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    This section represents a preparatory approach to the mainsubject. It aims at providing basic design data on which the designprocess will be based.

    1. Rocket turbopumps an overview

    Generally speaking, a turbopump is simply composed of a pump(sometimes two pumps) and a driving turbine, both mounted on the sameshaft or coupled via a reduction gearbox. It is designed in differentconfigurations and features according to the field of application. Common

    fields of applications are:

    Industrial projects. Turbojet engines. Ramjet motors. Rocket engines.

    Though of these various fields of applications, turbopumps have thereputation for mainly being used with rockets, replacing the place of the

    heavy pressurized tanks. So, the term " rocket turbopump" is common inthis field of literature. Using turbopumps in rocket engines greatlyimproves the engine performance as would be shown later. But, thedemand on attaining higher engine performance has imposed moresophistications on turbopump design.

    Rocket turbopumps have a reputation for being extremely hard to design to getoptimum performance. Whereas a well engineered pump can manage 70-90%efficiency, figures less than half that are common. Low efficiency may be acceptable insome applications, but inrocketry this is a severe problem. Turbopumps in rockets areimportant and problematic enough that launch vehicles using one have been described

    as a 'turbopump with a rocket attached'- up to 55% of the total cost has been ascribed tothis area

    Both centrifugal and axial pumps are used in rocket turbopumps butcentrifugal pumps have proven to be the best choice for rocketapplications due to their ability to generate higher discharge head in onestage. The high discharge head is a significant factor in selecting pumpsfor rocket engines. Nevertheless; axial pumps are also used in somerocket turbopumps as pumping units ( ), but they are commonly usedas 'inducers' for centrifugal pumps. The inducer plays an important role in

    http://en.wikipedia.org/wiki/Rocketryhttp://en.wikipedia.org/wiki/Rocketryhttp://en.wikipedia.org/wiki/Rocketry
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    rocket turbopumps where it raises the inlet pressure of the centrifugalpump enough to prevent excessive cavitation from occurring in thesuction side of the pumping lines.Cavitation is the main problem that limits pump speed whereas the pump

    speed is the independent variable to which the head rise is related by theform : head rise(H) shaft speed(N) squared. Therefore, cavitationavoidance is considered to be the pump key parameter in the design level.Other methods used to suppress cavitation in the suction side also exist,such as using independent boosting pumps ? before the main pump andslightly pressurizing the propellant tanks.1.2 Design arrangement

    The following arrangements are commonly used in rocket turbopumps?:1. Two pumps - direct drive with outboard turbine.The two pumps are mounted on the same shaft close to each other withturbine outboard. The turbine shaft goes through the fuel pump inlet,figure 1,1.a .

    2. Two pumps - direct drive with turbine in middleThe two pumps are mounted on the same shaft but withthe turbine placed in the middle. The common shaft goes through theturbine discharge manifold, figure 1,1.b .

    3. Gear driveThe two pumps are driven by a gearbox each with a separate shaft.The gearbox is driven by the turbine, figure 1,1.c .

    4. Two pumps two turbinesThe two pumps each is driven by a separate turbine. The driving gaseither flows in parallel or in series, figure 1,1.d .

    5. Four pumps four turbinesHere there are two main pumps and two booster pumps each with itsown driving turbine, figure 1,1.b .

    Common problems associated with centrifugal pumps include:

    1.excessive flow from the high pressure rim back to the low pressure inlet along thegap between the casing of the pump and the rotor

    2.excessive recirculation of the fluid at inlet

    3.excessive vortexing of the fluid as it leaves the casing of the pump

    http://en.wikipedia.org/wiki/Cavitationhttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://en.wikipedia.org/wiki/Vortexhttp://en.wikipedia.org/wiki/Cavitationhttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://var/www/apps/conversion/current/tmp/scratch19169/Figure%201,1.dochttp://en.wikipedia.org/wiki/Vortex
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    In addition, the precise shape of the rotor itself is critical.

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    1.3 Rocket engine features and performance parameters affecting thepumping system design:

    P

    T

    P T PT

    P

    TP

    T

    http://var/www/apps/conversion/current/tmp/Phd%20%202/jet%20engines/Rockets/Engines/Rocket%20engine.dochttp://var/www/apps/conversion/current/tmp/Phd%20%202/jet%20engines/Rockets/Engines/Rocket%20engine.doc
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    Even the design configuration of a rocket turbopump is highly dependenton the following engine requirements and factors which characterize therocket engine:

    The engine requirements for flow and pressure.

    The engine cycle or power cycle. The engine throttling requirements. The types of propellants used. The propellant inlet conditions.

