In late 1969 the USAF had indicated a preference for all-aluminium structures in the shuttle due to a titanium shortage. This requirement forced a move to non-metallic thermal protection systems, which at the time it was thought would weigh 15% less but cost 300% more. Thermal protection shingles for a titanium structure would weigh 2300 to 4500 kg less, but an aluminium structure would weight about 1800 kg more - meaning there was no essential weight difference between the two approaches. Therefore at the aluminium structure was accepted as a specification requirement. In retrospect it could hardly have been necessary to apply this requirement on a project where only a few flight vehicles were be built. It made the shuttle much more vulnerable to any breach of heat shield integrity and would lead to the death of the Columbia crew 35 years later. The resulting need for a non-metallic thermal protection system would also have enormous cost and schedule consequences for the actual program. Manufacturer: North American, Convair. LEO Payload: 20,500 kg (45,100 lb). to: 185 km Orbit. at: 28.50 degrees. Liftoff Thrust: 25,564.100 kN (5,747,038 lbf). Total Mass: 2,000,000 kg (4,400,000 lb). Core Diameter: 10.37 m (34.02 ft). Total Length: 98.00 m (321.00 ft).
Version: Saturn Shuttle.
A winged recoverable Saturn IC stage was considered instead of solid rocket boosters after the final shuttle design was selected. In yet another iteration of shuttle design studies, $ 2.8 million contracts were given in November 1971 to Grumman/Boeing, Lockheed, McDonnell-Douglas/Martin Marrietta, and North American Rockewell. The development costs for the Phase B Prime contracts had still been over the Nixon administration's budget cap, and still further ways to reduce development cost had to be found. The studies were to run through 15 March 1972 and study lower cost booster concepts, one of them a Saturn V first stage modified to serve as a flyback booster. The studiy concluded that a Saturn S-IC flyback booster would need a wing with at least a 700 sq m area, would be powered by five F-1 engines and have a gross mass of 1.6 million kg. Staging would take place at 6450 kph. The vehicle would be reusable, except for the F-1 engines. The use of expendable engines was considered a drawback. The study assumed a series burn, with the shuttle orbiter igniting at altitude. Liftoff Thrust: 33,737.900 kN (7,584,582 lbf). Total Mass: 3,161,710 kg (6,970,370 lb). Core Diameter: 10.06 m (33.00 ft). Total Length: 107.00 m (351.00 ft). Flyaway Unit Cost $: 1,020.500 million. in: 1985 unit dollars.
Shuttle LRB 1972.
Original design for a shuttle with liquid rocket boosters, completed in March 1972 as part of the shuttle design decision process In November 1971 Shuttle Phase B Double Prime studies were initiated. In yet another iteration of shuttle design studies, $ 2.8 million contracts were given to Grumman/Boeing, Lockheed, McDonnell-Douglas/Martin Marrietta, and North American Rockewell. The development costs for the Phase B Prime contracts had still been over the Nixon administration's budget cap, and still further ways to reduce development cost had to be found. The studies were to run through 15 March 1972 and study lower cost booster concepts, one of which was a fully recoverable stage but with a new pressure-fed engine The new-design pressure fed liquid propellant booster would be parachute-recovered and reused. Using liquid oxygen/kerosene propellants, each booster would be 9.93 m in diameter, 48.5 m long, and be equipped with two 612,000 kgf engines. Three boosters would be assembled in parallel, with the external tank for the shuttle orbiter atop the core booster. This design would have a gross lift-off mass of 2,626,000 kg. Another design used 4 x 475,000 kgf engines in each stage, resulting in a 2,394,000 kg vehicle. It was estimated the LRB's would cost $4.2 billion to develop, plus $ 8.9 billion to operate, making shuttle cost $275/kg to orbit. The study assumed a series burn, with the shuttle orbiter igniting at altitude. Liftoff Thrust: 36,010.000 kN (8,095,370 lbf). Total Mass: 2,626,000 kg (5,789,000 lb).
Version: EDIN05.
