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Titan 1: ICBM, built as back-up to Atlas, using two stages instead of one and a half, and conventional tank construction in lieu of balloon tanks. Was also to have been used for suborbital tests of X-20A Dynasoar manned space plane. For unknown reasons never refurbished for use as space launcher and scrapped after being replaced by Titan II in missile role in mid-1960's.
Titan 2: (1964) Also known as SLV-5. ICBM, used as Gemini launch vehicle in 1960's. When ICBM's retired in 1980's they were refurbished and new series of launches began.
Titan 3A: (1964) Also known as SLV-5A. Titan with Transtage third stage. Core for Titan 3C.
Titan 3C: (1965) Also known as SLV-5C. Titan 3A with five segment solid motors.
Titan 3B: (1966) Also known as SLV-5B. Standard Titan core with Agena upper stage. Found to be more cost effective and higher performance than using Transtage.
Titan 3BAS2: Configuration of Titan 3B proposed by Martin in mid-1960's. Titan 3B for deep space missions with Centaur upper stage, Algol strapons for liftoff thrust augmentation. Never flown.
Titan 3C7: Variant of Titan 3C with seven segment solid motors. Proposed by Martin for precise delivery of payloads beyond Titan 3C capacity into geosynch orbit. Never flown.
Titan 3D: (1971) Also known as SLV-5D. Titan 3C without transtage.
Titan 3L2: Variant of Titan with 15 foot Large Diameter Core, 2 x 7 segment strap-ons. Man-rated, optimized for delivery of heavy payloads into LEO. Never developed.
Titan 3L4: Variant of Titan with 15 foot Large Diameter Core, 4 x 7 segment strap-ons. Man rated, optimized for delivery of 40,000 pound manned payloads into 250 nm / 50 deg space station orbit.
Titan 3M: Man-rated launch vehicle designed for MOL and other missions of the 1970's. Malfunction Detection System initiated abort procedures during launch. Also suited for launch of 'bulbous and lifting body payloads'. 7 segment UA1207 motors developed but not used until Titan 4 in 1990's. Cancelled with MOL program in 1969.
Titan 3E: (1974) Also known as SLV-5E. Titan 3D with Centaur D upper stage. Used by NASA for deep space missions in 1970's.
Titan 34B: (1975) Stretched Titan core, originally developed for Titan 3M MOL, with Agena upper stage.
Titan 34D: (1982) Stretched Titan core with 5 1/2 segment solid rocket motors. IUS (Interim/Inertial Upper Stage) solid upper stages or used without upper stages. Confusingly called Commercial Titan 3 when marketed in 1990's.
Titan 4: (1989) Developed to handle military payloads designed for launch on Shuttle from Vandenberg before the USAF pulled out of the Shuttle program after the Challenger disaster. Further stretch of core from Titan 34, 7-segment solid rocket motors (developed for MOL but not used until 25 years later). Enlarged Centaur G used as upper stage (variant of stage designed for Shuttle but prohibited for flight safety reasons after Challenger). Completely revised electronics. All the changes resulted in major increase in cost of launch vehicle and launch operations.
Titan 4B: Titan 4 with Upgraded Solid Rocket Motors replacing UA1207. Developed to improve performance for certain missions, and reduce number of field joints in motor after Challenger and Titan 34D explosions involving segmented motors.

First launch: 1-Sep-1964 (18-Jun-1965 for strap-ons)
Number launched: 156 to end-1996
Launch sites: Cape Canaveral pads 40/41; Vandenberg AFB SLC-4E/4W
Vehicle success rate: 91.0% to end-1996
Success rate, past 25 launches: 84% to end-1996


