A study in contrasts. During a test firing, the liquid propellant inlet section of this experimental turbopump is coated with ice, while the gas inlet from the combustion chamber glows red hot. Such as the working demands on turbopump technology.

The need for turbopumps is directly related to mission velocity and payload requirements.

Engine Requirements

Liquid rocket engines are either pressure-fed or pump-fed, depending on the mission requirements. If the mission velocity and payload are low, the propellants are fed to the thrust chamber by pressurizing the vehicle tanks. As the mission velocity requirement increases, the chamber pressure must be increased to raise the thrust from each pound of propellant in order to increase the vehicle thrust-to-weight ratio. This justifies the added complexity of the turbopumps to minimize the vehicle tank weight.

The type of engine cycle selected also influences the turbopump requirements and configuration. Generally, three types of engine cycles have been used in liquid rocket engines: the gas generator cycle, the staged combustion cycle and the expander cycle. The engine cycle terminology refers to the source of energy to drive the turbine.

Gas Generator Cycle

In a gas generator cycle engine, the turbine flow is in parallel with the thrust chamber and is not used to develop thrust. Sufficient propellants to drive the turbine are removed from the pump discharge, combusted in the gas generator, and expanded through the turbine to atmospheric pressure. The required pump discharge pressure is established by the combustion chamber injection pressure and establishes the available pressure to drive the turbine. Hence, the required pump flow rate is equal to the combustion chamber flow plus the flow required to drive the turbine. With the turbine discharging to atmosphere, the available energy per pound of flow is large due to the large pressure ratio. Maximizing the turbopump efficiency and increasing the turbine operating temperature to the available material limits reduces the required turbine flow rate while increasing the engine specific impulse per second (thrust per pound of propellant used). Turbine temperature is generally selecte d based on a trade study of engine weight, turbine design complexity and specific impulse. Gas generator cycle engines, therefore, minimize the pump-required discharge pressures, maximize the pump-required flow rate and maximize the turbine operating temperature for a given combustion chamber pressure. Prior gas generator cycle engines built by Rocketdyne (now Rocketdyne Propulsion and Power, a part of The Boeing Company) included the Redstone, Thor, Jupiter, H-l, F-l, and the J-2. Current active gas generator cycle engines built by Rocketdyne include the Atlas, RS-27, and the XLR-132.

Staged Combustion Cycle

In the staged combustion cycle engine, the turbine flow is in series with the thrust chamber. Most of the fuel flow and enough of the oxidizer flow to accomplish the desired turbine temperature are removed from the pump discharge, combusted in a preburner, and expanded through the turbine to the combustion chamber injection pressure. The remainder of the oxidizer flow is added in the main combustion chamber to complete the combustion process. A staged combustion cycle engine maximizes the engine’s specific impulse by passing the turbine flow through the thrust chamber to develop thrust. This does, however, raise the turbine discharge pressure to a value in excess of the main combustion chamber injection pressure. The turbine power required to drive the pumps determines the required turbine inlet pressure and the required pump discharge pressures; pump flow rates are equal to the thrust chamber flow rates. Maximum turbopump efficiencies and maximum turbine temperatures minimiz e the required turbine inlet pressure and pump discharge pressures and, therefore, minimize the turbopump weight. A tradeoff between turbine temperature and pump discharge pressure is normally made to establish the turbopump requirements. The staged combustion engine, therefore, maximizes the pump discharge pressures, minimizes the pump flow rates and maximizes the turbine operating temperature. The Space Shuttle Main Engine (SSME) is a staged combustion cycle engine.

Three common engine types. Red denotes fuel flow, while green denotes oxidizer flow.

Expander Cycle

In the expander cycle engine, the turbine flow is also in series with the thrust chamber. However, the turbine available energy is limited only to the fuel flow which is preheated in the thrust chamber coolant passages instead of being combusted with liquid oxygen in the preburner. The expander cycle engine also maximizes the engine’s specific impulse for a given chamber pressure by passing all the propellants through the thrust chamber. The pump-required discharge pressure is equal to the thrust chamber injection pressure plus the turbine pressure drop. Energy available to drive the turbine limits the expander cycle engine to relatively low thrust chamber combustion pressures, moderate pump discharge pressures, and low turbine operating temperatures. The cryogenic Orbital Transfer Vehicle engine is an expander cycle engine.

Other Engine Factors

Other engine factors that significantly influence the turbopump configuration selection are the types of propellants, the propellant inlet conditions and the engine throttling requirements. Typical propellants include RP-1, LH2, LO2, MMH, NTO, and other liquids with wide density ranges and temperatures. The variations in density produce significantly different pump head rise (pressure) requirements and large differences in volumetric flow, i.e., low density propellants require a much higher head rise to develop the same discharge pressure (head rise=pressure rise/density, DH=DP/p). The variations in the combined propellant available energy have a significant influence on the turbine design.