In February 1976 this version of the shuttle was proposed. A single liquid rocket booster under the external tank would replace the two solid rocket boosters. In January 1976 the Sigma Corporation proposed use of a single liquid rocket booster under the external tank to replace the two solid rocket boosters. The pod would be powered by 3 to 4 F-1 engines or by three high performance liquid oxygen/kerosene engines being designed by the System Development Corporation. These developed either 308,000 kgf or 363,000 kgf at sea level, depending on the version selected. Both parallel and serial burns of the LRB and the Shuttle/External tank combination were studied. The external tank would be stretched and the orbiter strengthened to handle the increased payload available. However use of the booster would require the construction of new launch facilities.
The Boeing SDV Class I vehicle would lead to the Shuttle-C, using the shuttle aft fuselage with SSME engines to power a cargo canister into orbit. Boeing was awarded NASA contract NAS8-32398 on 28 July 1977 to study unmanned cargo derivatives of the shuttle. The Class I vehicle would be similar to the Shuttle-C, using the shuttle aft fuselage with SSME engines to power a cargo canister into orbit. The engine pod would be recovered, and the payload shroud would be 7 m in diameter and 29 m long (allowing payloads up to 20.9 m long. A stretched shroud would be 38 mlong and accommodate 30 m payloads. Both integral and separable engine pod variations were explored, resulting in payloads between 59,000 kg and 91,000 kg. A three-SSME version could but 91,000 kg into orbit and would have a gross liftoff mass of 922,000 kg. A four SSME version could put 102,000 kg into orbit with a gross liftoff mass of 975,000 kg. The propulsion module was to be designed for 300 flights. Development cost was estimated at $930 million, and first article cost for the engine pod $135 million. Cost of expendable items per flight was estimated at $4.23 million, with total cost per flight $14.136 million, or $1118 per kg to orbit, as opposed to the $1323/kg expected for the basic shuttle at that time. Class II designs used a payload fairing 8.31 m in diameter, with a single liquid rocket booster, and 4 SSME engines uprated to 308,000 kgf and using a 2.9:1 oxidiser/fuel ratio. These would be rated for 50 flights between overhaul. The booster would have a gross liftoff mass of 1,130,000 kg, and be capable of placing 91,000 to 135,000 kg into orbit. The new booster would cost $2.863 billion to develop, with the first article costing $241 million. Cost per flight was expected to be $13.437 million, $ 4.38 million of that for expendable items, resulting in a cost per pound to orbit of $593. Manufacturer: Boeing. LEO Payload: 91,000 kg (200,000 lb). to: 300 km Orbit. at: 28.50 degrees. Total Mass: 922,000 kg (2,032,000 lb).
Version: IHLLV. Orbital launch vehicle. Status: Study 1980. Other Designations: Interim Heavy Lift Launch Vehicle. Same concept as Shuttle C. Shuttle orbiter replaced by recoverable pod with shuttle main engines and payload cannister. Quick way for US to obtain heavy payload capability and reduce shuttle cost per kg to orbit by 3 X. LEO Payload: 77,000 kg (169,000 lb). to: 400 km Orbit. at: 28.00 degrees. Liftoff Thrust: 20,299.200 kN (4,563,442 lbf). Total Mass: 1,966,675 kg (4,335,776 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft). Flyaway Unit Cost $: 84.970 million. in: 1985 unit dollars. Martin Marietta SDV.
The Martin Marietta Class I SDV would lead to the Shuttle-C, using the shuttle aft fuselage with SSME engines to power a cargo canister into orbit. Martin studied cargo versions of the shuttle under two-phase contract NAS8-34183. Phase I was completed in July 1981, and Phase II on 10 February 1983. The Phase I study looked at four classes of Shuttle-Derived Vehicles:
Phase II studies refined these concepts somewhat:
Manufacturer: Martin. LEO Payload: 61,024 kg (134,534 lb). to: 300 km Orbit. at: 28.50 degrees.
Version: Shuttle LRB.
Shuttle with Liquid Rocket Boosters in place of Solid Rocket Boosters.
Liftoff Thrust: 21,107.400 kN (4,745,132 lbf). Total Mass: 1,575,493 kg (3,473,367 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft). Development Cost $: 1,629.000 million. in: 1985 average dollars.
