Variants of the "stage and a half" drive system (MA) of the Atlas rocket
 

Norbert Brügge, Germany

Unwilling to take the risk of building a multistage missile that might later prove unworkable, Convair built the Atlas rocket around its unique "stage-and-a-half" propulsion system.

This "stage-and-a-half" propulsion system in which three engines – two boosters and a sustainer engine – are fed by the same liquid oxygen/RP-1 (kerosine mixture) propellant tanks and all ignited at liftoff. During the first few minutes of flight, the boosters shut down and fall away (to save weight), while the sustainer continues burning.

The two booster engines LR-89 were developed by Rocketdyne. Combustion chamber and nozzle were made by nickel welded tubes. Characteristic of the MA-1, MA-2; MA-3 and MA-5 systems are different drive combinations of the two LR-89 engines:

  • MA-1:    Both engines are powered centrally by one set of turbo pumps with gas generator.

  • MA-2:    Both engines are powered by own set of turbopumps. All turbo pumps are driven by a single gas generator, which are located at one of the engines.

  • MA-3:    Both engines are powered by own set of turbo pumps and gas generator.

  • MA-5:    Both engines are powered by two sets of turbo pumps with a common gas generator. All turbo pumps are arranged at one of the engines.

  • MA-5A: Both engines (new RS-27 combustor) are powered by two sets of turbo pumps with a common gas generator. All turbo pumps are arranged at one of the engines.

A new development was the central engine LR-105 (Sustainer). The thrust was lower, but it was the first engine, which was specifically designed to operate in a vacuum. It had a higher outflow velocity than the two booster engines. The gas turbine with its own gas generator operates at a high rotational speed of 10,800 rev / min (Booster only 6.300 U / min). The central engine could be swung on gimbals in two axes by 3 degrees. The burning time of the LR-105 varied depending of the time of dropping the booster engines.


The variants of the MA drive system and their use.

  MA-0


     
              
  Convair X-11 (Atlas-A)
 

The Convair X-11 was the first testbed and was developed into the Atlas-A. The Atlas-A flights were powered by a single engine consisting of two thrust chambers XLR-89 fed by a single set of turbopumps.
The central sustainer engine XLR-105 still missing.

 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

XLR-89-1

MA-0

2x1

 

 

 

RP-1/LOX

 Booster

 

 

 

 

 

 

 

 

 -

-

 -

 -

 -

 Sustainer

 -

 -

 -

 -

-

 -

-

-

 

  MA-1


   
Convair X-12 (Atlas-B)

The Convair X-12 was the second, more advanced testbed for the Atlas rocket program. It was designed with 2 engines, the booster engine XLR-89 used on the predecessor X-11 plus a sustainer engine XLR-105. This combination of booster plus sustainer engines was designated the MA-1 engine system. MA-1 was used in Atlas-B and first series of  Atlas-C.


(X)LR-105 combustor

 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

LR-89-NA-3

MA-1

2x1       RP-1/LOX

 Booster

2.25 1,316.7 248.0 2432 1,497.3 282.0 2765 0.5414

LR-105-NA-3

1

     

 Sustainer

2.27 244.7 215.0 2108 352.2 309.0 3030 0.1166

 



XLR-89 engine



XLR-105 engine (no aspirator !)

 
  MA-2



LR-89-NA-5



LR-105-NA-5 exaust tube



Atlas-D/Mercury


Atlas-C/Able



 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

LR-89-NA-5

MA-2

2x1   8 3.92 RP-1/LOX

 Booster

2.25 1,334.5 248.0 2432 1,517.4 282.0 2765 0.5487
 LR-105-NA-5

1

  25 4.41

 Sustainer

2.27 253.1 215.0 2108 366.1 311.0 3050 0.1200

 

 
  MA-3


      


    
Atlas-E/F ICBM


    
    

 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

LR-89-NA-6

MA-3

2x1   8 4.04 RP-1/LOX

 Booster

2.25 1,467.9 256.0 2510 1,662.8 290.0 2844 0.5847
 LR-105-NA-6

1

  25 4.71

 Sustainer

2.27 253.1 215.0 2108 366.1 311.0 3050 0.1200

 



LR-89-NA-6


     
 

 
    
LR-105-NA-6; Openings in the aspirator are different

 

  MA-5


Atlas-G/Centaur



LR-105-NA-7



LR-89-NA-7
 (On this part: two sets of turbo pumps and a common gas generator)


 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

LR-89-NA-7.1

MA-5.1

2x1   8 4.12 RP-1/LOX

 Booster

2.25 1,645.9 258.0 2530 1,863.4 292.2 2865 0.6505
 LR-105-NA-7.1

1

  25 4.80

 Sustainer

2.27 268.7 220.4 2161 385.2 316.0 3099 0.1243

LR-89-NA-7.2

MA-5.2

2x1   8 4.12 RP-1/LOX

 Booster

2.25 1,679.2 259.1 2541 1,901.6 293.4 2877 0.6609
 LR-105-NA-7.2

1

  25 4.80

 Sustainer

2.27 269.1 220.0 2157 386.4 316.0 3099 0.1243

 

 
   MA-5A



 



Atlas-IIAS



RS-56 OBA combustor (RS-27 derivative)

 

Engine

Other
Design

Chamb.

Nozzle
Area Ratio

Press.
Exp. Ratio

Chamb. Press.
(MPa)

Propellants

Stage

Oxid.
Mix Rate

Thrust s.l.

Isp s.l.

Thrust vac.

Isp vac

Flow  Rate
(t/sec)

kN

sec

Ns/kg

kN

sec

Ns/kg

RS-56 OBA

MA-5A

2x1   8 4.71 RP-1/LOX

 Booster

2.25 1,906.0 262.1 2570 2,155.2 296.4 2907 0.7415
 RS-56 OSA

1

  25 5.07

 Sustainer

2.27 268.6 220.4 2161 385.2 316.0 3099 0.1243