Atlas "stage-and-a-half" drive system

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

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 SM-65A (Atlas-A)

The Convair SM-65A 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)

sec

Ns/kg

sec

Ns/kg

XLR-89-1

MA-0

2x1

RP-1/LOX

Booster

Sustainer

MA-1

Convair SM-65B (Atlas-B)

The Convair SM-65B 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 SM-65A 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)

sec

Ns/kg

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

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)

sec

Ns/kg

sec

Ns/kg

LR-89-NA-5

MA-2

2x1

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

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)

sec

Ns/kg

sec

Ns/kg

LR-89-NA-6

MA-3

2x1

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

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)
Engine	Other Design	Chamb.	Nozzle Area Ratio	Press. Exp. Ratio	Chamb. Press. (MPa)	Propellants	Stage	Oxid. Mix Rate	kN	sec	Ns/kg	kN	sec	Ns/kg	Flow Rate (t/sec)
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	MA-5.1	1		25	4.80	RP-1/LOX	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	MA-5.2	1		25	4.80	RP-1/LOX	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)

sec

Ns/kg

sec

Ns/kg

RS-56 OBA

MA-5A

2x1

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

5.07

Sustainer

2.27

268.6

220.4

2161

385.2

316.0

3099

0.1243