Modern Liquid Propellant Rocket Engines

Transcription

1 Modern Liquid Propellant Rocket Engines 2000 Outlook By B.T.C. Zandbergen TU-DELFT, MAY 5, 2000 i

2 PHOTO CREDITS: Rocketdyne, Snecma SA, DASA MBB, Pratt & Whitney, Aerojet, Royal Ordnance, and Kaiser Marquardt. ii

3 Preface This document intends to provide the (future) space engineer with a starting point for both practical liquid rocket engine selection and engineering. The text for this document originated from a lecture series on chemical rocket propulsion, which the author provided during the years at TU-Delft, Faculty of Aerospace Engineering. This document is intended as a lively document. Readers are invited to send their comments, amendments or recommendations to: B.T.C. Zandbergen, Space Systems Engineer (MSSE), TU-Delft, Faculty of Aerospace Engineering, Kluyverweg 1, 2629 HS Delft, The Netherlands or to bzandbergen@tudelft.nl Special thanks go to ir. G.W.R. Frenken of Stork Product Engineering BV, Amsterdam, The Netherlands, for proof-reading this publication and for providing many useful comments. The current document is a slightly updated version of the original document. Most important change is the addition of more motors in the moderate impulse section. iii

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5 Contents PREFACE... III CONTENTS... V LIST OF ACRONYMS... VI INTRODUCTION ENGINES FOR HIGH TOTAL IMPULSE LAUNCHER MISSIONS... 3 SPACE SHUTTLE MAIN ENGINE... 6 HM60 CRYOGENIC ROCKET ENGINE... 8 LE-7 CRYOGENIC ENGINE RD-170/171 AND RD RS RS-2200 LINEAR AEROSPIKE ENGINE ENGINES FOR MODERATE TOTAL IMPULSE LAUNCHER MISSIONS HM7 ENGINE VINCI ENGINE LE-5/5A RL10A ENGINES FOR LOW TOTAL IMPULSE LAUNCHER MISSIONS AESTUS 27,5 KN ENGINE EUROPEAN ADVANCED TECHNOLOGY ENGINE AESTUS II 46 KN ENGINE ORBITING MANEUVERING SYSTEM (OMS) ENGINE RS AJ10-118K/ ENGINES FOR REACTION CONTROL SYSTEMS BIPROPELLANT ENGINES CATALYTIC HYDRAZINE THRUSTERS COLD GAS THRUSTERS ABOUT THE DATA PRESENTED REFERENCES A. DEFINITION OF TERMS B. PROPELLANT PROPERTIES C. NOZZLE PROPERTIES D. ROCKET MOTOR PERFORMANCE CALCULATION; AN EXAMPLE v

6 List of acronyms A50 Aerozine 50 ATE Advanced Technology Engine CHT Catalytic Hydrazine Thruster CMC Carbon Matric Composite CNES French Space Agency DASA DaimlerChrysler Aerospace ESA European Space Agency GEO Geostationary Earth Orbit GG Gas Generator cycle GH2 Gaseous Hydrogen GOX Gaseous Oxygen GTO Geostationary Earth Orbit LOX Liquid Oxygen LH2 Liquid Hydrogen LRE Liquid Rocket Engine MMH Mono-Methyl-Hydrazine MON Mixed Oxides of Nitrogen NASA National Aeronautics and Space Administration NASDA NAtional Space Development Agency NTO Nitrogen Tetra-Oxide OMS Orbiting Manoeuvring System rpm Rounds per minute RCS Reaction Control System RD Raketnay Dvigatel (rocket motor) SC Staged Combustion cycle SSME Space Shuttle Main Engine STS Space Transportation System SEP Socièté Européene de Propulsion (Snecma SA, Snecma group, France) SL Sea Level TVC Thrust Vector Control TMC Thrust Magnitude Control USA United States of America US$ US dollar vi

7 Introduction In this document, a number of liquid propellant rocket engines are described to provide (future) space system/propulsion engineers with a starting point for practical liquid rocket engine selection and/or engineering. It is the task of the propulsion system to generate the propulsive force (thrust) needed to change the momentum of a vehicle. For a chemical rocket propulsion system, this thrust is generated in a thrust chamber or thrust generator using one or more chemical reactants to provide the required energy as well as the mass to be expelled. Liquid propellant rocket propulsion systems use liquid chemical reactants as the propellant. Besides the liquid propellant and 1 or more liquid rocket thrust chambers, the overall propulsion system includes a propellant feed system, a control system and a tankage system. A liquid propellant rocket thrust chamber is in principle a very simple device, which usually is made up of a cylindrically or spherically shaped reaction or combustion chamber 1 and a convergent-divergent (con-di) outlet referred to as nozzle. In the reaction chamber, the chemical energy from the liquid reactants is converted into thermal energy thereby creating a hot gas mixture. This hot gas mixture is accelerated by pressure forces to a high exhaust velocity in the specially shaped nozzle. The con-di shape of the nozzle is instrumental in generating a high pressure in the thrust chamber and the optimal conversion of the thermal energy of the hot gas mixture into kinetic energy of the jet exhaust. To allow high gas temperatures in the thrust chamber, most combustion chambers and nozzles are cooled to some extent. As (liquid) propellant various combinations of fuel and oxidiser are used, like hydrogen and oxygen, kerosene and oxygen or hydrazine and nitrogen-tetroxide. Such combinations of fuel and oxidiser are usually referred to as 'bipropellant'. Apart from bipropellants, sometimes a single fluid, referred to as 'monopropellant', is used. Chemical energy in that case stems from a decomposition reaction that occurs under the influence of a suitable catalyst. Monopropellants are for example hydrazine or hydrogen-peroxide, which decompose into mostly ammonia and water, respectively. Because of the high pressure in the combustion chamber a feed system is needed to pressurise and transport the propellant from the propellant tank(s) to the thrust chamber. In today s rocket engines, propellant pressurization is accomplished by either (turbo)pumps or by a high pressure gas that is released into the propellant tank(s), thereby forcing the propellants out of the tank(s). In space engineering, in case of high total impulse, short duration (up to hundreds of seconds), launcher missions, the choice is almost exclusively for pump-fed systems, whereas for low total impulse, long duration (typically years), orbital missions the choice is for using a high pressure gas. A control system ensures the proper flow of propellants to the thrust chamber. It includes amongst others: On/off valves that control whether the propellant is flowing or not; Check valves that prevent the fluids from flowing in the wrong direction; Fill and drain valves that allow for filling and emptying of the tank(s) when on ground; Pressure regulators that control the pressure in the tank(s); Filters that filter out contaminants; Transducers that provide information on pressures and temperatures. Generally, the thrust chamber together with the control system is referred to as the thruster. In case of turbopump-fed systems, one generally speaks of liquid rocket engines, meaning the assembly of thrust chamber(s), control system and turbopump(s). Liquid rocket engines may be of a single chamber or multi-chamber design, i.e multiple chambers fed by a single turbopump. First use of liquid propellant systems dates back to 1926 to the rocket pioneering days of Robert Goddard. In the 1940 s this was followed by the illustrious V-2, also known as the A(ggregate)-4, designed by Von Braun and his team, which clearly demonstrated the high 1 The reaction or combustion chamber is sometimes referred to as combustor. 1

8 performance capability of liquid propellant rocket propulsion systems. This was followed in 1957, by the first true space launch when the Soviets launched Sputnik I. This event spurred a surge in rocketry in the United States of America (USA) culminating in 1969 with the first manned lunar landing made possible by the design and development of the powerful Saturn V all liquid propellant rocket. Today, liquid propellant rocket propulsion systems still form the back-bone of the majority of space rockets and spacecraft allowing mankind to expand his presence into Space. An important advantage of liquid rocket propulsion systems compared to solid propellant rocket motors is that they can offer a much higher specific impulse 2 (typically 450 s compared to about 300 s) leading to a much lower propellant mass needed. Another advantage is that liquid propellant rocket systems offer a much higher degree of flexibility allowing for pulsed thrust over extended periods of time. This is e.g. advantageous for attitude and orbit control of spacecraft. On the other hand, liquid propulsion systems are more costly and have a lower reliability. For launcher applications [Andrews and Haberman, 1991] mention a reliability of 0,985-0,989 for liquid systems versus 0,998 for solid rocket motors with a cost level of US$ 0,090 per Ns of total impulse delivered versus 0,034 US$ 3 for solid rocket motors. In addition, because of low propellant densities, the use of liquid propulsion systems leads to larger sizes of the vehicle. To provide the (future) space engineer with a starting point for practical liquid rocket engine selection and engineering, this document describes some key liquid propellant rocket engines and thrusters for the 21 st century. In Chapter 1 attention is given to engines for high total impulse space launcher missions, with total impulse levels ranging from 50 MNs to MNs and thrust levels from 50 kn to several MN. Chapter 2 focusses on engines for moderate total impulse space launcher missions with total impulse ranging from MNs and thrust levels from 50 to 150 kn. In Chapter 3 attention is given to engines for low total impulse space missions like upper stage propulsion and orbit raising of spacecraft with total impulse levels in the range up to about 15 MNs. Finally, in Chapter 4 attention is given to engines for reaction control systems of space vehicles with total impulse levels up to about 1,5 MNs and with thrust levels well below 10 kn. All 4 chapters focus on the engine system and not on the associated tankage system. This is because burn times can be selected freely (as long as one stays below the engine's maximum allowed burn time), leading to different propellant mass and hence tank sizes. However, data will be provided that allow the designer to properly design the tankage system. Each chapter starts with a performance overview of some key engines for the next century. The performance data included (when available) are amongst others. thrust, maximum thrust duration, level of thrust control, specific impulse and/or mass flow, number of engine/thruster starts, propellants used, oxidiser-to-fuel mass mixture ratio, minimum impulse bit, reliability, and cost, see Annex A for a definition of terms used. Next, to provide insight in the various different designs and the design options that are available to the engine designer to influence the performances of a rocket engine, some key engines are discussed in more detail. Attention will be given to: Engine characteristics like type of engine, number of thrust chambers, (type of) propellants, propellant mass flow, engine mass, engine dimensions, expansion ratio, pump power, chamber pressure, inlet pressure, engine cooling, materials used and Development data like development period, development approach, development tests, etc. Finally, in Annexes B, C, and D some theoretical results are given that allow the reader to verify/estimate some engine performances, including engine thrust, and specific impulse. In addition, some data is provided that allow for determining propellant density and propellant volume flow (for a given burn time and mass flow rate). 2 Specific impulse is total impulse delivered per unit of propellant weight. 3 Correcting for inflation, the 1999 figures are 0,117 and 0,44 US$, respectively. 2

