Spaceflight Insider

The Hangar / Delta II

Delta II Photo

Photo Credit: Carleton Baile / SpaceFlight Insider

US Flag ULA Logo Delta Logo
Flight Record 153/155


The Delta II medium-lift launch vehicle was originally developed by McDonnell Douglas and is based on the Douglas Aircraft Company’s experience with the Thor intercontinental ballistic missile and the Delta launch vehicles. The first flight of a Delta II rocket took place in 1989. As of April 2017, the Delta II had 151 successful launches.

The Delta II has been used to launch a diverse array of payloads, including the following: interplanetary probes and rovers for NASA, Global Positioning System satellites for the U.S. Air Force, research and development satellites for the National Reconnaissance Office (NRO), Defense Advanced Research Projects Agency (DARPA) payloads, Missile Defense Agency payloads. The Delta II has also launched Earth-observing, science, communications, and imaging satellites for commercial and international customers, and also provided rideshare opportunities for secondary payloads belonging to universities and international organizations.

Vehicle Description

First Stage

The core of the Delta II rocket utilizes RP-1/LOX (rocket-grade kerosene/liquid-oxygen) to power a single Aerojet Rocketdyne RS-27 engine in its first stage. The RS-27A first flew in 1989, but its heritage traces back through the RS-27, which first flew in 1974. The RS-27A is a single-Start, fixed thrust liquid bi-propellant gas generator cycle main engine with two vernier engines, which provide vehicle roll control during flight as well as an additional 1,000 pounds-force (4,448 kN) of thrust to the main engine. The overall RS-27A engine system provides 200,000 pounds-force (889.6 kN) sea-level thrust upon liftoff.

The Delta II also has a Heavy configuration employing larger diameter GEM-46 solid strap-on rocket motors on the 7900-series vehicle to further improve payload performance. This configuration is designated as 7920H for two-stage missions and 7925H for three-stage missions.

Graphite Epoxy Motors

The Delta II’s standard solid rocket motor is the Orbital ATK Graphite Epoxy Motor (GEM), available in two variants: The 40-inch core diameter GEM-40, which is flown on the 732x, 742x, and 792x vehicle configurations; and the 46-inch core diameter GEM-46, which is flown on “Heavy” configurations and burns approximately 14 seconds longer than the GEM-40. Both types of GEMs are flown with a fixed nozzle that is canted 10 degrees outboard from the vehicle centerline.

The Delta II 792x vehicle configuration includes nine GEMs to augment first-stage performance. Six GEMs are ignited at liftoff, while the remaining three GEMs, with extended nozzles, are ignited in flight after burnout of the first six. The 732x and 742x vehicles include three or four GEMs respectively, all of which are ignited at liftoff.

Upper Stages

The second stage is powered by an Aerojet Rocketdyne AJ10-118K engine. This engine uses nitrogen tetroxide (N2O4) and Aerozine-50 storable propellants and has an ablative thrust chamber. The rated thrust is 9,850 lbf (43.8 kN) at altitude.

Typical two- and three-stage missions use two second-stage Starts, but the restart capability has been used as many as six times on a single mission, for a total of seven burns. During powered flight, the second-stage hydraulic system gimbals the engine for pitch and yaw control. A redundant attitude control system (RACS) uses nitrogen gas to provide roll control. The RACS also provides pitch, yaw, and roll control during unpowered flight.

Depending on payload requirements, the Delta II can use an optional spin-stabilized third-stage motor. The long-nozzle Orbital ATK Star 48B motor uses a high-energy, solid propellant and high-strength titanium case with forward and aft mounting flanges and multiple tabs for attaching external hardware. The submerged nozzle uses a carbon-phenolic exit cone and a 3D carbon-carbon throat. The third-stage Payload Attach Fitting (PAF) mates the third stage to the spacecraft.

The Delta II is capable of launching between 4,120 to 7,640 pounds (1,870 to 3,470 kg) to low-Earth orbit (LEO).


