Rocket Engines, concepts, experiments (past/present/future)

[RyanC]

So I found an engine program which wasn't on Astronautix by randomly reading documents on the US Space & Rocket Center's digital archives at UAH.

The document in question is:

Launch Vehicle Engines Project Development Plan MA001-A50-2H (1 January 1967)

And here's the link to download the 14 MB PDF from my site:

Link

It's spread across several places in the above document, but the concise tl;dr version is:

The C-1 Engine Project was intended to provide a 80-100 lbf pressure fed engine with an ablative or radiatively cooled nozzle capable of meeting the collective requirements of the following applications:

• Re-Entry Control for the Apollo Command Module
• Ullage settling for the S-IVB stage
• Reaction control for the S-IVB stage
• Reaction control for the Apollo Service Module
• Reaction control for the Apollo Lunar Excursion Module
• Extended mission requirements of Reaction Control Systems on AAP and post AAP flights.

Development began on 8 August 1964, with a six month competitive Definition/Demonstration phase from 5 March 1965 to September 1965 between TRW Systems Group and Reaction Motors Division of the Thiokol Chemical Corporation.

After winning the Definition/Demonstration phase, Reaction Motors Division of the Thiokol Chemical Corporation began the Development phase on 18 October 1965. At the time of the source [January 1967] this phase was to have continued for 21 months, and focused on the Apollo SM and LEM applications to allow a Flight Readiness Demonstration to be completed during the 15th program month and formal qualification to be completed at the end of the 21st program month.

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

Found another one on NTRS.

This one was the engine selected for General Electric's Apollo proposal, the Aerojet AJ10-133.

What's interesting is that it's a four chambered engine -- while reading the Redstone Missile Program histories, I found a mention of the AJ10-18, which was proposed for it, and it was a 4 chambered engine generating 160,000 lbf. It could very well be that Aerojet has been using the AJ10 system in various configurations since forever.

So onto the general specs:

Propellants: LOX/LH2
O/F Ratio: 5.0
Thrust (vac): 24,000 lbf at 430 ISP
Weight: 675~ lbs
T/W Ratio (vac): 35.55
Chamber Pressure: 65 psia
Expansion Area Ratio (ε = Ae/At): 35
Rated Burn Time: 137~ seconds at full thrust (can burn for 546 seconds on a single chamber)

Notes: Pressure-fed four chambered LH2 engine proposed by Aerojet for General Electric’s Apollo proposal. The system would have been insulated with SI-4 insulation, allowing it to operate for fourteen days as a sealed unit with minimal boil-off.

 

[Graham1973]

Re: Apollo-Era Rocket Engine Proposals

I'd classify both as very interesting and would like to suggest that the thread below about an Apollo era engine proposed by Bell Aerospace be merged into this thread.

http://www.secretprojects.co.uk/forum/index.php/topic,14880.0.html

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

Agree!
Pratt & Whitney RL20
Source Description: “A new approach to the design of a high pressure bell engine was also demonstrated in full scale model form by Pratt & Whitney.

Somewhat less exotic than Rocketdyne's approach, the P&W entry, nevertheless, exhibits some impressive innovations.
Designated RL-20, it is a restartable LOX/Hydrogen engine similar in size and general appearance to the J-2.

Aside from the high pressure combustion chamber, its main feature is a two position nozzle skirt extension.

With the skirt in a retracted position, the engine takes up less axial length in the stacked vehicle. The engine can be operated in this configuration using the regeneratively cooled stub nozzle to provide expansion for low altitude operation. At high altitude or after staging, the nozzle skirt would be extended (in 15 seconds) to increase its expansion ratio.

Bleed hydrogen is used to cool the extended skirt and is expanded to local ambient pressure in miniature nozzles at the large nozzle's outlet plane. Nearly as much thrust results from this effect as would be obtained in burning the hydrogen. Pump turbines are mounted between a preburner and the film-cooled (hydrogen) main chamber where final O2 is added. The preburner, main combustion chamber arrangement provides a progressive combustion effect which may reduce the possibility of combustion instability.”

References:
Saturn Improvement Studies – A Summary – Case 330 (28 October 1966)

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

AJ-550 was Aerojet's designation for their SSME proposal.