9 1. Engines for high total impulse launcher missions In this section, some key liquid rocket engines for high total impulse space launcher missions with total impulse levels ranging from 50 MNs to MNs and thrust levels from 50 kn to several MN are discussed. Typical such missions are first stage and booster applications. Two cases can be discerned for these types of engines: The required thrust is produced without the help of solid rocket boosters; and The required thrust is produced in conjunction with solid rocket boosters. In the first case, attention is on both high specific impulse and on high propellant density. The high specific impulse allows reducing propellant mass, whereas the high propellant density helps to limit the size of the propulsion system and more specific the propellant tanks and thus limits gravity losses. This mostly leads to the choice for oxygen-kerosene as the propellant combination with typical vacuum specific impulse of s and propellant density of kg/m 3 compared to s and kg/m 3 for oxygen-hydrogen. In the second case, attention can be fully on achieving high specific impulse, since in that case, the solid propellants provide sufficient density. In that case, the choice is obviously for oxygen-hydrogen. Because of the large propellant mass needed and hence the large tanks needed, pump-feeding is used almost exclusively. The use of pumps allows for a relatively benign tank pressure of about 3-5 bar, with the pumps providing the necessary pressure rise to allow feeding of the propellant into the combustion chamber. In addition, because of the high pressure attainable with pumps, it is much easier to attain a high expansion ratio and hence a high specific impulse. A major disadvantage of pump-fed engines, however, is the high complexity and the associated high development and production cost. A further disadvantage is a fairly complex starting procedure. Restartability is in principle possible, but not much done today. To drive the propellant pumps mostly turbines driven by high pressure gasses are used. This high-pressure gas can be generated in a number of ways, which each has its own (dis)advantages. We distinguish: Staged combustion: The turbo-pumps (pump-turbine combination) are driven from hot gases produced by a pre-combustion chamber or pre-burner. Parts of the propellants are burned at a low temperature, sufficiently low to allow driving a turbine. These hot gasses are led to the turbo-pump(s), where they drive the turbine. When leaving the outlet of the turbine, the gasses are fed to the thrust chamber, where the remainder of the fuel or oxidiser is burned and can contribute to the generation of thrust Gas Generator Cycle: This cycle is similar to the staged combustion cycle, but instead that the gases are fed to the thrust chamber, they are exhausted separately or injected at a point downstream of the nozzle throat, where the pressures balance. Coolant Bleed or Expander Cycle: Part of the coolant flow rate is heated to a high temperature (typically K) in a heat exchanger portion of the main thrust chamber. The heated coolant is expanded through high-pressure-ratio turbines. Highest specific impulse is achieved by the choice for the very complex staged combustion cycle. The gas generator cycle is somewhat less complex than the staged combustion cycle, but also the attainable specific impulse is somewhat less. Compared to the gas generator and staged combustion cycle, the bleed or expander cycle offers least complexity and mass. In the next sections, the detailed workings of a number of high total impulse space launcher engines is given, including the Main Engine of the Space Shuttle Orbiter, the HM-60 (or Vulcain) engine of the Ariane V first stage, the LE7 first stage engine of the Japanese H2 launcher, the RD 170 engine used on the Russian Zenith and Energia launcher, the RD-180 engine currently under development for Atlas III, and the RS-2200 linear aerospike engine under development for VentureStar. 3

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11 Table 1-1:Performance data of some typical liquid rocket engines SSME HM60 LE-7 RS-68 RD-170 RD-180 RS-2200 Propellants LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/Kerosene LOX/Kerosene LOX/LH 2 Engine cycle SC GG SC GG SC SC GG Vacuum thrust (kn) Specific impulse (s) 455, Overall mixture ratio; O/F (-) 6,0 5,3 6,0 2,63 2,72 5,5 Propellant density (kg/m 3 ) 1) Length (m) 4,24 3,00 3,40 5,18 4,0 3,8 Total dry mass (kg) Mission duty cycle (s) Life span (s) >2000 Max number of starts (-) > >20 Thrust/weight ratio (-) Throttle capability (%) No Yes Operational use (year) ? ? Reliability 0,999 0,9927 0,9935 0,999 5

12 Space Shuttle Main Engine Under contract to the National Aeronautics and Space Administration (NASA), the Rocketdyne division of Rockwell International developed NASA s Space Shuttle (Orbiter) Main Engine [2, 5, 6]. The Orbiter forms part of NASA s Space Transportation System (STS), which carries satellites into low Earth orbit, releases them for free flight operation or retrieves them from space. NASA s STS, furthermore, consists of two large solid rocket boosters providing thrust during the initial phases of the flight, and a large external tank carrying the liquid propellant needed for the SSME s. Figure 1: Space Shuttle Main Engine Each Orbiter is equipped with three SSME's, which are located in the aft section of the aft fuselage of the Orbiter. Each SSME is supplied with liquid oxygen and liquid hydrogen, which are stored in an external tank structure, separated by a common bulkhead (tandem configuration). At the launch pad the propellant tanks are pressurised with gaseous helium. The pressurised helium forces the liquid oxygen and hydrogen from the tanks through 43cm ducts. The 43cm LOX and LH 2 ducts in the aft fuselage of the Orbiter branch out into separate 30,5cm ducts to each SSME. After engine start, pressurisation of the propellant tanks is by vaporised oxygen and hydrogen extracted from the engines. The SSME is a gimballed, staged combustion cycle, single thrust chamber engine. Under nominal conditions, it nominally consumes 467 kg of oxygen and hydrogen per second at a mixture ratio of 6:1. The propellant burns at a high pressure (maximum 223 bar) to reduce engine envelope. The hot (3300 K) combustion gases are accelerated in a bell-type (clock-shape) nozzle with an area ratio of 77,5:1 and an exit diameter of about 2,27 m. Because of the high gas temperatures in the thrust chamber, the thrust chamber is regeneratively cooled using hydrogen fuel as the coolant. Propellant feeding is through a high-pressure pump following a lowpressure pump. The latter is to prevent cavitation in the former. The hydrogen pump operates at rpm and provides 51,2 MW of power. The oxygen pump provides 18,6 MW of power. When the ignition circuit is closed, the main oxygen valve opens and the oxygen is pumped by the pumps. Part of the oxygen flow that leaves Figure 2: SSME engine schematic the high-pressure pump is used to drive the turbine of the low-pressure oxygen pump. The turbine exhaust is returned to 6

13 the low-pressure oxygen circuit. The oxygen flow that leaves the high-pressure pump is tapped and part is pumped through two regulating valves to the injectors of the pre-burners. Here it is mixed with the hydrogen and burned in a fuel-rich mixture at about 1000 K. The hot exhaust drives the turbines of the high-pressure turbo-pumps and is ducted through the hot-gas pipe to the main injector. Here the gas is mixed with the rest of the pressurised oxygen flow and led into the main combustion chamber. After opening the hydrogen flow control valve, the liquid hydrogen is pumped through the cooling channels of the combustion chamber and nozzle after which it is led to the two preburners. A small part of the flow pumped by the high-pressure pump is tapped off and used to drive the low-pressure turbo-pump of the hydrogen circuit. When leaving the turbine of the low-pressure pump, the hydrogen is ducted through the double wall of the hot gas pipe and main injector (thereby cooling them) into the main combustion chamber. Combustion pressure in the pre-burner is about 2 times higher ( bar) than in the main combustion chamber. This is to make sure that, after the pressure drop that occurs in the turbines, the turbine exhaust gas still has a sufficiently high pressure to enter the main combustion chamber. Because of the high pressure in the pre-burners, this requires an equally high pressure for the pre-burner propellants. For the hydrogen this pressure is even higher (415 bar) because of pressure losses in the cooling ducts. Each SSME has an engine controller based on a digital computer, which monitors the engine parameters such as pressure and temperature fifty times every second and automatically adjusts engine operation for the required thrust and constant mixture ratio. The system maintains a record of engine operating history for maintenance purposes to ensure a flight-ready engine and extend total engine life. Thrust magnitude control is achieved through mass flow control, which in turn is controlled by the high-pressure pumps turbine mass flow. Thrust vector control is by gimballing the thrust chamber, which allows for a maximum deflection of 11 degrees. Each SSME has a mass of about 3170 kg and develops N of thrust at sea level and N in vacuum with a capability to throttle between 67%-109% of the nominal thrust. The specific impulse at sea level is 363 sec and in vacuum 455,2 sec. The SSME is designed for a large number of missions. It has a total lifetime of 7,5 hours, including 55 ignitions. Until 1994, all SSME s together have been exposed to about 2000 starts and stops and seconds of operation both on the ground and in flight. With the exception of 1 sensor failure that caused early shutdown of one engine, the SSME has operated flawlessly in more than 60 Shuttle missions. With more than 180 engine missions (3 engines per mission), the percentage of demonstrated SSME reliability is 0,999 and the three engine cluster is 0,994. SSME unit cost is about 45 4 M$ (2000). In Table 1-1 some performances of the SSME are summarised. Presently, the SSME is being upgraded (block IIA) to incorporate a larger throat and to reduce the number of components. The larger throat area should reduce wear of the engine because of a lower operating pressure of maximum 207 bar. In addition, because of the larger throat, the nozzle area ratio is reduced to 69:1, reducing vacuum specific impulse slightly to 452s. Engine mass does increase from 3170 to 3390 kg M$ (1992). Cost in year 2000 has been estimated using an inflation factor of

14 HM60 Cryogenic Rocket Engine The SEP 5 -developed HM60 cryogenic rocket engine or Vulcain [7] propels the first stage of the core vehicle of the European developed Ariane 5 rocket launcher. This launcher is capable of launching 20 tons in low Earth orbit or 9700 kg in geostationary transfer orbit and consists of a two-staged core vehicle and two large solid rocket boosters. The HM60 is a gimballed, gas-generator fed, single thrust chamber engine, see Figure 1. Each second, about 255 kg (214,5 kg oxygen and 40,5 kg hydrogen) of propellant flows from the propellant tanks to the engine through two 185 mm diameter lines. Approximately 244 kg/s flows into the combustion chamber via the main injector at a mixture ratio of 6,0. Combustion takes place at a pressure of 105 bar, resulting in a combustion temperature well in excess of 3000K. To allow testing at sea level, the combustion chamber includes a divergent part with an area ratio of 5,8:1. To withstand the high temperature, the combustion chamber is regeneratively cooled by most of the hydrogen fuel. The rest (0,4 kg/s) is tapped off to provide pressurisation of the hydrogen tank. The hot exhaust gases are expelled through a belltype nozzle extension with an expansion ratio of 45 and a nozzle exit diameter slightly less than 1,76 m. The nozzle is dump cooled by 5,3% (1,8 kg/s) of the Figure 3: HM60 engine total hydrogen flow. The feed assembly provides some 200 to 600 litres per second of liquid oxygen and hydrogen. The high pressure needed for injection is supplied by two independent turbopumps. The hydrogen pump operates at rpm and provides 11,8 MW of power, whereas the oxygen pump operates at rpm, providing 3,7 MW of power. These turbopumps are driven by hot gas from a common gas generator. This gas generator burns a mixture of oxygen and hydrogen (mixture ratio of 0,9) at a combustion pressure of about 80 bar, giving a combustion temperature of 910 K. The propellants needed for the gas generator are tapped off from the main propellant supply lines, see figure 3 (Hydrogen in red, oxygen in yellow and hot gas in blue). Total estimated mass flow rate is 8,9 kg/s. Of this, approximately 4,7 kg/s is hydrogen. The exhaust of the turbines also participates in the propulsion process by expanding through two secondary nozzles located on either side of the main nozzle. Since the engine is of a gimballed design, the feed lines are equipped with flexible joints. Engine start-up is achieved by a solid gas generator (turbopump starter), which generates a high-pressure hot gas stream during a few seconds. This hot gas drives the turbines and through the turbines the oxidiser and fuel pump. Control valves are actuated to make sure that the propellants can flow when the pumps are activated. Once the propellants start flowing, part of the propellants is tapped of and fed to the gas generator, where they are mixed. Ignition of the gas mixture is Figure 4: HM60 flow schematic by a pyrotechnic system. Ignition of the gas mixture in the main chamber is obtained from a pyrotechnic igniter situated in the centre of the injector. 5 SEP is now part of Snecma S.A., which in turn is part of the Snecma group. 8

