SLS Family Credit - © Mark Wade

In the mid-1950's, US Air Force-funded studies identified the optimum long-term solution for space launch. The studies indicated the desirability of segmented solids for a first stage to achieve low cost, high reliability and flexibility of basic booster size by adding or subtracting segments. Studies further showed that oxygen-hydrogen propellants, with their very high specific impulse, were a preferred choice for upper stages, where mass was more important. This choice also resulted in minimum systems cost.

By stressing the concept of a single liquid stage with a single engine, it was felt that a high reliability for the over-all system could be achieved. Furthermore, by starting out from the very beginning with the concept that this was to be a standardised vehicle for a wide variety of space missions, it was felt that a basically good design could be achieved which would be useful for '…at least five, and perhaps ten years…' as a work-horse booster.

In fact the design philosophy of this 'Space Launching System' was decades ahead of its time. Due to short-term funding restrictions, the Air Force selected adding solid rocket boosters to the Titan 2 ICBM for its SLV-4 space launch requirement. If they had gone forward with the Space Launching System, they would have fielded the equivalent of the Ariane 5, Delta 4, or Atlas 5 forty years earlier. Such a system would probably still be in use today, and in no need for replacement.

The Air Force had sponsored key technology development in the late 1950's to prepare for the SLS. This included the successful testing of a Titan LR-87 engine to burn liquid oxygen and hydrogen in 1958-1960, and initial contracts for test of 100 inch segmented solid rocket boosters in 1959-1960.

The smallest member of SLS was sized to boost the Air Force's Dynasoar manned spaceplane into low earth orbit. The 'A' liquid core stage was limited to 14 feet diameter to allow rail transport. The 100 inch diameter of the solids were set by the limitation of Aerojet's heat-treat facility. This had the capability to put 9 tonnes of payload into low earth orbit.

To support the Air Force's project Lunex moon landing project, 'B' and 'C' Lox/LH2 stages were conceived. These were 25 feet in diameter, and their lateral solid boosters 180 inches/15 feet in diameter. Both would have to be transported by ship.

A decision to proceed with the Space Launching System in July 1961 would have resulted in first flight of the A vehicle in mid-1964, first manned launch in mid-1965, first launch of the BC super-booster in mid-1966, and the first manned landing on the moon in late 1967. Instead NASA was given the Apollo program, the Titan 3 was developed for the Air Force's launch needs, and an opportunity to build a flexible launch system that would still be in use today was lost.

SLS A-410.

SLS A-410 Credit - © Mark Wade

Status: Study 1961. Other Designations: Space Launching System A-410.

The smallest identified member of the SLS family, selected to place the Air Force Lunex lunar lander re-entry vehicle in a low earth orbit for initial tests, was the A-410. This consisted of the 'A' Lox/LH2 stage supplemented by 100-inch diameter solid fuel booster rockets.

LEO Payload: 9,070 kg (19,990 lb). to: 560 km Orbit. at: 28.00 degrees. Liftoff Thrust: 7,553.000 kN (1,697,981 lbf). Total Mass: 420,000 kg (920,000 lb). Core Diameter: 7.62 m (24.99 ft). Total Length: 27.00 m (88.00 ft). Span: 18.60 m (61.00 ft).

SLS AB-825.

SLS AB-825 Credit - © Mark Wade

Status: Study 1961. Other Designations: Space Launching System AB-825.

The AB-825 represented a medium launch vehicle of the USAF 1961 Space Launching System family. The AB-825 would have conducted earth orbit tests of partially-fuelled Lunex lunar lander stages, and also have boosted the Lunex manned glider on circumlunar test flights. It consisted of the 'A' stage and 'B' stages with 180 inch diameter short-length solid fuel booster motors.

LEO Payload: 39,460 kg (86,990 lb). to: 560 km Orbit. at: 28.00 degrees. Payload: 10,890 kg (24,000 lb). to a: translunar trajectory. Liftoff Thrust: 14,700.000 kN (3,304,600 lbf). Total Mass: 1,000,000 kg (2,200,000 lb). Core Diameter: 7.62 m (24.99 ft). Total Length: 55.00 m (180.00 ft). Span: 16.76 m (54.98 ft).