Also, Aerojet carried out ground testing of actual hardware for the early Space Shuttle RCS systems (back when it was known as APS).

Way back when STS was to be a two stage to orbit fully reusable system with internal LOX/LH2 tanks in the Orbiter, they proposed using two AJ10A-3-3 as the OMS and a entirely new 1,000~ lbf class Gaseous Oxygen/Gaseous Hydrogen thruster with an ISP of around 376.5 for orbital changes.

It was actually pretty ingenious. You would've had unused propellants boiling off inside the orbiter during the up to 30 day orbital period; and instead of having to vent them to maintain tank pressure, why not burn them as part of the attitude control system?

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19720064115_1972064115.pdf​

​

​

Aerojet AJ-1200 Pressure Fed Booster Engine, circa 1972.​

​

​

Quick Summary:​

​

​

Propellants: LOX/RP-1​

O/F Ratio: 2.4​

Thrust (sl): 1,200,000 lbf at 237.8 ISP at NPL​

Thrust (vac): 1,466,000 lbf at 289.8 ISP at NPL / 1,020,000 lbf at 288.1 ISP at 70% PC​

Weight: 17,484 lbs dry​

T/W Ratio (sl): 68.63​

T/W Ratio (vac): 83.84​

Chamber Pressure: 250 psia​

Expansion Area Ratio (ε = Ae/At): 5​

Engine Length: 256 inches​

Engine Diameter: 160 inches​

Nozzle Diameter: 155 inches outside diameter​

​

​

Notes: Designed for a minimum of 100 uses with a service life of 15,000 seconds. Designed to operate at two thrust levels – Normal Power Level (NPL) and a reduced thrust level of approximately 70% PC of NPL during "Max Q" transition points.​

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

Pratt & Whitney RL200 (1960)

Propellants: LOX/LH2
Thrust (Vac): 200,000 lbf
Total Installation Length: 177 inches
Engine Machinery Length: 65 inches
Nozzle Length: 94 inches
Nozzle Diameter: 90 inches

Notes: This engine was Pratt & Whitney’s proposal for the 200K LH2 Second Stage Engine contract for Apollo that was eventually won by Rocketdyne’s J-2.

Essentially, the RL200 was a “big” RL10; utilizing the shunt expander cycle, but modified with the addition of a separate gear driven inducer pump to reduce the launch vehicle’s job regarding propellant supply.

P&W submitted their proposal in mid-March 1960 for the RL200 to NASA, with a total development cost estimated to be $138 million. Development cost estimates for the other two competitors were $66 million for Aerojet and $44 million for Rocketdyne.

Apparently in the late 1990s, the RL200 designation was reused by P&W for a speculative engine design.

References:
Advanced Engine Development at Pratt and Whitney: The Inside Story of Eight Special Projects, 1946-1971 by Dick Mulready

 

[RyanC]

Re: Apollo-Era Rocket Engine Proposals

This *might* be Aerojet's 200K LH2 proposal, but I'm not paying $15 to find out.

 

[RyanC]

The goal of the first J-2X program, which was operational from 1964 to 1968, were to simplify the engine, the stage equipment associated with pre-launch operations, reduce engine ground support equipment. In addition, the engine was to have higher thrust and ISP to enable greater operational flexibility. All these goals were to be achieved without significant modifications to the S-II or S-IVB stages.


To achieve these goals, the J-2X program investigated the following areas of improvement over the J-2:

• Higher Chamber Pressures

• Extendible Nozzles
• There were three extensible nozzle concepts evaluated within the ground rules of adding no more than 750 pounds to engine weight, being able to fit into the S-IC/S-II or S-II/S-IVB interstages, and having a 2 second (or less) actuation time. They were the:

• Airmat: An inflatable extension skirt design that is fabricated from a woven wire structure manufactured by the Goodyear Aerospace Corporation. This concept features a nozzle skirt wall fabricated from two layers of conventionally woven, stainless steel wire yarn with the layers being loosely tied together by a number of interwoven wire yarn strands 10.2 cm (4.0 in.) in length. These strands hold the two wire-cloth layers from expanding outward when pressurizing gas is fed between them. The resulting effect when pressurized is a 10.2-cm (4.0-in.) thick wall that exhibits a high resistance to bending moments when fabricated into the integral cone frustum of the nozzle extension. When unpressurized, the Airmat extension is flexible and can be folded or rolled around the existing J -2 nozzle for storage within the S-IVB interstage. The inflation gas for pressurizing the wall structure comes from the J -2 turbine exhaust products that are normally dumped into the main exhaust stream through a ring of orifices around the basic J-2 nozzle.
  