    Having these engine requirements and features been well decided , theturbopump configuration can be selected based on optimizing the pumpsfor each propellant, the turbine for the drive gas available energy, and themechanical design arrangement for reliability, weight and producibility

    considerations.Therefore, each of these factors would be discussed in the subsequentsections and the suitable selection is to be made either based onassumptions or on results of calculations. This will lead to the preliminaryengine design suggested in the proposal.

    1.3.1 The engine requirements for flow and pressure

    On the basics of mass conservation, the mass flow rate of exhaust gasesof the engine is the same as the propellants mass flow rate delivered bythe pumps to the engine combustion chamber. Therefore, the enginethrust as one of the engine requirements is a function of the mass flowrate of propellants.?Also, for a specified engine and pair of propellants, the pressure of thecombustion chamber can be increased by increasing the mass flow rate ofthe propellants? which leads to a better engine performance. ?Putting in mind that there will be considerable friction losses in the flowlines of the propellants expressed as pressure drops, and that there is the

    injection requirement for pressure difference, the pumps should deliverpropellants at a discharge pressure considerably greater than the pressureof the combustion chamber.

    Consider equations (1.1), (1.2) and (1.3) bellow:

    Below is an approximate equation for calculating the net thrust of a rocket engine

    1.1

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    Where:

    exhaust gas mass flow

    effective exhaust velocityactual jet velocity at nozzle exit plane

    flow area at nozzle exit plane (or the plane where the jet leaves the nozzleif separated flow)

    static pressure at nozzle exit planeambient (or atmospheric) pressure

    impulse thrust is equal to the product of the propellant mass flow rate and theexhaust gas ejection speed. The ideal exhaust velocity is given by

    where k is the specific heat ratio, R' is the universal gas constant (8,314.51 N-m/kg mol-K in SI units, or 49,720 ft-lb/slug mol-oR in U.S. units), Tc is thecombustion temperature, M is the average molecular weight of the exhaustgases, Pc is the combustion chamber pressure, and

    1.3.2 The engine cycle

    The engine cycle terminology refers to the source of energy to drivethe turbine. There are three types of engine cycles have been used inliquid rocket engines:

    The gas generator cycle. The staged combustion cycle. The expander cycle.

    In the gas-generator cycle (figure 1.1), a small percentage of thepropellants (typically 3 to 7 percent) is bled off the main flow and sent to

    a separate small combustion chamber where the gas is generated andthen expanded through the turbine of the turbopump.The hot gas is theneither dumped overboard or sent into the main nozzle downstream. Thegas generator must burn propellants at off-optimal mixture ratio to keepthe temperature low for the turbine blades. Of course this leads to a smallreduction in performance Thus, this cycle is appropriate for moderatepower requirements but not high-power systems, which would have todivert a large portion of the main flow to the less efficient gas-generatorflow.

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    As in most rocket engines, some of the propellant in a gas generator cycle is usedto cool the nozzle and combustion chamber, increasing efficiency and allowing

    higher engine temperature.

    Figure 1.1 The gas generator cycle

    In a staged combustion cycle ( Figure 1.2 ), the propellants are burned instages. Like the gas-generator cycle, this cycle also has a smallcombustion chamber called the preburner, which also generates gas forthe turbopump turbine. The preburner as a first stage combustor, receivesand burns a small amount of the oxidizer which is bled from the mainflow, and all the amount of the fuel, thus producing a fuel-rich hot gas

    mixture that is mostly unburned vaporized propellant. This hot gas is thenpassed through the turbine, injected into the main chamber as a secondstage combustor to burn again with the remaining oxidizer. Theadvantage of this cycle over the gas-generator cycle is that all of thepropellants are burned at the optimal mixture ratio in the main chamberand no flow is dumped overboard which leads to higher engineperformance. It suits high power applications.

    The staged combustion cycle is often used for high-power applications. The higherthe chamber pressure, the smaller and lighter the engine can be to produce the

    same thrust. Development cost for this cycle is higher because the high pressurescomplicate the development process. Further disadvantages are harsh turbine

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    conditions, high temperature piping required to carry hot gases, and a verycomplicated feedback and control design.

    Staged combustion was invented by Soviet engineers and first appeared in 1960.In the West, the first laboratory staged combustion test engine was built inGermany in 1963.

    .

    Figure1.2 Staged combustion cycle

    In the expander cycle(Figure 1.3), the small combustion chamberfound in the other two cycles is completely eliminated. It makes use ofthe heat extracted from the engine to power the turbopump turbine. Itis similar to the staged combustion cycle in the fact that no portion ofthe combusted propellants is dumped over board and all of thepropellants are burned at the optimal mixture ratio; but anothercategory of this cycle is found which uses only a portion of the fuel todrive the turbine where the turbine exhaust is dumped overboard toincrease the turbine pressure ratio and power output.

