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THE ATLAS FAMILY
MX-774: Project MX-774 inaugurated by AAF with Consolidated-Vultee to study rocket capabilities with an ICBM as a final objective. Limited funds permitted a few test launches. These rockets demonstrated technologies that woud later be applied to the Atlas.
Atlas A: First test model of Atlas ICBM. Two booster engines, no sustainer, dummy warhead. Also known as X-11. 50% reliability in 8 flight tests.
Atlas B: (1958) Also known as X-12. First all-up version of Atlas ICBM, with jettisonable booster engines and single engine sustainer on core. '1 1/2' stage launch vehicle.
Atlas Able: (1959) Atlas with upper stage based on Vanguard second stage.
Atlas D: (1959) Also known as LV-3B. First operational version of Atlas ICBM and used as launch vehicle for Project Mercury.
Atlas Agena A: (1960) Also known as LV-3A. Agena originally called 'Hustler', based on engine for cancelled rocket-propelled nuclear warhead pod for B-58 Hustler bomber .
Atlas Agena B: (1961) Also known as SLV-3. Improved, enlarged Agena upper stage.
Atlas Centaur C: (1962) Also known as LV-3C. First test verison of Atlas with Centaur upper stage.
Atlas Agena D: (1963) Also known as SLV-3A. Further improved and lightened Agena upper stage.
Atlas Centaur D: (1963) Also known as SLV-3C; SLV-3D. Fully developed version of Atlas with Centaur upper stage.
Atlas E/F: (1966) Final operational versions of Atlas ICBM. Differed in guidance systems. Deployed as missiles from 1960 to 1966. After retirement, the ICBM's were refurbished and used over twenty years as space launch vehicles.
Atlas G/H/I: (1983) Atlas-Centaur launch vehicles using stretched, uprated Atlas core. Atlas H flown a few times without Centaur.
Atlas I: (1990) Commercial Atlas with upgraded core and Centaur upper stage.
Atlas II: (1991) Commercial Atlas using stretched core with uprated boosters and stretched Centaur upper stage. Verniers replaced by hydrazine thruster modules.
Atlas IIA: (1992) Enhanced version of Atlas II.
Atlas IIAS: (1993) Atlas IIA but with solid rocket motor strap-ons.
Atlas IIIA: (2000) Development of Atlas using Russian engines in place of booster/sustainer group used on all previous models. First stage couples unique Atlas balloon tanks and high performance Glushko engines.
Atlas IIIB: (2002) Atlas IIIA 1st stage with stretched 1 or 2-engine Centaur upper stage.
Atlas V: (2002) Based on the 3.8 m diameter Common Core Booster (CCB) powered by a single RD-180 engine and incorporating a stretched Centaur upper stage. Variants use from 0 to 5 solid rocket motor strap-ons. 400-series uses standard 4.2 m diameter Atlas payload fairing, 500-series uses larger 5.4 m diameter fairing.
ATLAS CENTAUR FAMILY RECORD
First launch: 8-May-1962
Number launched: 151 to end-2004
Launch sites: Cape Canaveral pads 36A/B; Vandenberg AFB SLC-3E from 1998
Vehicle success rate: 91.39% to end-2004
Success rate, past 25 launches: 100% to end-2004
ATLAS I SPECIFICATIONS
First launch: 25-Jul-1990
Last launch: 25-Apr-1997
Number launched: 11
Launch sites: Cape Canaveral pad 36B
Principal uses: medium-class telecom and metsat payloads into GTO
Vehicle success rate: 72.7% to end-2000
Performance:
LEO (185 km, 28.5o): 5,900 kg medium fairing, 5,700 kg large
GTO (167 x 35,786 km, 27.0o): 2,375 kg medium fairing, 2,255 kg large
Earth escape: 1,520 kg medium fairing, 1,400 kg large
Availability: typically four launches/year per pad. Launch typically 24-30 months following contract
Cost: about $60 million with delivery into GTO
Number of stages: 2-1/2 (booster engines burn in parallel)
Overall length: 42.0 m with medium fairing, 43.9 m with large
Principal diameter: 3.05 m
Launch mass: 163,900 kg with medium fairing, 164,290 kg with large
Launch thrust: 1,953 kN sea level
Guidance: Honeywell's Inertial Navigation Unit mounted on Centaur's forward equipment module performs the inertial guidance and attitude control computations for both Atlas & Centaur. Some initial Atlas 1s retained the existing Honeywell inertial unit and Teledyne flight control computer but subsequent vehicles incorporate a Honeywell ring laser
gyro INU + Gulton digital data acquisition unit, saving 36 kg and enhancing reliablity.
