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Veronique: French Veronique sounding rocket launched from Columb Bechar to an altitude of 104 km.
Europa I: (1964) Europe's first space launcher. First stage British Blue Streak IRBM, second stage French, third stage German. All orbital launch attempts failed due to unreliability of third stage. Project cancelled after withdrawal of British support and replaced by Ariane.
Europa II: (1971) As Europa I but with solid propellant fourth stage.
Diamant A: (1965) First French indigenous satellite launch vehicle.
Diamant B: (1970) Upgraded Diamant with liquid stage; antecedent of Ariane.
Ariane 1: (1979) First successful European commercial launch vehicle, developed from L3S Europa launch vehicle replacement design.
Ariane 2: (1983) As Ariane 1 but with stretched third stage.
Ariane 3: (1984) As Ariane 2 but with solid rocket motor strap-ons.
Ariane 4: As Ariane 2 with stretched first and third stages. See Ariane 40, 42L, 44L, 42P, 44P, 44LP for figures on varous combinations of solid and liquid propellant strap-on motors.
Ariane 44LP: (1988) Ariane 4 with 2 liquid rocket + 2 solid rocket strap-ons.
Ariane 44L: (1989) Ariane 4 with 4 liquid rocket strap-ons.
Ariane 40: (1989) Ariane 4 with no strap-on motors. A fully fueled Ariane core cannot lift off the ground without strap-on liquid or solid motors. When Ariane 4 is launched in this configuration, the propellant tanks of the first and second stages are not completely filled.
Ariane 42P: (1990) Ariane 4 with 2 solid rocket strap-ons.
Ariane 44P: (1991) Ariane 4 with 4 solid rocket strap-ons.
Ariane 42L: (1993) Ariane 4 with 2 liquid rocket strap-ons.
Ariane 5: (1996) Completely new design.

First launch: 24-Dec-1979
Number launched: 83 thru Feb-1996
Launch sites: Kourou, French Guiana; ELA-1 (Ariane 1-3; site closed Dec-1989), ELA-2 (Ariane 2-4), and ELA-3 (completed 1995 for Ariane 5)
Vehicle success rate: 91.57% thru Feb-1996
Success rate, past 25 launches: 92.0% thru Feb-1996


First launch: 15-Jun-1988
Number launched 55 thru Feb-1996
Launch sites: Kourou ELA-2
Principal uses: delivery of 1,900-4,800 kg payloads into GTO for apogee kick motor insertion into GEO
Vehicle success rate: 94.55% thru Feb-1996
Dedicated 7o GTO: 1,900 kg (A40; H10+ 2,020; H10-3 2,105), 2,600 kg (A42P; H10+ 2,740; H10-3 2,930), 3,000 kg (A44P; H10+ 3,290; H10-3 3,465), 3,200 kg (A42L; H10+ 3,350; H10-3 3,480), 3,700 kg (A44LP; H10+ 4,030; H10-3 4,220), 4,200 kg (A44L; H10+ 4,460; H10-3 4,820)
Dual 7o GTO: 1,700 kg (A40), 2,400 kg (A42P), 2,600 kg (A44P), 2,800 kg (A42L), 3,300 kg (A44LP), 3,800 kg (A44L)
Sun-synchonous: 2,740 kg (A40), 3,400 kg (A42P), 4,100 kg (A44P), 4,500 kg (A42L), 5,000 kg (A44LP), 6,000 kg (A44L)
LEO 7o: 4,600 kg (A40), 6,000 kg (A42P), 6,500 kg (A44P), 7,000 kg (A42L, 44LP, 44L - structurally limited)
Escape: 2,600 kg for 5 km2/s2 missions
Availability: 12 launches per year
Cost: $40-50 million for 1,200-1,600 kg (A40), $55-65 million for 2,000-2,500 kg (A42P), $65-80 million for 2,500-3,000 kg (A44P), $90-110 million for >3,000 kg (A42L and above)
Number of stages: 3 + up to 4 strap-ons
Overall length: 56.35-60.13 m depending on fairing/payloads
Principal diameter: 3.80 m
Launch mass: 245,000 kg (A40), 324,000 kg (A42P), 356,000 kg (A44P), 363,000 kg (A42L), 421,000 kg (A44LP), 484,000 kg (A44L)
Launch thrust: 2,720 kN (A40), 3,944 kN (A42P), 4,060 kN (A42L), 5,140 kN (A44P), 5,270 kN (A44LP), 5,400 kN (A44L)
Guidance: Matra Marconi Space inertial in stage 3 Vehicle Equipment Bay, with Sextant laser gyro providing attitude data from stage 1 separation. Some 600 parameters are telemetered on 2.1 GHz. Ariane follows pre-programmed attitude profile during first stage burn.