15 Mixture ratio of the engine is controlled through the use of a hot-gas control valve, which varies the power of the LOX turbo-pump turbine. Mass flow of the engine and mixture ratio of the gas generator is controlled through the use of the injection/control valves of the gas generator. The engine computer insures all the checking and control functions, using transducers mounted on the engine and in the tanks. The control valves and the engine gimballing systems are hydraulically actuated. A hydraulic pump driven from the LOX turbopump shaft is used for all the hydraulic power needed on the stage. The other valves on the engine are actuated from a helium pressure of 23 bar. Early development of the HM60 started in the late 1970 s when studies made in Europe concerning launch systems needed for the 1990 s, showed the necessity of developing a highthrust LOX/LH 2 rocket engine. In 1978 the French Space Agency (CNES) and the Socièté Européene de Propulsion (SEP) conducted preliminary studies of a 500kN thrust LOX/LH 2 engine in order to evaluate the required technology effort. During the period, a 800 kn vacuum thrust HM60 engine emerged to power the second stage of a two-stage Ariane V launcher concept. At that time, three engine cycles were considered, including the heat exchanger cycle, staged combustion cycle, and gas generator cycle. These early studies resulted in 1981 with the selection of the gas generator engine cycle with only a single set of turbopumps driven by a single gas generator. The reasons for selecting this cycle were as follows: The heat exchanger cycle was found to provide insufficient thermal energy to drive the turbines for a chamber pressure of 100 bar; The development cost for the staged combustion engine was found to be 25 % higher than for the gas generator engine, and the production costs 20% higher Experience on the gas generator cycle was available from the HM7 third stage cryogenic LOX/LH 2 engine for Ariane I to IV, see later in this document. For the other two cycles no experience was available. Actual development of the HM60 started in 1984 as a joint venture between 12 European countries (including The Netherlands) with more than 37 industrial companies involved under supervision of ESA. At that time, the HM60 was already assigned to Ariane s first stage and up-rated to produce more than 1 MN of vacuum thrust. Initial plans called for the development to be completed by 1991 to allow a first flight of the Ariane V in Due to some technical difficulties, the latter date shifted to 1996 with total development costs of the HM60 engine of about $ 1.3 billion (1990). Since then, a total of 3 flights have been made in the period until Nominal characteristics of the HM60 (Vulcain-1) for a geostationary transfer orbit mission are a vacuum thrust of 1075 kn, a vacuum specific impulse of 430 s, an overall oxidiser/fuel mixture ratio of 5,3:1. Engine dry mass is 1700 kg, length is 3,00 m and maximum diameter is 1,76 m. In addition, one ignition in flight is possible and the engine is recoverable and reusable for future recoverable launchers. Burn time is 580 s, depending on the mission. Engine projected (mature) reliability is 0,9927. Since 1995, Vulcain enhancements are being worked upon that should lead to a 2040 kg Vulcain-2 version with a vacuum thrust level of 1,35 MN and a specific impulse of 434s. Changes considered include a change in overall mixture ratio to 6,1, an increase in combustion pressure to 115 bar, an increased expansion ratio of 58,5 with an enlarged nozzle exit diameter of 2,15 m and instead of dumping the turbine exhaust gases overboard, they will be used to cool part of the nozzle extension. The propellant supply comprises an improved high speed liquid hydrogen pump ( rpm) delivering 14 MW of power and a low speed liquid oxygen pump ( rpm) delivering 5 MW of power. First flight of this new improved version is expected in

16 LE-7 Cryogenic engine The LE-7 engine [10] propels the first stage of the core vehicle of Japan s H-2 launch vehicle. The H-2 vehicle is developed by Japan s National Space Development Agency (NASDA) and consists of the core vehicle with two large (1,5 MN) solid rocket boosters attached to it. The 256 ton H-2 is capable of placing a 2,2 ton payload into a geo-stationary orbit or a 4 ton payload into geo-stationary transfer orbit. Figure 5: LE-7 engine The LE-7 is a gimballed, staged combustion cycle cryogenic engine. Liquid oxygen and hydrogen are combusted in the combustion chamber at a mixture ratio of 6,0 and a pressure of 147 bar. Expansion of the hot combustion gases is through a bell-type nozzle extension with a nozzle exit diameter of 1,90 m and a 60:1 (geometric) expansion ratio. Total mass flow is estimated at 246,5 kg/s. Before combustion in the main combustion chamber, the propellants are first partially burned in a pre-burner at a fuel-rich mixture ratio of 0,81 (25,6 kg/s liquid oxygen and 31,6 kg/s liquid hydrogen) and a pressure of 240 bar giving a temperature of about 970 K. From the pre-burner, the gases are routed to drive the turbopumps and then combined with more oxygen in the main combustion chamber to burn. Cooling of the thrust chamber is through regenerative cooling using hydrogen fuel as the coolant. After cooling, part of the gaseous fuel is led to the pre-burner to burn with part of the total oxygen flow and part is used to pressurise the hydrogen propellant tank. Helium is used to pressurise the oxygen tank. Feeding of the propellant to the engine is by turbopumps. The liquid hydrogen pump has a discharge pressure of about 319 bar at a flow rate of 40 kg/s. Rotational speed of the pump is rpm and shaft power is 24 MW. The oxygen pump has two outlets and provides the oxygen for the pre-burner as well as the oxygen that directly goes to the main combustion chamber. Discharge pressure is 327 and 213 bar respectively. Pump rotational speed is rpm and shaft power is about 6,5 MW. The LE-7 is mounted at the centre of a crossbeam structure at the rear end of the engine section, Gimballing of the engine is accomplished through an hydraulic actuator system. Start-up of the LE-7 occurs with the propellant fed to the engine by tank pressure. Ignition is by an oxygen/hydrogen igniter which is initiated by an electrical exciter. Total mixture ratio of the main chamber igniter is 6,0 and total mass flow rate is 1,34 kg/s, which is about 0,5% of the main chamber mass flow rate. Development of the LE-7 by Mitsubishi for NASDA was authorised in A total of two engineering model engines and as much as 7 prototype engines were planned. With these engines, 146 firings and a total of 3986 sec of burn time had been conducted as of the end of May 1991 including two full duration test firings. Development cost of the LE-7, originally budgeted at about 593 M$, is 781 M$ (1991). This increase is attributed to technical difficulties related to miss-timing of engine start up early in the development program (early 90 s). The development program ended with the first flight in In the period until 1998, 5 more flights have been conducted. The LE-7 produces a vacuum thrust of 1078 kn and has a vacuum specific impulse of 446 sec (4405 Nsec/kg). Sea level thrust is 910 kn. The engine can be gimballed ±7. It has a single start capability and a mission duty cycle of 346 sec. Engine mass is 1714 kg giving a vacuum thrust to mass ratio of about 64. The engine has a total length of 3,4 m and a maximum diameter of 1,90 m. Engine life is in excess of 2000 sec or 20 starts. 10

17 RD-170/171 and RD180 The Russian RD-170/171 6 engine, designed by Glushko, is the world s most powerful multichamber rocket engine ever flown. It is used to power the first stage of the two stage Russian Zenit (SL-16) rocket and the four strap-on boosters of the Russian Energiya (SL-17) launcher. A development two-chamber version referred to as RD-180 has been selected for use by the Lockheed Martin Evolved Expendable Launch Vehicle (EELV) family, Atlas III and Sea Launch vehicle boosters. The RD-170 [12,14] is a staged combustion cycle, multi-chamber, rocket engine burning oxygen and kerosene as propellants. It is made up of 4 thrust chambers with a single pre-burner, see figure. Each of the 4 thrust chambers consists of a combustion chamber, injector, and nozzle and is mounted on a two-axis gimbal system allowing for pitch and yaw control. Thrust chamber length is 2,26 m and the combustion chamber internal diameter is 0,38 m. Nozzle expansion ratio is 36,9 and the nozzle exit diameter is 1.43 m. Combustion pressure and temperature is 250 bar and 3676 K, respectively. The thrust chambers are regeneratively cooled to cope with the high combustion temperature. All Figure 6: RD-170 engine the kerosene fuel is used to cool the nozzle and combustion chamber before entering through the injector. The feed system consists of both high and low-pressure propellant pumps. The highpressure fuel and oxidiser pumps are driven by a single axis turbine, which is driven by the high-pressure (300 bar) hot (675K) pre-burner exhaust. A hydraulic turbine driven with kerosene supplied by the main (high-pressure) fuel pump drives the low-pressure fuel pump. The low-pressure oxidiser pump is driven with the oxygen-rich pre-burner gas supplied downstream of the high-pressure turbo-pump unit. RD-170/171 development by the Russian Gas Dynamics Laboratory (currently NPO Energomash Design Bureau) started in 1974 with a first flight in 1985 (sub-orbital on Zenit). Since that time, the RD-170/171 has developed as a very reliable system with a reliability of 0,999 for one time use. More than 804 engine flight tests with an accumulated duration of s have been conducted. Also 22 engines have been flight tested during the Zenit and Energiya flights. The engines are manufactured in batches of five: one is tested through 3 life cycles and is subsequently disassembled and inspected, and the other four go through a single acceptance test. Acceptance checkout is accomplished through check firing of each flight unit. In 1993, NPO, in partnership with Pratt & Whitney Space Propulsion, started development on a two-thrust-chamber RD-170 derivative referred to as RD-180. It shares many (about 80%) components with the RD-170, with only a new main turbo-pump and boost pumps. Development took about 42 months with a first firing in the USA in July The RD-170/171 burns liquid oxygen and kerosene at an oxidiser to fuel mass ratio of 2,63. Total vacuum thrust is 7,91 MN (7,26 MN at sea level) with a throttle range from 100%- 56%. Vacuum specific impulse is s and at sea level s. Burn time is s. Each chamber can be gimballed over ±8º along either one (RD-170) or two different axes (RD-171). The RD-170/171 is produced both as an expendable and reusable unit; the latter version can make up to 17 flights. Overall engine diameter and length is 4 m and 3,78 m respectively. Engine mass is 9750 kg. Details of the RD-180 are given in Table RD stands for Raketnay Dvigatel, which translates into rocket motor. 11