The propellant inlet condition, which is expressed as the pump-inlet net positive suction pressure (NPSP=propellant inlet total pressure-propellant vapor pressure), dictates the pump’s suction performance requirements, The pump suction performance requirement is its ability to operate at the available NPSP without detrimental cavitation.

The engine throttling requirements define the range of flow and discharge pressure that the turbopump must deliver with stable operation. The engine start and shutdown characteristics must also be considered to prevent unstable turbopump operation due to cavitation or stall.

The SSME HPF turbopump

The Mark 49F, used in the OTV

The Mark 15F, used with the J-2.

Configuration Selection

With the engine requirements established, the turbopump configuration is selected based on optimizing the pumps for each propellant, the turbine for the drive gas available energy, and the mechanical design arrangement for life, weight and producibility considerations.

Pumps for engines with similar density fuel and oxidizer propellants such as RP-1/LOX and similar discharge pressure requirements will typically be optimum at approximately the same speed. This permits the fuel and oxidizer pumps to be placed on a common shaft and driven by a common turbine (Redstone, Atlas, RS-27, F-1, and XLR-132). Maximum pump speed is generally limited by the suction performance requirements to avoid cavitation. Optimum turbine efficiency requires a certain pitchline velocity which is a product of the shaft speed and the turbine diameter. The minimum weight turbine has the highest speed and smallest diameter within the structural and mechanical arrangement limitations.

When the Atlas booster and sustainer turbopumps were designed, the speed of the pumps and the turbine were optimized independently and linked together with a speed reduction gear box, This required the development of a highly loaded gear train to minimize the turbopump weight, but was considered the best design selection based on suction performance, turbine performance and material technology at that time.

When the F-1 turbopump was designed, canted inducer technology had been developed to increase the pump suction performance capability. This permitted designing the pumps and turbine to operate at the same speed on a common shaft and eliminated the need for a 60,000-hp reduction gearbox, which was probably not feasible anyhow.

The J-2 was the first gas generator cycle engine to use liquid hydrogen (LH2) as the fuel and liquid oxygen (O2) as the oxidizer. The low-density liquid hydrogen introduced the need to operate the fuel pump at a much higher speed than the LO2 pump in order to develop the high head required. High solidity inducer technology had been developed which permitted optimizing the LH2 pump at a higher speed and driving the pumps with separate turbines. The turbines were arranged in series to best utilize the large pressure ratio available energy and maximize the turbine efficiencies at their respective speeds.

Selecting a 3,000-psi chamber pressure and staged combustion cycle for the SSME to maximize the specific impulse significantly increased the turbopump requirements compared to the F-1 or J-2 engines. Adding preburner and turbine pressure drops in series with the high combustion chamber pressure resulted in discharge pressure requirements of 8,500 psia and 7,000 psia for the LO2 and LH2 pumps, respectively.

Propellant tank pressures were also minimized to optimize the Space Shuttle vehicle weight. The combination of low inlet pressures (low NPSP) and high required discharge pressures introduced the need for separate boost pumps to optimize the turbomachinery weight. The low pressure fuel turbopump and low pressure oxidizer turbopump receive the propellants at low NPSP and raise their pressures sufficiently to optimize the high pressure fuel and oxidizer turbopumps at high speed. The added complexity of four turbopumps is justified to optimize the turbomachinery weight and maintain suction performance margin for safe engine operation. The combination of high pump discharge pressure and flow requirements, combined with high horsepower turbines driven by high-pressure hydrogen-rich steam, have made the SSME turbopumps a significant advancement in the state of the art for rocket engine turbomachinery.

With the inlet diameter selected, the shaft speed is selected to limit the inducer tip speed to approximately 550 ft/sec. The tip speed limit is for controlling the tip vortex cavitation energy, which is a function of tip speed to the sixth power. The blade thickness must also increase with increased tip speeds to react the centrifugal and pressure loading. This reduces the flow passage area and, therefore, lowers the suction performance. The pump suction specific speed is expressed as

Turbine efficiency is shown as a function of blade velocity and gas spouting velocity ratios

Suction performance improves over a 40 year period.

A convenient parameter which reflects the difference in pump geometry characteristics is the specific speed which is a function of the shaft speed, volumetric flow, and required headrise:

Pump Types. top to bottom: inducer, high specific speed;
axial, intermediate specific speed; centrifugal low specific speed.

Low specific speed pumps are typically centrifugal with head coefficients (y) ranging from 0.4 to 0.7, which is a function of the impeller blade discharge angle. Intermediate specific speed pumps are typically mixed or axial flow with head coefficients, y, ranging from 0.4 to 0.2 per stage; and high specific speed pumps require only an inducer to generate the required head.

The head requirements for high-density fluids such as RP-1 and LO2 can be generated with a single stage centrifugal pump, with the impeller diameter well within aluminum and nickel-base alloy steel structural limits. Head requirements for low-density fluids such as LH2 are very high and typically require several stages to develop. An axial flow main pumping element was selected for the J-2 LH2 pump because of its intermediate specific speed and narrow throttling range requirements. The 200,000-foot head requirement for the SSME HPFTP dictated a three-stage centrifugal pump with the impellers operating at 2,000 ft/sec tip speed. Titanium, which has a higher strength- to-weight ratio than the high-strength nickel-base alloys, was required for the high tip speed.