    It is worth mentioning that this cycle works with fuels having a lowboiling point and can be vaporized easily such as methane andhydrogen, which have a low boiling point and can be vaporizedeasily. Also, the power available to the turbine is limited by the heattransfer to the fuel which makes this cycle appropriate for small to

    midsize engines.

    http://en.wikipedia.org/wiki/Expander_cyclehttp://en.wikipedia.org/wiki/Expander_cycle
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    Figure 1.3 The expander cycle

    The type of engine cycle selected also influences the turbopump requirements andconfiguration..

    1.3.3 The engine throttling requirements.

    Rockets can be throttled by controlling the propellant rate .Therefore,the engine throttling requirements define the range of flow and dischargepressure that the turbopump must deliver with stable operation.

    1.3.4 The types of propellants used

    Although both the propellants used in any liquid rocket engine arephysically in the liquid phase, still the pump and turbine selection isaffected by the type of the propellant used. That is due to the fact thatpropellants are found in wide density ranges and thermal properties.The variations in density lead to different pump head rise requirementsand large differences in volumetric flow. For example, lower densitypropellants require a much higher head rise to develop the same dischargepressure (pressure rise = head rise density), which implies higher tip

    speed of the pump impeller (tip speed (gravity acceleration headrise))

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    Also, the differences in the propellants available energy have a significantinfluence on the turbine design. Table 1 shows common propellantcombinations used forliquid propellant rockets and their propertiesaffecting the turbopump design.

    Pair ofpropellants

    Density(g/cm3)or specificgravity

    Vapor pressure Boilingpoint(K)

    Heat ofcombustion

    Red fumingnitricacid(RFNA)&

    kerosene orRP-1

    1.549(273K)

    0.78-0.81

    0.807(290K)

    0.0027

    RFNA&

    UDMH 0.865(228K)Dinitrogentetroxide&

    UDMH,MMH and/orHydrazine

    1.447(293K)

    Liquidoxygen&

    kerosene

    1.23(77.6K) 0.0052(88.78K)

    Hydrogen

    Peroxide&Alcohol orRP-1

    0.424(111.5K) 0.101(117K) 111.6

    Chlorinepentafluoride&Hydrazine

    1.005(293K) 0.0014(293K)

    http://en.wikipedia.org/wiki/Bipropellant_rockethttp://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/MMHhttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Alcoholhttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/Bipropellant_rockethttp://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/MMHhttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Alcoholhttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Hydrazine
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    .

    RFNA andkerosene orRP-1 RFNA andUDMH Dinitrogen tetroxide andUDMH,MMH and/orHydrazine Liquid oxygen and kerosene orRP-1 Liquid oxygen and liquid hydrogen Hydrogen Peroxide and alcoholorRP-1 Chlorine pentafluoride & Hydrazine

    1.3.5 The propellant inlet conditions

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

    Pumps

    Pump configuration is based on the requirements derived from theengine system. Inlet conditions (NPSP), discharge pressure, flow rate,and operating range must all be satisfied. A parametric analysis is

    performed to select the best speed, diameter and number of stages compatible with the turbineand mechanical design considerations.

    The pump inlet diameter is generally selected based on the available NPSP. Test experiencehas been accumulated on inducers to correlate their suction performance as a function of theNPSP (generally expressed as NPSH), the fluid inlet meridional velocity (Cm), and the inducerflow coefficient (f).

    The inducer diameter (inlet area) is selected to limit the fluid meridional velocity (Cm) so thatthe available NPSH/Cm2 /2g is equal to or greater than 3 velocity heads for water, 2 for LO2and 1 for LH2 . Variation in the empirical limit accounts for the difference in thermodynamicsuppression head between water, LO2 and LH2. As the available inlet pressure and NPSH aredecreased, the inducer diameter must be increased in order to decrease the fluid velocity (cm)and maintain NPSH/cm2/2g equal to the velocity head limit. The limit is also a function of theinducer flow coefficient, which is defined as the meridional velocity divided by the inducer tipspeed:

    http://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/MMHhttp://en.wikipedia.org/wiki/MMHhttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Hydrogenhttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Alcoholhttp://en.wikipedia.org/wiki/Alcoholhttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/RFNAhttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/Dinitrogen_tetroxidehttp://en.wikipedia.org/wiki/UDMHhttp://en.wikipedia.org/wiki/MMHhttp://en.wikipedia.org/wiki/Hydrazinehttp://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Kerosenehttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Liquid_oxygenhttp://en.wikipedia.org/wiki/Hydrogenhttp://en.wikipedia.org/wiki/Hydrogen_Peroxidehttp://en.wikipedia.org/wiki/Alcoholhttp://en.wikipedia.org/wiki/RP-1http://en.wikipedia.org/wiki/Chlorine_pentafluoridehttp://en.wikipedia.org/wiki/Hydrazine
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    With the inlet diameterselected, the shaft speedis selected to limit theinducer tip speed to

    approximately 550 ft/sec.The tip speed limit is forcontrolling the tip vortexcavitation energy, which isa function of tip speed tothe sixth power. The bladethickness must alsoincrease with increased tipspeeds to react thecentrifugal and pressureloading. This reduces theflow passage area and,therefore, lowers thesuction performance. Thepump suction specificspeed is expressed as:

