Pulsed combustion engine

Abstract

A turbine engine has a case with an axis. A fan is mounted for rotation about the axis. A turbine is mechanically coupled to the fan to drive rotation of the fan about the axis. A number of compressor/turbine units are downstream of the fan and upstream of the turbine along a core flowpath. A number of compressors are coupled to the compressor/turbine units to receive air and deliver combustion gas to drive the turbine.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic partial longitudinal sectional view of a turbofan engine.

FIG. 2 is a cutaway view of the engine of FIG. 1.

FIG. 3 is a schematic partial longitudinal sectional view of an alternative engine.

FIG. 4 is a front schematic view of a second alternative engine.

Like reference numbers and designations in the various drawings indicate like elements.

DETAILED DESCRIPTION

FIG. 1 shows a turbofan engine 20 having central longitudinal axis 500, a case 22, and a core 24. The case 22 defines a duct 26 extending from an upstream inlet 28 to a downstream outlet 30. Of an inlet airflow 510 entering the duct, a fan 32 drives a bypass portion 512 and a core portion 514 along respective bypass and core flowpaths through the duct. The exemplary fan 32 has two blade stages and two interspersed vane stages. The blade stages may be supported on a shaft 34.

As is described in further detail below, the exemplary engine 20 also includes a circumferential array of compressor/turbine units 38, a combustor section 40 (e.g., circumferential array of combustors 41), and a turbine section 42. Other components (e.g., an augmentor and an exhaust nozzle) may also be present. FIG. 2 shows further details of exemplary positions of the exemplary compressor/turbine units 38 and combustors 41.

The core airflow 514 is divided by ducts 44 into branching portions directed to the compressor sections 50 (e.g., centrifugal compressors) of each of the units 38. Rotation of the impeller of the section 50 is driven by the turbine of the turbine section 52 (e.g., a radial turbine) of the associated unit 38. The units 38 thus compress the flow 514 into compressed flows 516 directed to the combustor section 40. In the combustor section 40, the compressed air is mixed with a fuel flow 518 and combusted to form combustion gas 520. The gas 520 is directed to the turbine of the turbine section 52 where it is partially expanded to extract the work to compress the flow 514.

From the unit 38, the partially expanded combustion gas flow 522 is directed to the turbine section 42. For example, the turbine sections 52 of the various units 38 may be coupled to a common discharge manifold 60 feeding an upstream/inlet end of the turbine section 42. As the flow 522 passes through the turbine section 42 it is further expanded and discharged as a flow 524. The exemplary flow 524 is directed via a manifold duct 62 to merge with the bypass flow 512 and form a combined flow 526. This combined flow may ultimately be discharged from the outlet 30.

In the exemplary engine of FIG. 1, the blade stages of the turbine section 42 are co-spooled with the fan on the shaft 34. The positioning of the turbine section 42 forward of the combustor section 40, along with the generally forward flow through the turbine section 42 facilitates a short shaft 34 and a longitudinally compact engine. The configuration also hides the moving/hot surfaces of the turbine section 42 from line-of-sight exposure through the outlet. This may be advantageous for low observability properties including radar return and infrared signature.

FIG. 1 shows further details of the exemplary combustor section 40. FIG. 1 shows an inner member 80 within an outer member 82. The airflow 516 is received through an associated conduit 84 to a volume or space 86 between the inner and outer members. There may be a circumferential array of the inner members 80 (one for each combustor 41). In some variations, the outer member 82 may be a single outer member containing all or more than one of the inner members (e.g., an annular outer member). In other variations, there may be a circumferential array of the outer members 82, each containing an associated one of the inner members 80.

The exemplary inner member 80 has an aft end 90 and a fore end 92. The exemplary inner member 80 has a first frustoconical wall portion 94 diverging forward from the aft end 90. The wall portion 94 is foraminate allowing the inflow of air. In the exemplary combustor, a fuel injector 100 may be positioned at the aft end to introduce the fuel flow 518. An igniter 102 (e.g., a sparkplug) may be positioned to ignite the fuel air mixture to cause combustion. The divergence of the wall portion 94 helps facilitate a deflagration-to-detonation transition.

The exemplary inner member 80 has a second wall portion 110 forward of the portion 94. A convergent wall portion 112 is downstream of the portion 110. An outlet conduit 114 connects the inner member 80 to the associated turbine section 52. Individual coupling of the combustors to at least the turbine section 52 prevents crosstalk between the discharge ends of the combustors. This is relevant where the combustors are operated out-of-phase so that the combustion gas discharged by one combustor is not ingested by another.

Inlet decoupling is less critical. Thus, there may be a common outer member 82 defining a common inlet plenum. In yet other embodiments, each combustor may be coupled to receive air from the compressor section 50 of one unit 38 while discharging gases to the turbine section 52 of another unit.

FIG. 3 shows an alternative configuration with a long shaft 34′ connecting a turbine section 42′ to the fan. The exemplary turbine section 42′ is aft of the combustor section and receives combustion gases from the compressor/turbine unit array through a manifold 160′ directing the combustion gases generally aftward and radially inboard of the combustors. The discharged combustion gases and bypass air mix relatively downstream.

