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US Space Launching System

PMN1

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http://www.astronautix.com/lvfam/sls.htm

How accurate do you judge this statement?

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.
 

McTodd

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Very interesting, but frankly, this rings alarm bells...

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... 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.

'Segmented solids'? Would they have used O-rings, like the Shuttle's?

Sounds like a recipe for 28 January 1986 happening 20-odd years early...
 

Michel Van

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the SLS Program schedule of USAF is pure Fantasy

they think no limit on money, man power and you can send 3 man in LUNEX 1967 to Moon.

reality is another
alone the SYS A-388
Program schedule Program start 1961, see the first test launch at November 1963
10 test flight (5 without payload)
first Operational Flight is November 1964.
(source: Book Dyna-Soar by Robert godwin page 145)

allone for development of J-2 main engine needed Rocketdyne 7 years
October 1959 to January 1966 (and 203 test burns from 1965 to 1966)
almost 2 year to late ::)

and opportunity to build a flexible launch system ?
YES
take one SLS A-388 SRB put a Centaur on top you got small Sat launcher
with SLS A-388 launch 9 ton in orbit (like Dyna-Soar)
put 2 more SRB to that you go SLS A-410 more Payload.
use the corestage of A-388 as upperstage for SLS A-825

and cheaper ! because standart part production.

if USAF and NASA had taken SLS instat Saturn Ib - V
it would still launch US Astronaut into Space.

McTodd said:
Very interesting, but frankly, this rings alarm bells...

'Segmented solids'? Would they have used O-rings, like the Shuttle's?

Sounds like a recipe for 28 January 1986 happening 20-odd years early...
i think so, in case of SLS with Segmented solids to have malfunction, take this exampel:
Shuttle Launches: 122. Failures: 2. (one with SRB)
Ariane 5 Launches: 28. Failures: 2. (no SRB problems)
Titan IIIC Launches: 36. Failures: 5.
but Dyna-Soar and Apollo had Launch escape system, the Shuttle NOT.
 

Just call me Ray

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McTodd said:
Very interesting, but frankly, this rings alarm bells...

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... 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.

'Segmented solids'? Would they have used O-rings, like the Shuttle's?

Sounds like a recipe for 28 January 1986 happening 20-odd years early...

Not necessarily. There was nothing fundamentally wrong with the O-Rings except that they have a certain climatic envelope, and on 28 January, 1986, NASA made the decision to go despite that envelope being busted due to pressure to maintain their launch schedule. Ultimately, that decision was made on a very poor understanding of linear vs. exponential statistics (they used a linear equation and thought the temperatures were well within range, when it turns out they should've used an exponential equation and, well....)
 

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McTodd said:
Very interesting, but frankly, this rings alarm bells...

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... 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.

'Segmented solids'? Would they have used O-rings, like the Shuttle's?

Sounds like a recipe for 28 January 1986 happening 20-odd years early...

There's nothing inherently unsafe about the O-rings. But the logic that you can quickly develop boosters of varying size by playing Legos with booster segments is fallacious. Just look at the 5-segment SRB program. Only the casings and aft skirt are going unchanged from the Shuttle SRB, and it's going to cost ~$3B to develop.
 

PMN1

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I've noticed a section in the Astronautix article on CEV that states that lifting-body and winged configurations that could have been recovered with horizontal landings were possible for Apollo but time was against their development - how did they compare to the USAF glider?
 

Michel Van

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PMN1 said:
I've noticed a section in the Astronautix article on CEV that states that lifting-body and winged configurations that could have been recovered with horizontal landings were possible for Apollo but time was against their development - how did they compare to the USAF glider?
thats lift-to-drag ratio http://en.wikipedia.org/wiki/Lift-to-drag_ratio

USAF
Dyna Soar L/D Hypersonic: 1.50 http://www.astronautix.com/craft/dynasoar.htm
PRIME L/D Hypersonic: 1.0 http://www.astronautix.com/craft/prime.htm
NASA
Apollo R-3 L/D Hypersonic: 0.70 http://www.astronautix.com/craft/apollor3.htm
Apollo M-1 L/D Hypersonic: 0.52 http://www.astronautix.com/craft/apollom1.htm
fly Hardware
HL-10 L/D Hypersonic: 1.30 http://www.astronautix.com/craft/hl10.htm
Apollo L/D Hypersonic: 0.30 http://www.astronautix.com/craft/apolocsm.htm
 

hesham

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Hi,

Here are the individual documents on each of the SLS vehicles.
 