    This is a measure of the pump's ability to operate at low inlet head (NPSH) without cavitation(formation of vapor bubbles) sufficient to cause head loss. A 50% NPSH margin is generallyselected during the design process for long-life rocket engine applications. Cavitation, inaddition todecreasing thepump dischargepressure andefficiency due to the

    formation of vaporbubbles, can causesignificant structuraldamage when thevapor bubblescollapse (implode),particularly withhigh-density fluids.Pratt & WhitneyRocketdyne'sinducer technologydevelopment hasbeen a key state-of-

    the art advancementfor increasing thepump speed,decreasing theturbopump weightand increasing the safe operating life. The double entry back-to-back pump was selected for theHPOTP in the SSME in order to increase the shaft speed by 2 and stay within the tip speedlimit, while maintaining the required total inlet flow area.

    Required pump head, which is a function of the required discharge pressure, the available inletpressure, and the propellant density [DH=(Pd - P in.) / r], is the major factor in selecting thepump configuration. The head coefficient (y=DH / U2 /g) is a function of the pump type andestablishes the required pumping element diameter and number of stages to develop the

    required pump head for a given shaft speed. The main pumping element may be a centrifugal,mixed, or axial flow type.

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

    Suction performance improves over a 40 year period.

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

    Low specific speed pumps are typically centrifugal with head coefficients (y) ranging from 0.4 to0.7, which is a function of the impeller blade discharge angle. Intermediate specific speedpumps are typically mixed or axial flow with head coefficients, y, ranging from 0.4 to 0.2 perstage; 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 asingle stage centrifugal pump, with the impeller diameter well within aluminum and nickel-basealloy steel structural limits. Head requirements for low-density fluids such as LH2 are very highand typically require several stages to develop. An axial flow main pumping element wasselected for the J-2 LH2 pump because of its intermediate specific speed and narrow throttlingrange 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, whichhas 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 influencethe shaft speed and specific speed selected. Maximum pump efficiency can generally bedeveloped in the 2,000 to 3,000 specific speed range. Small flow rate pumps are generally lessefficient than large flow rare pumps because the clearance and surface finish related lossescannot be scaled with size.

    1. Engine design

    As stated in the proposal, a preliminary engine design is to be carried firstso that the subsequent steps of the project are provided by the necessarydata such as the discharge head of each pump, propellants flowrate, .. .. .ectRegarding the engine, the design parameters which directly affect thepumping system design can be classified into tow types:

    (1) Parameters which are assumed by the designer.(2)Parameters which are obtained as results of design calculations.

    Table (1) below explains some parameters and working situations of each

    type :

    Assumed parameters Calculated parameters

    Combustion chamber pressure Propellant flow rate as exhaustEngine nozzle exit pressure Mixture ratioPropellants Fuel flow ratePower cycle of the pumping system Oxidizer flow rate

    Table (1) some assumed and calculated parameters of the engine.

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    Each assumed parameter, working situation or whatever, will bediscussed in the following sections and an assumption will be stated at theend of each section.

    1.1 Combustion chamber pressure

    Chamber pressures can range from about 7 to 250 atmospheres.(8)

    From equation (1.22) we see that high chamber temperature and pressure, andlow exhaust gas molecular weight results in high ejection velocity, thus high

    thrust(8)

    Although the engine performance improves much by increasing the

    combustion pressure (8,10) it is believed that after certain levels ofpressure the improvement in the performance is insignificant. The trendof the curve of figure (1) refers to this fact.

    Rockets can usually be throttled down to an exit pressure of about one-third of ambientpressure (often limited flow separation in nozzles) and up to a maximum limitdetermined only by the mechanical strength of the engine

    In practice, the degree to which rockets can be throttled varies greatly, but most rocketscan be throttled by a factor of 2 without great difficulty; the typical limitation iscombustion stability, as for example, injectors need a minimum pressure to avoidtriggering damaging oscillations (chugging or combustion instabilities); but injectorscan often be optimised and tested for wider ranges

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    (Rocket engine features and performance parameters affecting thepumping system design:

    Turbopumps are centrifugal pumps which are spun by gas turbines and are used to raisethe propellant pressure above the pressure in the combustion chamber so that it can beinjected and burnt. Turbopumps are very commonly used with rockets, but ramjets andturbojets also have been known to use them. The drive gases for the turbopump isusually generated in separate chambers with off-stochiometric combustion and therelatively small mass flow is dumped either through a special nozzle, or at a point in themain nozzle; both cause a small reduction in performance. In some cases (notably theSpace Shuttle Main Engine) staged combustion is used, and the pump gas exhaust isreturned into the main chamber where the combustion is completed and essentially no

    loss of performance due to pumping losses then occurs.