Optimizing the pump efficiency, which is a measure of the work-out/work-in, can also influence the shaft speed and specific speed selected. Maximum pump efficiency can generally be developed in the 2,000 to 3,000 specific speed range. Small flow rate pumps are generally less efficient than large flow rare pumps because the clearance and surface finish related losses cannot be scaled with size.

Turbines

The turbine must supply the required power to drive the pump utilizing the drive gas provided by the selected engine cycle. Overall performance of the turbine depends upon three variables: the available energy content per pound of drive gas, the blade tangential velocity (U), and the number of turbine stages. The available energy for the turbine pressure ratio can be expressed as an ideal velocity, C. The turbine velocity ratio, U / C, is used to empirically characterize these two variables versus the turbine efficiency.

The ideal velocity can be distributed between the turbine stages with either a pressure- compounded or velocity-compounded design. The major difference between these two turbine designs is where expansion occurs in the stationary blade rows. For the velocity-compounded turbine, all the expansion occurs in the first stationary blade row, while for a pressure compounded turbine, the expansion is distributed between the stationary blade rows. For high U/C designs, the turbine efficiency can be further improved by having some of the expansion (reaction) take place in the rotor blades. The design selection is made to maximize the turbine efficiency and minimize the weight compatible with the selected shaft speed. In general, when a direct drive turbopump configuration is selected, the shaft speed is less than optimum for the turbine and additional stages must be added to utilize the available energy.

Blade tip diameter is selected to optimize the U/C for efficiency within the blade height-to- diameter performance limits and within the tip speed structural limits. If the blade height- to-diameter ratio gets too small, the tip clearance and secondary flow losses become large, decreasing the turbine efficiency. The tip speed structural limit is based on the centrifugal pull that can be carried at the base of the blade airfoil for the selected material and is generally expressed as allowable annulus area x N2 versus temperature. Partial admission turbines are selected when the shaft speed is too slow and the blade height-to-diameter ratio becomes too small to obtain the desired U/C. The blade diameter is increased to increase U and the arc of admission is decreased to maintain the blade height at an acceptable height-to-diameter ratio.

Mechanical Design

Turbopump mechanical design is a compatible compromise involving many contributing factors, including power transmission, rotordynamics, axial thrust balance, bearings, dynamic seals and thermal considerations.

The shaft diameters, splines, and couplings must all be large enough to transfer the torque, which is a function of speed and horsepower. This establishes the shaft diameters and the bearing inner race minimum allowable diameter, depending on the axial stations selected along the shaft. The rotor critical speeds are a function of the rotor mass distribution and spring rate, the bearing locations and spring rate, and the housing stiffness. For rolling element bearings, the product of the inner diameter and the shaft speed (DN) is used as a measure of the bearing internal loading. Empirical life limits have been established for DN as a function of the propellant or lubricant used to cool the bearing.

With these interacting limitations, the bearing locations are selected to keep the operating speed range clear of critical speeds and minimize the bearing bore diameter to maximize bearing life.

From a critical speed standpoint, inboard bearings decrease the bearing span and increase the first critical speed so long as the overhang does not exceed approximately one-half the span length. However, this means that the power torque must pass through the bearing bore which maximizes the bearing diameter and DN when the bearing is placed between the pump and turbine.

In this series demonstrating the effects of cavitation note the gradual formation of small bubbles as inlet pressure (NPSH) is decreased. These eventually implode, causing damage
to the turbine metal. The final photograph indicates the severity of the effect.

With improvements in suction performance, higher operating speeds and smaller bearings, an active balance piston was introduced to react the rotor thrust. Orifices sensitive to the axial position regulate the pressure over an area of the rotor sufficient to balance the axial thrust. A separate disc was utilized in the J-2 LH2 turbopump for a double-acting balance piston. The back-to-back double-entry main impeller is utilized as the balance piston in the SSME HPOTP, while the third-stage impeller is utilized as the balance piston in the SSME HPFTP. Duplex pairs of preloaded angular contact ball bearings that float axially permit the balance piston to function in these turbopumps.

The development of hydrostatic bearings to center the rotor raises the DN limit associated with rolling contact bearings and increases the turbopump life potential. Hydrodynamic lift-off seals are also being developed to eliminate the life limits of face-riding seals. The combination of these two technologies removes some of the interactive constraints in the mechanical design selection process.

The rocket engine turbopump, in addition to being a high energy/weight ratio machine, must be designed to operate with the pump at cryogenic conditions and the turbine at high temperature. This requires design concepts that provide thermal growth flexibility while reacting large torques, separating loads and external ducting loads. Radial pins, axial cylinders and curvic couplings are typically used to perform this function.