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OM

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...Those drawings look a *lot* like what Mark Wade has on his site.
 

RyanCrierie

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This is from a briefing given by the USAF's Ballistic Missile Division (AFBMD) to industry on 14 September 1960 that I found on the NRO's FOIA site:

----------------------------------------

True Military Operational Space launching Vehicle Concept
(PHOENIX concept vehicle)
  • Recoverable, low cost, booster stage
  • Low cost, simplified, throw-away, high energy, final stage
  • Recoverable, low cost, second stage
...
D. Launch Vehicle Systems Required for Military Space Systems (PHOENIX)
Major Henderson

The technical and operational characteristics of military space booster systems are quite different than those for ballistic missile systems or scientific payload boosters. Ballistic missiles are conceived to be in a state of instant readiness for launch during a period of hostilities. As a result, a launch facility with ground support equipment is provided for each missile. Missile facilities are designed to withstand initial attack and reliably conduct one or more launchings as soon as possible upon command. Ballistic missiles costs are amortized over a long period during which the actual vehicles comprise part of our deterrent force. Cost effectiveness is based upon the destructive effect of nuclear payloads on enemy targets.

With few exceptions, launch operations for Space systems will be conducted during peacetime. Facilities must be designed for efficient and economical conduct of many launches. Invulnerability to attack is of lesser importance and, instead of a minimum reaction time, a flexible, and predictable action time is needed. Frequent operational launches will be needed by a variety of space weapon systems. Hence, space launch vehicle systems must be flexible—able to launch a variety of payloads for a variety of purposes—and amenable to routine operations. They should be simple and capable of routine launches by military crews. The paramount need, however, is to reduce the overall cost of launch vehicle systems and launch operations by one order of magnitude or more. Otherwise, this nation cannot afford the number of launchings necessary to provide an effective military capability in Space.

These requirements are the basis of the PHOENIX concept. This concept poses a requirement for a launch vehicle system, including vehicles, facilities, support equipment and services, logistics, and operational routines which minimize overall system launch costs. A study is underway at AFBMD to determine what type of launch vehicle system will best fulfill these requirements. The key figure of merit for comparing, candidate systems is cumulative program cost divided by total payload weight or, in simple terms, dollars per useful pound in orbit.

The first phase of the study, now underway, will be complete by the end of 1960. Parametric design studies are being conducted on a number of attractive vehicle design approaches. Complete system concepts are being laid out including development and operational facilities, and logistical/operational support needs as veil as the vehicles themselves. Complete development, test, production, logistic and operational programs are being outlined. Working closely with the Aerospace Corporation, personnel from RAND are performing cost-sensitivity analysis and are costing out the complete programs. For the first phase, studies are being conducted for vehicles of 800,000 pounds first stage thrust, optimized for performance in a 300 nautical mile eastward orbit. The size and orbit are selected for convenience in study and not because of any specific requirement for 800,000 pounds launch thrust. Specific missions and payloads are not being considered in Phase I. The objective of Phase I is to enable selection of the two or three candidate vehicle systems approaches which can best fulfill requirements.

The second Phase of our study is to be completed by the end of August, 1961. Using technical scale factors and cost factors derived during Phase I, candidate systems will be analyzed over the complete size range of interests. All mission and payload planning data available will be collected and analyzed according to launch rates, time periods, and equivalent 300 nautical mile orbit payload capability required. Design optimization of all parts of each system for economy will be completed in as much detail as possible. We should be able to identify the best system, lay out a complete program to develop it, and prove its potential at the end of the study.

Our plans are tied to the objective achieving an operational capability with the first PHOENIX vehicle by late 1965. A part of the propulsion applied research program has been oriented to provide data needed for our analysis and to develop the basic technologies required should more advanced propulsion techniques prove critical to obtaining the PHOENIX objective of maximum overall reduction in costs.
 