ATLAS I STAGE 1
Engines: Rocketdyne MA-5 propulsion system of two booster, one sustainer and two vernier single-start liquid bipropellant engines
Length: 22.16 m
Diameter: 3.05 m
Dry mass: 7,882 kg (including 3,646 kg booster section)
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 137,530 kg
Thrust: 1,953 kN SL
Burn time: 156 s boosters, 266 s sustainer
Attitude control: engines are gimbaled hydraulically to provide 3-axis control during burn
Separation: Atlas uniquely incorporates two booster engines fired in parallel with the central sustainer until the base section is jettisoned by the release of pneumatically-actuated latches at about T+156 s/5.5 g longitudinal acceleration. Separation is ensured by eight solid propellant retros around the base firing angled at 40o to the vertical to prevent spacecraft contamination.
INTERSTAGE ADAPTER
The 477 kg, 3.96 m long, 3.05 m diameter ISA supports the Centaur until separation at about 268 s is effected by a flexible linear shaped charge around the forward circumference. Construction is an aluminum skin/stringer and frame.
CENTAUR (ATLAS VERSION)
Engines: two P&W RL10A-3-3A cryogenic multiple start engines
Length: 9.15 m
Diameter: 3.05 m
Dry mass: 1,700 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 13,790 kg
Thrust: 146.8 kN vac
Burn time: 408 s for single burn direct ascent, or 312 + 93 s for dual-burn parking orbit ascent
Attitude control: Engines are gimbaled for 3-axis control during burn; 12 x 27 N hydrazine thrusters provide 3-axis control during coast, spin-up for payload separation and collision avoidance maneuver.
FAIRING/PAYLOAD ACCOMODATION
Two fairing designs are available for spacecraft protection during ascent: 4.19 m diameter, 12.22 m long, 2,005 kg; or 3.30 m diameter, 10.36 m long, 1,375 kg mass. Usable diameters are 3.65 m and 2.92 m, respectively. Both employ an aluminum skin/stringer/frame structure and non-contaminating pyro separation bolts for jettison in halves at about 205 s prior to sustainer engine cutoff when heating rate has reduced to 1,135 W/m2. On the pad, air conditioning can provide a 15-25oC environment around the spacecraft.
Acceleration load: 5.5 g maximum longitudinal, 0.4 g lateral
Acoustic load: maximum 138.9 dB overall
ATLAS II SPECIFICATIONS
First launch: 7-Jul-1991
Last launch: 16-Mar-1998
Number launched: 10
Launch sites: Cape Canaveral pads 36A/B, Vandenberg AFB SLC-3E from 1998 for access to Sun-synchronous polar and 63.4o orbits for military and Earth observation satellites
Principal uses: delivery of DSCS-3 satellites into GTO; future: high inclination missions
Vehicle success rate: 100% to end-2000
Performance:
LEO (185 km, 27.0o Canaveral): 6,780 kg medium fairing, 6,580 kg large
LEO (185 km, 90.0o VAFB): 5,510 kg large
GTO (160 x 35,786 km, 28.5o): 2,950 kg medium fairing, 2,810 kg large
Earth escape: 2,000 kg large fairing
Number of stages: 2-1/2 (booster engines burn in parallel)
Overall length: 46.8 m with medium fairing, 47.4 m with large
Principal diameter: 3.05 m
Launch mass: 187,170 kg with medium fairing, 187,560 kg with large
Launch thrust: 2,159 kN sea level
Guidance: as Atlas I
ATLAS II STAGE 1
Engines: Rocketdyne MA-5A single-start liquid bipropellant consisting of two booster engines and one sustainer
Length: 24.9 m
Diameter: 3.05 m
Dry mass: 10,282 kg (including 4,187 kg booster section)
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 156,260 kg
Thrust: 2,159 kN SL
Burn time: 169 s boosters, 277 s sustainer
INTERSTAGE ADAPTER
The ISA is similar to that on Atlas I but two hydrazine thruster modules similar to those on Centaur provide roll control; mass is 545 kg.
