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*HN03 or N204 oxidizer with N2H4) MMH, or UDMH as fuels (see Chapter 7). TC = thrust chamber, C = chamber wall, N = nozzle convering section wall and throat region walls, E = walls at exit region of diverging section of nozzle, I = injector face, F = fuel cooled (regenerative), R = radiation cooled, U = uncooled, T = transient heat transfer or heat sink method (ablative material).

Eleelrodeposited outer shell

Milled cooling channel

Hot combustion gas

Eleelrodeposited outer shell

Milled cooling channel

Hot combustion gas

Estimated temperature distribution for above construction

P^iji-b-^ Isotherms (solid lines)

Dashed lines indicate direction of heat flux

Hottest areas

FIGURE 8-17. Enlarged cross section of thrust chamber's regenerative cooling passages for two types of design.

for small thrusters with many pulses, also the fatigue of valve seats. Therefore, there is a maximum limit on the number of firings that such a thrust chamber can withstand safely, and this limits its useful life (see Refs. 8-7 and 8-8).

For radiation cooling, several different carbon materials have worked in a reducing, fuel-rich rocket atmosphere. At other gas mixtures they can oxidize at the elevated temperatures when they glow red or white. They can be used at wall temperatures up to perhaps 3300 K or 6000 R. Carbon materials and ablative materials are used extensively in solid propellant rocket motors and are discussed further in Chapter 14.

For some small radiation-cooled bipropellant thrusters with storable pro-pellants, such as those of the reaction control thrusters on the Space Shuttle Orbiter, the hot walls are made of niobium coated with disilicide (up to 1120 K

or 2050 R). To prevent damage, a fuel-rich mixtures or film cooling is often used. Rhenium walls protected by iridium coatings (oxidation resistant) have come into use more recently and can be used up to about 2300 K or 4100 R (see Ref. 8-9). Other high temperature materials, such as tungsten, molybdenum, alumina, or tantalum, have been tried, but have had problems in manufacture, cracking, hydrogen embrittlement, and excessive oxidation.

A small radiation-cooled monopropellant thruster is shown in Fig. 8-16 and a small radiation cooled bipropellant thruster in Fig. 8-18. This thruster's injection has extra fuel injection holes (not shown in Fig. 8-18) to provide film cooling to keep wall temperatures below their limits. This same thruster will also work with hydrazine as the fuel.

Until recently it has not been possible to make large pieces of carboncarbon material. This was one of the reasons why large nozzle sections and integral nozzle-exit-cone pieces in solid motors were made from carbon phenolic cloth lay-ups. Progress in manufacturing equipment and technology has now made it possible to build and fly larger c-c pieces. A three-piece extendible c-c nozzle exit cone of 2.3 m (84 in.) diameter and 2.3 to 3 mm thickness has recently flown on an upper-stage engine. This engine with its movable nozzle

Solenoid operated fuel valve

Unlike doublet pattern injector with additional film coolant injection holes near periphery

Solenoid operated fuel valve

Unlike doublet pattern injector with additional film coolant injection holes near periphery

Solenoid operated oxidizer valve

Combustion chamber with integral nozzle throat, rhenium, coated with iridium

Solenoid operated oxidizer valve

Mounting flange and injector assembly

Combustion chamber with integral nozzle throat, rhenium, coated with iridium

Upper nozzle exit section, niobium with disiiicide coating

Lower nozzle exit section, titanium

551.94 mm

Thrust 100 Ibf

Chamber pressure ~ 140 psia Nozzle area ratio 250 to 375 Specific impulse up to 323 sec Mass 10.5 ibm

FIGURE 8-18. Radiation-cooled reaction control thruster R-4D-15 uses nitrogen tetroxide and monomethylhydrazine propellants. The large nozzle area ratio allows good vacuum performance. It has three different nozzle materials, each with a lower allowable temperature (Re 4000°F; Nb 3500°F; Ti 1300°F. (Courtesy of KaiserMarquardt Company.)

extension is shown in Fig. 8-19, its parameters are listed in Table 8-1, and its testing is reported in Ref. 8-4.

The material properties have to be evaluated before a material can be selected for a specific thrust chamber application. This evaluation includes physical properties, such as tensile and compressive strengths, yield strength, fracture toughness, modulus of elasticity (for determining deflections under load), thermal conductivity (a high value is best for steady-state heat transfer), coefficient of thermal expansion (some large thrust chambers grow by 3 to 10 mm when they become hot, and that can cause problems with their piping connections or structural supports), specific heat (capacity to absorb thermal energy), reflectivity (for radiation), or density (ablatives require more volume than steel). All these properties change with temperature (they are different when they are hot) and sometimes they change with little changes in composition. The temperature where a material loses perhaps 60 to 75% of its room temperature strength is often selected as the maximum allowable wall temperature, well below its melting point. Since a listing of all the key properties of a single material requires many pages, it is not possible to list them here, but they are usually available from manufacturers and other sources. Other important properties are erosion resistance, little or no chemical reactions with the pro-pellants or the hot gases, reproducible decomposition or vaporization of ablative materials, ease and low cost of fabrication (welding, cutting, forming, etc.), the consistency of the composition (impurities) of different batches of each material (metals, organics, seals, insulators, lubricants, cleaning fluids), and ready availability and low cost of each material.

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