18 RS-68 The Rocketdyne RS-68 engine [20] is the first major US engine to emerge from a "blank piece of paper" since the introduction of the Space Shuttle Main Engine in The RS-68 is intended for use on the first stage of the Delta IV medium plus launcher family 7, which consists of a core vehicle with 2-4 solid rocket motors for a GTO payload between 4,7-6,7 tons. The RS-68 is a liquid oxygen and liquid hydrogen booster engine. The RS-68 is a single chamber, gas-generator cycle engine. Its thrust chamber consists of the main injector, the combustion chamber and a nozzle and connects to a thrust frame residing in the top half of the engine. This thrust frame serves as structural support and transmits the loads from the thrust chamber to the vehicle. The combustion chamber is of a regenerative design using liquid hydrogen fuel as the coolant. Combustion chamber pressure is about 96 bar at 100% thrust (or 58 bar at 60% thrust). The nozzle is a four piece ablative bell-type nozzle using the same technology as used for the nozzle of the Space shuttle solid rocket booster. It has a nozzle expansion ratio of 21,5:1. The RS-68 uses a basic gas generator cycle, with gas generator discharges used to drive the turbo-pumps. The rest of the propellant then ends up in the main injector. There are two turbine exhaust appendages. One is used to route gas through a nozzle for roll control. The second is an exhaust coming out of the LOX turbo-pump; it passes through a heat exchanger and then goes overboard. Gaseous oxygen is provided to pressurise the Figure 7: RS-68 liquid oxygen tank. The turbo-machinery is mounted on the thrust frame, along with the gas generator. Primary development of the RS-68 started mid Design emphasis has been on simplicity and ease of fabrication, e.g. by reducing the number of components used, instead of high performance. Component testing started as early as November At that time, a first full-scale main injector was hot-fire tested - the same injector that's on the first assembled engine - along with a representative section of the combustion chamber at almost full thrust. More recently, a comprehensive series of gas generator tests has been conducted. Already in January 1998, only 28 months after the start of the detailed design, an entire engine has been assembled and a first 100% power level test achieved, thereby bringing the total accumulated test duration at 300 seconds. Total development cost until that time being $54 million compared to about $700 million for the SSME. Following prototype testing, engineering and manufacturing development is planned, ending with a critical design review. At that point, the flight design will be "frozen" and the full-scale development program will be underway. Two certification engines will be built and one for use in a full-upstage test series with the common booster core in early Then the engine will be certified for flight. The intent is to run about 300 tests, a cumulative 30,000 seconds. First flight of the RS-68 engine on a Delta IV is planned for Preliminary performance data of the RS-68 indicates a sea level thrust of 2,89 MN (vacuum thrust level of 3,31 MN) and a sea level specific impulse of 365 sec (versus 410 sec in vacuum). Thrust chamber mixture ratio is 6,0. Throttling capability is 60%-100%. Engine mass is 6597 kg and engine height is 5,18 m. 7 This launcher family is under development by Boeing s Space & Communications group. 12

19 Production planning aims at an initial production rate of 30 engines per year, which can grow to 40 engines per year after Rocketdyne plans call for a total number of engines until These engines will all be hot-fire tested at a test stand for acceptance. 13

20 RS-2200 Linear Aerospike Engine The RS-2200 Linear Aerospike Engine [15] is considered by Rocketdyne for use on the Lockheed Martin VentureStar. To this end, an experimental (XRS-2200) version is currently being developed by Rocketdyne for use on the Lockheed Martin X-33 advanced technology demonstrator vehicle. The XRS-2200 is a reduced scale linear aerospike engine which burns oxygen and hydrogen in the mass mixture ratio of 5,5. Combustion and intial expansion takes place in a linear array of small thrust chambers assembled around the base of a central truncated (shortened) linear spike. Thereafter, a linear aerospike nozzle takes over, where the hot gas expands along the spike on one side and into free air on the other side. This allows for a much higher degree of Figure 8: XRS-2000 overexpansion than for a conventional nozzle without the risk of flow separation. Total (calculated) mass flow is about 277 kg/s of which about 43 kg/s hydrogen and 234 kg/s oxygen. Combustion takes place at a pressure of 58,8 bar and initial expansion is up to a nozzle area ratio of 58. Propellant supply is by turbopumps that receive their energy from a gas generator derived from the Saturn J-2 upper stage engine. The gas generator exhaust gases are used to fill up the base at the engine. The engine structure is integral to the vehicle, offering installed weight benefits.. One-axis vectoring is accomplished by running the top and bottom nozzles at different thrusts. This is used for pitch and roll control during flight. Throttling between engines is used for yaw control. The advantage of this concept is a reduction of engine mass, because there is no need for mechanical gimbals. On the other hand, major re-design is necessary when using another vehicle. Ignition of the thrust cells is accomplished by a combustion wave ignition system (CWIS). Pilot fuel and oxidizer propellants are distributed to the thrust cells and ignited using a combustion wave. The ignited pilots then provide the ignition source for the thrust cells. This method uses a pressurized ground helium supply to spin-start the turbopumps and initiates the flow of pressurized propellants to the Gas Generator (GG) and thrust cells. Then a spark ignition system ignites the GG and CWIS. Up to date, over $500 million has been invested in linear and annular aerospike engines and previous full-size ( lbf and lbf) versions of the engine have accumulated 73 tests and over 4000 s of operation. In the beginning of 2000 the longest test run at 100% power to date was set at 125 seconds. The test also marked the first demonstration of plus or minus 15% thrust vector control. The test also demonstrated engine operation at varied power levels and tested different mixture ratios. XRS-2200 nominal performances are a vacuum thrust of 1,2 MN and specific impulse of 428,2 s. At sea level these values are 0,9 MN ( lbf) and 338,3 s, respectively. Engine length is 2,286 m and height is 3,40 m. The width of the engine is also 2,286 m. Thrust-toweight ratio is about 35. For the full size RS-2200, using new materials and a much higher chamber pressure of about 155 bar, the figures in vacuum are a thrust of 2,2 MN and a specific impulse of 455 s and at sea level 1,9 MN and 347 s. Thrust-to-weight ratio (based on vacuum thrust) is expected to be 84, which leads to an engine mass of 2670 kg. 14

21 2. Engines for moderate total impulse launcher missions This Chapter aims to provide an overview of some key engines for moderate total impulse space launcher missions with total impulse ranging from MNs and thrust levels from 50 to 150 kn. Typical such engines are used for upper (second, third and sometimes even fourth) stage propulsion. Attention for moderate total impulse engines for launcher applications is on high gravimetric specific impulse rather than high volumetric specific impulse. This is, because these engines operate at altitudes, where drag losses are not as important as for the first stage and booster engines discussed in the foregoing section. High specific impulse, like for high total impulse engines, is achieved by the selection of oxygen-hydrogen as the propellant combination and the use of a pump-fed feed system. However, since total impulse level are not as extreme as for first stage and booster applications and to allow for a less complex (more reliable) starting procedure with the possibility of engine restart, currently engine working pressure is much more moderate (in the range of about bar) and the complexity of the feed cycle is limited to the gas generator or the bleed/expander cycle. Typical vacuum specific impulse and propellant density levels for these engines are in the range of s and kg/m 3. In the next sections, the detailed workings of a number of moderate total impulse engines is given including the European HM7 and the more recent VINCI engine, the Japanese LE-5, and various versions of the US RL10. The performances of these engines are summarised in Table 2-1 on the next page. 15

22 Table 2-1: Performance data of some typical liquid rocket engines for moderate total impulse missions HM7A HM7B VINCI LE-5 LE-5A RL10 RL10A-3-3A RL10A-4N Engine cycle Gas generator Gas generator Expander Gas generator Expander Expander Expander Expander Propellants LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 LOX/LH 2 Vacuum thrust (kn) 61,6 62, ,7 73,4 88,9 Vacuum specific impulse (s) 441,4 444, ,4 448,9 Overall mixture ratio (-) 4,43 4,56 5,8 5,5/5,6 5,0 5,0 5,5 Propellant density (kg/m 3 ) / Total mass flow rate (kg/s) 14,2 14,4 33,8 23,1 26,9 16,8 Length (m) 1,81 2,01 4,2 2,65 2,65 1,78 Maximum diameter (m) 0,938 0,992 2,1 1,65 1,65 0,9 1,65 Life span (s) Mission duty cycle (s) (332+21) Dry mass (kg) ,5 168 Thrust/weight ratio (-) 42,2 40,9 32,9 40,0 50,6 51,9 54,2 54,0 Restartable (yes/no) No No Yes Yes Yes No No Yes 1 st flight (yr) Development cost (M$) 108 Development period (yr) ) Estimated density. See for density values annex B. 16

23 HM7 engine The HM7 cryogenic engine [2, 21] powers the third stage of the Ariane 1-4 launch vehicles. These vehicles are capable of delivering payload masses of up to 4200 kg into geo-stationary transfer orbit or 7000 kg in to low-earth-orbit (Ariane 44L). The HM7 is a single thrust chamber, cryogenic, gas-generator cycle engine. It generates thrust by the high-speed ejection of gases that are produced by the combustion of LOX and LH 2 at a mixture ratio of 5,14 and a pressure of about 36 bar. These hot gases are accelerated in a bellshaped nozzle with an exit diameter of 0,992 m. Propellant supply is ensured by a two-shaft turbo-pump, which comprises: A high speed liquid hydrogen pump ( rpm), driven directly by a turbine, that boosts hydrogen pressure from 3 to 55 bar; And a low speed liquid oxygen pump ( rpm), driven through a gearbox, to boost LOX pressure from 2 to 50 bar. Turbine power is 380 kw, which is delivered by hot combustion gases produced in a gas generator. This gas generator burns LOX and LH2 at a rate of 0,25 kg/s at a mixture ratio of 0,87 tapped of at the pump outlets. The mixture burned in the gas generator is hydrogen-rich, which limits gas temperature ( K) and keeps the combustion gases reducing instead of oxidising, thus protecting the turbine blades. After driving the turbopumps, the drive gases are exhausted over board through a separate nozzle. Propellant Figure 9: HM7B supply to the combustion chamber is controlled by pneumatic injection valves, which are operated by electrovalves. A regenerative cooling system is used for the combustion chamber and throat assembly; i.e. most of the liquid hydrogen flow is routed through channels in the combustion chamber wall where it is heated to a temperature of 150 K. About 0,15 kg/s of hydrogen is used for nozzle dump cooling where they escape through 242 micronozzles (1 for each cooling tube) for improved specific impulse purposes. To start the engine, a solid propellant turbo-pump starter is used to activate the turbine. Upon activation of the turbine, the propellants start to flow to the thrust chamber. Ignition of the main chamber is insured by a pyrotechnic igniter that brings the propellants to a temperature of 1000K within a very short time (50 milliseconds). Ignition of the gas generator is also by pyrotechnic igniter. Figure 10: HM7B Flow schematic Development of the HM7 engine by SEP (part of Snecma S.A.) started in The development set out based in part on the knowledge gained with the 40 kn thrust HM4 engine developed and tested from Flight qualification of the very first HM7 version took place in This was followed in 1983 with the qualification of the HM7B, which differed from the initial HM7 in a slightly improved specific impulse, due to a slight change in mixture ratio and a 165 seconds longer mission duty cycle. By the end of the qualification program for the HM7 engine, prior to the second flight, 15 engines have accumulated a total of sec. of operation in 150 separate tests. As of June 1995, 111 HM7 engines have been built, with a cumulated total of nearly seconds of operation, including seconds in flight. 17