ATLAS II STAGE 2
Engines: as Atlas I
Length: 10.1 m
Diameter: 3.05 m
Dry mass: 2,053 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 16,780 kg
Thrust: 146.8 kN vac
Burn time: typically as Atlas I
FAIRING/PAYLOAD ACCOMODATION
As Atlas 1
ATLAS IIA
Atlas IIA was a commercial derivative of the Atlas II developed for the US Air Force. Higher performance RL10A-4-1 engines replaced the Atlas II's RL10A-3-3A engines. RL10A-4-1 engines are offered with or without extendable nozzles, which increase the engine's specific impulse providing additional performance if required. The upgraded second stage increased LEO capacity to 7,280 kg (185 km, 28.5o), and GTO performance to 3,039 kg. 23 Atlas IIA's were launched with a 100% success rate; first launch was 10-Jun-1992, last launch 4-Dec-2002.
ATLAS IIA STAGE 2
Engines: two P&W RL10A-4-1 cryogenic multiple start engines
Length: 10.1 m
Diameter: 3.05 m
Dry mass: 1,840 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 16,930 kg
Thrust: 198.4 kN vac
Burn time: 370 s
ATLAS IIAS
The Atlas IIAS solid thrust augmented version enhances the IIA's performance by adding four Castor 4A strap-ons. The strap-ons increase LEO capacity to 8,618 kg (185 km, 28.5o), and GTO performance to 3,719 kg (167 x 35,786 km, 27.0o). To end-2004, 30 Atlas IIAS's have been launched with a 100% success rate; first launch was 15-Dec-2002.
ATLAS IIAS STRAP-ONS
Length: 11.16 m
Diameter: 1.02 m
Mass at ignition: each 11,600 kg
Propellant: TP-H8299 HTPB polymer, 20% aluminum
Propellant mass: each 10,100 kg
Thrust: each 433.7 kN SL average
Specific impulse: 237.8 s SL
Burn time: 55 s
Burn sequence: one pair of strap-ons ignite at launch, burnout at 54.7 s, and are jettisoned at 77.1 s. The second pair is air-lit at 60 s, burnout at 115.3 s, and are jettisoned at 117.2 s.
ATLAS III SPECIFICATIONS
The Atlas III consists of two versions, the IIIA and IIIB. Both versions are based on a first stage incorporating 3.05 m diameter Atlas balloon tanks and a single RD-180 engine. The IIIA uses a Centaur upper stage with a single RL10A-4-1 engine. The IIIB uses a stretched Centaur upper stage for enhanced performance powered by either one (SEC) or two (DEC) RL10A-4-1 engines. The specifications given below apply to the Atlas IIIA; where the Atlas IIIB's characteristics differ they are included in parentheses.
First launch: 24-May-2000
Number launched: 6 to Feb-2005
Launch sites: Cape Canaveral pad 36
Principal uses: delivery of single payloads to LEO or GTO
Vehicle success rate: 100% to Feb-2005
Performance:
LEO (185 km, 28.5o): 8,686 kg (10,759 kg) with LPF, 8,641 kg (10,718 kg) with EPF
GTO (167 x 35,786 km, 27.0o): 4,060 kg (4,500 kg) with LPF, 4,037 kg (4,477 kg) with EPF
Number of stages: 2
Overall length: 51.9 m (53.6 m) with LPF, 52.8 m (54.5 m) with EPF
Principal diameter: 3.05 m
Launch mass: 218,127 kg to 218,295 kg (222,237 kg to 222,585 kg), + payload
Launch thrust: 3,827 kN sea level
Guidance: inertial, from Centaur upper stage
ATLAS III STAGE 1
Engines: P&W/NPO Energomash RD-180 with two gimbaled chambers
Length: 28.91 m
Diameter: 3.05 m
Dry mass: 13,725 kg
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 183,200 kg
Thrust: 3,827 kN SL
Burn time: 184 s (182 s)
INTERSTAGE ADAPTER
The 465 kg, 4.45 m long, 3.05 m diameter ISA supports the Centaur until separation is effected by a flexible linear shaped charge around the forward circumference. Construction is an aluminum-lithium skin/stringer and frame.
