Rocket Engine Specifications
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Societe Europeene de Propulsion (SEP) Engines

Credits for providing information to
Jane's Space

Vinci engine data by Jens Lerch

  • France
  • France

    Societe Europeene de Propulsion (SEP)
    24 Rue Salomon de Rothschild
    BP 303, f-92150 Suresnes Cedex

    Large Lquid Propulsion Division
    BP 802, F-27207 Vernon.

    Solid Propulsion & Composites Division
    Les Cinq Chemis, Le Hallian
    F-33165 St.Medard-en-Jalles

    Small Propulsion & Equipment Division
    F-77550 Moissy-Cramayel

    Control Thrusters

    Liquid AOCS Thrusters

    Dry Mass 0.320/0.355kg
    Length 107.7/145.1 mm
    Mounting Method fixed
    Start Temperature both 180 celsius
    Thrust 3.5/15.4N vacuum at max pressure
    Isp 230/232 sec vacuum
    Number Of Hot Pulses 380,000/116,000
    Number Of Starts 9000/13,500
    Total On Time 40/17 hours
    Valve Power Supply both 27 V & 6 W
    Heater Power Supply 5.2/6 W at 34 V
    Total Impulse 520/615kNs 

    Mage Solid Propellant Motors/Mage 1S Specifications

    First Flown July 2, 1985, Ariane V14, Giotto
    Number Flown 2, to end of 1993
    Mass 448kg full-loading, 362kg max off-loading, 38 kg empty mass, 37kg burnout mass
    Length 128.8cm
    Max Diameter 76.6 cm
    Propellant CTPB 16-12, cylindrical in shape, mass fraction 0.915 full loading (410kg), 0.906 max off loading (328 kg)
    Thrust (kN, mean vacuum): 23.73 fully loaded, 19.0 max off loaded
    Isp 295 sec vacuum
    Burn Time 50 sec
    Nozzle Materials SEPCarbinox carbon-carbon
    Nozzle Exit Diameter 58.2 cm (expansion ratio 45)
    Nozzle Casing Materials kevlar 49 filament wound; EPDM rubber insulation
    Igniter pyrogen head-end


    Ariane Viking Engines (Viking 5C & 6 specifications)

    Applications Ariane 4 first stage propulsion (four Viking 5C), Ariane 4 PAL liquid strap-on 
    propulsion (single Viking 6, specifications as for Viking 5C but with different layout for PAL accommodation)
    First Flown June 15, 1988 on Ariane V22-401
    Number Flown 136xViking 5C, 76xViking 6, to end of 1993
    Dry Mass 826 kg
    Length 287.3 cm
    Max Diameter 99.0 cm
    Mounting Method gimballed for pitch/roll/yaw control
    Oxidizer NTO at 173.3 kg/s (Viking 6 173.7 kg/s)
    Fuel UH25 at 101.9 kg/s (Viking 6 101.5 kg/s)
    Mixture Ratio (O/F): 1.70 (Viking 6 1.71)
    Turbopump 10,000rpm, 2500kW power rating.  Single-shaft turbopump driven by a gas 
    generator consuming 1.2kg/s propellants.  A third pump adds water to limit turbine operating temperatures.  A main flow control unit maintains thrust at a pre-determined reference level by adjusting the relative flows of the three liquids entering the gas generator.  Another control loop adjusts the thrust chamber prssure by controlling gas production and regulating the turbopump.
    Thrust 752kN vacuum, 678kN sea level
    Isp 278.4 sec vacuum
    Expansion Ratio 10.5
    Nozzle Materials cobalt alloy with SEPHEN (phenolic resin/silica fiber) throat
    Combustion Chamber Pressure 58 atm
    Combustion Chamber Temperature ~3000 celsius
    Combustion Chamber Cooling Method UDMH propellant film supplied through additional channels on the injector's lower
    Engine Cycle open gas generator
    Ignition  hypergolic
    Burn Time 209 sec Viking 5B, 143 sec Viking 6

    Viking 4B Specifications

    Applications Ariane 2-4 stage 2 engine
    First Flown August 4, 1984
    Number Flown 53 to end of 1993
    Dry Mass 826 kg
    Length 350.9 cm
    Max Diameter 170.0 cm
    Mounting Method gimballed for pitch/yaw control
    Oxidizer N2O4 at 175.0 kg/s
    Fuel UH25 at 103.0kg/s
    Mixture Ratio (O/F): 1.70
    Thrust 805kN vacuum
    Turbopump as Viking 5B/6
    Isp 295.5 sec
    Expansion Ratio 30.8:1
    Burn Time 125 sec
    Combustion Chamber Pressure 58.5 atm