24 The HM7 model has been introduced on Ariane 1 in 1979, but subsequent launcher versions have utilised the HM7B, offering a 265 s longer burn duration. The HM7B produces a vacuum thrust of 62,2 kn. Its vacuum specific impulse is 444,6 sec. Overall mass flow rate is 14,29 kg/s and includes 2,56 kg/s of liquid hydrogen and 11,72 kg/s of oxygen giving an overall mixture ratio of 4,56. Engine length is 2,013 m and maximum diameter is 0,992m. Engine mass is 155 kg. The engine can be gimballed. 18

25 VINCI engine In May 1999, at a meeting of the European space ministers the development green light was given for the VINCI engine [22]. This engine is intended to power the upper stage of Ariane 5 starting in It will give Ariane 5 a geo-stationary transfer orbit payload capacity of kg. The VINCI engine is a single chamber, expander type, closed-cycle engine. The engine burns 28,5 kg liquid oxygen and 4,5 kg liquid hydrogen per second at a pressure of about 60 bar. Mass mixture ratio of the propellants is 5,8. Acceleration of the hot combustion gases takes place in an extendable nozzle 8 with a maximum nozzle exit diameter of 2,10 m and a nozzle area ratio of 280:1. The upper part of the nozzle is regeneratively cooled. The lower part is made of composite material and is radiatively cooled. Propellant supply is ensured by two turbopumps that are driven in series (hydrogen pump first) by the hydrogen coolant heated by thermal exchange through the walls of the combustion chamber (expander cycle). Pump characteristics are: A high speed liquid hydrogen pump ( rpm) developing 1,8 MW of power to ensure a hydrogen mass flow to the combustion chamber of 5 kg/s; A low speed liquid oxygen pump ( rpm) developing 0,3 MW of power for an oxygen mass flow to the Figure 11: VINCI engine combustion chamber of 28,8 kg/s. The oxidiser is injected directly from the turbopump into the chamber. The hydrogen fuel is first used to cool the nozzle and to drive the turbopumps. Two valves control propellant supply to the combustion chamber, while two other turbine valves regulate turbine power, which in turn regulates mixture ratio (variable from 5,5-6,5) and thrust. The engine will be 5 times restartable and will have a sparktorch igniter. Development green light for the VINCI engine was given at the meeting of European space ministers in May Prime contractor of the program is SNECMA, who leads a team of European companies. Development work will benefit from SEP s work on the Veda demonstration programme. First test firings are planned for 2001 with a first flight in Preliminary performance figures for the VINCI engine indicate a vacuum thrust of 155 kn (with the possibility of operating at a reduced thrust level of about 100 kn) and a vacuum specific impulse of 464 s (nozzle in fully extended position). Overall mixture ratio is 5,8. The engine has a re-start capability in flight. Length and maximum diameter of the engine are 4,20 m and 2,10 m (nozzle in extended position), respectively. Engine mass is 480 kg. 8 The nozzle has been made extendable to limit the dimensions of the stage 1-2 interstage. 19

26 LE-5/5A The LE-5 engine [2, 23, 24] has been developed as a second stage engine of the Japanese H1 rocket launcher. A slightly modified version (LE-5A) currently propels the second stage of the successor of the H1, the H2 rocket launcher. As such, it has a restart capability that allows for two separate firings of the engine. The first firing generates speed and then the engine is shut down temporarily before being relit at high altitude, again generating speed. As the timing of both first and second firings may be freely selected, the LE-5A offers more flexibility in the choice of trajectory and orbit than non-restartable engines. The LE-5 is a multi-start, single chamber, pump-fed, gas generator cycle, cryogenic engine, which burns liquid oxygen and hydrogen at a pressure of about 35 bar. The thrust chamber assembly consists of a combustion chamber and a nozzle extension. The Niobium 200 combustion chamber has a characteristic length of 84 cm and is regeneratively cooled by the hydrogen fuel. The expansion ratio of the chamber is 8,48 so that it can be ground tested at sea level conditions without flow separation in the nozzle. The 70 kg nozzle (extension) is designed as a truncated bell nozzle with an expansion ratio of 140. It is made of austenitic stainless steel and is dumb cooled by the hydrogen fuel. A torch igniter, burning 22 g/s gaseous oxygen and hydrogen at a mixture ratio of 1, ensures proper ignition of the combustion in the main combustion chamber. The feed assembly consists of a separate fuel pump and an oxidiser pump, both driven by a turbine, valves, a gas generator with igniter and a control box and an auxiliary turbine for the gimbal system. Propellant supply is ensured by two turbopumps. The turbines are driven in Figure 12: LE-5A engine series by combustion gases, which are produced in a gas generator burning LOX and LH2 at a nominal mixture ratio of 0,85 and a pressure of 26 bar. The LOX and LH2 are tapped of at the pump outlets. Ignition is like for the main chamber by a gaseous oxygen-hydrogen igniter. The relatively cool turbine exhaust gases are injected back into the nozzle near the end to add to the cooling of the nozzle. To start up the engine, it uses a coolant bleed or expander bleed start, wherein gaseous coolant with a temperature of about 140 K drives the two turbopumps in series. The gaseous coolant is tapped of downstream of the main chamber regenerative cooling jacket and results from cooling of the main chamber during the initial combustion phase. In this phase, tank pressure fed propellants burn in the main combustion chamber. Only after reaching a preprogrammed thrust level, the gas generator is ignited and the engine progresses to a steady gas generator cycle. The LE-5A engine is about equal to the LE-5 except that the gas generator cycle based feed system has been replaced by a hydrogen bleed system, where hot coolant gases drive the turbines. To ensure sufficient energy in the coolant flow, it is tapped off from a point close to the nozzle exit where the coolant attains a temperature of about 1000 K. This design change led to a reduction in mass of about 10 kg and a somewhat higher chamber pressure of 38.5 bar compared to the 35 bar of the LE-5. A second change incorporated is a slightly bigger 20

27 throat area to allow a higher mass flow rate, without increasing the chamber pressure too much. Development of the LE5 was initiated in 1976 and took about 10 years with a first flight of an LE5 on 13 August Development included 3 prototype and 5 flight type engines, and 3 engines for qualification. During development over 480 hot starts and 32000s of total operation time were achieved. Total development costs have been reported to amount to $108 million (1986). The major performances of the LE 5 engine are a vacuum thrust of 100 kn and corresponding specific impulse of 452 s. Engine mass is 255 kg, length is 2,65 m and a largest diameter of 1,65 m, expansion ratio is 140 and a combustion pressure of 36,8 bar and a mixture ratio of 5.5. In addition, the engine has a multi-start capability. Major performances of the LE-5A engine show a slightly higher thrust (121,5 kn). In addition, because of the removal of the gas generator, the specific impulse has increased with about 10s even though the expansion ratio has decreased somewhat by an increased throat diameter. 21

28 RL10A Pratt & Whitney s RL10A [18-19]was the first operational liquid oxygen/hydrogen engine in the world. It fired for the first time in 1959 with a first flight in Current versions of the engine are used on the Centaur stages in both the Atlas and Titan 4 programs. Future use is foreseen for Atlas 2A and 2AS and possibly even future advanced (re-usable) launchers. Figure 13: RL10 engine The RL10A is a pump-fed, expander cycle, rocket engine burning liquid oxygen and hydrogen at a pressure of 32,2 bar (20,4 bar for the original RL10A and 32,2 for the current RL10A-3-3A version). Engine start-up is by using an igniter. The 3-3A thrust chamber assembly consists of a regeneratively cooled combustion chamber and a Columbium nozzle extension with an exit diameter of 1,65 m. The nozzle expansion ratio is 61. The thrust chamber is cooled by the LH 2 fuel, which flows from the propellant tanks to the thrust chamber walls. On its way, the LH 2 fuel heats up and gasifies. This gaseous fuel is then used to drive the propellant pumps. The feed assembly consists of a separate fuel and oxidiser pump, valves, and an engine control unit. Both pumps are driven by a single turbine, which in turn is driven by the expanding (hence expander cycle) gaseous fuel that results from the cooling of the engine. A gearbox reduces the speed of the oxidiser pump. By end 1989, 176 RL10 s of all types had flown in space, accumulating 286 firings with a total of more than 20 hrs of operation, without a single failure. Further orders called for a total of 134 engines for the Atlas and Titan 4 programs. In 1990 Pratt &Whitney started on a 4N version for the Atlas 2A launcher. This version provides a higher vacuum thrust of up to 88,9 kn with 448,9 s specific impulse at a mixture ratio of 5,5. This is accomplished by increasing the chamber pressure up to 38,4 bar and by incorporating an extending nozzle, which increases the nozzle area ratio from 61 to 84. In addition, a multi-start capability has been incorporated allowing insertion of satellites into their final orbit. In 1992, development of an A-5 version for use on a third scale model of a future vertical take-off and landing single stage to orbit launcher has been initiated. This version will be throttle able from 30% - 100% and have an extendable nozzle allowing low altitude operation as well as high-altitude operation like the original A-3 version. At the lower thrust rating, the sea level specific impulse would be 380,5 sec and the chamber pressure 9,8 bar. The values for 100% thrust are 373 s and 32,8 bar chamber pressure. Without extension, the engine has a nozzle expansion ratio of 4,28 and a length of 1,07 m. Engine mass with nozzle extension is 143 kg, only about 2,5 kg heavier than the 3A version. The original RL10A (RL10A-1) generated 66,7 kn vacuum thrust with 412 s specific impulse. The current version (RL10A-3A) generates 73,4 kn with 444,4 sec specific impulse. Engine dry mass is 140,5 kg, length is 1,78 m and largest diameter 1,65m and the oxidiser-to-fuel mass mixture ratio is 5. 22