The Delta II’s numerous configurations are based on the types and numbers of engines, boosters, and stages aboard. Its designations appear thus: 7###-##

The Delta II 7000-series owes its “7” to the RS-27A rocket engine used in the launch vehicle’s first stage.

The second digit in the designation is determined by the number of GEMs attached to the Delta II’s first stage. The GEMs used usually number 3, 4, or 9.

The third digit designates the second-stage Aerojet Rocketdyne AJ10-118K engine package.

The fourth digit designates the type of third stage:

  • 0: No third stage
  • 0H: No third stage, “Heavy” configuration with GEM-46
  • 5: Orbital ATK Star 48B solid motor
  • 5H: Orbital ATK Star 48B solid motor, “Heavy” configuration with GEM-46
  • 6: Orbital ATK Star 37FM solid motor

The numbers following the hyphen (-) signify the payload fairing type:

  • 9.5: 9.5 feet (2.9 m) diameter, 27.8 feet (8.5 m) length
  • 10: 10 feet (3 m) diameter, 29.1 feet (8.9 m) length
  • 10L: 10 feet (3 m) diameter, 30.4 feet (9.2 m) length

Example: Delta II 7925 would be a Delta II core first stage with nine GEM strap-on SRMs, an AJ10-118K second stage engine, Star 48B third stage, and a 10-foot diameter by 29.1-foot long payload fairing.

Mission Profile

Launch Sites

Depending on the specific mission requirement and range safety restrictions, the Delta II 7300-, 7400-, and 7900-series vehicles could be launched from either the Eastern Range or Western Range launch site.

Eastern Launch Site: The Delta II’s Eastern Range launch site for Delta II was Space Launch Complex 17 (SLC-17), launch pads A and B, at CCAFS in Florida. This site could accommodate flight azimuths in the range of 65 to 110 degrees, with 95 degrees being the most commonly flown. The final Delta II flight from SLC-17 was conducted in 2011, with all remaining flights to take place from Vandenberg Air Force Base.

Western Launch Site: The Western Range launch site for Delta II is Space Launch Complex 2 (SLC-2) at Vandenberg Air Force Base in California. Flight azimuths in the range of 190 to 225 degrees are currently approved by the 30th Space Wing, with 196 degrees being the most commonly flown.

Launch Profiles

7300 Series: At liftoff, the first stage RS-27A engine and three solid rocket motors (SRMs) ignite. The SRMs are jettisoned following burnout. The main engine continues to burn until main engine cutoff (MECO) at propellant depletion.

7400 Series: At liftoff, the first stage RS-27A engine and four strap-on solid rocket motors are ignited on the ground. The remaining vehicle sequence of events is approximately the same as that of the 7300 series.

7900 Series: In launches from both the Eastern and Western Ranges, the first stage RS-27A main engine and six of the nine strap-on solid rocket motors ignite on the ground at liftoff. Following burnout of these six SRMs, the remaining three are ignited. Once vehicle and range safety constraints have been satisfied, the six spent SRMs automatically jettison in sets of three. The second set of motors jettisons one second after the first set. The remaining three SRMs jettison approximately 3 seconds after burnout. The main engine then continues to burn until MECO.

7900H Series: The 7900H-series Delta II is available in both two- and three-stage configurations only for launches from CCAFS. With the exception of the solid-rocket motor burn durations (which are approximately 14 seconds longer), the sequence of liftoff events is approximately the same as that of the 7900-series vehicle.

The remainder of the two- and three-stage mission profiles for the 7300-, 7400-, and 7900-series vehicles are nearly identical. Eight seconds after MECO, the first stage separates and is dropped into the ocean; the second stage ignites five seconds later. The payload fairing is jettisoned early in the second-stage flight, once an acceptable free-molecular-heating rate has been reached.

In the typical two-stage mission, the second stage burns for approximately 340 to 420 seconds, after which second-stage engine cutoff (SECO-1) occurs. The vehicle then follows a Hohmann transfer trajectory to the appropriate LEO altitude. Near the transfer orbit’s apogee, the second stage restarts and completes its burn to inject the payload into the desired orbit. Approximately 250 sec. after second-stage engine cutoff (SECO-2), the second stage separates once the spacecraft achieves the proper separation attitude.

