US3546880A - Compressors for gas turbine engines - Google Patents

Compressors for gas turbine engines Download PDF

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US3546880A
US3546880A US857264A US3546880DA US3546880A US 3546880 A US3546880 A US 3546880A US 857264 A US857264 A US 857264A US 3546880D A US3546880D A US 3546880DA US 3546880 A US3546880 A US 3546880A
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impeller
compressor
flow
axial
air
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US857264A
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Pierre G Schwaar
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Avco Corp
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Avco Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D21/00—Pump involving supersonic speed of pumped fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F02C3/145—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chamber being in the reverse flow-type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D19/00—Axial-flow pumps
    • F04D19/02—Multi-stage pumps
    • F04D19/024—Multi-stage pumps with contrarotating parts

Definitions

  • the disclosure illustrates a gas turbine engine having an improved compressor comprising a three-stage axial ow compressor and a counterrotating supersonic mixed dow compressor immediately adjacent the downstream end of the axial flow compressor.
  • the axial flow cornpressor and the mixed flow impeller are rotated at speeds that cause the air discharged from the axial flow compressor to enter the mixed flow impeller at a relative velocity greater than sonic.
  • Shock waves are then se't up in the inlet portion of the impeller which reduces the relative Velocity to a subsonic value and increases the pressurization of the air.
  • An alternate embodiment utilizes a stator vane assembly between the axial flow compressor and the mixed ow impeller.
  • the stator vane assembly is shaped to minimize the velocity variation from the hub to the tip of the mixed fiow impeller entrance.
  • the present invention relates to compressors and more specifically to compressors for use in gas turbine engines.
  • van axial flow compressor commonly used in large gas turbine engines. It is well known in the art that this compressor has a high degree of efficiency and is capable of producing a very high pressure ratio if enough compression stages are provided. However, the large number of stages increases the complexity and the cost of the axial compressor to a level that practically negates its use in a low cost, small gas turbine.
  • the centrifugal compressor has been extensively used.
  • high pressure ratios are achieved by rotating the impeller at such a speed that the air is discharged from the impeller at a supersonic velocity level. It is then necessary to provide a diffuser which reduces the velocity of the air to a subsonic value by a shock wave process and further reduces the velocity to the lower level used in the combustion portion of the engine.
  • a counterrotating bladed impeller having a generally axially directed inlet fiow path receives air from the axial flow compressor.
  • the rotational speed of the bladed impeller is sucient to produce a relative velocity between the discharge air and the impeller inlet greater than the speed of sound, thereby producing shock waves adjacent the inlet of the impeller to reduce the absolute velocity of air below sonic and pressurize it.
  • the impeller has an outlet flow path spaced radially outward from the inlet whereby air is substantially radially accelerated and diffused for further increase of its pressure. As a result, the flow area over which the air flow is supersonic is minimized.
  • a compressor of the above general configuration has a series of stator vanes positioned between the axial flow compressor and the bladed impeller.
  • the stator vanes are shaped to turn the flow discharged at the bladed impeller in such a way that the radial variation in relative velocity from the impeller to the tip of the blades on the impeller is minimized.
  • FIG. 1 is a longitudinal cross section of a gas turbine engine having a compressor unit embodying the present invention
  • FIG. 2 is a view taken on lines 2 2 of FIG. l and illustrating a feature of the compressor of the engine of FIG. l;
  • FIG. 3 is a fragmentary longitudinal view of a portion of a compressor, such as that shown in FIG. l, illustrating an alternate embodiment of the present invention
  • FIGS. 3, 4, 5, 6 and 7 are velocity triangles showing the Mach number and direction of the flow through the compressor of FIG. l;
  • FIG. 8 is a fragmentary view of an alternate embodiment of a compressor constructed in accordance with the present invention.
  • FIGS. 9, l0, 1l and 12 are velocity triangles showing the Mach number and direction of the fiow through the compressor of FIG. 8.