29 3. Engines for low total impulse launcher missions Low total impulse launcher systems provide the total impulse needed to perform orbit insertion, orbit transfer from low Earth orbit to a higher orbit (for example geostationary Earth orbit), orbit rendezvous and de-orbit of spacecraft. Typical characteristics for such applications are a total impulse level of about 1 MNs up to about MNs, the need to operate after having spent some time in space and intermittent propulsion. In view of the above considerations, most low total impulse engines are of a pressure-fed design, using an earth storable, hypergolic 9, bipropellant. Pressure feeding offers the advantage of a relatively simple feed system, which makes them low cost and very reliable, compared to pump-fed engines. A disadvantage though is that inlet feed pressures are limited to fairly modest levels (up to bar) in order to limit tank pressure and to ensure sufficient tank life. This leads to a less compact engine. To reduce the storage volume required for the pressurant, it is stored in a separate tank at an initial storage pressure of up to about 350 bar. When operating, a pressure regulator regulates the pressure down to a level acceptable for the engine and ensures a constant feed pressure (regulated pressure feed system). Recently, a start has been made to develop a pump-fed unit for such applications. This way, we have the benefit of a higher specific impulse and a smaller engine envelope, but at the expense of higher engine cost and reduced reliability; the latter due to a more complex engine and engine start-up. The future should indicate whether such engines are really viable for these applications or not. In the next sections, the detailed workings of a number of low total impulse engines is given including the Ariane 5 Aestus engine, the Space Shuttle OMS, the Aerojet AJ10-118K and the European ATE engine. The first three engines are all pressure-fed, whereas the latter is of a pump-fed design. The performances of these engines are summarized in Table Self-igniting. 23

30 Table 3-1: Performance data of some LRE s for low total impulse missions OMS Aestus (L7) Aestus II RS-72 ATE AJ10-118K Fuel/oxidiser MMH/NTO MMH/NTO MMH/NTO MMH/NTO MMH/NTO A-50/NTO Cycle Pressure-fed Pressure-fed Pump-fed Pump-fed Pump-fed Pressure-fed Thrust (kn) 26,7 27, , ,4 Throttling capability (%) No No No Thrust Vector Control (deg) ± 6º pitch ± 7º yaw ± 4º ± 6º ± 6º ± 15º Fixed Specific impulse (s) , ,5 Overall mixture ratio (-) 1,65 2,0 2,05 2,05 2,0 1,9 Propellant density (kg/m 3 ) Max. single burn time (s) ? 500 Cumulative life span 15 hrs 100 min See above Number of missions > Number of starts/mission Multiple Multiple > Engine mass (kg) ,2/57,9 124,5 Thrust/weight ratio (-) 23,3 25,5 31,7 36,6 27,4/35,2 35,5 Overall length (m) 1,956 2,2 2,2 2,286 >1,4 2,7 Maximum diameter (m) 1,168 1,27 0,38 1,53 Production cost 2000 (M$) 1,6 >3,5 1) On Transtage, the engine restarts minimum 3 times, but essentially the number of restarts is unlimited. 24

31 Aestus 27,5 kn engine The Aestus 27,5 kn [2, 25] engine is a storable bipropellant medium thrust class rocket engine used on Ariane 5 s L7/9.7 upper stage. Figure 14: Aestus engine The Aestus engine uses the earth-storable hypergolic propellant combination of NTO and MMH, which are burned at a pressure of 10 bar. It uses MMH regenerative chamber cooling to an area ratio of 10:1. The nozzle is radiation cooled and has an area ratio of 83,3:1 with a nozzle exit diameter of 1270 mm. Maximum nozzle wall temperature does not exceed 1400 K, whereas, the combustion chamber wall temperature remains below 750K. The engine is of a pressure-fed design and operates in the regulated mode for the largest part of its mission. Nominal oxidiser and fuel mass flow rate is 5,89 and 2,87 kg/s, respectively. The engine assembly is gimballed by pitch and yaw electromechanical actuators attached by struts at the combustion chamber and at the stage propellant tank. Engine control is by two hydraulic valves, each activated by an electrically controlled pilot valve. These pilot valves in turn are controlled from a control box. Engine instrumentation provides the necessary inputs for the control box. Development of the Aestus engine by MBB (now DaimlerChrysler Aerospace) started in Up to now, a total of three development engines and 5 qualification engines have been produced. Qualification testing has included performance mapping, vacuum starts, operating limits, stability, life cycle, life duration, blow down capability, abnormal operating, environmental, propellant valve and gimbal tests. For this more than 800 firings with a total burn time of 6000 s have been conducted. Recently, DaimlerChrysler, has teamed with Boeing Rocketdyne to develop the Aestus II as part of the Perfo 2000 program, see entry on Aestus II. The baseline L7 version of the Aestus engine develops a nominal vacuum thrust of 27,50 kn. Specific impulse is 320s at a nominal oxidiser to fuel mixture ratio of 2,05. Engine mass is 110 kg and it fits in a envelope with a length of 2,2 m and a diameter of 1,27 m. Maximum gimbal capability in both pitch and yaw direction is 8º of which 4º during operation and 4º by mechanical adjustment. The engine is designed for 20 starts and 6000 sec of cumulative firing duration. During the last 100 seconds of the mission, the engine is operated in a blowdown mode 10, where the chamber pressure is reduced by 15% at engine cut-off. After each firing, MMH passages downstream the engine valve is purged with regulated helium delivered by the propulsion subsystem s helium pressurisation system. European Advanced Technology Engine In 1986, ESA identified the need for an advanced high performance 20 kn rocket engine intended for use as both an upper stage engine for Ariane 5 (to replace the original L5 Aestus engine) and in an orbital propulsion module. This engine is commonly referred to as the European Advanced Technology Engine or ATE [26]. Two thrust chamber versions are under study, a conventional metallic one and a ceramic one. 10 In blow down mode, the engine is operated purely on the pressurant that is left in the propellant tank (pressurant is no longer added). In that case, the feed pressure will drop as the propellant tank gets depleted. This is due to the larger volume available for the pressurant. In that case also mass flow will change and because the effects differ for fuel and oxidiser, one must reckon with a change in mixture ratio. 25

32 The ATE is a gimballed pump-fed bipropellant unit. The thrust chamber assembly consists of the combustion chamber with an injector and a nozzle extension (expansion cone) and a gimbal. It uses MMH regenerative chamber cooling to an area ratio of 10:1. The nozzle has an area ratio of 81:1 and is radiation cooled. The propellants are burned at a nominal oxidiser to fuel mixture ratio of 2,0 and a pressure of 10 bar. The engine assembly is gimballed by pitch and yaw electromechanical actuators attached by struts at the combustion chamber and at the stage propellant tank. The feed system consists of a precombustor, a single shaft turbo-pump, and a control system. An oxidiser-rich precombustor has been selected to produce the hot gas needed to drive the turbo-pump. The turbo-pump is a single-axis design, with oxidiser and fuel pumps at each shaft end and a turbine driving the axis in the middle. A low turbine entry temperature of 779 K, compared to a maximum turbine blade temperature of 1079 K, has been selected to keep the turbine expansion ratio low and to avoid excessive pump delivery pressure. Starting the turbine is by helium until rpm is achieved. Shaft power is about kw, depending on the type of turbine used. The engine control system provides a reliable start and shutdown sequence, stable operation and health monitoring of the engine. The ATE develops a vacuum thrust of 20 kn using an earth-storable hypergolic propellant combination of MON-3 (Mixed Oxides of Nitrogen) and MMH. The nozzle has an area ratio of 81:1 and is radiation cooled. The propellants are burned at a nominal oxidiser to fuel mixture ratio of 2,0 and a pressure of 90 bar. Fiat, Royal Ordnance and Volvo (Sweden) have performed initial studies for ESA. These studies indicate that the engine could be developed in about 8 years at a cost of 340 M$ (2000). Production costs are estimated at 1,6 11 M$ (2000). Aestus II 46 kn engine As part of the Ariane 5 performance improvement program, DaimlerChrysler, who manufactures the Aestus engine has teamed up with Boeing Rocketdyne to develop the Aestus II. The prime modification of the engine is the elimination of the pressure feeding of the baseline engine in favor of a gas generator driven turpopump allowing much higher chamber pressures (50 bar instead of 10 bar for the original Aestus engine). The Aestus II engine will also feature a 280:1 area ratio nozzle instead of the original 83:1. To keep engine size limited, the throat area will be reduced with about a factor 2. These modifications will significantly increase the thrust and specific impulse of the engine, while keeping engine size moderate. For performance details, see Table 3-1. Orbiting Maneuvering System (OMS) engine Aerojet s OMS engine [1, 2] provides the thrust needed to perform orbit insertion, circularisation, transfer, rendezvous and de-orbit of the USA Space Shuttle Orbiter. For this reason, each Orbiter is equipped with two OMS engines, which are located at the aft end of the Orbiter on the left and right sides of the aft fuselage in what is referred to as the Orbital Manoeuvring System/Reaction Control System (OMS/RCS) pods. 11 Cost data given based on 1990 data using an inflation factor of

33 The OMS engine uses the Earth-storable hypergolic bipropellant combination of NTO and MMH. NTO and MMH mass flow rate is 5,37 and 3,25 kg/s, respectively. The propellants are burned in the engine s regeneratively cooled combustion chamber at a pressure of about 9 bar. Expansion takes place through a radiation cooled, Columbium alloy nozzle, with an area ratio of 55:1 and an exit diameter of 116,8 cm. As coolant of the combustion chamber, the MMH fuel is used. The propellant feed system is of a pressure-fed design using gaseous helium stored in a high-pressure tank to provide pressure to the propellant tanks. Engine inlet pressure is estimated at bar. The engine assembly is gimballed by two pitch and yaw electromechanical actuators. The OMS engine has been derived from Aerojet s Apollo Service propulsion system. The first OMS demonstration tests were completed 1972/73, with full development starting in The first prototype engine was delivered in February 1977 for extensive testing. Qualification firings took place in 1979 with more than 270 firings using a single engine and a total accumulated burn time of s. First flight took place in 1981, more than 7 years after start of full development. In 1986, a contract has been awarded to develop an uprated OMS featuring a pump-fed design. This design, however, has not seen production until today. Each OMS engine produces N (or about 26,7 kn) of vacuum thrust. Vacuum specific impulse is 316 s and the oxidiser-to-fuel ratio is 1,65. The engine has a mass of 118 kg and it fits in an envelope of size 1,96 m x 1,17m (diameter). Each engine is designed to be reusable for 100 missions and capable of sustaining 1000 starts (10/mission) with 15 hours of cumulative firing. The gimbal actuation system provides multi-axis gimballing of plus or minus 8 degrees. RS-72 The 55,4 kn RS-72 rocket engine is a commercial joint-development program between Boeing Rocketdyne and DaimlerChrysler Aerospace to provide an advanced engine that addresses the increasing payload and launch vehicle upper stage requirements in both the American and European markets. The RS-72 is a pump-fed, gas generator cycle engine. Its design is based on the flight proven Aestus baseline engine. Increased performance is achieved by integrating a gas-generator driven turbopump allowing a chamber pressure of 61,7 bar. A high expansion ratio nozzle ensures a high specific impulse. Detailed performances are given in Table 3-1. AJ10-118K/138 Aerojet s 12 AJ10-118K [2, 27, 28] engine has flown for than 38 years as second stage engine on McDonnell Douglas (now Boeing) Delta. The engine has also been utilized as the third stage engine for the U.S. Air Force s Titan rocket. In the past, it also has been flying paired (as the AJ10-138) on the Martin Marietta 13 Titan 3 Transtage upper stage. In 1996, Aerojet has been awarded a multi-million dollar follow-on contract to deliver 40 Delta II second stage liquid rocket engines, including launch support, through Aerojet, a segment of GenCorp. 13 Martin Marietta has since then merged with Lockheed to form Lockheed Martin. 27