First launch: 8-Apr-1964
Number launched: 16 orbital to Jul-1996, plus 2 suborbital
Launch sites: Cape Canaveral pad 19 (deactivated), Vandenberg AFB SLC-4W
Principal uses: currently small-medium payloads into low polar orbits; ten manned Gemini launches 1965-66
Vehicle success rate: 100% to Jul-1996
LEO (185 km): 2,177 kg polar, 2,360 kg 63.5o, 3,175 kg 28.6o
Sun-synchronous (546 km polar): 3,028 kg with Star 37 kick stage
GTO: 1,043 kg with SSPS upper stage
Number of stages: 2 + optional kick stages
Overall length: 31.4 m
Principal diameter: 3.05 m
Launch mass: 153,700 kg + payload
Launch thrust: 1,913 kN sea level
Guidance: Delco Electronics Titan 4 digital inertial mounted on stage 2

Engines: refurbished gimbaled paired Aerojet LR87-AJ-5 single-start hypergolic
Length: 21.4 m
Diameter: 3.05 m
Dry mass: 4,220 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 118,300 kg
Thrust: 1,913 kN SL
Burn time: 158 s

Engines: refurbished Aerojet LR91-AJ-5 single-start hypergolic single chamber
Length: 12.2 m
Diameter: 3.05 m
Dry mass: 2,860 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 28,440 kg
Thrust: 445 kN vac
Burn time: 175 s

3.05 m diameter, 6.1-9.2 m Titan 34D-type aluminum skin/stringer payload fairing, providing a 2.83 m diameter, 9.1 m long envelope.


The Titan 2 Basic is simliar to the 2G configuration but with minimal modification to the ICBM. It incorporates new wiring and avionics, and offers the option of an attitude control system. Instead of the McDonald Douglas fairing, it uses the existing re-entry vehicle fairing.


The Titan 2S Solid Thrust Augmented version enhances the 2G's performance by adding 2-8 Castor 4A strap-ons. This requires a stage 1 extension to provide attach points. The strap-ons increase LEO polar orbit capacity to 3,700 kg, depending on inclination (3,540 kg, 185 km, 99o); performance to 1,100 km circular orbits >1,900 kg

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


First launch: 1-Jan-1990
Number launched: 4 to end-1996
Launch sites: Cape Canaveral pad 40
Principal uses: delivery of Shuttle-class payloads to LEO for perigee kick motor insertion into GTO
Vehicle success rate: 75.0%
LEO (148 x 259 km, 28.6o): 14,334 kg dual payload, 14,742 kg single payload
GTO (26.4o): 1,279 kg with PAM-D, 1,851 kg with PAM-D2 dual carrier, 4,944 kg with IUS, 4,990 kg with TOS single carrier
Space Station transfer: 11,700 kg with no upper stage
Lunar delivery: 3,400 kg with TOS
Venus delivery (type II): 2,600 kg with TOS
Mars delivery (type II): 2,430-2,600 kg with TOS
Availability: 33 months after contract go-ahead; potentially 2-3 flights/year
Cost: $110 million for a dedicated flight; $100,000 reservation deposit required; reflight insurance offered at 10% premium
Number of stages: 2 + 2 strap-ons
Overall length: 47.3 m dual carrier, 44.06 m single carrier
Principal diameter: 3.05 m
Launch mass: 680,000 kg
Guidance: Delco Systems Operations inertial, incorporated in stage 2 with Olin Aerospace's 178 N MR-107 hyrazine attitude control thrusters

Length: 27.57 m
Diameter: 3.11 m
Mass at ignition: each 250,387 kg
Propellant: UTP-30001 B solid
Propellant mass: each 210,630 kg
Thrust: each 6,227 kN vac average
Burn time: 113.7 s
Steering: secondary ignition through 24 nozzles into exhaust of nitrogen tetroxide from 3,630 kg capacity tank (8.5 m long, 1.1 m diameter), although typically less than half consumed. Injection adds about 17.8 kN thrust to each motor.
Separation: jettison at 116 s by eight 20 kN solid motors firing for 1 s