    Ariane HM-7B

    Applications Ariane stage 3
    First Flown August 4, 1984 on Ariane V10
    Number Flown 60 HM-7/7B to 1993, including 3 failures
    Dry Mass 155 kg
    Length 201.3 cm
    Max Diameter 99.2 cm
    Mounting Method gimballed for pitch/yaw control
    Oxidizer LO2
    Fuel LH2
    Mixture Ratio (O/F): 4.76:1 (gas generator 0.87) for H10 stage, 4.66 H10+, 5.40 H10-3
    Turbopump 60,500 rpm, 380kW single turbine powered by gas generator requiring 0.25 kg/s propellants.  Single-stage hydrogen pump raises pressure from 3 atm to 55 atm; single-stage oxygen pump 2.5-50 atm.  Gas generator exhaust temp. 800-900K
    Isp sec vacuum, 446 H10, 446.1 H10+, 445 H10-3
    Thrust in vacuum, 62.3kN H10/H10+, 63.8kN H10-3
    Cooling Method hydrogen is heated to 150K during cooling of combustion chamber wall along 
    128 longitudinal channels for injection in gaseous form.  The nozzle cooling system employs 242 helical tubes; 0.15 kg/s of hydrogen is dumped overboard. 
    Expansion Ratio 83.1:1
    Combustion Chamber Pressure 35.5 atm
    Combustion Chamber Ignition pyrotechnic, comprising Isolite 1431 Rauline grain, with redundant initiators, providing 
    two flame jets angled at 45 degrees to propellant flow. 
    Materials stainless steel
    Burn Time sec 735 H10, 760 H10+, 790 H10-3


    Applications Ariane 5 upper stage ESC-B
    Schedule start of development in 1999, first test firings in late 2001, first flight in 2005
    Dry Mass 480 kg
    Length 420 cm
    Max Diameter 210 cm
    Mounting Method gimballed for pitch/yaw control
    Oxidizer LO2
    Fuel LH2
    Mixture Ratio (O/F): 5.5:1 to 6.5:1
    Turbopump Separate fuel/oxidiser expander type closed cycle turbopumps.
    Hydrogen turbopump: 85,000 rpm, 1800 kW
    Oxygen turbopump: 18,000 rpm, 300 kW
    Isp 464 sec vacuum
    Thrust 155 kN vacuum
    Cooling Method Upper part of nozzle is regeneratively cooled, extension made of composite material is radiatively cooled..
    Combustion Chamber Pressure 60 atm
    Combustion Chamber Ignition electric igniter, restartable in flight 
    Remarks will likely use Snecma developed nozzle extension of RL-10B-2 

    Applications Ariane 5 core stage
    First Flown 1995
    Dry Mass 1475 kg
    Length 300 cm
    Max Diameter 176 cm
    Mounting Method gimballed +/-6 degrees for pitch/ yaw control
    Propellant LOX at 225.6kg/s into chamber, LH2 at 36.4kg/s into chamber; total propellant flow 271.5 kg/s includes 9kg/s for gas generator, attitude control thrusters, nozzle dump cooling and pressurization.
    Mixture Ratio (O/F) 6.2:1 in chamber, 0.94:1 gas generator.  Valves are pneumatically operated and provide only two mixture ratio settings.
    Turbopumps Separate fuel/oxidiser turbopumps. Hydrogen pressure raised to 160 atm by 34,500 rpm 11.9 MW (range 30,340-39,330rpm & 7.7-16.9 MW) 240kg 2-stage centrifugal pump; oxygen to 150 atm by 130kg 13,700rpm 3.3MW
    (range 11,300-14,880rpm & 1.8-4.3 MW) 1-stage pump
    Thrust 1145kN vacuum, 900kN sea level
    Isp 430 sec vacuum
    Expansion Ratio 45:1
    Combustion Chamber Materials stainless steel
    Combustion Chamber Injector Inconel 718 516-element coaxial
    Combustion Chamber Pressure 110 atm
    Combustion Chamber Cooling Method Regenerative hydrogen through 360 longitudinal channels in copper-silver-zirconium alloy liner.
    Nozzle Diameter 176.0cm
    Nozzle Mass 170kg
    Nozzle Cooling Hydrogen circulated through helical tubes and dumped into exhaust through 228 ports.
    Burn Time 580 sec (design life: 6000 sec +20 starts)

    Rocket Engine Specifications

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