Co

Oxidizer thrust vectoring control valve (2 places)

Oxidizer turbopump

Thrust cell (10 ea. bank)

Fuel turbopump

Thrust cell (10 ea. bank)

Engine base

Oxidizer turbopump

Fuel turbopump

Nozzle ramp (cooled)

Heat exchanger

Oxidizer thrust vectoring control valve (2 places)

Engine base

Oxidizer turbopump

Powerpack isolation valve (LOX)

Thrust cell (10 ea. bank)

Fuel turbopump

Oxidizer thrust vectoring control valve (2 places)

Oxidizer turbopump

Thrust cell (10 ea. bank)

Fuel turbopump

Nozzle ramp (cooled)

Thrust, sea level/vacuum, Ibf 206,200/266,000 Mixture ratio 02/H2 5.5

Specific impulse, SL/vac.(sec) 340/429 Dimensions, (in.), aft end 46 W x 88 L

Chamber pressure, psia 854 forward end 133 W x 88 L

FIGURE 8-13. Side view and oblique top view of the XRS-2200 aerospike linear rocket engine with 20 thrust cells and two curved fuel-cooled ramps. (Courtesy of The Boeing Company, Rocketdyne Propulsion and Power.)

Propellant enters through

Propellant enters through

FIGURE 8-14. Pressure profile and flow pattern along the ramp of an aerospike nozzle.

FIGURE 8-14. Pressure profile and flow pattern along the ramp of an aerospike nozzle.

and the ambient air stream, creating further expansion waves. At high altitude there are no compression waves emanating from the ramp and the expanding diverging flow exerts a decreasing pressure on the ramp area and behaves very similarly to the diverging nozzle flow in a bell-shaped conventional nozzle exit section. The contact locations, where the compression waves touch the ramp surface, have higher local heat transfer than other areas on the ramp surface; these locations move downstream as the vehicle gains altitude. The wave patterns and the pressure distribution on the spike or ramp surface can be determined from computerized fluid dynamics programs or a method of characteristics program.

The advantages claimed for a linear aerospike engine are these: (1) compared to the axisymmetric rocket engine, it fits well into the trailing edge of a winged or lifting body type vehicle and often has less engine and structural mass; (2) it has altitude compensation and thus operates at optimum nozzle expansion and highest possible performance at every altitude; (3) differential throttling of certain sets of individual thruster modules allows pitch, yaw, and roll control of the vehicle during powered flight, as explained in Chapter 16. There is no gimbal joint, no movement of the nozzle, no actuators, and no actuator power supply or strong structural locations for actuator side loads; (4) the truncated aerospike is short and requires less vehicle volume and structures; and (5) the engine structure can be integrated with the vehicle structure, avoiding a separate vehicle structure at or near the engines. The disadvantages include the lack of proven flight experience, proven reliability and performance validation (which are expected to happen soon), and a larger-than-usual surface area subject to high heat transfer.

Low-Thrust Rocket Thrust Chambers or Thrusters

Spacecraft, certain tactical missiles, missile defense vehicles, and upper stages of ballistic missiles often use special, multiple thrusters in their small, liquid propellant rocket engines. They generally have thrust levels between about 0.5 and 10,000 N or 0.1 to 2200 lbf, depending on vehicle size and mission. As mentioned in Chapter 4, they are used for trajectory corrections, attitude control, docking, terminal velocity control in spacecraft or ballistic missiles, divert or side movement, propellant settling, and other functions. Most operate with multiple restarts for relatively short durations during a major part of their duty cycle. As mentioned in Chapter 6, they can be classified into hot gas thrusters (high-performance bipropellant with combustion temperatures above 2600 K and Is of 200 to 325 sec), warm gas thrusters such as monopropellant hydrazine (temperatures between 500 and 1600 K and Is of 18 to 245 sec), and cold gas thrusters such as high-pressure stored nitrogen (200 to 320 K) with low specific impules (40 to 120 sec).

A typical small thruster for bipropellant is shown in Fig. 8-15 and for hydrazine monopropellant in Fig. 8-16. For attitude control angular motion these thrust chambers are usually arranged in pairs as explained in Section 4.6. The same control signal activates the valves for both units of such a pair. All these small space rocket systems use a pressurized feed system, some with positive expulsion provisions, as explained in Section 6.3. The vehicle mission and the automatic control system of the vehicle often require irregular and frequent pulses to be applied by pairs of attitude control thrust chambers, which often operate for very short periods (as low as 0.01 to 0.02 sec). This type of frequent and short-duration thrust application is also known as pulsed thrust operation. For translation maneuvers a single thruster can be fired (often in a pulsing mode) and the thrust axis usually goes through the center of gravity of the vehicle. The resulting acceleration will depend on the thrust and the location of the thruster on the vehicle; it can be axial or at an angle to the flight velocity vector.

