Turbopump Feed Systems And Engine Cycles

The principal components of a rocket engine with one type of turbopump system are shown in the simplified diagram of Fig. 1-4. Here the propellants are pressurized by means of pumps, which in turn are driven by turbines. These turbines derive their power from the expansion of hot gases. Engines with turbopumps are preferred for booster and sustainer stages of space launch vehicles, long-range missiles, and in the past also for aircraft performance augmentation. They are usually lighter than other types for these high thrust, long duration applications. The inert hardware mass of the rocket engine (without tanks) is essentially independent of duration. Examples can be seen In Figs. 6-1 and 6-9 and also in Refs. 6-1, 6-2, and 6-6. For aircraft performance augmentation the rocket pump can be driven directly by the jet engine, as in Ref. 6-12. From the turbopump feed system options depicted in Fig. 6-2, the designer can select the most suitable concept for a particular application.

An engine cycle for turbopump-fed engines describes the specific propellant flow paths through the major engine components, the method of providing the hot gas to one or more turbines, and the method of handling the turbine exhaust gases. There are open cycles and closed cycles. Open denotes that the working fluid exhausting from the turbine is discharged overboard, after having been expanded in a nozzle of its own, or discharged into the nozzle of the thrust chamber at a point in the expanding section far downstream of the nozle throat. In closed cycles or topping cycles all the working fluid from the turbine is injected into the engine combustion chamber to make the most efficient use of its remaining energy. In closed cycles the turbine exhaust gas is expanded through the full pressure ratio of the main thrust chamber nozzle, thus giving a little more performance than the open cycles, where these exhaust gases expand only through a relatively small pressure ratio. The overall engine performance difference is typically between 1 and 8% of specific impulse and this is reflected in even larger differences in vehicle performance.

Figure 6-9 shows the three most common cycles in schematic form. Reference 6-13 shows variations of these cycles and also other cycles. The gas generator cycle and the staged combustion cycle can use most of the common liquid propellants. The expander cycle works best with vaporized cryogenic hydrogen as the coolant for the thrust chamber, because it is an excellent heat absorber and does not decompose. The schematic diagrams of Fig. 6-9 show each cycle with a separate turbopump for fuel and for oxidier. However, an arrangement with the fuel and oxdizer pump driven by the same turbine is also feasible and sometimes reduces the hardware mass, volume, and cost. The "best" cycle has to be selected on the basis of the mission, the suitability of existing engines, and the criteria established for the particular vehicle. There is an optimum chamber pressure and an optimum mixture ratio for each application, engine cycle, or optimization criterion, such as maximum range, lowest cost, or highest payload.

In the gas generator cycle the turbine inlet gas comes from a separate gas generator. Its propellants can be supplied from separate propellant tanks or can be bled off the main propellant feed system. This cycle is relatively simple; the pressures in the liquid pipes and pumps are relatively low (which reduces inert engine mass). It has less engine-specific impulse than an expander cycle or a staged combustion cycle. The pressure ratio across the turbine is relatively

Flight ArcOpen Cycle Turbine Turbopump
FIGURE 6-9. Simplified diagrams of three engine cycles for liquid propellant rocket engines. The spirals are a symbol for an axisymmetric cooling jacket where heat is absorbed.

high, but the turbine or gas generator flow is small (1 to 4% of total propellant flow) if compared to closed cycles. Some early engines used a separate mono-propellant for creating the generator gas. The German V-2 missile engine used hydrogen peroxide, which was decomposed by a catalyst. Typically, the turbine exhaust gas is discharged overboard through one or two separate small low-area-ratio nozzles (at relatively low specific impulse), as shown schematically in Fig. 1^1 and in the Vulcain engine or RS-68 engine listed in Table 10-3. Alternatively, this turbine exhaust can be aspirated into the main flow through openings in the diverging nozzle section, as shown schematically in Fig. 6-9. This gas then protects the walls near the nozzle exit from high temperatures. Both methods can provide a small amount of additional thrust. The gas generator mixture ratio is usually fuel rich (in some engine it is oxidizer rich) so that the gas temperatures are low enough (typically 900 to 1350 K) to allow the use of uncooled turbine blades and uncooled nozzle exit segments. The RS-68 rocket engine, shown in Fig. 6-10, has a simple gas generator cycle. This engine is the largest liquid hydrogen/liquid oxygen rocket engine built to date. As can be seen from the data in the figure, with a gas generator cycle the specific impulse of the thrust chamber by itself is always a little higher than that of the engine and the thrust of the thrust chamber is always slightly lower than that of the engine.