Engines: gimbaled paired Aerojet LR87-AJ-11
Length: 24.0 m
Diameter: 3.05 m
Dry mass: 7,000 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 109,700 kg
Thrust: 2,340 kN vac
Burn time: 160 s, igniting at 118 s
Attitude control: gimbaled main engines for pitch/roll; forward boattail carries four external modules, each with three nozzles, providing 3-axis control. 102 kg hydrazine, thrust 80-133 N. The thrusters also provide DV corrections and payload spinup prior to deployment
Separation: at 270 s

Engines: gimbaled Aerojet LR91-AJ-11, ignited at 269 s
Length: 9.85 m
Diameter: 3.05 m
Dry mass: 2,900 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 28,600 kg
Thrust: 467.0 kN vac
Burn time: 225 s (269-494 s after launch)
Separation: stage 2 can provide spin of up to 2 rpm in addition to optional spin table performance; payload spring-separated at 0.6 m/s

10.4 m long, 3.95 m internal diameter aluminum honeycomb/graphite epoxy fairing. The standard fairing sits atop the Payload Carrier in dual or single form. In single form, it accommodates the payload and fairing, creating a payload unit 12.69 m long. With dual carrier, the fairing is separated in halves at 280 s during stage 2 burn, when atmospheric heating has reduced to 1,135 W/m2; the forward payload is deployed 67 min after launch, the 5.59 m high aft carrier cover is jettisoned forward as a unit at 148 min for payload #2 deployment 7:50 min later. Standard Titan 34D 3.05 m diameter fairings are also available, as are spin tables providing up to 70 rpm.
Acoustic load: 142 dB peak after 10 s and during transonic flight (~54 s)
Vibration: typically 4.2 g rms for 60 s
Acceleration, longitudinal: 2.8 g strap-on burn peak, 4 g stage 1 peak, 2.5 g stage 2 peak. Payload design requirement for 6 g
Pad environmental control: 9.4-37oC, 30-50% humidity


First launch: 14-Jun-1989
Number launched 15 to end-1995
Launch sites: Cape Canaveral pad 41 for LEO/GEO missions (plus pad 40 from 1992), Vandenberg SLC-4E for NUS version from 1991 (capacity 2/year). Canaveral pad 41 was used in the 1970s for Titan 3E Centaur D-1T launches of deep space probes.
Principal uses: large military payloads into Sun-synchronous orbit from Vandenberg; large military payloads such as elint, DSP and Milstar into GEO from Canaveral, BMDO tests in lower orbits (no upper stage), NASA deep space launches
Vehicle success rate 93.3%
LEO (28.6o): 17,770 kg NUS+SRM, 21,900 kg NUS+SRMU
GEO: 4,545 kg with Centaur+SRM, 5,773 kg with Centaur+SRMU, 2,364 kg with IUS+SRM, 2,860 kg with IUS+SRMU
Sun-synchronous (VAFB): 14,090 kg NUS+SRM
Molniya (12 h, 925 x 39,450 km, 63o) from CC: IUS/SRM 3,645 kg (via 185 km, 55o), IUS/SRM 3,866 kg (via 185 km, 57o), IUS/SRM 4,189 kg (via 185 km, 63o), IUS/SRMU 4,933 kg (via 185 km, 55o), IUS/SRMU 5,253 kg (via 185 km, 57o)
Deep space: 4,545 kg at 20.5 km2/s2 (SRM)
Availability: no commercialization plans
Number of stages: (2 or 3) + 2 strap-ons
Overall length: 63.14 m for Centaur, 53.99 m for IUS, 53.99/57.04 m for NUS
Principal diameter: 3.05 m
Launch mass: Centaur: 868,644 kg (SRM), 939,301 kg (SRMU); IUS: 910,018 kg (SRM), 924,515 kg (SRMU); NUS: 906,937 kg (SRM), 910,018 kg (SRMU)
Guidance: GM's Delco Systems Operations digital interial. 4B's standardized avionics from vehicle #24 will use Honeywell system: 3 Honeywell GG 1342 RLGs, 3 Sundstrand QA 3000 accelerometers, 1750A 1 MIPS processors, 25,000 h MTBF, 30 kg, 0.03 m3, 110 W. 38 kg 250 Ah SAFT lithium-thionyl chloride battery