On a typical three-stage mission to geosynchronous transfer orbit (GTO), the first burn of the second stage places the payload into a 100-nautical-mile (185 km) circular parking orbit inclined at 28.7 degrees. The second stage restarts once the vehicle coasts to a position near the equator. Following SECO-2, the third stage spins up, separates, and burns to achieve GTO. At apogee, the spacecraft’s propulsion circularizes the orbit to GEO. The inclination may be removed or apogee altitude raised to optimize satellite lifetime, depending on mission requirements and spacecraft mass.

After payload separation, the Delta second stage restarts to deplete any remaining propellants and/or to move the stage to a safe distance away from the spacecraft.

The Delta II’s second stage can restart multiple times, enabling it to provide flexible orbital trajectories or the ability to launch multiple spacecraft, if needed.

Vehicle Status

The Delta II Program has launched 153 times with a success rate of 98 percent.

The U.S. Air Force used the Delta II to launch GPS satellites from 1989 to 2008. However, once the last of the second-generation satellites was launched, the USAF Medium Launch contract ended. The end of the contract increased the price per launch of Delta II, though some of the costs were mitigated by the Air Force’s elimination of staff needed to meet a 40-day call-up window for the rocket.

Despite the end of the Air Force contract, NASA put Delta II on its Launch Services contract in 2011 to support the agency’s science launches.

NASA ordered additional Delta II launches in 2012 to support the Orbiting Carbon Observatory (2014), Soil Moisture Active Passive/SMAP (2016), Joint Polar Satellite System/JPSS (2017), and ICESat-2 (2018) missions.

Beyond 2018, ULA is planning to phase out the Delta II, along with the Delta IV Medium in the 2018–2022 time frame, when ULA’s Next Generation “Vulcan” launch vehicle is expected to enter the market.


Launch History
Status Operational
Launch sites Cape Canaveral AFS SLC-17
Vandenberg AFB SLC-2
Total Flights 153*
Successes 151*
Partial Failures 1* (KoreaSat – August 5, 1995)
Failures 1* (GPS IIR – January 17, 1997)
First Flight Delta 6000: February 14, 1989 (GPS II-1)
Delta 7000: November 26, 1990 (GPS IIA-1)
Delta 7000H: July 8, 2003 (MER-B)
Notable Payloads Mars Pathfinder
Mars Exploration Rovers
Phoenix Mars Lander
GPS II Constellation

The Delta II has launched 46 GPS satellites for the U.S. Air Force, as well as numerous commercial payloads. The vehicle has been a workhorse for NASA’s science missions, including the Mars rovers Spirit and Opportunity, Deep Space 1, Mars Pathfinder, Fermi/GLAST, Genesis, Phoenix Mars Lander, Stardust, the twin GRAIL spacecraft, and NPOESS Preparatory Project.