    Ramjet turbopumps use ram air expanding through a turbine.

    14 Turbopump

    From Wikipedia, the free encyclopedia

    As the name suggests, a turbopump comprises basically two main components: arotodynamicpump and a driving turbine, both mounted on the same shaft.

    An axial turbopump designed and built for the M-1 rocket engine

    A turbopump can refer to either of two types ofpumps: centrifugal, where the pumping

    is done by throwing fluid outward at high speed; or axial, where alternating rotating andstatic blades progressively raise the pressure of a fluid.

    Axial flow pumps have small diameters, and are used for this reason in duct jet engines,but give relatively modest pressure increases, and multiple compression stages areneeded. Centrifugal pumps are far more powerful, but physically larger.

    Turbopumps operate in much the same way as turbo units for vehicles. Higher fuelpressures allow fuel to be supplied to higher-pressure combustion chambers for higherperformance engines.

    Contents

    http://var/www/apps/conversion/current/tmp/Phd%20%202/jet%20engines/Rockets/Engines/Rocket%20engine.dochttp://en.wikipedia.org/wiki/Space_Shuttle_Main_Enginehttp://en.wikipedia.org/wiki/Staged_combustionhttp://en.wikipedia.org/wiki/Pumphttp://en.wikipedia.org/wiki/Pumphttp://en.wikipedia.org/wiki/Turbinehttp://en.wikipedia.org/wiki/M-1_(rocket_engine)http://en.wikipedia.org/wiki/Pumphttp://en.wikipedia.org/wiki/Axial_flow_pumphttp://en.wikipedia.org/wiki/Centrifugal_pumphttp://en.wikipedia.org/wiki/Combustion_chamberhttp://en.wikipedia.org/wiki/File:M-1_rocket_turbopump.JPGhttp://en.wikipedia.org/wiki/File:M-1_rocket_turbopump.JPGhttp://var/www/apps/conversion/current/tmp/Phd%20%202/jet%20engines/Rockets/Engines/Rocket%20engine.dochttp://en.wikipedia.org/wiki/Space_Shuttle_Main_Enginehttp://en.wikipedia.org/wiki/Staged_combustionhttp://en.wikipedia.org/wiki/Pumphttp://en.wikipedia.org/wiki/Turbinehttp://en.wikipedia.org/wiki/M-1_(rocket_engine)http://en.wikipedia.org/wiki/Pumphttp://en.wikipedia.org/wiki/Axial_flow_pumphttp://en.wikipedia.org/wiki/Centrifugal_pumphttp://en.wikipedia.org/wiki/Combustion_chamber
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    [hide]

    1 Historyo 1.1 Early developmento 1.2 Development from 1947to 1949

    2 Centrifugal turbopumps 3 Axial turbopumps 4 Complexities of centrifugalturbopumps 5 Driving Turbopumps 6 See also 7 External links

    8 References

    [edit] History

    [edit] Early development

    Turbopumps were originally developed for fire fighting (pumping water at high ratesand pressures to put out fires). The initial breakthrough for turbopumps used in rocketmotors occurred under Dr. Walter Thiel, during the development of the V2 in Germany.

    Prior to Dr. Thiel's work, pressurized tanks had been used. The early rocket turbopumpswere slightly modified turbopumps originally intended for pumping water. Usingturbopumps in rockets was a breakthrough; the power of the rocket motors wasincreased by an order of magnitude, making the lifting of heavy loads practical.

    [edit] Development from 1947 to 1949

    The principal engineer for turbopump development at Aerojet was George Bosco.During the second half of 1947, Bosco and his group learned about the pump work ofothers and made preliminary design studies. Aerojet representatives visited Ohio StateUniversity where Florant was working onhydrogen pumps, and consulted Dietrich

    Singelmann, a German pump expert at Wright Field. [51] Bosco subsequently usedSingelmann's data in designing Aerojet's first hydrogen pump.

    By mid-1948, Aerojet had selected centrifugal pumps for both liquid hydrogen andliquid oxygen. They obtained some German radial-vane pumps from the Navy andtested them during the second half of the year.

