Sundstrand Corporation

International Harvester Company

Colin Rodgers

Jack R Shekleton

Roy W Vershure, Jr

Anthony C Jones

Robert G Thompson

Malcolm J McArthur

John P Archibald

Nipulkumar Shah

In a gas turbine engine having gaseous oxidant delivered to a turbine from a source such as a pressurized bottle via a conduit or duct, a compact annular combustor is provided which offers enhanced performance through incorporation of means for uniformly distributing portions of the oxidant utilized as a diluent, and for convective cooling of engine components.

Turbine including a stored energy combustor

Difficulties in starting a turbine engine, particularly ones having a small volume combustor (83) with a low dome height, may be avoided by a method including the steps of a) forming the turbine engine compressor (48) of a low density, high temperature resistant ceramic material, b) directing pyrotechnic gases at a temperature of at least about 3,000.degree. F. against the blades (48) of the compressor, c) continuing the step of directing until the turbine rotor (44) is rotating at at least about 60 percent of rated speed and d) combusting fuel in the combustor (83) of the turbine engine to maintain the rotor (44) rotating at at least about 60 percent of rated speed.

Starting of a small turbojet

Surge protection is achieved in a turbine engine intended to provide varying quantities of bleed air and having two compressor stages (48, 36) by providing the inlet (56) of the first stage (48) with variable inlet guide vanes (58, 60). The first stage compressor (48) is a high specific speed, single stage centrifugal compressor whose stable operating characteristics are highly sensitive to inlet guide vane geometry, thereby allowing variations in such geometry to be employed to prevent compressor surge.

Surge protected gas turbine engine for providing variable bleed air flow

Inefficiencies in the operation of auxiliary power units producing electric and hydraulic power along with bleed air resulting from independence in the demand for power may be minimized in a system including at least one bleed air consuming system (84, 86), an electrical generator (92), a hydraulic pump (90), and a turbine wheel (10, 20) for driving the generator (92) and the pump (90). Also included is a combustor (32) for receiving fuel and compressed air and producing gases of combustion to drive the turbine wheel (10, 20). A rotary compressor (10, 14) is coupled to the turbine wheel to be driven thereby. The compressor (10, 14) is a radial discharge compressor with an impeller having blades (14) extending from an inlet end (40) to a discharge end (42) and a shroud (18) to the blades (14). Two axially spaced annular ports (44, 46; 44, 50) are adjacent the discharge end (42). The one of the ports (44, 46) nearer the shroud ( 18) is connected to the bleed air consuming system (84, 86) with the other of the ports (44, 50) being connected to the combustor (32) to provide compressed air thereto. Variable inlet guide vanes (76) are adjacent the inlet ends (40) and extend only partially across the inlet (74) thereto so that varying the position of the vanes (76) will vary the volume of air discharged into only one of the ports (44, 46; 44, 50).

Auxiliary power unit for use in an aircraft

The diameter of a gas turbine engine may be reduced through the use of a construction including a monorotor (44) journaled for rotation and mounting compressor blades (48) at one end (46) and turbine blades (56) at an opposite end (54). An axial diffuser (96) receives the discharge from the compressor blades (84) and is contained in a plenum (98). A reverse flow, annular combustor (83) is disposed about the rotor (44) and has a dome (100) whose height is a minor fraction of the axial length of the combustor (83) and which discharges to an annular turbine nozzle (77) having axial turbine nozzle blades (76).

Small diameter gas turbine engine

In order to establish a larger dome height to thereby increase combustion volume within the same turbine engine envelope, the inner wall region of an annular combustor is formed to define a thrust nozzle which in turn provides an enhanced combustion zone. The turbine engine will include a rotary compressor and a turbine wheel coupled to the rotary compressor for driven movement thereof along with an annular nozzle proximate the turbine wheel for directing gases of combustion thereat. The annular combustor will be disposed about the turbine wheel and defined by a radially extending wall region connecting the inner wall region to an outer wall region such that the annular combustor has a combustion zone upstream of the annular nozzle and a dilution air zone intermediate the annular nozzle and the combustion zone. The turbine engine will operate such that the annular combustor combusts fuel from a fuel source and air from the rotary compressor to generate the gases of combustion which are directed at the turbine wheel by the annular nozzle for driven movement of the turbine wheel and the rotary compressor. By reason of the thrust nozzle being axially adjacent the turbine wheel and radially inwardly of the annular combustor, the turbine engine has a generally annular expanded combustion zone relative to the dilution zone within the same turbine engine envelope.