Shuttle II. Orbital launch vehicle. Status: Study 1988. In May 1988 NASA Langley studied a new-technology approach to improving the shuttle's payload capability. The design would allow 9,000 to 18,000 kg of additional payload to be carried in an external payload container or in the orbiter. Composite material technology would be used in a substantial rebuild of the shuttle orbiters. New, lighter RSI tiles would be used, and a new SSME jointly designed by Pratt and Whitney and Aerojet would replace the Rocketdyne motor. The new SSME would deliver 304,000 kg while weighing only 3770 kg versus 4688 kg for the Rocketdyne engine. Electric actuators would replace hydraulic actuators inside the shuttle, In all, the new orbiter would be 16% lighter at SECO. Manufacturer: NASA Langley. LEO Payload: 40,000 kg (88,000 lb). to: 400 km Orbit. at: 28.00 degrees.
Version: Low Cost Cargo Vehicle. Orbital launch vehicle. Status: Design 1990. This variant of the Shuttle C was envisioned for delivery of liquid hydrogen and liquid oxygen to orbit. It would have no payload module, just a thrust structure for two SSME's. 74,000 kg of unused propellant within the external tank could be delivered to orbit for transfer to space station propellant farms or interlunar/interplanetary spacecraft.
NASA Marshall design for a cargo version of the shuttle system. The shuttle orbiter would be replaced by an unmanned recoverable main engine pod. The same concept was studied earlier as the Interim Heavy Lift Launch Vehicle (IHLLV) and as the Class I Shuttle Derived Vehicle (SDV). The Phase I two-SSME configuration would have a payload of 45,000 kg to low earth orbit. Design carried to an advanced phase in 1987-1990, but then abandoned when it was found the concept had no cost advantage over existing expenable launch vehicles. After the Challenger accident, NASA believed there would be a need for a 14 per year shuttle launch rate. But such a rate would be unachievable given the existing orbiter flow times and the reduced number of orbiters available. The Shuttle-C (Cargo) would allow such a rate to be achieved. After Challenger, NASA studied development of a new unmanned Heavy Lift Launch Vehicle with the Air Force. In the discussions NASA insisted that any HLLV be a Shuttle Derived Vehicle. This led to the USAF pulling out of the shuttle program, and developing the Titan IV for the short term, and beginning a study of a new ALS Advanced Launch System for the long term. The USAF did not believe low-cost access to space was possible using man-rated systems. In the aftermath of the kerfuffle a Joint NASA-DoD-USAF steering group was set up to monitor development of the ALS and Shuttle-C. In August 1987 NASA Houston officially began Shuttle-C studies through assignment of a task team. In November 1987 nine-month study contracts were let to Martin-Marietta, United Space Boosters Inc, and Rockwell for definition of a Shuttle-Derived Vehicle. Phase I of the studies was to determine the optimum vehicle configuration, and Phase II was to completely define the selected vehicle. Concept studied included:
The Class I SDV was again found to be the best solution. It was estimated it could deliver payload to orbit at a cost of $4400/kg, as opposed to $1720 for the Delta II, $ 1800 for the Titan IV, or $ 3400 for the shuttle. As a result a Request for Proposal was issued for the Expendable Cargo Element - a payload fairing for the Shuttle-C to be mounted on the side of the external tank. This could accommodate 4.6 m x 22 m payloads weighing up to 47,000 kg and would be delivered into a 407 km / 28.5 deg orbit for docking with Space Station Freedom. The system could also deliver 52,000 kg to a 300 km / 28.5 deg orbit. The CE (Cargo Element) was equipped with 2 Space Shuttle Main Engines, and 2 Orbital Maneuvering System pods. The payload would either be released attached to the planned Orbital Maneuvering Vehicle transfer stage, or an OMV already in orbit would dock with the CE and take the payload away; or the OMS itself would be used to put the payload in its final orbit, release it, then retrofire to return the CE to earth. In February 1988 the industrial team Rockwell - Martin Marietta - Boeing - Teledyne - Intermetrics - United Space Boosters was awarded follow-on18-month study contract NAS8-37144. During the course of the study NASA canceled the OMV, and the design had to be modified to handle in-orbit