A long duration and high reliability liquid apogee engine for satellites

doi:10.1016/j.actaastro.2004.05.030

Acta Astronautica

Volume 55, Issues 3–9, August–November 2004, Pages 401-408

https://doi.org/10.1016/j.actaastro.2004.05.030 Get rights and content

Abstract

The Chinese unique $490 N$ liquid apogee engine, Model FY-25, is described in detail in this paper. The engine, which was developed by Shanghai Institute of Space Propulsion (SISP) and used to provide thrust for apogee injection maneuver required for the telecommunication satellite, has been flight qualified for use in 6 satellites using storable nitrogen tetroxide (NTO) and monomethylhydrazine (MMH) propellants. It has the characteristics such as long operation duration, high reliability and stable performance. The thrust chamber made from NbHf10-1M niobium alloy has been successfully demonstrated with a nominal vacuum specific impulse of $3030 N s / kg$ at nozzle area ratio of 154:1 and chamber pressure of $0.68 MPa$ . The maximum operating temperature is 1360°C at the throat. The engine has demonstrated the capability of 26 accumulated restarts and about $26 110 s$ accumulated burning time with single unit. The investigation to improve the specific impulse, including adoption of secondary combustion assembly, Re/Ir and C/SiC composite materials combustion chamber is described in detail. Both the rhenium/iridium and the C/SiC are advanced materials that can withstand higher temperature than the conventional NbHf10-1M niobium alloy.

Introduction

The liquid apogee engine controlled by two electrically driven valves is used to provide thrust for apogee injection maneuver required for geostationary satellites. FY-25 made by SISP uses the earth storable propellants of NTO and MMH. The research and qualification tests demonstrates a nominal Isp of $3030 N s / kg$ based on thrust level of $490 N$ under inlet feed pressure of $1.40 MPa$ with the mixture ratio of 1.65 and burning time more than $26 000 s$ .

With the development of the new generation satellites of large capacity, long-life and high reliability, the new demand to satellite propulsion system has been required. It is estimated that every $1 s$ increase in value of Isp will result in a propellant mass benefit of about $3 kg$ for a satellite with an initial mass of $2200 kg$ after every $6000 s$ firing time of the apogee engine. If the firing time extends to $16 000 s$ , the propellant mass benefit will be $8.5 kg$ . This is of great important to prolong the life or increase the loads mass of spacecraft.

In the fields of liquid rocket engine, the current research for improving the Isp performance of the engine focuses on two directions. One is to develop new material which can withstand high temperature. The other is to ameliorate the thrust chamber structure, such as the secondary combustion augmented thruster which was developed by TRW with regeneration cooling tunnel and without refractory metals.

Higher performance can only be achieved by running the engines at higher operating temperature, which implies unacceptable operating life restrictions for the state-of-art material system of NbHf10-1M niobium alloy with disilicide coatings. In order to improve performance and reliability of the bipropellant liquid rocket engine, the technology has been continuously upgraded for material systems with higher temperature capabilities and has been introduced to engines for high altitude simulation test.

Two kinds of advanced material systems are being developed. One is rhenium/iridium and the other is C/SiC composite material. The hot firing test of the $10 N$ Re/Ir thrusters shows that the operation temperature is 1800°C, the specific impulse reaches $2903 N s / kg$ . The results indicate that both the rhenium substrate with iridium coatings and the C/SiC composite material have potential to improve the performance of satellite engines.

FY-25 and the investigations on the high performance bipropellant liquid rocket engine are illustrated here.

Section snippets

Overview of FY-25

FY-25 $490 N$ liquid apogee engine is shown in Fig. 1. The mass of the engine is approximately $4.2 kg$ . A summary of FY-25 liquid apogee engine characteristics is tabulated in Table 1.

The engine assembly is composed of three subassemblies as followings:

•
Two G4800 propellant control solenoid valves.
•
A 8 on 8 unlike doublet propellant injectors with integral fuel film cooling.
•
A thrust chamber made of NbHf10-1M niobium alloy material.

Propellant control valves

The two solenoid valves, Model G4800, are respectively mounted to the

$490 N$ Secondary combustion test assembly

Fuel cooling film, which is about 26.5% of the total fuel mass flow, is used to keep FY-25 chamber at an acceptable operating temperature below 1350°C. The cool boundary layer throughout the chamber implies that there is unmixed, unreacted fuel exiting from the nozzle, which may result in a performance loss and be a potential source of spacecraft contamination. A new design, incorporating of a secondary mixing device or boundary layer trip (BLT) within the combustion chamber, has been developed

Conclusion

FY-25 $490 N$ liquid apogee engine is developed for satellite application, long duration development and qualification tests show that the engine has a stable performance of $3030 N s / kg$ , temperature distribution of less 1350°C, life of more than $26 000 s$ and has high reliability characteristics.

A second generation of engines was decided in order to improve the performance and the operation range for future application. Based on the investigations on $10 N$ rhenium/iridium and $150 N$ ceramic matrix

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