The effects of the pressure pulses from the individual combustors is minimized by operation out-of-phase with each other. Exemplary firing frequency may be in the vicinity of 50-300 Hz and may vary considerably depending on the scale/size of the engine and resulting impact on combustor section geometry and volume. Various phase combinations are possible, including firing in opposed pairs to limit wobble. Exemplary fan spool speeds are 2000-20000 revolutions per minute (RPM), more narrowly 6000-12000 RPM. Exemplary speeds for the units 38 are 5000-50000 RPM, more narrowly 20000-35000 RPM as an approximation for the 6000-12000 RPM fan spool speeds under steady-state conditions.

Many variations are possible. For example, the combustors take a variety of forms, including shapes, positions, and orientations. FIG. 4 shows an exemplary configuration wherein eight combustors are grouped in two groups concentrated on respective left and right sides of the engine. This creates a wide but small height package which may be advantageous for integration into the airframe of an aircraft (e.g., a fighter aircraft, unmanned aerial vehicle, or missile).

One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. For example, the details of any particular application will influence the configuration of the combustor. Various features of the combustor may be fully or partially integrated with features of the turbine or the compressor. If applied in a redesign of an existing turbine engine, details of the existing engine may implement details of the implementation. The rotating combustor may alternatively be used in applications beyond turbine engines. Accordingly, other embodiments are within the scope of the following claims.

Claims

1. A turbine engine comprising: a case having an axis;a fan mounted for rotation about the axis;a turbine mechanically coupled to the fan to drive rotation of the fan about the axis;a plurality of compressor/turbine units downstream of the fan and upstream of the turbine along a core flowpath; anda plurality of combustors, coupled to the plurality of compressor/turbine units to receive air and deliver combustion gas to drive the turbine.

2. The engine of claim 1 wherein: the compressor/turbine units each have a centrifugal compressor section and a coaxial radial turbine section.

3. The engine of claim 1 wherein: the compressor/turbine units each have a compressor section having an axial inlet and radial outlet and a coaxial turbine section having a radial inlet and axial outlet.

4. The engine of claim 1 wherein the plurality of compressor/turbine units is a circumferential array of compressor/turbine units; andthe plurality of combustors is a circumferential array of combustors.

5. The engine of claim 1 wherein: each of the plurality of compressor/turbine units is uniquely associated with a single one of the combustors and vice versa and coupled so that:the compressor of the compressor/turbine unit delivers the air to the associated combustor; andthe turbine of the compressor/turbine unit receives the combustion gas from the associated combustor.

6. The engine of claim 1 wherein: said turbine is an axial turbine and receiving the combustion gas from all of the compressor/turbine units.

7. The engine of claim 6 wherein: the axial turbine is co-spooled with the fan.

8. The engine of claim 1 wherein: there are at least eight of the compressor/turbine units and at least eight of the combustors.

9. The engine of claim 1 wherein: the combustors are non-rotating.

10. The engine of claim 1 wherein each combustor comprises: an outer wall;an inner wall, within the outer wall;a space between the inner and outer walls coupled to receive air from at least an associated one of the compressor/turbine units;a fuel injector positioned to introduce fuel to the air; anda space within the inner wall coupled to deliver combustion gas to at least an associated one of the compressor/turbine units.

11. The engine of claim 1 wherein: flow of the combustion gas to the plurality of compressor/turbine units is axially toward the fan.

12. A method for operating a turbine engine comprising: directing air from a fan to a plurality of compressor/turbine units;compressing the air in the compressor/turbine units;directing the air to a plurality of combustors;combusting the air with fuel in the combustors to produce combustion gas;extracting work from the combustion gas in the compressor/turbine units to drive said compressing;directing the combustion gas from the compressor/turbine units to a turbine; andextracting work from the combustion gas in the turbine to drive rotation of the fan.

13. The method of claim 12 wherein: the compressing is centrifugal; andthe extracting work from the combustion gas in each of the compressor/turbine units comprises passing the combustion gas radially inward toward an axis of the unit.

14. The method of claim 12 further comprising: directing the combustion gas from the turbine to join a bypass flow of air from the fan.

15. The method of claim 14 wherein: a mass flow ratio of the flow of the air delivered to the combustors to the bypass flow is between 1:1 and 1:3.

16. The method of claim 12 wherein: the combusting is a pulsed combusting.

17. The method of claim 16 wherein: the combusting comprises detonation.

18. The method of claim 16 wherein: the pulsed combusting in each of the combustors has a frequency of 50-300 Hz.

19. The method of claim 16 wherein: the combusting comprises operating respective ones of the combustors out of phase with each other.

20. The method of claim 12 used in aircraft propulsion.

21. The method of claim 12 wherein: the compressor/turbine units are running at a speed of at least 20000 RPM; andthe fan rotates at a speed of 6000-12000 RPM.