SLS BC-2720.

SLS BC-2720 Credit - © Mark Wade

Status: Study 1961. Other Designations: Space Launching System BC-2720.

The BC-2720 was the member of the SLS family selected to boost the Air Force Lunex lunar lander on a direct lunar trajectory. This would have used four 180 inch solid rocket boosters strapped around an the 'C' Lox/LH2 core vehicle. The core would have required either 12 J-2 engines or 2 M-1 engines. The translunar injection third stage was the 'B', with a single J-2 engine.

NASA, not the Air Force, received the task of going to the moon and building large boosters. But the Saturn series was abandoned, while the BC-2720 resembles very much the Shuttle and Russian Energia configurations. Perhaps if things had been different...

BC Vehicle Launch Facility Considerations

Investigation of the launch pad requirements for a launch rate of two per month indicated that from 4 to 6 launch pads would be necessary depending on the launch site location and the means available for handling the booster. There were no existing launch pads capable of handling this vehicle, nor were there facilities capable of conducting static testing of the "C" booster and the launch of the complete Lunar Transport Vehicle. It was considered possible that by combining the capabilities for both static firing and launch in two of the pads required, a significant cost saving could be gained and an accelerated test program effected. This would provide a capability for the launch of the "C" booster with or without solid boost during R&D flight test and for early test missions of the Lunex Re-entry Vehicle. The development and flight test of the "B" booster was planned at Cape Canaveral during the development program.

It was assumed in the Lunar Transport Vehicle study that the manufacture of all boosters and the payload would be accomplished at existing factories. New and added facilities and equipment such as large forming brakes, special welding jigs, fixtures and machines, and large processing facilities would be required. In plants of sufficient size these facilities and equipment could readily be installed. Further investigation comparing the relative economics of manufacture at the launch site versus manufacture at existing facilities was required to insure an economical choice.

Assemblies having a diameter exceeding 12 feet or weighing over 200,000 pounds could not be transported over United States railways. A load of 78,000 pounds was considered to be the limit over selected highway routes. Inasmuch as both the "B" and "C" boosters of the Space Launch Systems had diameters in excess of 14 feet, transport from manufacturing plant to the launch site would have to be by barge. The large quantities of boosters and the special environmental protection required suggest that specially designed barges be constructed to transport these assemblies. Harbours and docking facilities would be required near the manufacturing facility and at the launch site.

By locating the launch facilities at or near Cape Canaveral for an easterly launch significant savings may be effected. The use of existing administrative capabilities, personnel housing, assured tracking facilities, and technical support areas would provide a saving in costs and in lead-time required for construction of support facilities. Similar gains could be made by locating launch facilities at Point Arguello for polar launch. This did not mean that Cape Canaveral and Point Arguello were the only reasonable locations for the launch site. In fact, by extending the Atlantic Missile Range in a westerly direction across the Gulf of Mexico it was conceivable that a launch site in the vicinity of the Corpus Christi Naval Air Complex would provide the full use of Atlantic Missile Range facilities with minimum overfly of foreign land masses. Likewise, extension of the Atlantic Missile Range in a northerly direction to the coast of South Carolina would provide a similar accommodation.

The logistic support for the launch rate as indicated in the study dictated that new propellant manufacturing plants be constructed at the launch site. Existing propellant manufacturing plants were inadequate and the launch rates mentioned would use the full capacity of a separate propellant manufacturing facility.

a. Propellant use rates for a 2 per month launch rate are estimated as follows:

  (1) Liquid Hydrogen manufacture: 50 tons per day.

  (2) Liquid Hydrogen storage at launch pad: 1.5 million pounds.

  (3) Liquid Oxygen/Nitrogen Manufacture: 120 tons per day.

  (4) Liquid Oxygen storage at launch pad: 4 million pounds.

Barges would be required for transport of boosters from the manufacturing plant to the launch complex.