In the expander cycle most of the engine coolant (usually hydrogen fuel) is fed to low-pressure-ratio turbines after having passed through the cooling jacket where it picked up energy. Part of the coolant, perhaps 5 to 15%, bypasses the turbine (not shown in Fig. 6-9) and rejoins the turbine exhaust flow before the entire coolant flow is injected into the engine combustion chamber where it mixes and burns with the oxidizer (see Refs. 6-2 and 6-14). The primary advantages of the expander cycle are good specific impulse, engine simplicity, and relatively low engine mass. In the expander cycle all the propellants are fully burned in the engine combustion chamber and expanded efficiently in the engine exhaust nozzle.

This cycle is used in the RL10 hydrogen/oxygen rocket engine, and different versions of this engine have flown successfully in the upper stages of several space launch vehicles. Data on the RL10-A3-3A are given in Table 10-3. A recent modification of this engine, the RL10B-2 with an extendible nozzle skirt, can be seen in Fig. 8-19 and data on this engine are contained in Table 8-1. It delivers the highest specific impulse of any chemical rocket engine to date. The RL10B-2 flow diagram in Fig. 6-11 shows its expander cycle. Heat absorbed by the thrust chamber cooling jacket gasifies and raises the gas temperature of the hydrogen so that it can be used to drive the turbine, which in turn drives a single-stage liquid oxygen pump (through a gear case) and a two-stage liquid hydrogen pump. The cooling down of the hardware to cryogenic temperatures is accomplished by flowing (prior to engine start) cold propellant through cooldown valves. The pipes for discharging the cooling propellants overboard are not shown here, but can be seen in Fig. 8-19. Thrust is regulated by controlling the flow of hydrogen gas to the turbine, using a bypass to maintain constant chamber pressure. Helium is used as a means of power boost by actuating several of the larger valves through solenoid-operated pilot valves.

In the staged combustion cycle, the coolant flow path through the cooling jacket is the same as that of the expander cycle. Here a high-pressure pre-combustor (gas generator) burns all the fuel with part of the oxidizer to provide high-energy gas to the turbines. The total turbine exhaust gas flow is injected into the main combustion chamber where it burns with the remaining oxidizer. This cycle lends itself to high-chamber-pressure operation, which allows a small thrust chamber size. The extra pressure drop in the precom-bustor and turbines causes the pump discharge pressures of both the fuel and the oxidizer to be higher than with open cycles, requiring heavier and more

Four leg structural support

Fuel iurbopump

High pressure oxygen line

Oxygen turbopump

Oxygen tank pressurization heat exchanger

Turbine-»

exhaust nozzle

Main fuel valve

Roll control nozzle using turbine exhaust

Gimbal actuator

Oxidizer valve

Regenerative cooled thrust chamber with nozzle throat

Ablative lined bell-shaped nozzle extension

Roll control nozzle using turbine exhaust

Gimbal actuator

Oxidizer valve

Regenerative cooled thrust chamber with nozzle throat

Ablative lined bell-shaped nozzle extension

High pressure oxygen line

Oxygen turbopump

Oxygen tank pressurization heat exchanger

Turbine-»

exhaust nozzle

Main fuel valve

Four leg structural support

Fuel iurbopump

Parameter

Thrust chamber

Engine

Specific impulse at sea level (max.), sec

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Responses

  • geronimo
    What is the use of tank pressurization in turbo pump feed system?
    8 months ago

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