Pairs of PAP strap-ons are carried by the 42P/44P/44LP variants (44P two pairs).
Length: 12.05 m
Diameter: 1.07 m
Mass at ignition: each 12,660 kg
Propellant: CTPB 1613
Propellant mass: each 9,500 kg
Thrust: each 650 kN SL average
Burn time: 33 s (igniting at launch, 4.2 s after the main engines)
Separation: two spring pairs (66 kN forward, 59 kN aft) ensure 6o/s rotation away from base at 5 m/s separation rate. Separation in pairs 89 s after launch (42P), 80 + 70.3 s (PAP 1/3 + 2/4, 44P) and 67 s (44LP)

Pairs of PAL strap-ons are carried by the 42L/44L/44LP variants (44L two pairs).
Engine: single SEP Viking 6, canted at fixed 10o
Length: 18.60 m
Diameter 2.22 m
Dry mass: each 4,550 kg
Oxidiser: nitrogen tetroxide
Fuel: UH25 (UDMH + 25% hydrazine)
Propellant mass: each 39,000 kg (A44L/44LP), 37,000 kg (A42L)
Thrust: each 670 kN SL at launch, 753 kN vac average
Burn time: 142 s (igniting with main engines, 3.4 s before launch)
Separation: the PALs are pyrotechnically released at 149.1 s/37.5 km after burnout at 143.6 s and separated by six solid BPD motors identical to the stage 1 retrorockets

Designation: L220
Engine: four SEP Viking 5, canted at fixed 3.8o
Length: 28.39 m (including 3.31 m interstage 1/2)
Diameter: 3.80 m
Dry mass: 17,515 kg
Oxidizer: nitrogen tetroxide
Fuel: UH25 (UDMH + 25% hydrazine)
Propellant mass: 167,500 kg (A40), 217,200 kg (A42P), 201,000 kg (A42L), 227,100 kg (A44P/44LP/44L)
Thrust: 2,718 kN SL
Burn time: 150 s (A40), 196 s (A42P), 181 s (A42L), 205 s (A44P/44LP/44L)
Attitude control: engine gimballing
Separation: stage 1/2 separation is commanded by the VEB when stage 1 thrust decays to 50%. The four stage 2 acceleration solids fire, followed 2 s later by the interstage explosive cord + stage 1's eight BPF solid retros. Stage 2's Viking 4 ignites after a further 1 s, its solids burn out 4 s later and are jettisoned with a 4 s delay

Designation: L33
Engine: single SEP Viking 4
Length: 11.61 m
Diameter: 2.60 m
Dry mass: 3,400 kg
Oxidizer: nitrogen tetroxide
Fuel: UH25 (UDMH + 25% hydrazine)
Propellant mass: 34,600 kg
Thrust: 798 kN vac
Burn time: 125 s
Attitude control: pitch/yaw by engine gimballing; roll by two sets of 50 kN thrusters mounted on rear skirt fed from Viking's gas generator. Guidance begins about 10 s after stage 1 separation (stage 1 executes pre-programmed profile)
Separation: at about 344 s, 147.4 km, 5,379 m/s, assisted by two solid deceleration rockets; sequence begins when stage 2 has added required delta V

Specifications are given in H10/H10+/H10-3 order.
Designation: H10/H10+/H10-3
Engine: single cryogenic open cycle SEP HM-7B
Length: 10.73 m/11.05 m/11.05 m
Diameter: 2.60 m
Dry mass: 1,200 kg/1,240 kg/1,240 kg, excluding interstage 2/3
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 10,800 kg/11,140 kg/11,860 kg
Thrust: 63 kN vac/63.2 kN vac/64.8 kN vac
Burn time: 720 s/750 s/780 s, until desired GTO has been achieved in single direct ascent burn beginning about 347 s after launch
Attitude control: the Roll & Attitude Control System (SCAR) maintains roll control during main engine burn (engine gimballing provides pitch/yaw) and 3-axis following cut-off. SCAR consists of two sets of diametrically opposed triple thrusters fed with gaseous hydrogen; total impulse >15 kNs. Two aft-facing nozzles depressurize the LOX tank after cut-off and are sequenced as part of the attitude control system and collision avoidance maneurver following payload separation. Spin stabilization of about 10 rpm is provided for payload deployment

Matra Marconi Space's 520 kg 1.037 m high VEB performs guidance, data processing, sequencing, telemetry, tracking and destruction function. A digital flight control system replaced Ariane 2/3's analog equipment.