delivery and release of the payloads. The final report of the study envisioned two generations of Shuttle-C's. Generation 1 would be an expendable CE with a 4.6 m x 24.7 m cargo bay, using two SSME's, and capable of deliverying 45,000 kg to orbit. The CE would have an empty weight of 31,750 kg, using the shuttle thrust structure, and be fitted with shuttle engines and computers at the end of their useful lives. Three to four flights per year could be accomplished using the expendable approach, while NASA believed it had a requirement for 10 to 12 per year. Generation 2 would have a new-design recoverable CE, powered by 3 SSME's, and capable of delivering 77,000 kg in a 7.3 m x 29.3 m volume. In early 1989 the study contract was extended by one year, including consideration of use of the Centaur G-Prime from the Titan IV as an upper stage. It was determined that the MPTA-098 structural article built in the 1970's for the SSME development program could be used as the Shuttle-C prototype. At the end of the study NASA decided that the development cost for Shuttle-C would be $ 1.8 billion. The Office for Technology Assessment estimated the cost as only $985 million, a rare case indeed where NASA made a high-ball estimate. Cost per launch would be $424 million for the Generation 1 design, or $9350 per kg. 14 Space Shuttles and 10 Shuttle-C's could be launched per year using existing Kennedy Space Center facilities. However such a rate would quickly exhaust the supply of surplus SSME engines. Therefore new-build engines would have to be purchased at a cost of $38 million each, or a total of $500 million per year. When this cost was taken into consideration, Shuttle-C was more expensive than the USAF Titan IV - therefore, NASA concluded, there was no reason to develop it. The decision was taken in 1990 to cancel Shuttle-C. LEO Payload: 77,000 kg (169,000 lb). to: 400 km Orbit. at: 28.00 degrees. Liftoff Thrust: 20,299.200 kN (4,563,442 lbf). Total Mass: 1,966,675 kg (4,335,776 lb). Core Diameter: 8.70 m (28.50 ft). Total Length: 56.00 m (183.00 ft). Flyaway Unit Cost $: 84.970 million. in: 1985 unit dollars.
Shuttle C Block II.
In August 1989 NASA studied a version of the Shuttle-C with two Advanced Solid Rocket Mortors (ASRM's) in place of the standard RSRM's. This would increase the payload by 4500 kg, but also require use of a new 10 m x 30 m payload module.
Manufacturer: Martin. LEO Payload: 81,500 kg (179,600 lb). to: 407 km Orbit. at: 28.50 degrees.
Version: Shuttle LRB 1989.
In July 1989 a NASA Langley/George Washington University joint study was made of various Liquid Rocket Booster configurations. A parametric trade analysis looked at the optimum solution. This was found to be twin oblique-wing recoverable boosters. Staging at Mach 3 would allow them to be built of aluminium rather than higher-temperature materials. Standard SSME's would power each booster. Two optimum configurations were identified. One was a 2-engine orbiter and two 3-engine boosters (2-3-3). In this case the orbiter would be 45 m long, and the booster 35.4 m long. Payload would be 16,800 kg, and gross liftoff mass 880,000 kg. Staging would be at 86 seconds after launch at 26,000 m altitude, earth orbit insertion 516 seconds after launch. The 3-4-4-configuration would result in a 50.3 m long orbiter and 40.5 m long booster. Payload would be 31,750 kg, gross liftoff mass 1,316,000 kg. In either case the orbiter would be equipped with a 4.6 m x 9.2 m cargo bay and a crew of two. A 5-3-3 configuration, using 3 x liquid oxygen/kerosene engines in each booster was also studied. This would have a payload of 38,000 kg to orbit. However the study concluded that having the same propellants and engines in the orbiter and booster was an advantage that outweighed other considerations. Manufacturer: George Washington Univeristy.
Version: Ares Mars Direct.
The Ares launch vehicle was designed for support of Zubrin's Mars Direct expedition. It was a shuttle-derived design taking maximum advantage of existing hardware. It would use shuttle Advanced Solid Rocket Boosters, a modified shuttle external tank for handling vertically-mounted payloads, and a new Lox/LH2 third stage for trans-Mars or trans-lunar injection of the payload. Ares would put 121 tonnes into a 300 km circular orbit , boost 59 tonnes toward the moon or 47 tonnes toward Mars. Without the upper stage 75 tonnes could be placed in low earth orbit.