  • FIG. 1 shows a gas turbine engine 10 comprising an annular housing 12 having an inlet 14 for passage of air to a compressor unit 16.
  • the compressor unit 16 has rotating portions which pressurize air and discharge it through an annular ow path to a combustion unit, generally indicated by reference numeral 18.
  • the pressurized air is then mixed with fuel injected into the combustion unit 18 by a fuel nozzle 20 and the resultant mixture is ignited by well-known means to generate a hot gas stream which is discharged across a turbine unit 22, wherein a portion of the energy of the hot gas stream is extracted to drive the counterrotating portions of the compressor unit 16 through interconnecting shafts 24, 26.
  • the hot gas stream may be discharged through an exhaust nozzle to produce a propulsive thrust.
  • a power turbine unit 28 extracts a major portion of the energy available from the hot gas stream to drive a rotatable output shaft 30 which is used as the prime power output for the engine 10.
  • the engine 10 incorporates an improved compressor unit 16, as described in detail below, together with the cooperating engine components.
  • the compressor' unit 16 has an upstream axial flow portion which comprises a rotor 32 having rst and SeC- ond stages of circumferentially positioned compressor blades 34 and 36, respectively secured thereto by suitable retaining means (not shown).
  • the rotor 32 is telescoped onto a shaft and held against a shoulder 49 by a retaining nut 42.
  • the aft portion of the shaft 40 is integral with a rotor 46, having a third stage of circumferentially positioned compressor blades 48.
  • the aft portion of the shaft 40 also has a splined opening 41 which is telescoped over a splined portion of the interconnecting shaft 24.
  • a rst series of circumferentially positioned compressor stator vanes 38 extend from the housing 12 into an annular space between the adjacent stages of blades 34 and 36.
  • a second series of circumferentially positioned compressor vanes extend from the housing 12 radially inward between the second and third stages of compressor blades 36 and 48, respectively.
  • the compressor blades S0 ⁇ are fabricated to provide a relatively rigid supporting structure for an inner annular structural member 52.
  • the structural member 52 is secured to an annular bearing sump chamber 54.
  • a pair of bearings 56 and 58, used to journal the shaft 40 are mounted in the sump chamber 54.
  • the ends of the sump chamber S4 are sealed by suitable annular seal assemblies 60 which may be of the friction seal or laybrinth seal type.
  • the downstream portion of the compressor unit 16 is comprised of a rotatable impeller comprised of a hub 62 which forms the inner bounds of a generally annular flow path thereacross.
  • the outer bounds of the flow path is then defined by the adjacent portion of the engine housing 12.
  • a series of generally radially extending blades 64 are secured to the outer periphery of the hub 62 by suitable retaining means (not shown).
  • the hub 62 is telescoped over the interconnecting shaft 26 and urged against a shoulder 68 by a retaining nut 66.
  • the end of the shaft 26 adjacent the hub 62 is journaled by means of a bearing assembly 70 which is mounted in a generally annular sump chamber 72.
  • the sump chamber 72 is sealed at its forward end by a seal assembly 74.
  • the shaft 26 may be used for accessory drive purposes by securing a bevel pinion gear 76 over the shaft 26 to engage a bevel gear 78.
  • the gear 78 is secured to a shaft 80 which in most instances would be journaled in a generally annular support structure 82 extending from the sump chamber 72.
  • the annular support structure 82 is secured at its outer periphery to the inward edges of a plurality of circumferentially positioned diffuser vanes 84 which receive the air discharged from the impeller blades 64.
  • a generally annular entrance duct 86 is secured to the aft ends of the diffuser vanes 84 and provides a flow path for air into the combustion unit 18.
  • the combustion unit 18 comprises an annular chamber 88 which may be formed integrally with the engine housing 12, as herein illustrated.
  • a combustor 90 is positioned in the chamber 88.
  • the combustor 90 may take the form herein illustrated of a reverse-flow cannular combustor assembly.