There is a performance degradation with decreasing pulse duration, because propellants are used inefficiently during the buildup of thrust and the decay of

Columbium chamber with disilicide coating (0.003 in. thick)

Columbium chamber with disilicide coating (0.003 in. thick)

2 Propellant valves

FIGURE 8-15. This radiation-cooled, insulated vernier thruster is one of several used on the Reaction Control System of the Space Shuttle vehicle for orbit stabilization and orientation, rendezvous or docking maneuvers, station keeping, deorbit, or entry. The nozzle is cut off at an angle to fit the contour of the vehicle. Performance data are given in Table 6-3. Operation can be in a pulse mode (firing durations between 0.08 and 0.32 sec with minimum offtime of 0.08 sec) or a steady-state mode (0.32 to 125 sec). Demonstrated life is 23 hr of operation and more than 300,000 starts. (Courtesy of Kaiser Marquardt Company and NASA.)

2 Propellant valves

FIGURE 8-15. This radiation-cooled, insulated vernier thruster is one of several used on the Reaction Control System of the Space Shuttle vehicle for orbit stabilization and orientation, rendezvous or docking maneuvers, station keeping, deorbit, or entry. The nozzle is cut off at an angle to fit the contour of the vehicle. Performance data are given in Table 6-3. Operation can be in a pulse mode (firing durations between 0.08 and 0.32 sec with minimum offtime of 0.08 sec) or a steady-state mode (0.32 to 125 sec). Demonstrated life is 23 hr of operation and more than 300,000 starts. (Courtesy of Kaiser Marquardt Company and NASA.)

thrust, when they operate below full chamber pressure and the nozzle expansion characteristics are not optimum. The specific impulse suffers when the pulse duration becomes very short. In Section 3-5 the actual specific impulse of a rocket operating at a steady state was given at about 92% of theoretical specific impulse. With very short pulses (0.01 sec) this can be lower than 50%, and with pulses of 0.10 sec it can be around 75 to 88%. Also, the reproducibility of the total impulse delivered in a short pulse is not as high after prolonged use. A preheated monopropellant catalyst bed will allow performance improvement in the pressure rise and in short pulses.

One way to minimize the impulse variations in short pulses and to maximize the effective actual specific impulse is to minimize the liquid propellant passage volume between the control valve and the combustion chamber. The propellant flow control valves for pulsing attitude control thrust chambers are therefore often designed as an integral part of the thrust chamber-injector assembly, as shown in Fig. 8-15. Special electrically actuated leakproof, fast-acting valves with response times ranging from 2 to 25 msec for both the opening and closing operation are used. Valves must operate reliably with predictable characteristics for perhaps 40,000 to 80,000 starts. This in turn often requires endurance proof tests of 400,000 to 800,000 cycles.

Shower head

FIGURE 8-16. Typical hydrazine monopropellant small thrust chamber with catalyst bed, showing different methods of injection.

Shower head

FIGURE 8-16. Typical hydrazine monopropellant small thrust chamber with catalyst bed, showing different methods of injection.

Liquid storable bipropellants such as N204-monomethylhydrazine are used when high performance is mandatory. Some have used ablative materials for thrust chamber construction, as in the Gemini command module. The Space Shuttle small thrusters use radiation cooling with refractory metals, as shown in Fig. 8-15. A radiation cooled thruster is shown later in Fig. 8-18. Carbon materials made of woven strong carbon fibers in a matrix of carbon have also been used in radiation-cooled bipropellant thrusters.

Hydrazine monopropellant thrusters are used when system simplicity is important and moderate performance is acceptable. They have a nontoxic, clear, clean exhaust plume. Virtually all hydrazine monopropellant attitude control rockets use finely dispersed iridium or cobalt deposited on porous ceramic (aluminum oxide) substrate pellets 1.5 to 3 mm in diameter as a catalyst. Figure 8-16 shows a typical design of the catalyst pellet bed in an attitude control rocket designed for pulse and steady-state operation meeting a specific duty cycle. Each injection hole is covered with a cylindrical screen section which extends into a part of the catalyst bed and distributes the hydrazine propellant. Fig. 8-16 also shows other successful types of hydrazine injector. Several arrangements of catalyst beds are employed; some have spring-loading to keep the pellets firmly packed. Hydrazine monopropellant thrust units range in size from 0.2 to 2500 N of thrust; the scaling procedure is empirical and each size and design requires testing and retesting. The amount of ammonia decomposition, shown in Fig. 7-3, can be controlled by the design of the catalyst bed and its decomposition chamber.