34 Figure 15: AJ10-118K engine without nozzle extension The AJ10 engine is a pressure-fed engine that uses the Earth-storable hypergolic bipropellant combination of NTO and A Mass flow rate is 9,1 kg/s of NTO and 4,76 kg/s of A50. Combustion takes place in an ablatively cooled combustion chamber at a pressure of 8,9 bar. Ignition is on contact of the fuel and oxidiser (hypergolic propellant). Expansion is through a radiation-cooled nozzle with an area ratio of 65:1 and a throat diameter of 0,187 m. Engine mounting is fixed with no means of thrust vector control. The propellants are fed to the engine by helium gas under high pressure. The AJ engine family originated in the 1950s when Aerojet developed the AJ-10 second stage engine for NASA s Thor rocket. When NASA changed the program s name to Delta in 1960, Aerojet replaced the AJ-10 with the AJ A series of improved versions followed, with the AJ10-118K introduced in the mid-1970s with first flight in Since that date over K engines have flown on either Delta or Titan. The current AJ10-118K based engine produces a nominal vacuum thrust of 43,38 kn. Vacuum specific impulse of the 118K version is 320,5 s (315 for AJ version) and the oxidiser-to-fuel ratio is 1,9. Its mass is about 124,5 kg (108 for AJ version due to a lower area ratio of 40:1), length is about 2,7 m and maximum diameter (nozzle exit) is 1,53 m. Maximum burn time is 500 s with an essentially unlimited number of starts. Engine year 2000 costs are estimated at about 3,5 15 M$ based on a series of 25 engines. 14 A50: Aerozine 50, a mixture of hydrazine and unsymmetrical dimethyl hydrazine. 15 Based on a 1990 cost figure of 2,5 M$. Year 2000 cost has been determined using an inflation factor of

35 4. Engines for reaction control systems A reaction control system 16 (RCS) provides the propulsive thrust necessary for orbit circularisation, orbit manoeuvring, and orbit and attitude control of a spacecraft (launchers, satellites, deep space probes, etc.). Important requirements for a RCS are that it must be able to cope with long mission duration. Typical mission duration for e.g. GEO telecommunication satellites is up to 15 years. Another requirement stems from that the disturbances in space are relatively small and the need for precise pointing. This requires a pulsed mode of operation with several 1000 s of pulses needed over the life of the spacecraft and small total impulse bits per pulse. Finally, because of the small disturbances, total impulse levels required are very modest ranging between 0,1 and 1,5 MNs (depending on mission, mission duration and vehicle mass). In view of the above requirements, all RCS s of today are of a pressure-fed design, using storable propellants. Pressure feeding offers the advantage of a relatively simple feed system compared to pump-fed engines. This makes them low cost and very reliable. A disadvantage though is that, in order to ensure sufficient tank life, inlet feed pressure and hence combustion pressure is limited to fairly modest levels (inlet feed pressures are limited up to about bar) leading to a relatively large and heavy engine. Most engines allow two modes of pressure feeding, a constant-pressure (regulated) mode and a simple blowdown pressurization mode, see the Section on low total impulse engines. For reasons of simplicity, RCS s mostly use fixed thrusters (no means of TVC). However, to allow for attitude and orbit control, usually a network of thrusters (directed in various directions) is used. For example, OLYMPUS employs a network of 22 thrusters and ITALSAT 16 thrusters [27]. Propellants used include earth storable, hypergolic 17 bipropellant or (anhydrous) hydrazine (a monopropellant) and cold gas 18. The bipropellants are mostly used for high performing RCS. For lower performing RCS or if ultra-fine pointing with very low thrust levels is needed, use is made of hydrazine or cold-gas. The latter also allows for a clean exhaust. In the next sections a number of RCS thrusters are discussed and performance figures are given for steady state operation. For pulsed operation, performance values generally are lower 19. For example, specific impulse for pulsed operation of hydrazine thrusters can be about 70-90% of the steady state value, depending on the duty cycle [46]. 16 Sometimes the term attitude and orbit control system (AOCS) is used. 17 Hypergolic refers to that the propellant is self-igniting. 18 Cold gas systems are not chemical thrusters in the real sense of the word, since the energy needed for thrust generation does not come from a chemical reaction. However, like in chemical rocket motors, the thrust is generated by expansion of a high-pressure gas in a nozzle. 19 The first pulse of a pulse train is always inefficient due to that the cold thruster absorbs most of the heat generated and the gases leaving the engine remain relatively cold. Succeeding pulses, depending on the time interval, will reach higher efficiency. 29

36 Bipropellant engines Bipropellant RCS engines or thrusters have been used on a large variety of spacecraft, like many of today s high performing communication satellites (e.g. ASTRA, INTELSAT and INMARSAT), but also on the Space Shuttle and the Apollo and Gemini return capsules. Bipropellant RCS engines use mostly nitrogen tetroxide (NTO; N 2 O 4 ) or Mixed Oxides of Nitrogen (MON) as oxidiser and hydrazine (N 2 H 4 ) or a derivative like monomethyl hydrazine (MMH; N 2 H 3 CH 3 ) as fuel. These bipropellants are earth storable and will selfignite once they are intimately mixed in the combustion chamber. Bipropellant engines can be operated from either a constant-pressure (regulated) feed system or in a simple blowdown pressurization mode. In the latter mode, one must in this case not only take into account that during operation feed pressure and thrust decay as propellant is consumed, but also a change in mixture ratio and hence in energetic properties of the propellant. This is because of different changes in mass flow for the oxidiser and fuel. Bipropellant engines typically take about 1 year to qualify and 6-12 months to produce depending on model and quantity. In the following sections, some bipropellant RCS engines are discussed in some more detail. A performance overview of these engines is given in Table 4-1. Kaiser-Marquardt R-40/40B 4 kn thruster [34] The model R-40 B bipropellant engine is designed to provide perigee and orbit adjust forces for satellites. A slightly modified engine (R-40) is used to provide for small velocity changes along the axis of the Orbiter as well as attitude control (pitch, yaw, and roll). The R-40 single thrust chamber; pressure-fed engine uses MMH and NTO and more recently MON-3 as the propellants. These propellants are burned at a nominal oxidiser-to-fuel ratio of 1,65 and chamber pressure of about 10,5 bar in the engine s combustion chamber. The combustion chamber is made of columbium (C-103) and incorporates radiation and film cooling limiting wall temperature to maximum 1375 K. Film cooling is provided by a thin layer (a film) of fuel. Expansion takes place through a radiation cooled, convergent-divergent nozzle of which several versions are available with different expansion ratios (up to 160) as well as long and short scarf nozzles. The latter is to allow easy integration in the streamline shape of the Space Shuttle. Nozzle exit diameter (160 expansion ratio version) is about 0,65 m. Nominal system inlet (feed) pressure is bar with a demonstrated capability to operate at inlet pressures between 10-27,5 bar. An electromechanical valve controls the propellant flow. Valve power is typically VDC. An R-40 is able to provide a nominal vacuum thrust of 3870 N and vacuum specific impulse of about :1 expansion ratio. By adapting the inlet (feed) pressure, this thrust can be adapted within a range between N. The thruster is capable of providing a maximum total impulse of 92 MNs and has a demonstrated life performance of s steady state and approximately cycles. Pulse width can be as low as 40 ms providing a minimum impulse bit of about 150 Ns. When equipped with a long nozzle, the engine has a mass of 13,6 kg (10,25 kg for an engine with a 20:1 area ratio), a maximum diameter of 0,66 m and a length of about 1,15-1,30 m. 30

37 Atlantic Research Corporation (ARC) 20 LEROS 1 engine [32] The ARC LEROS 1 is a high-performance 467 N thruster designed and developed to fulfil the orbit maneuvering requirements (apogee kick) of GEO telecommunications satellites using dual mode propulsion systems, 21 and has a.o. been applied on ASTRA 1B satellite. The LEROS 1 engine is of a single chamber pressure-fed design. It is made up of a film cooled thrust chamber made of columbium C103 alloy with a silicide-based coating, a 150:1 Columbium expansion cone with the same coating as the combustion chamber and a titanium injector assembly with film coolant orifices. It operates with the propellants MON and MMH, which are burned at a nominal chamber pressure of 6,2 bar. Maximum chamber temperature is 1633 K (1360ºC). An electromechanical actuated propellant control valve controls propellant flow. Propellant inlet (feed) pressure is 14,5 bar, but can be adapted over a range of 10,9-19,2 bar. Development of LEROS 1 occurred in the period First flight was in March 1991, when 2 engines successfully placed an ASTRA 1B satellite into geosynchronous orbit. Nominal performances are a thrust of 467 N and a specific impulse of 314 s at a propellant mixture ratio of 0,77. By selecting a different inlet (feed) pressure, the thrust can be adapted over a wide range, with only minor changes in specific impulse. Demonstrated maximum single burn time is 3000 s. Overall length of the thruster is 0,66 m and its maximum diameter is 0,294 m (determined by the nozzle exit diameter). Thruster (assembly) mass is 4,2 kg including 0,91 kg for the propellant control valve. Kaiser-Marquardt R1E engine [35] The Marquardt model R1E engine is a vernier engine providing 110 N of thrust. This engine is a/o used for attitude (pitch, yaw, and roll) fine control of the Space Shuttle when the Orbiter is above about 21-km altitude. To this end, the Space Shuttle is equipped with a total of 6 such engines divided over three RCS pods (two at the back and 1 in the nose of the vehicle). The R-1 engine is a pressure-fed single thrust chamber design. It operates with the hypergolic propellant combination NTO and MMH. Propellant injection is through a titanium single doublet injector. The cylindrically shaped combustion chamber, providing a minimum expansion ratio of 26:1, is made of Columbium C103 alloy with a silicide coating. Film and radiation cooling allows an operating temperature of 1475 K. A radiation cooled nozzle extension provides for an overall expansion ratio of Figure 16 R1E thruster 100:1 with an exit diameter slightly less than 0,152 m. Engine start-up is by an electrical signal to an on/off valve. Valve power is about VDC. Propellant inlet (feed) pressures are in the range 6,8-27,2 bar with a nominal value of 15 bar. 20 Formerly Royal Ordnance (Great Britain). 21 A dual mode propulsion system is a bi-propellant satellite propulsion system in which the fuel component is hydrazine instead of e.g. MMH. With the use of hydrazine, it becomes possible to use monopropellant catalytic hydrazine thrusters in the attitude control system whilst retaining the performance advantages and operational duty cycle flexibility of a bipropellant apogee engine. An important advantages of using catalytic hydrazine thrusters is an increased reliability and a cleaner exhaust. 31