  • the cannular burners briefly comprise a series of generally axially extending perforated cans positioned around the chamber 88. They are interconnected circumferentially by a tube and discharge the hot gases into a piece 92 which turns the ow to an aft direction and effects the transition from the circular can outlets 91 into the full turbine entrance annulus in front of nozzle 94.
  • the turbine nozzle 94 generally comprises a series of radially extending vanes 96 secured to the engine housing 12 and to the inner portion of the transition piece 92.
  • a turbine rotor 98 is positioned adjacent the turbine nozzle 94 and has a plurality of radially extending turbine blades 100 which form a first stage of the turbine unit 22.
  • the turbine rotor 98 is secured by suitable means to the shaft 26 at a ange 102, integral therewith.
  • the aft portion of the shaft 26 is journaled by a bearing assembly 104 secured in the annular sump chamber 72.
  • a suitable seal 4 assembly 106 seals the aft portion of the sump chamber 72.
  • a plurality of circumferentially positioned turbine vanes 108 are secured to the engine housing 12 downstream of the turbine blades 100.
  • the inner ends of the vanes 108 connect with an annular duct member 110 having a suitable gas seal 112 between the duct member 110 and the flange 102. of the shaft 26.
  • a second turbine rotor 114 is positioned downstream of the turbine vanes 108 and has secured thereto a plurality of turbine blades 116.
  • the rotor 114 is integral with the shaft 24 and is journaled by means of a bearing assembly 118, supported by a bearing sump chamber 120.
  • the sump chamber is sealed at its forward end by a seal 122 and secured to a generally annular support member 124.
  • the support member 124 is connected at its outer edge to a series of structural vanes 128 extending to the engine housing 12.
  • the counterrotating axial and centrifugal compressor sections are directly driven by two mechanically independent turbine rotors 100 and 116, which thus rotate in opposite directions. It is clear, how-ever, that counterrotation of the two compressor sections also can be achieved by a gear driven by a single turbine rotor.
  • the axial compressor has a flow velocity at exit of the last axial rotor stage 48 that is subsonic relative to the fixed engine housing 12 (absolute velocity).
  • the counterrotating impeller 62 causes the ow velocity of air entering the impeller to be supersonic relative to the impeller (relative velocity).
  • Shock waves are generated in the entrance region of the impeller blading, which decelerate the flow to a subsonic velocity level relative to the impeller.
  • the blading 64 in this eutrance region of the impeller is characterized by an essentially constant, or slightly increasing, mean radius, and by a moderate camber angle, so that flow separation induced by shock-boundary layer interaction is minimized.
  • the flow is gradually turned radially outwards and toward the relative axial direction, thereby experiencing a further pressure increase by centrifugation and additional relative deceleration.
  • the type of impeller illustrated herein is generally termed a mixed flow impeller.
  • the shape of the impeller blades differs from the usual mixed flow configuration in that there is relatively little or no turning of the flow in the first portion of the blades, i.e., the portion where the supersonic flow is decelerated to a subsonic velocity level by the shock wave process, as shown in FIG. 2.
  • the essentially axial flow channel in the rst portion of the impeller is designed so that there is relatively little or no increase of the relative flow passage area through the rst portion of the blading, in order to stabilize the shock configuration and to ensure maximum efliciency of the compression by the shock process.
  • the impeller is preferably designed with a mixed axial-radial flow path, that is with an annulus channel forming an angle with the axial direction substantially smaller than 90 at the impeller discharge diameter.
  • FIGS. 4, 5, 6 and 7 illustrate exemplary flow conditions that exist in a typical compressor embodying the present invention. It should be noted, however, that the description of these flow ⁇ conditions is not intended to limit the scope of the present invention but to merely enable a clearer understanding of the concepts involved in its operation.
  • the velocities illustrated are for a compressor which has a transonic axial flow portion having three stages. While it is not necessary to utilize a transonic axial iiow compressor in the invention, the use of a transonic compressor enables a reduction in the number of stages necessary to achieve a given pressure ratio.