    By the end of 1948, Aerojet had designed, built, and tested a liquid hydrogen pump (15cm diameter). Initially, it usedball bearings that were run clean and dry, because thelow temperature made conventional lubrication impractical. The pump was firstoperated at low speeds to allow its parts to cool down to operating temperature. Whentemperature gauges showed that liquid hydrogen had reached the pump, an attempt wasmade to accelerate from 5000 to 35 000 revolutions per minute. The pump failed and

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    examination of the pieces pointed to a failure of the bearing, as well as the impeller.After some testing, super-precision bearings, lubricated by oil that was atomized anddirected by a stream of gaseous nitrogen, were used. On the next run, the bearingsworked satisfactorily but the stresses were too great for the brazed impeller and it flewapart. A new one was made by milling from a solid block of aluminum. Time was

    running out, as the contract had less than six months to go. The next two runs with thenew pump were a great disappointment; the instruments showed no significant flow orpressure rise. The problem was traced to the exit diffuser of the pump, which was toosmall and insufficiently cooled during the cool-down cycle so that it limited the flow.This was corrected by adding vent holes in the pump housing; the vents were openedduring cool down and closed when the pump was cold. With this fix, two additionalruns were made in March 1949 and both were successful. Flow rate and pressure werefound to be in approximate agreement with theoretical predictions. The maximum

    pressure was 26 atmospheres and the flow was 0.25 kilogram per second.

    Today the Space Shuttle Main Engine's turbopumps spin at over 30,000 rpm, delivering

    150 lb of liquid hydrogen and 896 lb of liquid oxygen to the engine per second. [1]

    [edit] Centrifugal turbopumps

    In centrifugal turbopumps a rotating disk throws the fluid to the rim

    Most turbopumps are centrifugal - the fluid enters the pump near the axis and the rotoraccelerates the fluid to high speed. The fluid then passes through a diffuser which is a

    progressively enlarging pipe, which permits recovery of the dynamic pressure. Thediffuser turns the high kinetic1 energy into high pressures (hundreds ofbaris not

    uncommon), and if the outletbackpressureis not too high, high flow rates can beachieved.

    [edit] Axial turbopumps

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    Axial compressors

    Axial turbopumps also exist - in this case the axle has essentially propellers attached tothe shaft and the fluid is forced by these parallel with the main axis of the pump.Generally, axial pumps tend to give much lower pressures than centrifugal pumps, anda few bar is not uncommon. They are however still useful - axial pumps are commonlyused as 'inducers' for centrifugal pumps, which raise the inlet pressure of the centrifugal

    pump enough to prevent excessivecavitation from occurring therein.

    [edit] Complexities of centrifugal turbopumps

    Turbopumps have a reputation for being extremely hard to design to get optimumperformance. Whereas a well engineered and debugged pump can manage 70-90%efficiency, figures less than half that are not uncommon. Low efficiency may beacceptable in some applications, but in rocketry this is a severe problem. Turbopumpsin rockets are important and problematic enough that launch vehicles using one have

    been caustically described as a 'turbopump with a rocket attached'- up to 55% of the

    total cost has been ascribed to this area.Common problems include:

    1. excessive flow from the high pressure rim back to the low pressure inletalong the gap between the casing of the pump and the rotor2. excessive recirculation of the fluid at inlet3. excessive vortexing of the fluid as it leaves the casing of the pump

    In addition, the precise shape of the rotor itself is critical.

    [edit] Driving Turbopumps

    Steam turbine powered turbopumps do exist and are employed when there is a source ofsteam, e.g. theboilers ofsteam ships. Nowgas turbines are usually used whenelectricity or steam is not available and place or weight restrictions permit the use ofmore-efficient sources of mechanical energy.

    One of such cases are rocket engines which need to pump fuel and oxidizerinto theircombustion chamber. This is necessary for largeliquid rockets, since forcing the fluidsor gases to flow by simple pressurizing of the tanks is often not feasible: The high

    pressure needed for the required flow rates would need strong and heavy tanks.