Contractor: UTC CSD for heavyweight steel casing SRM (Solid Rocket Motor) versions; Alliant for SRMU (Solid Rocket Motor Upgarde) lightweight filament-wound composites
Length: 34.43 m (SRM), 34.25 m (SRMU)
Diameter: 3.11 m (SRM), 3.20 m (SRMU)
Mass at ignition: each 342,800 kg (SRM), each 349,600 kg (SRMU)
Propellant: PBAN (SRM), HTPB (SRMU)
Propellant mass: each 295,500 kg (SRM), 344,400 kg (SRMU)
Thrust: each 7.117 MN vac average (SRM), 7.652 MN vac average (SRMU)
Burn time: 126.5 s (SRM), 145 s (SRMU)

Engines: gimbaled paired Aerojet LR87-AJ-11A single start pressure fed
Length: 26.38 m
Diameter: 3.05 m
Dry mass: 8,000 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 170,000 kg
Thrust: 2,434 kN vac
Burn time: 186 s (ignites at 116 s)
Separation: stage 2 hot separtion

Engines: gimbaled Aerojet LR91-AJ-11A single-start liquid propellant pressure fed
Length: 9.94 m
Diameter: 3.05 m
Dry mass: 4,800 kg
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant mass: 38,400 kg
Thrust: 472.0 kN vac
Burn time: 240 s

Centaur and IUS are the principal upper stage choices but a range of spinning solids is possible; NUS flies the LEO missions. Centaur performs three burns, to: achieve low Earth parking orbit, establish GTO and inject into GEO. The 2-stage IUS is released in LEO and performs a stage 1 burn for GTO, followed by the stage 2 injection into GEO.
Designation: Centaur or IUS (designated Titan 401 when Centaur carried, 402 for IUS)
Type: liquid (Centaur), solid (IUS)
Engines: P&W RL10-3-3A (Centaur), CSD Orbus (IUS)
Length: 8.94 m (Centaur), 5.00 m (IUS)
Diameter: 4.51 m (Centaur), 2.90 m (IUS)
Mass at ignition: 26,000 kg (Centaur), 16,200 kg (IUS)
Centaur propellants: 23,000 kg liquid oxygen/hydrogen
IUS propellant: 9,818 kg stage 1, 2,722 kg stage 2
Stage thrust: 146.8 kN vac (Centaur), 202.8 kN stage 1/82.3 kN stage 2 (IUS)
Burn time: IUS: 152 s for stage 1 GTO insertion, 289 s for stage 2 GEO insertion; Centaur: 617 s total
Attitude control: 3-axis both Centaur/IUS

Construction: isogrid aluminum 6061
Length: 17.08 m NUS, 26.23 m Centaur, 17.06 m IUS (15.25, 20.1, 23.2 m versions also available)
Diameter: 5.09 m
Mass: 4,033 kg NUS, 6,073 kg Centaur, 4,026 kg IUS
Separation: low-explosive detonating fuse along seams and 12 explosive bolts divides fairing into three sections in 0.2 s.

The specifications given below apply to Titan 3 5-1/2 segment SRMs; where Titan 4's 7-segment characteristics differ they are included in parentheses.
First flown: Oct-1982 on Titan 34D, Jan-1990 on commercial Titan 3 (Titan 4 7-segment 1989)
Number flown: 15 sets on Titan 34D + four sets on commercial Titan 3 to end-1995 (15 sets Titan 4)
Length: 27.57 m (34.43 m), each segment 3.28 m
Diameter: 3.11 m (3.11 m)
Mass at ignition: 238,000 kg (316,600 kg)
    type: 84% solids PBAN (UTP-300 1B)
    shape: forward segment has fins; other are tubular, tapering from 98.3 cm thick to 85.6 cm aft for desirable tail-off characteristics
    mass fraction: 0.885 (0.846)
Propellant mass: 210,600 kg (268,100 kg)
Burn time: 113.7 s (119.5 s)
Thrust: 6,227 kN (7,117 kN) vac average
Specific impulse: 265.2 s (272.0 s) vac
Total impulse: 545 MNs (706.8 MNs) vac
Pressure: 58 atm max (joints designed for 72.5 atm)