Three basic lengths of Contraves' 25 mm-thick aluminum alloy honeycomb two-piece fairings are available: 8.6 m (type 01, 740 kg, 60 m3), 9.6 m (type 02, 800 kg, 70 m3) and 11.1 m (type 03, 86 m3, available on special request). The fairing halves are separated by pyrotechnic cord and piston when the heating flux has reduced to 1,135 W/m2, at about 285 s/115.3 km during stage 2's burn. The base clampband is released first. Several standard payload adapters are available.
Acceleration load: 4.0 g at end stage 1 burn, 4.1 g at end stage 2, 1.6 g end stage 3 (for A40); 7.3 g rms max vibration
Acoustic load: 142 dB integration at launch + transonic
Thermal load: max 500 W/m2 radiated by fairing/Spelda
Ariane 3 used the Sylda for dual launches, encapsulating the lower satellite and mounting the forward payload on its front face. Ariane 4 introduced the larger aluminum honeycomb Spelda as part of the fairing assembly. The lower 3.97 m diameter, 2.78 m high cylinder (Short Spelda) is attached to the VEB and topped by an interface for the upper payload. Short Spelda masses 408 kg, provides a 34 m3 payload envelope, and can carry a 2,800 kg forward load. Separation is achieved by pyrotechnic line charge around the circumference and six springs. The 458 kg Long Spelda adds a further 1 m cylindrical section, provides 42 m3 payload envelope, and can carry a 1,900 kg upper load. The 1.78 m-long Mini-Spelda can carry 24 m3 payloads of 400-1,000 kg. This 337 kg unit can handle a 3,200 kg forward load. The 350 kg Mini-Spelda+300 version stretches Mini-Spelda's cylinder by 300 mm for 26 m3 payloads.


First launch: 4-Jun-1996
Number launched 2 thru end-1997
Launch sites: Kourou ELA-3
Vehicle success rate: 50% thru end-1997
GTO (7o): 6,800 kg single or 5,970 kg double payload version
Sun-synchronous (800 km, 98.6o): 10,000 kg
Space Station transfer (70 x 300 km, 51.6o): 18,000 kg
Lunar transfer: 4,450 kg dedicated
Shared GTO mission: 400 kg lunar orbiter + 200 kg-payload lunar lander
Number of stages: 2 + 2 strap-ons
Overall length: 45.71-51.37 m
Principal diameter: 5.40 m
Launch mass: 746,000 kg
Launch thrust: 11.4 MN sea level
Guidance: provided by the Vehicle Equipment Bay encircling stage 2

Designation: P230
Length: 31.16 m
Diameter: 3.05 m
Empty mass: each 40,000 kg
Propellant: PCA + HTPB + aluminum solid
Propellant mass: each 237,700 kg
Thrust: each 5,250 kN SL at launch
Burn time: 132 s
Separation: at about 55 km

Designation: H155
Engine: single cryogenic gas generator cycle SEP Vulcain
Length: 30.7 m
Diameter: 5.40 m
Dry mass: 12,600 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 156,200 kg max
Thrust: 900 kN SL at launch, 1,145 kN vac
Burn time: 580 s
Attitude control: engine gimbaled ±6o for pitch/yaw control

Designation: L9-7
Engine: single DASA Aestus storable propellant, open cycle, pressure-fed, reignitable
Length: 3.3 m
Diameter: 3.94 m
Dry mass: 1,190 kg
Oxidizer: nitrogen tetroxide
Fuel: monomethyl hydrazine
Propellant mass: 9,700 kg
Thrust: 27.5 kN vac
Burn time: 1,100 s
Attitude control: engine gimbaled ±6o for pitch/yaw control

Matra Marconi Space's VEB commands missions from its position encircling stage 2 and incorporates an attitude control system, providing roll control during stage 1/2 burns and 3-axis control after burnout. The 5.4 m diameter, 1.56 m high, 1,400 kg unit carries 70 kg of hydrazine.

Contraves provides two standard 5.40 m diameter fairings to accommodate 4.57 m diameter payloads: 12.70m/2.3 t short and 17 m long units. The fairing jettisons at about 191 s/106 km during stage 1 burn when aerodynamic heating has reduced to 1,135 W/m2. The 880 kg Speltra, the equivalent of Ariane 4's Spelda, increases the length of the upper composite by 5.66 m and permits the launch of double payloads. Standard payload adapters are available.
Acceleration load: 4.2 g at end P230 burn, 3.4 g at end EPC burn, 0.4 g at end EPS burn; lateral acceleration does not exceed 0.25 g
Acoustic load: 142 dB at launch/transonic
Thermal load: max 1,000 W/m2 radiated by fairing + VEB

Application: Ariane 5 strap-on boosters
First flown: 4-Jun-1996
Length: 26.77 m (stage 31.16 m)
Diameter: 3.05 m
Mass at ignition: 268,700 kg
    type: HTPB 68% ammonium perchlorate, 18% aluminum, 14% PBHT liner
    shape: main section cylindrical, forward segment 15 rectangular slots
    mass fraction: 0.884 for 237.7 t loading
Propellant mass: 237,700 kg
Burn time: 132 s
Thrust: 5,250 kN SL at launch; 4,931 kN vac average, 6,637 kN vac max
Specific impulse: 271 s vac
Total impulse: better than 630 MNs
Pressure: 59.9 atm max