38 Nominal performances of the R1E thruster are a thrust of 110 N and a specific impulse of 280 s at a propellant mixture ratio of 1,65 and a 15 bar feed pressure. Nominal mass flow rate is 0,0256 kg/s of NTO and 0,0354 kg/s of MMH. By selecting/adapting the feed pressure, the mass flow and hence the thrust can be adapted in between N. Maximum demonstrated burn time is s. Minimum thrust duration is of the order of 0,08 s (80 ms) giving a minimum impulse bit of 0,89 Ns. The thruster s length and maximum diameter is 0,312 m and 0,152 m, respectively. Thrust chamber length is 0,20 m Thruster assembly mass is 4,26 kg including 0,59 kg for the propellant valves. Atlantic Research Corporation (ARC) 22 LEROS 20 thruster [33] The ARC LEROS 20 is a high-performance 22 N thruster designed and developed to fulfil the attitude and orbit control requirements of GEO telecommunication satellites using an all bipropellant unified propulsion system. The single chamber LEROS 20 operates with the hypergolic propellants Mixed Oxides of Nitrogen (MON) and MMH. These propellants are burned in a radiation and film cooled thrust chamber made of columbium (sometimes referred to as Niobium) C103 alloy with a silicide-based coating at a nominal chamber pressure of 9 bar. Maximum combustion chamber temperature is 1433 K Figure 17: Leros-1 engine (1160 C). Propellant injection is through a titanium doublet-type of injector assembly with film cooling orifices. The hot combustion gases are expanded in a bell-type expansion nozzle, also made of Columbium C103, with a 180:1 expansion ratio and a 0,06 m exit diameter. Engine start-up is accomplished by an electrical signal opening an electromechanical flow control valve followed by hypergolic ignition. Propellant feeding is through high gas pressure. Demonstrated propellant feed pressures are in the range 11,7-20,7 bar with a nominal pressure of 15,9 bar. Mixture ratio can be adapted over a range of 1.2 to 2,1. Development of the LEROS 20 thruster started in 1990 by Royal Ordnance (British Aerospace) with qualification in Nominal performances are a thrust of 22 N and a specific impulse of 296 s, which are realized at a propellant mixture ratio of 1,65 and a propellant inlet pressure of 15,9 bar. Demonstrated operational life is over seconds of steady state firing with a maximum single burn time of 7200 s. Maximum cycle life is over pulses. The thruster s overall length is 0,22 m and its maximum diameter is 0,07 m. Thruster mass is 0,73 kg. Kaiser-Marquardt R6 bipropellant thruster [35] The model R6 bipropellant rocket engine is designed to provide attitude control and stationkeeping forces. This engine is used on a great number of satellites including OLYMPUS, ITALSAT, ARABSAT, INSAT, and GOES. The R6 is a derivative of an earlier thruster developed by Marquardt in the early 1960 s. It uses the hypergolic propellant combination of NTO (recently also MON-3) and MMH, which are 22 Formarly Royal Ordnance (GB). 32

39 burned in a Columbium combustion chamber at a pressure of about 6,9 bar and a combustion temperature of about 1575 K. Nominal mass flow rate is 5 g/s of NTO and 3 g/s of MMH. The engine is pressure-fed with nominal inlet pressure in the bar range. Maximum inlet pressure is about 27 bar and minimum about 6 bar. The Columbium nozzle has a convergent-divergent design with an exit diameter of 55,9 mm and an expansion ratio of 100:1. The standard R6 thruster is fully radiation cooled (C-version). A slightly modified version exists wherein the combustion chamber is also film cooled, allowing longer operation times. To locate the thruster inside a satellite, heat shield cone should be provided. An electromechanical (solenoid-type) valve operating in the VDC range controls the propellant flow. Valve power is VDC. Start-up is by electrical signal to the valve followed by hypergolic ignition. Propellant injection is through a Titanium single doublettype of injector. Qualification testing of this thruster has included the demonstration of nearly cycles and 16 hours of continuous firing time. 3 versions of the R6 engine are currently available, including a baseline radiation cooled 22 N version (C version), a radiation and film cooled 22 N version (D version) for extended firing time and a 10 N version (version C-2.2). Typical steady state vacuum impulse is 289 s at a propellant mixture ratio of 1,65. The C version has a mass of 0,67 kg and is capable of cycles with a maximum single burn time of s time (compared to up to s single burn time for the D version). Minimum impulse bit is of the order of 0,115 Ns. Engine length and maximum diameter is 251,7 mm and 55,9 mm, respectively. SEP 20 N engine [38] SEP in 1988 started the design of a new generation of small high-performance bipropellant engines for orbit and attitude control of the European Hermes space plane 23 and of satellites based on the use of carbon matrix composite (CMC) materials. The SEP 20 N engine burns a hypergolic mixture of NTO and MMH in a ceramic matrix carbon thrust chamber allowing for a maximum wall temperature of about 1870 K (1600 o C) without a need for an anti-oxidation coating. The CMC combustion chamber and nozzle are manufactured in two parts with a junction at an area ratio of 18:1. As CMC material, SEPCARBINOX (carbon fibers, silicon carbide matrix) and CERASEP (silicon carbide fibers, silicon carbide matrix) have been selected. The propellants are injected into the chamber through a doublet-type of injector of a conventional metallic design. An expanding graphite gasket is used to ensure a leak tight connection between the metallic injector and the CMC chamber. The engine is of a pressure-fed design with typical combustion chamber pressures of about 8 bar. Total propellant mass flow rate is about 7 g/s of which about 2,7 g/s MMH. Propellant flow is controlled (on/off) using an electromechanical valve. This allows a pulsed mode of operation. Propellant is self-ignitable (hypergolic). Thermal insulation is implemented to prevent a too high heat flow from the chamber to the mounting plate. The SEP 20 N engine provides a nominal vacuum thrust of 20 N and specific impulse in the range of 290/295 s. Nominal mixture ratio is 1,65. The minimum impulse bit that can be delivered is 0,05 Ns and the total number of pulses is with over 200 cold starts (thermal cycles). Total firing time is about 18 hours. Development of the SEP 20 N engine started in 1988 leading to first engine firing in To date, more than 400 thermal cycles and 7000 hot pulses under vacuum have been demonstrated. 23 Development of the Hermes space plane was cancelled in

40 Table 4-1: Performances of typical thrusters used for spacecraft attitude and control applications [27-41] Engine Manuf. Application Vac. Life Cycle Spec. Propellant Mixture Engine Chamber Expansion Thrust Span Life Impulse Ratio Mass Pressure Ratio [N] [s] [cycles] [s] [-] [kg] [bar] [-] S4 DASA RCS 4, MON/MMH - 0, R-2B Marquardt RCS 4,5 >6000 > NTO/Hydrazine 1,65 0, RS-45 Rocketdyne RCS 4, NTO/MMH 1,6 0,73 4,8 175 S10/1 DASA RCS 10, > MON/MMH 1,64 0, R-6C Marquardt RCS 22, > NTO/MMH 1,6 0,67 6,8 100 Leros 20 ARC RCS 22, > NTO/MMH 1,65 0, RS-43 Rocketdyne RCS 22, NTO/MMH 1,6 0,62 6,9 150 R-43 Marquardt RCS 67, NTO/Hydrazine 1, R-1E Marquardt Orbit adjust/rcs > NTO/MMH 1,65 4, RS-25 Rocketdyne NTO/MMH 1,6 0,96 6,9 40 S400/1 DASA Kick motor/rcs MON/MMH 1,64 2,8 7,2 102 RS-42 Rocketdyne NTO/MMH 1,6 2,32 9,7 150 Leros 1 ARC Kick motor MON/Hydrazine 0,8 4,2 6,2 150 R-4D Marquardt Kick motor/rcs > NTO/MMH 1,65 3,76 6,9 164 R-42 Marquardt Orbit adjust MON/MMH 1,65 4, RS-21 Rocketdyne Deep space NTO/MMH 1,52 8, RS-14 Rocketdyne 1401, ,8 NTO/MMH 1,6 8,8 8,5 30 RS-28 Rocketdyne NTO/MMH 1,63 12,7 13,8 40 S3K DASA MON/MMH 1,6-2,1 14, R-40B Marquardt Orbit adjust > NTO/MMH 1,65 13,6 10,5 160 RS-41 Rocketdyne NTO/MMH 1,63 68,95 13, ) Presumably a value achieved under pulsed conditions. 2) A-50 or Aerozine-50 is a mixture of hydrazine and Unsymmetrical Di-Methyl Hydrazine (UDMH). 34

41

42 Catalytic hydrazine thrusters Catalytic hydrazine thrusters have been used on amongst others ESA s Orbital Test Satellite for telecommunications, GEOS, EXOSAT, SPOT,ISO and various METEOSAT satellites for meteorology [27]. Catalytic Hydrazine Thrusters (CHT s), see photo on this page, derive the energy necessary to produce thrust from the decomposition of hydrazine into ammonia, hydrogen and nitrogen. To decompose the hydrazine, a catalyst is needed. One such catalyst is Shell 405 developed by the Shell development company in the USA. It consists of a porous, high surface area, aluminum oxide support, which is impregnated with finely divided iridium, which is contained in a catalyst bed situated in the decomposition (reaction) chamber. According to Sutton [46], the best conditions result by using a catalyst bed that is mm long, a chamber pressure of bar and a hydrazine loading of 0,015-0,060 g/s per mm 2 of surface area (perpendicular to the flow direction) of catalyst bed. A CHT Figure 18: Catalytic hydrazine engines furthermore consists of a nozzle to accelerate the decomposition gases, a heat barrier to prevent excessive heating of the environment, and an electro-mechanically activated flow control valve. The latter regulates through on/off regulation the flow of hydrazine to the thruster. As material for the thrusters, mostly titanium or stainless steel are used. Both materials offer good compatibility with hydrazine. As such, these materials are also applied in the construction of the necessary valves, connecting pipework, tanks, etc. CHT s, like bipropellant RCS thusters (see previous section), can be operated from either a constantpressure (regulated) feed system or in a simple blowdown pressurization mode. Operational temperatures for these thrusters range in between 4-71 C. CHT s are available in a wide range of thrust levels between 0,75 N and 450 N and with steady state specific impulse levels ranging from about sec. In pulsed mode, specific impulse values (and hence thrust values) are typically between 70-90% of the calculated steady state value, depending on the duty cycle. This decrease is mostly due to that initially most of the heat generated is absorbed by the catalyst bed. To prevent such losses, sometimes catalyst bed heaters are used to preheat the catalyst. Minimum impulse bits that can be achieved are in the range 0,005-0,036 Ns. Table 4-2 lists some CHT s used for spacecraft propulsion and some of their important performances. Typical CHT cost ranges from about $ for a 1N thruster to about $ for a 20 N thruster. 36