  As in any woven product, the Airmat exhibits a matrix of fine openings between the woven yarn filaments that make up the two wire-cloth face layers. This porosity is utilized on the inner nozzle face, next to the hot main chamber combustion flow, to allow the turbine exhaust pressurizing gas to escape from the nozzle extension and perform a transpiration cooling function in the process. The external wall is sealed to prevent escape of the pressurizing gas.
  
  Actuation of the Airmat extension is automatic, with turbine exhaust flow providing the driving force to deploy the folded material to the fully extended position. Overall system weight estimates, including the turbine exhaust ducting and manifold, nozzle attachment fixtures, etc., indicate that the Airmat extension assembly will weight 204.12 kg (450 lb) .

• Aerobell: This nozzle extension concept required a single piece, truncated cone fabricated from refractory materials. It was to be mechanically actuated from its stowed position about the basic J -2 nozzle cone by eight pneumatic motors driving collapsible struts. These collapsible struts were to be fabricated from flexible steel tubing that would fold into a flat ribbon about a storage drum and expand into a rigid columnar strut upon being unrolled from the drum. Nozzle cooling was to be accomplished by dumping the turbine exhaust flow in at the extension attachment plane through a 360-deg circumferential slot. This flow would provide a protective film coolant boundary layer along the extension wall. This concept was estimated to have total assembly weight of 249.48 to 272.16 kg (550 to 600 lb).

• Telescoping: The telescoping extension concept composed of five truncated cone segments that nested one within another in the stowed position about the J-2 basic nozzle. These segments were linked together by six sets of spring-loaded scissor arm mechanisms. The extension was to be stored with the spring-loaded arms compressed so that the deployment command released them to actuate the conical segments to the fully extended position. In reaching the extended position, the scissor arm assemblies moved past center and mechanically locked into a compression-carrying member that was capable of transmitting the extension thrust into the basic nozzle. Cooling would be by radiation or film using the turbine exhaust products in a manner similar to the Aerobell concept. A circumferential seal was required at each of the four joints between the five extension segments, and a total assembly weight of 192.78 kg (425 lb) was predicted.

• Due to the premature cancellation of the J-2X program, an immediate downselect of nozzle extension concepts was forced.
  Because the Aerobell concept was similar in concept to the XLR-129 nozzle extension program already under research by the USAF, it was eliminated.
  The telescoping extension concept had already been modeled in full scale on a nonoperational J-2 engine to demonstrate the actuation of the spring driven scissor arms. Due to the complexity of the multiple section telescoping assembly, this concept was also eliminated, leaving only the Airmat extension for further study.
  The Airmat extension would have been attached to the existing engine nozzle at the 27.5:1 expansion ratio point, and upon inflation, would have extended it to an expansion ratio of 48:1. Payload gains for the S-IVB stage were estimated to be 3,200 pounds from this effort.
  Two full scale J-2 Airmat nozzle extensions were fabricated; one to 48:1, the other to a more conservative 41.3:1 extension. It was planned to fire a J-2 with first the 41.3:1 unit then the 48:1 unit at Arnold Engineering Development Center, but J-2 engine firings at that facility were prematurely canceled before the nozzle extensions could be fabricated. It was decided to use gaseous nitrogen on a non operational J-2 to test the nozzle extension process.
  Ultimately, the lessons learned from the experimental J-2X were applied to the production-capable J-2S design.

 

[J.A.W.]

Here, a U.S. post-war appraisal of Nazi rocket science, good enough for SS-Johnnies such as W. von Braun
to get an Operation Paperclip-to-Apollo gig..

http://www.cdvandt.org/Wunderwaffen-file-11110.pdf

 

[RyanC]

JAW's

"Here, a U.S. post-war appraisal of Nazi rocket science, good enough for SS-Johnnies such as W. von Braun
to get an Operation Paperclip-to-Apollo gig.."

Reposted with attachment to keep it from getting "lost"