  • the impeller decelerates the supersonic ow to a subsonic Velocity level in its front portion by a shock process which, depending upon the entrance Mach number MW1, may take the form of a single-shock, a multishock, or a so-called pseudo-shock configuration.
  • the above axial compressor and mixed ow impeller provide an extremely high pressure ratio for a relatively simple configuration. It is to be noted that the supersonic flow that is necessary to achieve high pressure ratios in a radial type compressor has been shifted from the diffuser entrance region to the impeller entrance region. This greatly reduces the area over which the (dow is supersonic and, accordingly, reduces the friction losses associated with supersonic flow. It is also to be noted that because a substantial portion of the compression work is accomplished by the shock process in the inlet portion of the impeller, the radial distance of the discharge relative to the inlet can be minimized for a given impeller pressure ratio. Furthermore, the lower ow velocity at the discharge of the impeller enables the use of a more eflicient dituser.
  • FIG. 8 illustrates the last stage 48' of an axial iiow compressor.
  • a plurality of stator vanes 150 are mounted at their outer edge to the engine casing 12' and at their inner edges to an annular channel member 152, defining the inner bounds of the annular ow path across the stator vanes 150.
  • the channel shaped member 152 is adapter to provide a seal in cooperation with an extension 154 of the rotor 40 for the axial ilow compressor.
  • the vanes 150 are shaped at their radially inner end to impart no turning of the air discharged from the axial flow compressor towards an axial direction.
  • the leading edge of vane 150 is approximately in line with the Mach number Vector Mv2 of the air discharged from the last stage of the axial ilow compressor.
  • the trailing edge of vane 150 turns the flow slightly towards a tangential direction. This enables the relative angle of the flow at the entrance to bladed impeller 62' to be uniform over the radial extent of the annular iiow path.
  • the vanes 150 are cambered, however, so that their tip section imparts a substantial turning towards the axial direction, as shown in FIG. 10. It is illustrated in this gure that the leading edge of the outer end of vanes 150 is approximately in line with the Mach number vector MV2 of air discharged from the last stage of the axial flow compressor. The blade is turned, however, so that the downstream end of the blade points towards an axial direction.
  • the camber of vanes 150 ⁇ produces Mach numbers into the entrance of the impeller 62' as shown in FIG. 11. It is apparent from these figures that the tangential Mach number component of air entering the centrifugal compressor at the tip of blades 64 has been reduced by the camber of the outer end of vanes 150. This minimizes the increased relative Mach number of the air caused b-y the increased tangential velocity component of the blades 64 at their inlet tip. This results in relative entrance Mach numbers substantially equal.
  • the radially inner section of the vanes 150 is shaped to turn the ow towards the tangential direction in a sense opposite to that of the turning at the outer portion of the vanes 150.
  • the combustion unit 18 is of the reverse flow type, which enables a substantial reduction in axial length of the interconnecting shafts 24 and 26.
  • the reverse flow combustion unit 18 additionally enables a substantial reduction in axial length for the engine 10 which greatly facilitates its use in applications where axial length is at a premium.
  • FIG. 3 there is illustrated a modiiied impeller ⁇ 62 which is secured to a shaft 26', journaled by a bearing assembly 70.
  • this impeller there is shown a rst series of radially extending blades forming an inlet portion of the impeller and a second series of generally radially extending blades 142 forming an outlet portion of the impeller 62.
  • a series of circumferentially positioned vanes 144 are secured to an engine housing 12 and extend between the adjacent series of blades 140 and 142.
  • the blades 140 receive the relative supersonic tiow from the axial iiow compressor and causes the air ow to be shocked down, thereby increasing its pressure.
  • the vanes 144 receive the flow and turn it towards an axial direction so that the blades 142 may be shaped to produce a greater amount of work on the air and accordingly increase the pressure ratio to a greater extent for a given hub and flow path configuration.