LEO Payload: 121,200 kg (267,200 lb). to: 300 km Orbit. at: 28.50 degrees. Payload: 47,200 kg (104,000 lb). to a: trans-Mars trajectory. Associated Spacecraft: Mars Direct. Total Mass: 2,194,600 kg (4,838,200 lb). Core Diameter: 10.00 m (32.00 ft). Total Length: 82.30 m (270.00 ft).
Version: Shuttle Z.
Shuttle-Z was Shuttle-C on steroids, the ultimate development of the shuttle to be used to put Mars expeditions into orbit. It would use 4 SSME's, and a third stage with 181,000 kg of propellant powered by 1 SSSME. But such designs would require new handling facilities due to the extra height of the vehicle.
LEO Payload: 87,500 kg (192,900 lb). to: 407 km Orbit. at: 28.50 degrees. Shuttle Chronology 1968 October 30 - Phase A Space Shuttle studies. NASA began the design, bidding, and source selection process leading to a single national space shuttle. At the beginning the design was known by the same nomenclature previously used by the USAF - Integrated Launch and Re-entry Vehicle (ILRV). The development program was seen as: Phase A: Advanced Studies; Phase B: Project Definition; Phase C: Vehicle Design; and Phase D: Production and Operations. Four contractors or contractor teams were to be selected in Phase A; two contractors or teams for Phase B; and then a single contractor for Phases C and D (which were later combined). NASA Houston and Huntsville jointly issued the Request for Proposal for eight-month Phase A ILRV studies. The requirements were for 2,300 to 23,000 kg of payload to be delivered into a 500-km altitude orbit. The re-entry vehicle should have a cross range of at least 725 km (NASA persisted in this requirement even though it knew the USAF needed more). General Dynamics, Lockheed, McDonnell-Douglas, Martin Marietta, and North American Rockwell all were invited to bid. The Space Shuttle Main Engine competition was run in parallel with the main shuttle development project, and also had four phases. Oversight for this program came from the USAF Space Division and its subcontractor, the Aerospace Corporation. Despite promising classified work on linear and conventional aerospike engines at the time, NASA dictated that the design had to use a conventional bell nozzle. February 1969 - Space Shuttle Phase A contracts Following evaluation of proposals submitted against the October 1968 request for proposal, NASA issued Advanced Design contracts for the shuttle to General Dynamics, Lockheed, McDonnell Douglas, and North American Rockwell. Martin Marietta did not receive a contract but was allowed to continue using company funds. Rocketdyne and Pratt & Whitney were selected for the Phase A, advanced study phase of the competition. The same basic engine (combustion chamber and turbomachinery) was to be used in both stages of the planned two-stage fully-recoverable shuttle. The orbiter would be equipped with a two-position deployable nozzle, with expansion ratios of 58:1 for the low altitude portion of the ascent, and 120:1 with the extension deployed for the vacuum portion of the flight to orbit. The engine was to have a thrust of 270,000 kgf in vacuum, 235,000 kgf at sea level, and be throttleable from 73% to 100% of the rated thrust. The engine for the booster was to use a 5:1 ratio expansion nozzle, producing 227,000 kgf at sea level. Pratt & Whitney seemed to have a clear lead in this portion of the competition, having produced the XLR-129-P-1, a prototype high-pressure Lox/LH2 engine under USAF contract. This produced 188,000 kgf using a smaller fixed nozzle. Most of the shuttle bidders proposed use of this engine in their Phase A vehicle designs. The Space Task Group put together to run the shuttle design process was composed of various agencies of the federal government. Each group favoured differing basic configurations for the shuttle, reflecting controversies extending back over ten years to the time of DynaSoar development. Faget at NASA Houston favoured a straight-wing orbiter, the bottom surface being essentially a cross shape cut out of the spherical section of one of the Apollo or Mercury heat shields he had designed. This had minimal cross range, but was supposed to have the advantages of minimum weight and good subsonic glide performance. NASA Langley and Edwards AFB favoured a lifting body, based on the HL-10 shape under test there. This had supposed weight advantages over a winged vehicle, more cross range than Faget's straight wing, but less cross range than a delta wing. USAF Flight Dynamics Laboratory and Draper Laboratories favoured a swept delta wing spaceplane, like the