    http://en.wikipedia.org/wiki/Cavitationhttp://en.wikipedia.org/wiki/Cavitationhttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=6http://en.wikipedia.org/wiki/Rocketryhttp://en.wikipedia.org/wiki/Vortexhttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=7http://en.wikipedia.org/wiki/Steam_turbinehttp://en.wikipedia.org/wiki/Boilerhttp://en.wikipedia.org/wiki/Boilerhttp://en.wikipedia.org/wiki/Steam_shiphttp://en.wikipedia.org/wiki/Gas_turbinehttp://en.wikipedia.org/wiki/Gas_turbinehttp://en.wikipedia.org/wiki/Spacecraft_propulsionhttp://en.wikipedia.org/wiki/Fuelhttp://en.wikipedia.org/wiki/Oxidizerhttp://en.wikipedia.org/wiki/Combustion_chamberhttp://en.wikipedia.org/wiki/Combustion_chamberhttp://en.wikipedia.org/wiki/Liquid_rockethttp://en.wikipedia.org/wiki/Liquid_rockethttp://en.wikipedia.org/wiki/File:Axial_compressor.gifhttp://en.wikipedia.org/wiki/File:Axial_compressor.gifhttp://en.wikipedia.org/wiki/Cavitationhttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=6http://en.wikipedia.org/wiki/Rocketryhttp://en.wikipedia.org/wiki/Vortexhttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=7http://en.wikipedia.org/wiki/Steam_turbinehttp://en.wikipedia.org/wiki/Boilerhttp://en.wikipedia.org/wiki/Steam_shiphttp://en.wikipedia.org/wiki/Gas_turbinehttp://en.wikipedia.org/wiki/Spacecraft_propulsionhttp://en.wikipedia.org/wiki/Fuelhttp://en.wikipedia.org/wiki/Oxidizerhttp://en.wikipedia.org/wiki/Combustion_chamberhttp://en.wikipedia.org/wiki/Liquid_rocket
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    Ramjet motors are also usually fitted with turbopumps, the turbine being driven eitherdirectly by external freestream ram air or internally by airflow diverted from combustor

    entry. In both cases the turbine exhaust stream is dumped oeditTurboexpander

    [edit] External links Book of Rocket Propulsion

    Turbopumps for Liquid Rocket Engines FromRocketdyne'sEngineering Journal of Power Technology

    [edit] References

    1. ^ Hill, P & Peterson, C.(1992) Mechanics and Thermodynamics of

    Propulsion. New York: Addison-Wesley ISBN 0-201-14659-2

    Retrieved fromhttp://en.wikipedia.org/wiki/TurbopumpCategoriesTurbinesPumpsVacuum

    13

    The function of the rocket engine turbopump is to receive the liquid propellants from the vehicletanks at low pressure and supply them to the combustion chamber at the required flow rate andinjection pressure. The energy to power the turbine itself is provided by the expansion of highpressure gases, which are usually mixtures of the propellants being pumped.

    The turbopump configuration is highly dependent on the engine cycle and the enginerequirements for flow and pressure. Various turbopump configurations will be

    discussed at a conceptual design level, utilizing the disciplines involved in the designselection process, the "state of the art" of rotating machinery when each design tookplace, and the engine cycle requirements that influence the turbopump design

    The type of engine cycle selected also influences the turbopump requirements andconfiguration. 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 cycleterminology refers to the source of energy to drive the turbine.

    ther Engine Factors

    Other engine factors that significantly influence the turbopump configuration selection are thetypes of propellants, the propellant inlet conditions and the engine throttling requirements.

    http://en.wikipedia.org/wiki/Ramjethttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=9http://books.google.com/books?id=LQbDOxg3XZcC&pg=PA383&lpg=PA383&dq=Rocket+turbine&source=web&ots=TwTkVGEPI_&sig=5RzKu_9lDqY6duEIos-apvUhMCw&hl=en&sa=X&oi=book_result&resnum=3&ct=result#PPA393,M1http://www.rocketdynetech.com/articles/turbopump.htmhttp://en.wikipedia.org/wiki/Rocketdynehttp://en.wikipedia.org/wiki/Rocketdynehttp://en.wikipedia.org/wiki/Rocketdynehttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=10http://en.wikipedia.org/wiki/Turbopump#cite_ref-0http://en.wikipedia.org/wiki/Special:BookSources/0201146592http://en.wikipedia.org/wiki/Special:BookSources/0201146592http://en.wikipedia.org/wiki/Ramjethttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=9http://books.google.com/books?id=LQbDOxg3XZcC&pg=PA383&lpg=PA383&dq=Rocket+turbine&source=web&ots=TwTkVGEPI_&sig=5RzKu_9lDqY6duEIos-apvUhMCw&hl=en&sa=X&oi=book_result&resnum=3&ct=result#PPA393,M1http://www.rocketdynetech.com/articles/turbopump.htmhttp://en.wikipedia.org/wiki/Rocketdynehttp://en.wikipedia.org/w/index.php?title=Turbopump&action=edit&section=10http://en.wikipedia.org/wiki/Turbopump#cite_ref-0http://en.wikipedia.org/wiki/Special:BookSources/0201146592
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    Typical propellants include RP-1, LH2, LO2, MMH, NTO, and other liquids with wide densityranges and temperatures. The variations in density produce significantly different pump headrise (pressure) requirements and large differences in volumetric flow, i.e., low densitypropellants require a much higher head rise to develop the same discharge pressure (headrise=pressure rise/density, DH=DP/p). The variations in the combined propellant availableenergy have a significant influence on the turbine design.