  • an axial flow compressor rotating in a first direction for discharging air at an absolute velocity lower than the speed of sound
  • a counter-rotating bladed impeller having a generally axial directed inlet portion defining a substantially undeflected flow path for receiving the air discharged from said axial compressor, the rotational speed of said bladed impeller being sufficient to produ a relative velocity between said discharge air and the impeller inlet portion greater than the speed of sound, thereby producing shock waves adjacent the inlet portion of said bladed impeller to reduce the absolute velocity of air owing therethorugh below sonic and pressurize it;
  • said bladed impeller having an outlet portion defining a flow path radially outward from said inlet portion whereby air is substantially radially accelerated and diffused for further increase of its pressure.
  • said bladed impeller comprises a hub defining the inner bounds of an annular flow path thereacross and a series of radially extending blades secured thereto, said blades having a configuration so that they define, in combination with said hub, flow paths having a relatively constant flow area in said inlet portion wherein said shock waves are produced and a diverging flow area adjacent the outlet portion thereof for diffusion of air.
  • said axial ow compressor is a transorn'c compressor having a minimum number of stages for pressurizing 4.
  • Apparatus as in claim 2 wherein:
  • said impeller comprises a mixed-flow impeller having a generally axially directed inlet portion and generally outwardly directed outlet portion;
  • said axial compressor imparts a substantial tangential velocity component to said air for discharge to said mixed flow impeller
  • chord of said impeller blades in their inlet portion is angled towards a tangential direction for generally parallel relative fiow of air from said axial flow compressor into the inlet of said impeller;
  • chord of said impeller blades adjacent the inlet portion has a minimum curvature along the portion wherein said shock waves are produced
  • the chord of said impeller blades downstream of the inlet portion wherein said shock waves are produced is curved to a generally axial direction at the discharge end of said impeller.
  • the flow path through said impeller blades is turned radially outward in its outlet portion with an angle substantially less than 90 degrees relative to the axis of said impeller;
  • said mixed-flow impeller is rotated at a speed sufficiently great so that a substantial amount of the pressurization of the air passing through said impeller is produced by the shock waves at the inlet thereto whereby the radial distance between the inlet and the outlet portions of said impeller is minimized.
  • Apparatus as in claim 4 further comprising:
  • annular diffuser positioned to receive the discharge from said impeller and having a diverging flow area whereby the pressure of the air discharged from the mixed fiow impeller is further increased.
  • a combustor means positioned to receive pressurized air from said diffuser, said combustor means including means for mixing said pressurized air with fuel and igniting the mixture to produce a relatively hot gas stream;
  • annular inlet nozzle having a series of radial vanes positioned downstream of said combustor means to receive and accelerate said hot gas stream;
  • annular chamber having the inlet from said diffuser positioned in a radially outward portion thereof and an outlet portion positioned radially inward of and adjacent said inlet portion;
  • At least one burner unit positioned in said chamber for mixing pressurized air with fuel and sustaining cornbustion thereof, said burner unit being positioned to direct the hot gas stream produced by combustion in a forward direction relative to the iiow through said compressor unit and said turbine means;
  • a rotatable hub defining the inner bounds of an annular flow path across said impeller
  • said apparatus further comprises:
  • stator vanes secured to said housing and extending towards said hub between said vanes;
  • stator vanes being shaped to turn the air passing thereacross so that the outlet vanes produce a maximum pressurization of said air.
  • stator vanes interposed between the downstream end of the axial ow compressor and the inlet of the counterrotating blade impeller for receiving air from the axial flow compressor and directing it to the inlet of the blade impeller, said stator vanes being shaped to produce a minimum turning to the flow discharged from the axial ow compressor for the radially inward portion of the vanes and to produce a substantial turning of the flow discharged from the radially outward portion of the axial flow compressor over the radially outward portion of the stator vanes for minimizing radial variations of relative Mach numbers at the inlet to the blade impeller.