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RyanCrierie

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AFSC Historical Publications Series 63-50-I "History of the X-20A Dyna-Soar: Volume I (Narrative)

says on the following pages:

xiv:
The Dyna-Soar Directorate of the Space System Division recommended employment of the Phoenix A388 space launch system for the Step IIA booster.

page 82:
The Dyna-Soar Directorate of the Space System Division, having the responsibility for developing boosters for System 620A, also made a recommendation on the Step IIA propulsion.

On 11 July, Colonel Joseph Pellegrini informed the Dyna-Soar office that his directorate favored employment of the projected Space Launch System A388. This proposal was an outgrowth of an SSD study on a Phoenix series of varying combinations of solid and liquid boosters to be used in several Air Force space missions. Phoenix A388 was to have a solid first stage, which could produce 750,000 pounds of thrust, and a liquid propellant second stage, using the J-2 engine.

Page 82's statement is footnoted, and the footnote is:

TWX, SSVS-10-7-3, Hqs. SSD to Hqs. ASD, 11 July 1961, Doc. 244.
 

RyanCrierie

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Mark Wade's source document dump page at Astronautix has this document:

Lunar Expedition Plan - LUNEX - May 1961, Headquarters Space Systems Division, Air Force Systems Command



Page 1.6:
A three-stage earth launch booster, referenced as a Space launching system. The first stage will use either LOX/LH2 with six million pounds of thrust or a solid fuel with an equivalent launch capability. The second and third stages will use LOX/LH2. The development of this space launching system is considered the pacing development item for the LUNEX program. Because of the magnitude of the booster program and the applicability of the booster to other programs, the plan for its development is being presented separately.

Page 1.8
CAPABILITIES DEVELOPED

The development of large boosters, rendezvous techniques and maneuverable Space vehicles, all required for the Lunar Expedition, will also provide a capability for many new and advanced space achievements. For example, the Space Launching System which will boost 134,000 pounds to escape velocity will boost approximately 350,000 pounds into a 300 nm orbit, or will launch a manned vehicle on a pass around either Mars or Venus.

Page 3.7
The Space Launching System boosters designated as A, AB and BC, and solids as required, will be needed as indicated and their payload capabilities are estimated as follows:

Booster --------------------- Payload
A 410 ---------------- 20,000 pounds (300 mile orbit)
AB 825 ---------------- 87,000 pounds (300 mile orbit)
AB 825 ---------------- 24,000 pounds (escape velocity)
BC 2720 --------------- 134,000 pounds (escape velocity)


Page 8.5
A modified Integrated Transfer Launch System is envisioned for the Lunar Transport Launch System. The size and weight of the Space Launching Vehicle, designated the BC 2720, precludes the transfer of the entire Lunar Transport Vehicle after assembly, but the Integrated transfer of upper stages and lower stages separately with a minimum mating and checkout on the launch pad may provide increased reliability and appreciable cost saving.

Page 8.6
In as much as both the "B" and "C" boosters of the Space Launch Systems have diameters In excess of 14 feet, transport from manufacturing plant to the launch site must be by barge.

Page 8.7 to 8.8
A modified Integrated-Transfer-Launch System is 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 precludes the transfer of the completely assembled Lunar Transport Vehicle from an integration building to the launch pad.
It is 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. (See Figures 8-1 and 8-2).

This can best be accomplished by a cliffside 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 may 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 is 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 is 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.



Page 8.9
[on TNT equivalence -- solids are 100% of weight in TNT equivalence, and it seems to be referring to the C2720 combination]

The TNT equivalent of one of the four segmented solid assemblies is 680,000 pounds.

Page A1.5
Space launching System
The complete system, including ground facilities, propellant manufacturing facilities, etc., as required to launch the boosters required for space operations.
 

RyanCrierie

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This is from:

AIAA-2001-2068
INFLATABLE STRUCTURES FOR DEPLOYABLE WINGS




Where it references:

Phoenix Space Launching System Study, Aerospace Corporation, 1962

and shows a diagram from it.

Some more search on google gets more information:

"Phoenix Space Launching System Study"
Phase I Final Report Volume III
Aerospace Corporation, 28 Jan 1962 (Confidential).
 

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RyanCrierie

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Bibliographic information found inside the following DTIC citiation:

http://www.dtic.mil/docs/citations/AD0408987

28. "Interim Report on Phoenix System Costs (U) - Volume I: Estimates and Procedures", Cost Analysis Department, RAND corporation, RM-2770-SSD, April 1961, CONFIDENTIAL.