ATLAS III STAGE 2
Engines: P&W RL10A-4-1 cryogenic multiple start engine (2 engines for DEC)
Length: 10.06 m (11.74 m)
Diameter: 3.05 m
Dry mass: 1,720 kg (1,930 kg SEC, 2,110 kg DEC)
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 16,930 kg (20,830 kg)
Thrust: 99.2 kN vac (198.4 kN vac DEC)
Burn time: 739 s (907 s SEC, 454 s DEC)
PAYLOAD FAIRING
The payload fairing consists of a boattail, cylinderical section, conical section, and spherical cap. Construction is an aluminum skin/stringer and frame with vertical, split-line longerons that allow the fairing to separate into bisectors for jettison. Separation is by pyro bolts and spring thrusters. Large payload fairing (LPF) is 2,087 kg, 12.2 m long, and 4.2 m diameter; extended payload fairing (EPF) is 2,255 kg, 13.1 m long, and 4.2 m diameter.
ATLAS V SPECIFICATIONS
The Atlas V launch system is based on the 3.8 m diameter Common Core Booster (CCB) powered by a single RD-180 engine and incorporating a stretched Centaur upper stage with either one (SEC) or two (DEC) RL10A-4-2 engines. Variants use from 0 to 5 strap-on solid rocket boosters, and either a 4.2-m or 5.4-m diameter payload fairing. Each Atlas V has a 3-digit vehicle naming designator; the first digit indicates the usable fairing diameter (4-m or 5-m), the second digit indicates the number of strap-on SRBs (0 to 5), and the last digit indicates the number of Centaur engines (1 or 2).
First launch: 21-Aug-2002
Number launched: 6 to Aug-2005
Launch sites: Cape Canaveral pad 41
Principal uses: delivery of single or double payloads to LEO or GTO
Vehicle success rate: 100% to Aug-2005
Performance:
LEO (185 km, 28.5o): 12,500 kg (402), 10,300 kg (502), 12,590 kg (512), 15,080 kg (522), 17,250 (532), 18,955 kg (542), 20,520 kg (552)
GTO (167 x 35,786 km, 27.0o): 4,950 kg (401), 5,950 kg (411), 6,830 kg (421), 7,640 kg (431), 3,970 kg (501), 5,270 kg (511), 6,285 kg (521), 7,200 kg (531), 7,980 kg (541), 8,670 kg (551)
(Quoted performance is with EPF for 400-series and short PLF for 500-series)
Number of stages: 2 + 0 to 5 strap-ons
Overall length: 400-series: 57.4 m w/LPF, 58.3 m w/EPF; 500-series: 59.7 m w/short PLF, 62.4 m w/medium PLF
Principal diameter: 3.81 m
Launch mass: 330,625 kg (401) to 470,107 kg (431) with EPF; 333,205 kg (501) to 566,297 (552) w/short PLF
Launch thrust: from 3,827 kN SL with 0 SRBs, to 10,632 kN SL with 5 SRBs
Guidance: inertial, from Centaur upper stage
ATLAS V SOLID ROCKET BOOSTERS (SRB)
Number used: 400-series: 0 to 3; 500-series: 0 to 5
Length: 19.5 m
Diameter: 1.55 m
Mass at ignition: each 46,494 kg
Propellant: HTPB solid
Propellant mass: each 42,630 kg
Thrust: each 1,245 kN vac average, 1,361 kN at ignition
Specific impulse: 275 s vac
Burn time: 94 s
Burn sequence: SRBs are ignited at launch; first three are jettisoned at 99 s followed by the next two at 100 s.