Application: Ariane 4 stage 1 (four Viking 5C), Ariane 4 PAL liquid strap-on (single Viking 6, specifications as for Viking 5C but with different layout for PAL accommodation)
First flown: 15-Jun-1988
Number flown: 212 x Viking 5C, 106 x Viking 6 to end-1995
Dry mass: 826 kg
Length: 2.87 m
Maximum diameter: 0.99 m
Mounting: gimbaled for pitch/yaw/roll control
Engine cycle: open gas generator
Oxidizer: nitrogen tetroxide, delivered at 173.3 kg/s (Viking 6 173.7 kg/s)
Fuel: UH25, delivered at 101.9 kg/s (Viking 6 101.5 kg/s)
Mixture ratio: 1.70 (Viking 6 1.71)
Turbopump: 10,000 rpm, 2,500 kW power rating. Single-shaft turbopump driven by gas generator consuming 1.2 kg/s propellants
Thrust: 752 kN vac, 678 kN SL
Specific impulse: 278.4 s vac
Expansion ratio: 10.5:1
Combustion chamber pressure: 58 atm
Combustion chamber temperature: about 3,000oC
Burn time: 209 s Viking 5C, 143 s Viking 6

Application: Ariane 2-4 stage 2
First flown: 4-Aug-1984
Number flown: 75 to end-1995
Dry mass: 826 kg
Length: 3.51 m
Maximum diameter: 1.70 m
Mounting: gimbaled for pitch/yaw control
Oxidizer: nitrogen tetroxide, delivered at 175.0 kg/s
Fuel: UH25, delivered at 103.0 kg/s
Mixture ratio: 1.70
Turbopump: as Viking 5C/6
Thrust: 805 kN vac
Specific impulse: 295.5 s vac
Expansion ratio: 30.8:1
Combustion chamber pressure: 58.5 atm
Burn time: 125 s

Application: Ariane stage 3
First flown: 4-Aug-1984
Number flown: 79 HM-7/7B to end-1995
Dry mass: 155 kg
Length: 2.01 m
Maximum diameter: 0.99 m
Mounting: gimbaled for pitch/yaw control
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Mixture ratio: 4.76 for H10 stage, 4.66 H10+, 5.04 H10-3
Turbopump: 60,500 rpm, 380 kW single turbine powered by gas generator requiring 0.25 kg/s propellants
Thrust: 62.3 kN vac H10/H10+, 63.8 kN vac H10-3
Specific impulse: 446 s vac H10, 446.1 s vac H10+, 445 s vac H10-3
Expansion ratio: 83.1:1
Combustion chamber pressure: 35.5 atm
Burn time: 735 s H10, 760 s H10+, 790 s H10-3

Application: Ariane 5 core stage
First flown: 4-Jun-1996
Number flown: 1 to end-1996
Dry mass: 1,475 kg
Length: 3.00 m
Maximum diameter: 1.76 m
Mounting: gimbaled ±6o for pitch/yaw control
Engine cycle: open gas generator
Oxidizer: liquid oxygen, delivered at 225.6 kg/s
Fuel: liquid hydrogen, delivered at 36.4 kg/s (+9 kg/s LOX/LH2 for gas generator, attitude control thrusters, nozzle dump cooling and pressurization)
Mixture ratio: 6.2 in chamber, 0.94 gas generator
Oxidizer turbopump: 130 kg mass, 13,700 rpm, 3.3 MW, 150 atm discharge pressure
Fuel turbopump: 240 kg mass, 34,500 rpm, 11.9 MW, 160 atm discharge pressure
Thrust: 1,145 kN vac, 900 kN SL
Specific impulse: 431.2 s vac
Expansion ratio: 45:1
Combustion chamber pressure: 110 atm
Burn time: 580 s (design life: 6,000 s + 20 starts)

Application: Ariane 5 stage 2
First flown: 4-Jun-1996
Number flown: 1 (launcher failed during stage 1 burn)
Dry mass: 111 kg
Length: 2.20 m
Maximum diameter: 1.26 m
Mounting: gimbaled ±6o by electromechanical actuators
Engine cycle: pump-fed
Oxidizer: nitrogen tetroxide, delivered at 5.89 kg/s
Fuel: MMH, delivered at 2.87 kg/s
Mixture ratio: 2.05
Thrust: 27.5 kN vac
Specific impulse: 324 s vac
Expansion ratio: 83.3:1
Combustion chamber pressure: 10.42 atm
Burn time: 1,100 s

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