Height 38.2 - 39.0 m
Diameter (Core) 2.4 m
Diameter (Payload Fairing) 2.9 - 3.0 m
Mass 151,700 - 231,870 kg
Stages 2-3
First Stage
Manufacturer Aerojet Rocketdyne
Engine RS-27A
Fuel RP-1/LOx
Length 26.1 m
Diameter 2.4 m
Gross Mass 101,800 kg
Thrust 889.6 kN
Specific Impulse 255 sec.
Burn time 265 sec.
Second Stage
Manufacturer Aerojet Rocketdyne
Engine AJ10-118K
Fuel Aerozine 50 / N2O4
Length 268.2 cm
Diameter 152.24 cm
Propellant Mass 5,443.1 kg
Total Mass 6,304.9 kg
Thrust 43,383 kN
Specific Impulse 320.5 sec.
Burn Time 431.6 sec.
Third Stage (7xx5)
Manufacturer Orbital ATK
Engine Star 48B
Fuel TP-H-3340
Length 203.2 cm
Diameter 124.5 cm
Propellant Mass 1,739 - 2,009 kg
Total Mass 1,863 - 2,134 kg
Thrust 76.1 kN
Specific Impulse 286 sec.
Burn Time 84.1 sec.
Third Stage (7xx6)
Manufacturer Orbital ATK
Engine Star 37FM
Fuel TP-H-3340
Length 168.9 cm
Diameter 93.5 cm
Propellant Mass 1,025 - 1,066 kg
Total Mass 1,107 - 1,148 kg
Thrust 54.8 kN
Specific Impulse 289.8 sec.
Burn Time 62.7 sec.
Solid Rocket Boosters
Manufacturer Orbital ATK
Engine GEM-40
Fuel Graphite Epoxy
Length 1,295.4 cm
Diameter 101.6 cm
Mass 13,101.1 kg
Thrust (Vacuum) 645 kN
Specific Impulse 245.5 sec.
Burn Time 62 sec.
Solid Rocket Boosters (Heavy)
Manufacturer Orbital ATK
Engine GEM-46
Fuel Graphite Epoxy
Length 1,465.6 cm
Diameter 116.8 cm
Mass 19,141.6 kg
Thrust (Vacuum) 885.2 kN
Specific Impulse 273 sec.
Burn Time 76 sec.
Two-Stage Capability
Variant LEO1 (CCAFS) [kg] LEO2 (VAFB) [kg] SSO3 (VAFB) [kg]
7320-9.5 / 7320-10 2,809 / 2,703 2,063 / 1,982 1,651 / 1,579
7420-9.5 / 7420-10 3,185 / 3,099 2,436 / 2,351 1,966 / 1,895
7920-9.5 / 7920-10 / 7920-10L 5,030 / 4,844 / 4,805 3,755 / 3,639 / 3,599 3,123 / 3,017 / 2,984
7920H-9.5 / 7920H-10 / 7920H-10L 6,097 / 5,959 / 5,899 Not available from VAFB

1 Low Earth Orbit = 100 nmi / 185 km circular, 28.7 inclination (from CCAFS)
2 Low Earth Orbit = 100 nmi / 185 km circular, 90.0 inclination (from VAFB)
3 Sun-Synchronous Orbit = 450 nmi / 833 km circular, 98.7 inclination (from VAFB)

Three-Stage Capability
Variant GTO1 (CCAFS) [kg] Interplanetary Transfer
Orbit2 (CCAFS) [kg]
Molniya Orbit3 (VAFB) [kg]
7325-9.5 / 7325-10 N/A4 / N/A4 N/A4 / N/A4 N/A4 / N/A4
7326-9.5 / 7326-10 934 / 898 629 / 604 636 / 611
7425-9.5 / 7425-10 1,110 / 1,073 805 / 779 N/A4 / N/A4
7426-9.5 / 7426-10 1,058 / 1,029 711 / 692 734 / 709
7925-9.5 / 7925-10 / 7925-10L 1,819 / 1,747 / 1,739 1,265 / 1,211 / 1,207 1,177 / 1,143 / 1,131
7926 / 7926-10 / 7926-10L 1,660 / 1,581 / 1,578 1,121 / 1,065 / 1,064 1,056 / 1,022 / 1,012
7925H-9.5 / 7925H-10 / 7925H-10L 2,171 / 2,123 / 2,102 1,508 / 1,474 / 1,460 Not available from VAFB
7926H / 7926H-10 / 7926H-10L 1,981 / 1,934 / 1,916 1,333 / 1,302 / 1,290 Not available from VAFB

1 GTO (Geosynchronous Transfer Orbit) = 100 19,323 nmi / 185 35,786 km, 28.7 inclination (from CCAFS)
2 Interplanetary Transfer Orbit = C3 0.4 km2/sec2, 28.7 inclination (from CCAFS)
3 Molniya Orbit = 200 21,649 nmi / 370 40,094 km, 63.4 inclination (from VAFB)
4 Not available; exceeds maximum allowable STAR 48B motor offload capability

Other Notes:
Star 48B uses a 3712A payload attach fitting with a mass of 100 lb (45.4 kg)
Star 37FM uses a 3724C payload attach fitting with a mass of 125 lb (56.7 kg)