ATLAS V COMMON CORE BOOSTER (CCB)
Engines: P&W/NPO Energomash RD-180 with two gimbaled chambers
Length: 32.46 m
Diameter: 3.81 m
Dry mass: 20,743 kg (21,173 kg for 55X configuration)
Oxidizer: liquid oxygen
Fuel: RP-1 hydrocarbon
Propellant mass: 284,089 kg
Thrust: 3,827 kN SL
Burn time: 236 s to 252 s
INTERSTAGE ADAPTER
The 400-series combines a conical CCB ISA and a short Centaur ISA. The total mass is 794 kg, overall length is 4.78 m, and the diameter is 3.83 m at the bottom by 3.05 m at the top. The 500-series combines a cylindrical CCB ISA and a large Centaur ISA. The total mass is 1,544 kg, overall length is 4.13 m, and the diameter is 3.83 m.
ATLAS V STAGE 2
Engines: one or two P&W RL10A-4-2 cryogenic multiple start engines
Length: 12.68 m
Diameter: 3.05 m
Dry mass: 1,914 kg SEC, 2,106 kg DEC
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 20,830 kg
Thrust: 99.2 kN vac SEC, 198.4 kN vac DEC
Burn time: 900 s SEC, 450 s DEC
PAYLOAD FAIRING
The 400-series utilizes the same payload fairings as the Atlas III. The 500-series uses 5-m short and medium payload fairings. The bisector fairings consist of sandwich construction with graphite epoxy face sheets and an aluminum honeycomb core. The fixed boattail is composite sandwich construction. Vertical separation is by a linear piston and cylinder activated by a pyrotechnic cord; horizontal separation is by as expanding tube shearing a notched frame, activated by a pyrotechnic cord. Short payload fairing (standard) is 4,085 kg, 20.7 m long, and 5.4 m diameter; medium payload fairing is 4,649 kg, 23.4 m long, and 5.4 m diameter. Air conditioning on the pad can provide 15-25oC.
ROCKETDYNE / MA-5
The MA-5 propulsion system comprises a YLR89-NA-7 dual-chamber liquid propellant booster engine, a YLR105-NA-7 single-chamber liquid sustainer and two YLR101-NA-15 verniers to control vehicle roll and provide final velocity and directional control following sustainer burnout. The single-start main engines are fed from the same propellant tanks but Atlas uniquely separates the two outer boosters some 172 s into flight, leaving the altitude configured sustainer to provide propulsion to near-orbital velocity.
Application: Atlas I propulsion system
First flown: 18-Nov-1961
Number flown: 215 MA-5, 152 MA-3, 118 MA-2 to end-1995
Dry mass: 1,423 kg for both boosters, 470 kg sustainer
Length: 3.40 m booster, 2.69 m sustainer (with aspirator)
Maximum diameter: 1.19 m booster, 1.17 m sustainer
Mounting: all thrust chambers are gimbaled
Engine cycle: gas generator
Oxidizer: liquid oxygen, delivered at 458 kg/s for booster, 89 kg/s for sustainer
Fuel: RP-1, delivered at 203 kg/s for booster, 39 kg/s for sustainer
Mixture ratio: 2.25 booster, 2.27 sustainer
Oxidizer turbopump: each 6,732 rpm, 1,678 kW, 68 atm discharge pressure for boosters; 10,568 rpm, 846 kW, 73 atm discharge pressure for sustainer
Fuel turbopump: each 1,116 kW, 68 atm discharge pressure for boosters; 508 kW, 73 atm discharge pressure for sustainer
Thrust: 1,882 kN for booster pair vac/1,681 kN SL; 374 kN sustainer vac/269 kN SL; verniers are 2,975 N each
Specific impulse: 292 s vac/259 s SL for boosters; 309 s vac/220 s SL for sustainer
Time to full thrust: 2.0 s
Expansion ratio: 8:1 boosters, 25:1 sustainer
Combustion chamber pressure: 44 atm booster/50 atm sustainer
Combustion chamber temperature: 3,316oC both booster/sustainer
Burn time: 167 s max booster, 368 s max sustainer
Verniers: each 24.1 kg mass, 2.24/2.97 kN SL/vac thrust, 172/231 s SL/vac Isp, 1.8 mixture ratio, 5.66 expansion ratio (exit diameter 9.65 cm), 391 s burn time
ROCKETDYNE / MA-5A
The MA-5 system was uprated for Atlas 2 by replacing the booster engines, now designated RS-56A, with Rocketdyne's RS-27. The side-mounted verniers were deleted; their roll control and final adjustment functions were assumed by thruster modules on the vehicle's interstage. Sustainer engine (RS-56SA) specifications are as for the MA-5 except that total oxidizer/fuel flow rates are 90/38 kg/s, respectively; only booster engine specifications are given below.