Hercules Aerospace (acquired by Alliant Techsystems in Mar-1995) was selected in Oct-1987 as the propulsion contractor for Titan's Solid Rocket Motor Upgrade program. Lightweight graphite composite materials replaced the steel used in current motor cases, and a high performance propellant similar to that developed for the Delta 2 strap-on is employed.
Status: delivery of the first flight set was made 1Q 1994; debut is expected 4Q 1996
Length: 34.25 m
Diameter: 3.20 m
Mass at ignition: 349,600 kg
    type: HTPB
    mass fraction: 0.985
Propellant mass: 344,400 kg
Burn time: 145 s
Thrust: 7,652 kN vac average

Configuration: twin fixed motors with individual turbo pump assemblies
Application: Titan 2 stage 1
First flown: 1962 ICBM; Sep-1988 orbital
Dry mass: 1,266 kg for full assembly
Length: 2.3 m
Maximum diameter: 1.1 m
Engine cycle: gas generator
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant flow rate: 754 kg/s
Mixture ratio: 1.93
Thrust: 1,913 kN SL paired
Specific impulse: 259 s SL
Expansion ratio: 8:1
Combustion chamber pressure: 53.7 atm
Burn time: about 165 s

Configuration: scaled-down version of stage 1 engine, featuring fixed single chamber
Application: Titan 2 stage 2
First flown: 1962 ICBM; Sep-1988 orbital
Dry mass: 472 kg
Length: 2.80 m
Maximum diameter: 1.68 m
Oxidizer: nitrogen tetroxide
Fuel: Aerozine-50
Propellant flow rate: 146.5 kg/s
Mixture ratio: 1.8
Thrust: 444.8 kN vac
Specific impulse: 312 s vac
Expansion ratio: 49.2:1
Combustion chamber pressure: 56.2 atm
Burn time: about 185 s

Application: Titan 3/4 stage 1
First flown: 1968 Titan 3, 1989 Titan 4
Dry mass: 1,874 kg (paired), 758 kg (single)
Length: 3.84 m to top of thrust structure, 3.23 m to top of turbopump assembly
Maximum diameter: 1.6 m
Mounting: gimbaled pair
Engine cycle: gas generator
Oxidizer: nitrogen tetroxide, delivered at 513 kg/s
Fuel: Aerozine-50, delivered at 268 kg/s
Mixture ratio: 1.91
Thrust: 2,340 kN vac paired
Specific impulse: 301 s vac
Expanion ratio: 15:1
Combustion chamber pressure: 55 atm
Burn time: about 200 s

Application: Titan 3/4 stage 2
First flown: 1968 Titan 3, 1989 Titan 4
Dry mass: 589 kg
Length: 281 cm
Maximum diameter: 163 cm (skirt outer diameter)
Mounting: gimbaled, turbine exhaust utilized for roll control
Engine cycle: gas generator
Oxidizer: nitrogen tetroxide, delivered at 97.0 kg/s
Fuel: Aerozine-50, delivered at 54.7 kg/s
Mixture ratio: 1.86
Thrust: 467 kN vac
Specific impulse: 316 s vac
Expanion ratio: 49.2:1
Combustion chamber pressure: 58.5 atm
Burn time: about 247 s

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

Boeing's 2-stage all-solid IUS is designed to deliver payloads of >2,000 kg directly into GTO from LEO following launch by either Shuttle or expendable Titan.
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