Mechanical, thermal, and chemical problems arise in designing a catalyst bed for igniting hydrazine, the more important of which are catalytic attrition and catalyst poisoning. Catalytic attrition or physical loss of catalyst material stems from motion and abrasion of the pellets, with loss of very fine particles. Crushing of pellets can occur because of thermal expansion and momentary overpressure spikes. As explained in Chapter 7, the catalyst activity can also decline because of poisoning by trace quantities of contaminants present in commercial hydrazine, such as aniline, monomethylhydrazine, unsymmetrical dimethylhydrazine, sulfur, zinc, sodium, or iron. Some of these contaminants come with the hydrazine and some are added by the tankage, pressurization, and propellant plumbing in the spacecraft. The high-purity grade has less than 0.003% aniline and less than 0.005% carbonaceous material; it does not contaminate catalysts. Catalyst degredation, regardless of cause, produces ignition delays, overpressures, and pressure spikes, decreases the specific impulse, and decreases the impulse duplicate bit per pulse in attitude control engines.

Figure 19-4 shows a combination of chemical monopropellant and electrical propulsion. Electrical post-heating of the reaction gases from catalysis allows an increase of the altitude specific impulse from 240 sec to about 290 or 300 sec. A number of these combination auxiliary thrusters have successfully flown on a variety of satellite applications and are particularly suitable for spacecraft where electrical power is available and extensive short-duration pulsing is needed.

Cold gas thrusters and their performance were mentioned in Section 6.8 and their propellants and specific impulses are listed in Table 7-3. They are simple, low cost, used with pressurized feed systems, used for pulsing operations, and for low thrust and low total impulse. They can use aluminum or plastics for thrusters, valves and piping. The Pegasus air-launched launch vehicle uses them for roll control only. The advantages of cold gas systems are: (a) they are very reliable and have been proven in space flights lasting more than 10 years; (b) the system is simple and relatively inexpensive; (c) the ingredients are nontoxic; (d) no deposit or contamination on sensitive spacecraft surfaces, such as mirrors; (e) they are very safe; and (f) capable of random pulsing. The disadvantages are: (a) engines are relatively heavy with poor propellant mass fractions (0.02 to 0.19); (b) the specific impulses and vehicle velocity increments are low, when compared to mono- or bipropellant systems; and (c) relatively large volumes.

Materials and Fabrication

The choice of the material for the inner chamber wall in the chamber and the throat region, which are the critical locations, is influenced by the hot gases resulting from the propellant combination, the maximum wall temperature, the heat transfer, and the duty cycle. Table 8-3 lists typical materials for several thrust sizes and propellants. For high-performance, high heat transfer, regen-eratively cooled thrust chambers a material with high thermal conductivity and a thin wall design will reduce the thermal stresses. Copper is an excellent conductor and it will not really oxidize in fuel-rich non-corrosive gas mixtures, such as are produced by oxygen and hydrogen below a mixture ratio of 6.0. The inner walls are therefore usually made of a copper alloy (with small additions of zirconium, silver, or silicon), which has a conductivity not quite as good as pure (oxygen-free) copper but has improved high temperature strength.

Figure 8-17 shows a cross section of a cooling jacket for a large, regenera-tively cooled thrust chamber with formed tapered tubes that are brazed together. The other fabrication technique is to machine nearly rectangular grooves of variable width and depth into the surface of a relatively thick contoured high-conductivity chamber and nozzle wall liner; the grooves are then filled with wax and, by an electrolyte plating technique, a wall of nickel is added to enclose the coolant passages (see Fig. 8-17). The wax is then melted out. As with tubular cooling jackets, a suitable inlet and outlet manifolds are needed to distribute and collect the coolant flow. The figure also shows the locations of maximum wall temperature. For propellant combinations with corrosive or aggressive oxidizers (nitric acid or nitrogen tetroxide) stainless steel is often used as the inner wall material, because copper would chemically react. The depth and width of milled slots (or the area inside formed tubes) vary with the chamber-nozzle profile and its diameters. At the throat region the cooling velocity needs to be at its highest and therefore the cooling passage cross section will be at its lowest.

The failure modes often are bulging on the hot gas side and the opening up of cracks. During hot firing the strain at the hot surface can exceed the local yield point, thus giving it a local permanent compressive deformation. With the cooldown after operation and with each successive firing, some additional yielding and further plastic deformation will occur until cracks form. With successive firings the cracks can become deep enough for a leak and the thrust chamber will then usually fail. The useful life of a metal thrust chamber is the maximum number of firings (and sometimes also the cumulative firing duration) without such a failure. The prediction of wall failures is not simple and Refs. 8-5 and 8-6 explain this in more detail. Useful life can also be limited by the storage life of soft components (O-rings, gaskets, valve stem lubricant) and,

TABLE 8-3. Typical Materials used in Several Common Liquid Propellant Thrust Chambers

Application

Propellant

Components

Cooling Method

Typical Materials

Bipropellant TC,

Oxygen-

C, N, E

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