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

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

    Configuration Selection

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

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

    Pumps

    Pump configuration is based on the requirements derived from theengine system. Inlet conditions (NPSP), discharge pressure, flow rate,and operating range must all be satisfied. A parametric analysis is

    performed to select the best speed, diameter and number of stages compatible with theturbine and mechanical design considerations.

    The pump inlet diameter is generally selected based on the available NPSP. Test experiencehas been accumulated on inducers to correlate their suction performance as a function of theNPSP (generally expressed as NPSH), the fluid inlet meridional velocity (Cm), and the inducerflow coefficient (f).

    The inducer diameter (inlet area) is selected to limit the fluid meridional velocity (Cm) so thatthe available NPSH/Cm2 /2g is equal to or greater than 3 velocity heads for water, 2 for LO2and 1 for LH2 . Variation in the empirical limit accounts for the difference in thermodynamicsuppression head between water, LO2 and LH2. As the available inlet pressure and NPSHare decreased, the inducer diameter must be increased in order to decrease the fluid velocity(cm) and maintain NPSH/cm2/2g equal to the velocity head limit. The limit is also a function ofthe inducer flow coefficient, which is defined as the meridional velocity divided by the inducertip speed:

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    With the inlet diameterselected, the shaft speedis selected to limit theinducer tip speed to

    approximately 550 ft/sec.The tip speed limit is forcontrolling the tip vortexcavitation energy, whichis a function of tip speedto the sixth power. Theblade thickness mustalso increase withincreased tip speeds toreact the centrifugal andpressure loading. Thisreduces the flowpassage area and,therefore, lowers thesuction performance.The pump suctionspecific speed isexpressed as:

    This is a measure of the pump's ability to operate at low inlet head (NPSH) without cavitation(formation of vapor bubbles) sufficient to cause head loss. A 50% NPSH margin is generallyselected during the design process for long-life rocket engine applications. Cavitation, inaddition to decreasing the pump discharge pressure and efficiency due to the formation ofvapor bubbles, can cause significant structural damage when the vapor bubbles collapse(implode), particularly with high-density fluids. Pratt & Whitney Rocketdyne's inducer

    technology development has been a key state-of-the art advancement for increasing the pumpspeed, decreasing the turbopump weight and increasing the safe operating life. The doubleentry back-to-back pump was selected for the HPOTP in the SSME in order to increase theshaft speed by 2 and stay within the tip speed limit, while maintaining the required total inletflow area.

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

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    Required pumphead, which is afunction of therequired dischargepressure, the

    available inletpressure, and thepropellant density[DH=(Pd - P in.) /r], is the majorfactor in selectingthe pumpconfiguration. Thehead coefficient(y=DH / U2 /g) is a

    function of the pump type and establishes the required pumping element diameter andnumber of stages to develop the required pump head for a given shaft speed. The mainpumping element may be a centrifugal, mixed, or axial flow type.

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

    Low specific speed pumps are typically centrifugal with head coefficients (y) ranging from 0.4to 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 perstage; 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 asingle stage centrifugal pump, with the impeller diameter well within aluminum and nickel-basealloy steel structural limits. Head requirements for low-density fluids such as LH2 are very highand typically require several stages to develop. An axial flow main pumping element wasselected for the J-2 LH2 pump because of its intermediate specific speed and narrow throttlingrange requirements. The 200,000-foot head requirement for the SSME HPFTP dictated athree-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, wasrequired for the high tip speed.

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

    Turbines

    The turbine must supply the required power to drive the pump utilizing the drive gas provided bythe selected engine cycle. Overall performance of the turbine depends upon three variables: the

    Suction performance improves over a 40 year period.

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    available energy content per pound of drive gas, the blade tangential velocity (U), and thenumber of turbine stages. The available energy for the turbine pressure ratio can be expressedas an ideal velocity, C. The turbine velocity ratio, U / C, is used to empirically characterize thesetwo 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 turbinedesigns is where expansion occurs in the stationary blade rows. For the velocity-compoundedturbine, all the expansion occurs in the first stationary blade row, while for a pressurecompounded turbine, the expansion is distributed between the stationary blade rows. For highU/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 turbineefficiency and minimize the weight compatible with the selected shaft speed. In general, when adirect drive turbopump configuration is selected, the shaft speed is less than optimum for theturbine 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 pullthat can be carried at the base of the blade airfoil for the selected material and is generallyexpressed as allowable annulus area x N2 versus temperature. Partial admission turbines areselected when the shaft speed is too slow and the blade height-to-diameter ratio becomes toosmall to obtain the desired U/C. The blade diameter is increased to increase U and the arc ofadmission is decreased to maintain the blade height at an acceptable height-to-diameter ratio.

    Ox

    pump F

    uel

    pmup

    Turbine

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    1