A modified Integrated Transfer launch System was envisioned for the Lunar Transport Launch System. This approach would allow the complete integration and checkout of the "B" booster together with the Lunar Transport Payload in a protected environment simultaneously with the assembly and checkout of the C2720 booster combination at the launch pad. The size and weight of the BC2720 Space Launching Vehicle precluded the transfer of the completely assembled Lunar Transport Vehicle from an integration building to the launch pad. It was feasible, however, to mate and integrate the "B" booster with the Lunar Transport Payload inside the protected environs of an integration building and, when completed, transfer the "B" booster and payload assembly to the launch pad for mating with the C2720 assembly.

This could best be accomplished by a cliff-side location or extending a ramp from the integration building to an elevation at the launch pad approximately equal to the height of the C2720 stage. The assembly and checkout of the C2720 vehicle could be accomplished in two ways depending on the specific location of the launch pad and its accessibility to navigable waters. For a launch pad having no direct access to navigable waters, the assembly and mating of the solid segmented motors to the "C" booster would be accomplished at the launch pad. The extended time necessary to accomplish this assembly and checkout accounts for the difference in the numbers of pads required. It was estimated that 6 launch pads would be needed for this plan.

For a launch pad having direct access to navigable waters, the assembly and mating of the solid segmented motors to the "C" booster could be accomplished at an interim integration building located some distance away from the launch pad. After assembly and checkout, the "C2720" combination would be transported by a barge to the launch pad and mated to the "B" booster and payload assembly. By using this approach it was estimated that 4 launch pads would be adequate for the 2 per month launch rate. Final confidence checks and integration of the booster and facility interface would be accomplished at the launch pad.

The TNT equivalent of vehicle propellants was estimated in the following manner. The TNT equivalent of the liquid propellants was taken at 60% of the total LOX/LH2 load for all stages. This was the figure currently used at Atlantic Missile Range for TNT equivalence for LOX/LH2. In this case, because of the great quantities of propellant involved, this degree of mixing was unlikely and the 60% figure would be conservative.

Solid propellants were taken at 100% of the propellant weight. It was also considered that detonation of the solid propellants may cause the subsequent detonation of liquid propellants and vice versa; but, the simultaneous detonation of all propellants was not likely to occur. This philosophy resolved to consideration of TNT equivalents of liquid propellants and solid propellants separately and they were not additive. The TNT equivalent of one of the four segmented solid assemblies was 680,000 pounds. The 60% TNT equivalent of the total liquid propellant load was approximately 1,300,000 pounds. Using the highest TNT equivalent (1,300,000 pounds) the inhabited building distance must be approximately 2 1/2 miles from the launch pad and minimum pad separation must be approximately 1 mile.

For an inhabited pad adjacent to a launch operation, pad separation would be 2 1/2 miles. It was obvious that the real estate problem would be extensive. For a coastal location of "C" launch pads up to 18 miles of continuous coast line would be required for a distance of 3 miles inland. These distances could be decreased by creating a buffer between the pads. Locating the launch pads in ravines or indentations in cliff aide launch locations might substantially reduce the land areas required. The selected location and orientation of the integration building and other support facilities to take best advantage of topography would do much to decrease distances and reduce costs.

The repeated launching of similar payloads in the Lunar Transport Launching System and the extended time between launches from each pad indicated that a central launch control for all pads might be desirable. To avoid analogue signal line driving problems and to allow greater distances than normal between the pads and the common blockhouse it was possible to use digital control for launch pad checkout and launch. Analogue to digital conversion would essentially be accomplished at each launch pad and transmitted to the blockhouse via digital data link.

With vertical mating, assembly and detailed checkout in the vertical assembly integration buildings, only gross, survey type testing or a simulated countdown and launch would be performed at the launch pad, since test and vehicle subsystem sequencing systems could be installed in both areas. Present day checkout methods, because of the many manual controls and long time spans involved, would not provide sufficient assurance of the high reliability of the complex integrated systems expected in the Lunar Transport Vehicle.

LEO Payload: 158,800 kg (350,000 lb). to: 560 km Orbit. at: 28.00 degrees. Payload: 60,800 kg (134,000 lb). to a: translunar trajectory. Liftoff Thrust: 30,000.000 kN (6,744,000 lbf). Total Mass: 2,600,000 kg (5,700,000 lb). Core Diameter: 7.62 m (24.99 ft). Total Length: 93.00 m (305.00 ft). Span: 16.76 m (54.98 ft).