Volume enhanced turbine engine combustion zone

Unstable operation of a gas turbine engine subject to variable shaft loading and bleed air loading may be avoided in a gas turbine (12, 22, 40) having first and second stage compressors (12, 22) on a common shaft (10) with a turbine wheel (40) and by providing the inlet (14) to the first stage compressor (12) with variable inlet guide vanes (16). Bleed air flow is controlled by a bleed air valve (60) connected to the interface (24) of the first stage compressor outlet (18) and the second stage compressor inlet (20). A controller (66) is operative to control the position of the inlet guide vane (16) in response to bleed air flow sensed by a sensor (58).

Integrated bleed load compressor and turbine control system

In order to provide an expansive central recirculating gas zone (40) in a combustor (10), and in a manner wherein the combustion chamber (20) is of compact volume, a stored energy combustor (10) comprises a vessel (12) having narrow, spaced-apart inlet and outlet ends (14, 16) interconnected by a wall (18) defining a relatively wise combustion chamber (20). The combustion chamber (20) is generally annular and a wall (18) defining the combustion chamber (20) includes an upstream wall region (18a) and a downstream wall region (18c) interconnected by a generally annular side wall region (18b). The inlet end (14) and outlet end (16) are generally tubular extensions of the vessel (12) leading to and from the combustion chamber (20) and oxidant is swirled in the tubular extension (14) leading to the combustion chamber (20) and directed in a swirling annulus into the combustion chamber (20) outwardly of a fuel injector (28). The combustion chamber (20) is formed such that at least the upstream wall region (18a) has an inner surface of gentle radius joining the generally annular side wall region (18b) to the tubular extension (14) of the vessel (12) leading to the combustion chamber (20). With this arrangement, the swirling oxidant annulus expands radially outwardly in a generally conical fashion within the combustion chamber (20) upon exiting the tubular extension (14) while sticking to at least the upstream wall region (18a) of the wall (18) defining the combustion chamber (20).

Stored energy combustor

The high cost of the fuel injection system in a gas turbine engine can be reduced by utilizing a low cost jet or nozzle (70) to spray fuel into the inlet (24) of a compressor (18). The resulting air-fuel mixture flows as a semi-discrete sector to a combustor (58) to be swirled therein and combusted to drive a turbine wheel (50).

Low cost fuel system for a gas turbine engine

Time lost in recharging compressed gas storage bottles on board aircraft is minimized with a fast recharge compressor (20) for rapidly providing compressed gas (30) having a housing (24) with a gas inlet (26), a shaft (46) within the housing (24) on which is mounted a turbine (34) and a multi-stage compressor (36). The turbine (34) is designed to run the shaft (46) and thereby the multi-stage compressor (36). Entering gas is divided into two streams and one stream (28) is directed through the turbine (34) while the other stream (30) is directed through the compressor stages (38,40,44). Heat is exchanged between the two streams by heat exchangers (52,54,56,57,60) after the expansion stream (28) has passed through the turbine (34) so as to cool the compression stream (30).

Fast recharge compressor

A power augmented gas turbine starter system which has a starter including a casing rotatably mounting a turbine wheel therewithin. A low pressure air supply communicates with the casing. A combustor is associated with the low pressure air supply for receiving air therefrom and supplying a flow of hot gas to the starter for augmenting the power of the starter. The air supply may be a bleed air conduit from an auxiliary power unit or from a main propulsion engine, and the combustor may be removably mounted directly in-line within the air supply conduit. The invention contemplates the use of a single auxiliary power unit and a single combustor for supplying a flow of hot gas to two or more gas turbine starters associated with their respective main propulsion engines. A cross-bleed circuit may be provided so that cross-bleed from one main engine can be used to start another main engine by providing bleed air for the other engine's gas turbine starter.

Gas turbine starter system with power augmentation

Surge protection is achieved in a turbine engine intended to provide varying quantities of bleed air and having two compressor stages 48, 36 by providing the inlet 56 of the first stage 48 with variable inlet guide vanes 58, 60. The first stage compressor 48 is a high specific speed, single stage centrifugal compressor whose stable operating characteristics are highly sensitive to inlet guide vane geometry, thereby allowing variations in such geometry to be employed to prevent compressor surge.

A narrow efficient operating island for a centrifugal compressor may be increased in a construction including a compressor section 18 mounting blades 22 surrounded by an annular shroud 38 by placing bleed passages 88 in the shroud 38 and angling them in the direction of flow, both axially (FIG. 1) and radially (FIG. 3).

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San Diego, CA, United States of America

Colin Rodgers