  • stator vanes are shaped at their radially outward portion to produce a substantial turning of the air ow from a tangential di- References Cited UNITED STATES PATENTS 2,689,681 9/1954 Sabatiuk 230--123 2,842,306 7/1958 Buchi 230-119X 3,037,349 6/1962 Gassmann 60-39.l6

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US857264A 1969-08-04 1969-08-04 Compressors for gas turbine engines Expired - Lifetime US3546880A (en)

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US3902314A (en) * 1973-11-29 1975-09-02 Avco Corp Gas turbine engine frame structure
US3918828A (en) * 1974-09-05 1975-11-11 Emerson L Kumm Flow control for compressors and pumps
US4147026A (en) * 1976-09-22 1979-04-03 Motoren-Und Turbinen-Union Munich Gmbh Gas turbine engine
US4592204A (en) * 1978-10-26 1986-06-03 Rice Ivan G Compression intercooled high cycle pressure ratio gas generator for combined cycles
US4611464A (en) * 1984-05-02 1986-09-16 United Technologies Corporation Rotor assembly for a gas turbine engine and method of disassembly
US4685286A (en) * 1984-05-02 1987-08-11 United Technologies Corporation Method of disassembly for a gas turbine engine
EP0402693A1 (de) * 1989-06-10 1990-12-19 Mtu Motoren- Und Turbinen-Union MàœNchen Gmbh Gasturbinentriebwerk mit Diagonal-Verdichter
US5074109A (en) * 1989-03-22 1991-12-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Low pressure turbine rotor suspension in a twin hub turbo-engine
US20110286836A1 (en) * 2010-05-24 2011-11-24 Davis Todd A Geared turbofan engine with integral gear and bearing supports
US20130224007A1 (en) * 2012-02-29 2013-08-29 Jose L. Rodriguez Mid-section of a can-annular gas turbine engine to introduce a radial velocity component into an air flow discharged from a compressor of the mid-section
US20130224009A1 (en) * 2012-02-29 2013-08-29 David A. Little Mid-section of a can-annular gas turbine engine with a radial air flow discharged from the compressor section
US20140003928A1 (en) * 2012-06-29 2014-01-02 Rolls-Royce Plc Spool for turbo machinery
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US20160138603A1 (en) * 2014-11-18 2016-05-19 Rolls-Royce North American Technologies, Inc. Split axial-centrifugal compressor
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US9982676B2 (en) 2014-11-18 2018-05-29 Rolls-Royce North American Technologies Inc. Split axial-centrifugal compressor
US10508546B2 (en) 2017-09-20 2019-12-17 General Electric Company Turbomachine with alternatingly spaced turbine rotor blades
US20200132306A1 (en) * 2018-10-25 2020-04-30 General Electric Company Combustor Assembly for a Turbo Machine
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WO2021170819A1 (fr) * 2020-02-28 2021-09-02 Safran Aero Boosters Sa Compresseur transsonique de turbomachine
US11118535B2 (en) 2019-03-05 2021-09-14 General Electric Company Reversing gear assembly for a turbo machine
US11143402B2 (en) * 2017-01-27 2021-10-12 General Electric Company Unitary flow path structure
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US11428160B2 (en) 2020-12-31 2022-08-30 General Electric Company Gas turbine engine with interdigitated turbine and gear assembly
CN114973902A (zh) * 2022-04-14 2022-08-30 西北工业大学 一种教学用航空发动机低压涡轮模型及装配方法
RU220991U1 (ru) * 2023-06-08 2023-10-12 Публичное акционерное общество "ОДК-Уфимское моторостроительное производственное объединение" Задняя опора ротора турбины низкого давления авиационного газотурбинного двигателя

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US2842306A (en) * 1952-04-30 1958-07-08 Alfred Buchi Turbine driven multi-stage blower or pump
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Cited By (44)

* Cited by examiner, † Cited by third party
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DE2032562A1 (de) 1971-02-18
SE356792B (xx) 1973-06-04
JPS5030854B1 (xx) 1975-10-04
GB1280113A (en) 1972-07-05

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