location along the thrust chamber profile. The number of channels or tubes will determine the shape of the cross section, ranging from deep and thin to almost square. The effect of varying the number of channels or channel dimensions and shape is shown in Table 8-6. The minimum inert mass of the cooling jacket and a low friction loss occur, when the shape (which varies axially throughout the jacket) is on the average close to a square. On the basis of analyses, as shown in the table, a 150-channel design has been selected for giving favorable cross section, reasonable dimensions for ease of fabrication, good cooling and often low thermal wall stresses.
Reinforcing bands have to be put on the outside of the tubes or channels to hold the internal gas pressure during operation, to contain the coolant pressures, which cause heated walls wanting to become round, and any surge pressures during the start transient or arising from water hammer in the lines. We assume a surge pressure of 50% above chamber pressure and a steel strength a of 120,000 psi. In the chamber the inside diameter is 16.7 cm (6.57 in.), the walls and channels are 0.3 cm thick, and the pressure is 700 psia or 4.826 MPa. If one band allows the reinforcing of a length of chamber of 3.0 in. the cross sectional area of that reinforcing band will be A = pDL/ila) = [700 x 1.5 x (6.57 + 0.3) x 3]/(2 x 120,000) = 0.0902 in. If the band were 1.0 in. wide, its thickness would be 0.09 in. and if it were 3 in. wide it would be 0.3 in. thick. Large nozzle exit sections have been observed to experience flutter or cyclic deformation, and therefore some stiffening rings may be needed near the exit.
The capacity of the fuel to absorb heat is approximately cpmfAT = 0.5 x 4.81 x 200 = 278,000 J/sec. The maximum AT is established by keeping the fuel well below its chemical decomposition point. This calculated heat absorption is less than the heat transfer from the hot gases. It is therefore necessary to reduce the gas temperature near the chamber and nozzle walls or to increase the heat absorption. This can be accomplished by (1) introducing film cooling by injection into the chamber just ahead of the nozzle, by (2) modifying the injection patterns, so that a cooler, fuel-rich thick internal boundary layer is formed, or (3) by allowing some nucleate boiling in the throat region. The analysis of these three methods is not given here. Item (2), supplementary cooling, is selected because it is easy to design and build, and can be based on extensive data of prior favorable experience. However it causes a small loss of performance (up to about 1 % in specific impulse).
Injector Design. The injector pattern can be any one of the several types shown in Figs. 8-3 and 8-^k For this propellant combination we have used both doublets (like and unlike), and triplets in the USA, and the Russians have used multiple hollow double posts with swirling or rotation of the flow in the outer annulus. Based on good experience and demonstrated combustion stability with similar designs, we select a doublet self impinging type stream pattern and an injector design similar to Fig 8^1. The impinging streams form fans of liquid propellant, which break up into droplets. Oxidizer and fuel fans alternate radially. We could also use a platelet design, like Fig. 8-5.
The pressure drop across the injector is usually set at values between 15 and 25% of the chamber pressure, in part to obtain high injection velocities, which aid in atomization and droplet breakup. In turn this leads to more complete combustion (and thus better performance) and to stable combustion. We will use 20% or 140 psi or 0.965 MPa for the injector pressure drop. There is a small pressure loss in the injector passages. The injection velocities are found from Eqs. 8-1 and 8-5. The equation is solved for the area A, which is the cumulative cross-section area of all the injection holes of one of the propellants in the injector face.
With rounded and clean injection hole entrances the discharge coefficient will be about 0.80 as shown in Table 8-2. Solving for the cumulative injection hole area for the fuel and the oxidizer flow gives 1.98 cm2 for the fuel and 4.098 cm2 for the oxidizer. A typical hole diameter in this size of injector would be about 0.5 to 2.5 mm. We will use a hole size of 1.5 mm for the fuel holes (with 90% of the fuel flow) and 2.00 mm for the oxidizer hole size, resulting in 65 doublets of oxidizer holes and 50 doublets of fuel. By using a slightly smaller fuel injection hole diameter, we can match the number of 65 doublets as used with the oxidizer holes. These injection doublets will be arranged on the injector face in concentric patterns similar to Fig. 8-^1. We may be able to obtain a slightly higher performance by going to smaller hole sizes and a large number of fuel and oxidizer holes. In addition there will be extra fuel holes on the periphery of the injector face to help in providing the cooler boundary layer, which is needed to reduce heat transfer. They will use 10% of the fuel flow and, for a 0.5 mm hole diameter, the number of holes will be about 100. To make a good set of liquid fans, equal inclination angles of about 25° are used with the doublet impingements. See Fig. 8-7.
Ignition. A pyrotechnic (solid propellant) igniter will be used. It has to have enough energy and run long enough to provide the pressure and temperature in the thrust chamber for good ignition. Its diameter has to be small enough to be inserted through the throat, namely 8.0 cm maximum diameter and 10 to 15 cm long.
Layout Drawings, Masses, Flows, and Pressure Drops. We now have enough of the key design parameters of the selected thrust chamber, so a preliminary layout drawing can be made. Before this can be done well, we will need some analysis or estimates on the manifolds for fuel and oxidizer, valve mounting provisions and their locations, a nozzle closure during storage, a thrust structure, and possibly an actuator and gimbal mount, if gimballing is required by the mission. A detailed layout or CAD (Computer Aided Design) image (not shown in this analysis) would allow a more accurate picture and a good determination of the mass of the thrust chamber and its center of gravity both with and without propellants.
Estimates of gas pressures, liquid pressures (or pressure drops) in the flow passages, injector, cooling jacket, and the valves are needed for the stress analysis, so that various wall thicknesses and component masses can be determined. Material properties will need to be obtained from references or tests. A few of these analyses and designs may actually change some of the data we selected or estimated early in this sample analysis and some of the calculated parameters may have to be re-analyzed and revised. Further changes in the thrust chamber design may become evident in the design of the engine, the tanks, or the interface with the vehicle. The methods, processes and fixtures for manufacturing and testing (and the number and types of tests) will have to be evaluated and the number of thrust chambers to be built has to be decided, before we can arrive at a reasonable manufacturing plan, a schedule and cost estimates.
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