Dynasoar, for maximum cross range on re-entry. Faget favoured a small net payload to orbit (6800 kg) while the other government centres favoured heavier payloads, at least 11,300 kg, and up to 29,500 kg. As in the case of earlier USAF ILRV studies, the Space Task Group had initially considered three categories of launch solutions. Class I used an existing expendable launch vehicle (the Titan 3MV or Saturn IB) and a reusable orbiter. Class II were 1.5 stage to orbit designs, using an orbiter vehicle and a drop tank. Class III were fully reusable two-stage-to-orbit designs. In contrast to the USAF studies, which favoured immediate development of a Class I vehicle, followed by a Class II vehicle, Task Group's preferred solution was to proceed immediately with a Class III vehicle. 1969 February 13 - Nixon forms Space Task Group Vice President Agnew was made chairman of the group, which was to formulate a Post-Apollo Space Program, providing policy direction for future American efforts after the moon landing. The Groups final report proposed three alternate future programs:
Nixon rejected all of the alternatives and wanted something even cheaper. 1969 April 21 - Space Shuttle Task Group formed The Director of Apollo Test in the NASA Hq. Apollo Program Office, LeRoy E. Day, was detailed to head the MSF Space Shuttle Task Group. The group would provide NASA with material for a report on the Space Shuttle to the President's Space Task Group. 1969 June 1 - Faget shuttle concept attacked The first report comes out attacking the Faget straight wing design. Another follows in November 1969; with the dispute becoming public with AIAA papers published in October 1970 and January 1971. These dissidents at other NASA centres calculated that a Faget orbiter was unsafe, as it could not withstand the re-entry thermal environment and aerodynamic stresses. NASA's Flight Research Center pushed a lifting body design, while the US Air Force noted that in any case the Faget design did not meet its cross-range requirements. Fall 1969 - No government approval for NASA's shuttle program NASA decided to take the minimum program proposed by the Space Task Group (just the space station and the shuttle), and then implement it over a very long period in phases. At first only a reusable space shuttle would be developed. When that was completed, work on a space station could start. However as of the fall of 1970, NASA was unable to obtain the Nixon administration's approval of even this limited program. 1969 September 11 - Two major directions were identified for NASA manned space flight in the next decade. Spacecraft: Skylab, Columbia. Flight: Skylab B. Two major directions were identified for manned space flight in the next decade. These were further exploration of the Moon, with possibly the establishment of a lunar surface base, and the continued development of manned flight in Earth orbit, leading to a permanent manned space station supported by a low-cost shuttle system. To maintain direction, the following key milestones were proposed: 1972 - AAP operations using a Saturn V launched Workshop 1973 - Start of post-Apollo lunar exploration 1974 - Start of suborbital flight tests of Earth to orbit shuttle - Launch of a second Saturn V Workshop 1975 - Initial space station operations - Orbital shuttle flights 1976 - Lunar orbit station - Full shuttle operations 1977 - Nuclear stage flight test 1978 - Nuclear shuttle operations-orbit to orbit 1979 - Space station in synchronous orbit By 1990 - Earth orbit space base - Lunar surface base - Possible Mars landing 1970 January 23 - NASA Houston in-house study of shuttle concepts The study was in an attempt to resolve disputes between the centres as to the best approach. Houston's Faget straight-wing two-stage vehicle was in competition with concepts from other centres - recoverable versions of Saturn boosters, and an advanced single-stage-to-orbit Aerospaceplane. Payload for the Faget vehicle was to be only 5,700 to 6,800 kg to low earth orbit, and the system was to be operational by the end of 1975, after the last Apollo flight. 1970 May 4 - DC-3 drop tests NASA conducted drop tests of a 1/10 scale model of Faget's 'DC-3' straight-wing shuttle design. The model was 4 m long, weighed 270 kg, and was dropped from 3,700 m altitude. Recovery was by parachute. 1970 June 1 - NASA completes Shuttle Phase A evaluations After over 200 man-years of NASA and contractor effort, the Agency reached the following conclusions at the end of Phase A:
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