I

Pressure

Velocity

■ Direction v. .. ' of motion Stationary row nßlHN^iir

Pressure

Velocity y3

Single-s two-row velocity compounded impulse turbine

Three-stage reaction turbine, -50% reaction

Single-s two-row velocity compounded impulse turbine

Isentropic velocity ratio ulc0

FIGURE 10-9. Top view diagram, pressure and velocity profiles, and efficiency curves for impulse and reaction type turbines. The velocity ratio is the pitch line velocity of the rotating blades u divided by the theoretical gas spouting velocity c0 derived from the enthalpy drop. Adapted with permission from Refs. 10-1 and 10-12.

Isentropic velocity ratio ulc0

FIGURE 10-9. Top view diagram, pressure and velocity profiles, and efficiency curves for impulse and reaction type turbines. The velocity ratio is the pitch line velocity of the rotating blades u divided by the theoretical gas spouting velocity c0 derived from the enthalpy drop. Adapted with permission from Refs. 10-1 and 10-12.

With some cycles the turbine exhaust gases pass through a De Laval nozzle at the exit of the exhaust pipe (see Fig. 1^1). The high turbine outlet pressure gives critical flow conditions at the venturi throat (particularly at high altitudes) and thereby assures a constant turbine outlet pressure and a constant turbine power which will not vary with altitude. Furthermore, it provides a small additional thrust to the engine.

Turbine Performance and Design Considerations. The power supplied by the turbine is given by a combined version of Eqs. 3-1 and 3-7:

The power delivered by the turbine PT is proportional to the turbine efficiency t}T, the flow through the turbine mT, and the available enthalpy drop per unit of flow Ah. The units in this equation have to be consistent (1 Btu = 778 ft-lbf= 1055 J). This enthalpy is a function of the specific heat cp, the nozzle inlet temperature Tu the pressure ratio across the turbine, and the ratio of the specific heats k of the turbine gases. For gas generator cycles the pressure drop between the turbine inlet and outlet is relatively high, but the turbine flow is small (typically 2 to 5% of full propellant flow). For staged combustion cycles this pressure drop is very much lower, but the turbine flow is much larger.

For very large liquid propellant engines with high chamber pressure the turbine power can reach over 250,000 hp, and for small engines this could be perhaps around 35 kW or 50 hp.

According to Eq. 6-12, the power delivered by the turbine PT has to be equal to the power required by the propellant pumps, the auxiliaries mounted on the turbopump (such as hydraulic pumps, electric generators, tachometers, etc.), and power losses in bearings, gears, seals, and wear rings. Usually these losses are small and can often be neglected. The effect of the turbine gas flow on the specific impulse of the rocket engine system is discussed in Sections 6.2 and 10.2. For gas generator engine cycles, the rocket designer is interested in obtaining a high turbine efficiency and a high turbine inlet temperature Tx in order to reduce the flow of turbine working fluid, and for gas generator cycles also to raise the overall effective specific impulse, and, therefore, reduce the propellant mass required for driving the turbine. Three-dimensional computer analyses of the gas flow behavior and turbine blade geometry have resulted in efficient blade designs.

Better turbine blade materials (such as single crystals which have been uni-directionally solidified) and specialty alloys can allow turbine inlet temperatures between 1400 K (or about 2050°F) and perhaps 1600 K (or 2420°F); these higher temperatures or higher gas enthalpies reduce the required turbine flow. Reliability and cost considerations have kept actual turbine inlet temperatures at conservative values, such as 1150 to 1250°F or about 900 to 950 K, using lower cost steel alloy as the material. The efficiency of turbines for rocket turbopumps is shown in Fig. 10-9. Maximum blade speeds with good design and strong high-temperature materials are typically 400 to 700 m/sec or about 1300 to 2300 ft/sec. Higher blade speeds generally allow an improvement in efficiency. For the efficiency to be high the turbine blade and nozzle profiles have to have smooth surfaces. Small clearances at the turbine blade tips are also needed, to minimize leakage.

The low efficiency in many rocket turbines is dictated by centrifugal pump design considerations, which limit the shaft speed for turbopumps in which the pump and turbine are mounted on a common shaft, as discussed in the next section. A low shaft speed together with minimum mass requirements, which prohibit a very large turbine wheel diameter, give a low blade speed, which in turn reduces the efficiency.

The advantage of increased turbine efficiency (less gas generator propellant requirement) can be realized only if the turbopump design allows high blade speeds. This can be achieved in rockets of medium and low thrust by gearing the turbine to the pumpshaft or by using pumps that permit high shaft speeds; in rockets of very high thrust the pumps have diameters and shaft speeds close to those of the turbines and can be mounted on the same shaft as the turbine. The power input to the turbine can be regulated by controlling the flow to the turbine inlet. This can be accomplished by throttling or by-passing some of the flow of the working fluid to the turbine and varying the turbine inlet pressure.

There is no warm-up time available in rocket turbines. The sudden admission of hot gas at full flow causes severe thermal shock and thermal distortion and increases the chances for rubbing between moving metal parts. The most severe stresses of a turbine blade often are thermal stresses; they come during the engine start when the leading edge is very hot but other parts of the blade are still cold. This and other loading conditions can be more severe in rocket turbines than in air-burning gas turbines.

For low-thrust engines the shaft speeds can become very high, such as over 100,000 rpm. Also, the turbine blade height becomes very short and friction losses can become prohibitive. In order to obtain a reasonable blade height we go to partial admission turbine designs. Here a portion of the turbine nozzles are effectively plugged or eliminated.

Gas Generators and Preburners

A gas generator is used as the source of hot gas (from combustion of propel-lants) for driving many of the turbines of turbopumps in a liquid rocket engine. Depending on the engine cycle, other sources of turbine drive gases are sometimes employed, as described in Section 6.3.

Gas generators can be classified as monopropellant, bipropellant, or solid propellant. Actually the basic design parameters for gas generators are similar to those for engine thrust chambers or solid rocket motors. The combustion temperature is usually kept below 1400 to 1600 K (or 2000 to 2400°F) by intentionally regulating or mixing the propellants in proportions substantially different from stoichiometric mixture, usually fuel rich. These lower gas temperatures allow uncooled chamber construction and prevent melting or limit the erosion or turbine blades. With monopropellants, such as hydrogen peroxide (H202) or hydrazine (N2H4), the flow is easily controlled and the gases are generated at predictable temperatures depending on the details of the catalyst and the gas generator design. In principle a gas generator looks like an uncooled rocket thrust chamber except that the nozzle is replaced by a pipe leading to the turbine nozzles.

Propellants supplied to the liquid propellant gas generators can come from separate pressurized tanks or can be tapped off from the engine propellant pumps. When starting pump-fed gas generators, the turbomachinery needs to be brought up to operating speed. This can be done by a solid propellant gas generator starter, an auxiliary pressurized propellant supply, or by letting the engine "bootstrap" itself into a start using the liquid column head existing in the vehicle tankage and feed system lines—usually called a "tank-head" start.

Gas generators have been used for other applications besides supplying power to rocket feed systems. They have a use wherever there is a need for a large amount of power for a relatively short time, because they are simpler and lighter than conventional short-duration power equipment. Typical applications are gas generators for driving torpedo turbines and gas generators for actuating airplane catapults.

In a staged combustion cycle all of one propellant and a small portion of the other propellant (either fuel-rich or oxidizer-rich mixture) are burned to create the turbine drive gases. This combustion device is called a preburner and it is usually uncooled. It has a much larger flow than the gas generators mentioned above, its turbines have a much smaller pressure drop, and the maximum pressure of the propellants is higher.

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