CE-20

From Wikipedia, the free encyclopedia
Jump to: navigation, search
CE-20
IndianCryoEngine25.JPG
Model of CE-20
Country of origin India
Date 2015
Designer LPSC, Indian Space Research Organisation
Manufacturer Hindustan Aeronautics Limited,[1]
Application Upper stage booster
Status Under Development
Liquid-fuel engine
Propellant LOX / LH2
Cycle Gas Generator
Configuration
Chamber 1
Nozzle ratio 100
Performance
Thrust (vac.) 200 kN (45,000 lbf)
Chamber pressure 6.000 MPa (60.00 bar)
Isp (vac.) 443 seconds (4.34 km/s)
Dimensions
Dry weight 588 kg

The CE-20 is a cryogenic rocket engine being developed by the Liquid Propulsion Systems Centre, a subsidiary of Indian Space Research Organisation. It is being developed to power the upper stage of the Geosynchronous Satellite Launch Vehicle Mk III.[2] It is the first Indian cryogenic engine to feature a gas-generator cycle.[3]

Overview[edit]

The CE-20 is the first Indian cryogenic engine to feature a gas-generator cycle.[4] The engine produces a nominal thrust of 200 kN, but has an operating thrust range between 180 kN to 220 kN and can be set to any fixed values between them. The combustion chamber burns liquid hydrogen and liquid oxygen at 6 MPa with 5.05 engine mixture ratio. The engine has a thrust-to-weight ratio of 34.7 and a specific impulse of 444 seconds (4.35 km/s) in vacuum. ISRO tested the CE-20 on April 28, 2015 at Mahendragiri test facility achieved on successful long duration hot test (635 seconds).[5] On July 16, 2015, CE-20 was successfully endurance hot tested for a duration of 800 seconds at ISRO Propulsion Complex, Mahendragiri. This duration is approximately 25% more than the engine burn duration in flight.[6] The CE-20 cryogenic engine was again hot-tested for a duration of 640 seconds at ISRO Propulsion Complex, Mahendragiri on 19 February 2016.[7]

Specifications[edit]

The specifications of the engine as listed on the LPSC handouts:[8]

  • Operating Cycle - Gas Generator
  • Propellant Combination - Liquid oxygen / Liquid hydrogen
  • Thrust Nominal (Vacuum) - 200 kN
  • Operating Thrust Range - 180 kN to 220 kN (To be set at any fix values)
  • Chamber Pressure (Nom) - 6 MPa
  • Engine Mixture ratio (Oxidizer/Fuel by weight) - 5.05
  • Engine Specific Impulse - 443 ± 3 seconds (4.344 ± 0.029 km/s)
  • Engine Burn Duration (Nom) - 595 seconds
  • Total Flow rate - 462 kg/s
  • Nozzle Area ratio - 100
  • Mass - 588 kg

See also[edit]

References[edit]

External links[edit]