29. "Interim Report on Phoenix System Costs - Volume II: Generalized Cost Model", Cost Analysis Department, RAND Corporation, RM-2770-SSD, April 1961, FOR OFFICIAL USE ONLY.

30. "Interim Report on Phoenix System Costs - Volume III: Calculations of the Generalized Cost Model:, Cost Analysis Department, RAND Corporation, RM-2770-SSD, September 1961, CONFIDENTIAL.

60. "The Phoenix Modular Booster Concept: Cost Estimates and Estimating Procedures (U)", prepared for Space Systems Division, AFSC, by RAND Corporation, R-395-SSD, February 1962, SECRET.

More information found in

http://www.dtic.mil/docs/citations/AD0627886

Cost Analysis Department, Interim Report on Phoenix System Costs Volume II: Generalized Cost Model, The RAND Corporation, RM-2770-SSD, April 1961 (For Official Use Only), 137 pp.

This memorandum describes a generalized mathematical model for estimating the cost of space and ballistic missile booster systems. The model considers "cradle-to-grave" costs. The model does not apply solely to the Phoenix system; with possible modification, it can be applied to costing systems employing different operating concepts. Cost elements are described within the category structure--R&D costs, nonrecurring operating costs, and operating costs. Included are the subroutines for computing costs of flight vehicles, ground support equipment, and military facilities.
 

archipeppe

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RyanCrierie said:
This is from:

AIAA-2001-2068
INFLATABLE STRUCTURES FOR DEPLOYABLE WINGS




Where it references:

Phoenix Space Launching System Study, Aerospace Corporation, 1962

and shows a diagram from it.

Some more search on google gets more information:

"Phoenix Space Launching System Study"
Phase I Final Report Volume III
Aerospace Corporation, 28 Jan 1962 (Confidential).


Is it related to ASSET prototypes?
The configuration, even minimal in this case, seems to be the same....
 

blackstar

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archipeppe said:
Is it related to ASSET prototypes?
The configuration, even minimal in this case, seems to be the same....

I don't think so. This looks like something that died as a study.
 

blackstar

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By the way, the newly-released NRO documents on Hexagon include one document that mentions (not by name) PRIME and ASSET and indicates that they could lead to future film return reentry vehicles for vehicles like Hexagon. This is the first time that I think we have a declassified document indicating a possible linkage between them, although it does not state that this is why PRIME and ASSET were flying, or that NRO was providing funding for them.
 

Michel Van

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blackstar said:
By the way, the newly-released NRO documents on Hexagon include one document that mentions (not by name) PRIME and ASSET and indicates that they could lead to future film return reentry vehicles for vehicles like Hexagon. This is the first time that I think we have a declassified document indicating a possible linkage between them, although it does not state that this is why PRIME and ASSET were flying, or that NRO was providing funding for them.

that's easy to explain why,
the film return capsules were caught in mid-air by an aircraft, over the Pacific Ocean.
under them were a unusual high density of Soviet fishing fleets...

A film return reentry vehicles on PRIME or ASSET, landing automatic at Vandenberg AFB, would please NRO.
So why was this not realized ? digital transmitting the Data from Spy Sat was it downfall.
 

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Michel Van said:
that's easy to explain why,
the film return capsules were caught in mid-air by an aircraft, over the Pacific Ocean.
under them were a unusual high density of Soviet fishing fleets...

A film return reentry vehicles on PRIME or ASSET, landing automatic at Vandenberg AFB, would please NRO.
So why was this not realized ? digital transmitting the Data from Spy Sat was it downfall.

The ocean is pretty big, and the Soviets never had advance warning when the capsules would come down. Certainly not enough to get a vessel into the right location.

The problem with land reentry is that there was the possibility of overshooting the target zone and landing in Mexico, or coming down in a populated area in the U.S. The early version of the GAMBIT-1 system was going to use land recovery, but they abandoned that in fall 1962. They were concerned about dropping pieces of the spacecraft over the northern US, and what might happen if the spacecraft came down in a populated area of the US or south of the border.

The technology for controlled reentry like PRIME was just too primitive at that time. I imagine that there was also a mass penalty.
 
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