Application: Atlas II propulsion system
First flown: 7-Nov-1991
Number flown: 20 to end-1995
Dry mass: 1,610 kg
Length: 3.43 m
Maximum diameter: 1.19 m
Mounting: all thrust chambers are gimbaled
Engine cycle: gas generator
Oxidizer: liquid oxygen, delivered at 505 kg/s
Fuel: RP-1, delivered at 224 kg/s
Mixture ratio: 2.25
Oxidizer turbopump: each 1,903 kW, 6,730 rpm, 70 atm discharge pressure
Fuel turbopump: each 1,362 kW, 75 atm discharge pressure
Thrust: total 2,100 kN vac, 1,890 kN SL
Specific impulse: 295 s vac, 263 s SL
Time to full thrust: 2.0 s
Expansion ratio: 8:1
Combustion chamber pressure: 48 atm
Combustion chamber temperature: 3,316oC
Burn time: 167 s max
P&W / RL10A-3-3A
Application: Centaur stage of Atlas & Titan
First flown: Nov-1963 (3-3A first flight Jun-1984)
Number flown: 246 to end-1995
Dry mass: 138 kg
Length: 1.78 m
Maximum diameter: 1.02 m
Mounting: gimbaled ±4o for pitch/yaw control
Engine cycle: expander
Oxidizer: liquid oxygen, delivered at 14.0 kg/s
Fuel: liquid hydrogen, delivered at 2.79 kg/s
Mixture ratio: 5.0
Oxidizer turbopump: 11.3 kg mass, 13,100 rpm, 88 kW, 45.6 atm discharge pressure
Fuel turbopump: 34 kg mass, 32,800 rpm, 76.2 atm discharge pressure
Thrust: 73.4 kN vac
Specific impulse: 444.4 s vac
Time to full thrust: typically 2.15 s
Expansion ratio: 61:1
Combustion chamber pressure: 32.2 atm
Combustion chamber temperature: 3,340oC
Burn time: about 600 s required on Titan 4 Centaur, engine qualified to 4,000 s
P&W / RL10A-4-1, RL10A-4-2
Application: Centaur stage of Atlas IIA & III (RL10A-4-1), and Atlas V (RL10A-4-2)
Dry mass: 168 kg
Maximum diameter: 1.53 m
Oxidizer: liquid oxygen, delivered at 19.0 kg/s
Fuel: liquid hydrogen, delivered at 3.45 kg/s
Mixture ratio: 5.5
Thrust: 99.2 kN vac
Specific impulse: 450.5 s vac
Expansion ratio: 84:1
Combustion chamber pressure: 39.0 atm
P&W-NPO ENERGOMASH / RD-180
Configuration: two gimbaled chambers
Application: Atlas III and Atlas V stage 1
First flown: 1999
Dry mass: 5,480 kg
Length: 3.56 m
Maximum diameter: 3.15 m
Engine cycle: staged combustion
Oxidizer: liquid oxygen, delivered at 916.5 kg/s
Fuel: kerosene, delivered at 337 kg/s
Mixture ratio: 2.72
Feed method: High-pressure turbopump assembly feeds both chambers
Thrust: 3,827 kN SL, 4,152 kN vac, throttleable 47-100%
Specific impulse: 311.3 s SL, 337.8 s vac
Expansion ratio: 36.4:1
Combustion chamber pressure: 253 atm
Burn time: 150 s