JPS5924260B2 - Variable cycle gas turbo fan engine - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION In the field of gas turbine engines, the concept commonly referred to as variable cycle engines has received considerable interest in recent years and is likely to receive significant effort in the future.

Although the specific characteristics required to classify an engine as a variable cycle engine have not yet been precisely defined, general characteristics have emerged.

In general, the term "variable cycle engine" refers to a dry type engine with a large shunt ratio, which is best operated at subsonic speeds (the term "dry type" refers to an engine that does not have an afterburner and does not increase thrust).
It refers to a hybrid engine that can operate with characteristics close to both a turbofan engine and a turbojet with afterburning, which is optimal for supersonic flight.

In a gas turbine engine, a large volume of air is first compressed by a fan installed in an annular duct.

A portion of this compressed air is typically ducted to the core engine, where it is further dashed, combusted with a fuel mixture, expanded in a turbine, and extracted from a nozzle to drive the compressor and fan. generates propulsion force.

The remaining air compressed by the fan is ducted around the core engine and exhausted through the nozzle.
Generates additional thrust.

The ratio of the flow rate bypassing the core engine to the flow rate passing through the core engine is called the bypass ratio.

The need for variable cycle engines has arisen because a particular engine needs to be operated in many ways or for various flight purposes.

For well-known reasons, engines with high shunt ratios are inherently quieter than engines of the same size (in terms of thrust) with lower shunt ratios, which may lead to more stringent federal standards for aircraft noise in the future. It is thought that this need will become even stronger.

At the same time, engines with high bypass ratios are not necessarily the optimal type for aircraft supersonic performance.

Modern aircraft that meet the requirements for multiple flight missions are powered by engines that are a compromise between high-side-ratio turbofans for subsonic speeds and thrust-enhancing turbojets for supersonic speeds.

As a result, a turbofan engine with a low to medium bypass ratio was developed.

The following are general characteristics desirable for variable cycle gas turbine engines:

1. Wide range of side road ratios for operation.

In other words, it is possible to perform operation with a large side road ratio and low specific impulse so that it can be operated economically at subsonic speed for a long period of time, and operation with a small side road ratio and high specific impulse for supersonic cruising.

2. A predetermined thrust force can be generated over a considerable range of air flow rate in order to facilitate adjustment of the intake air flow and to minimize the drag force of the inner bag.

3. The aforementioned changes can be made with minimal changes to the shape of the engine and the aerodynamic loads on the parts.

4. Able to change between each movement mode continuously and smoothly.

5 Mechanical and control simplicity.

The present invention provides a unique and simple arrangement that can achieve all of the above objectives.

The complexity of conventional approaches to variable cycle engines has hindered their development.

Traditionally using retractable fans, inefficient combinations of variable area turbines and variable pitch fans and compressors -
I was using it.

The reported type of combined engine is designed to utilize a pair of engines in cascade, with one or both working.

In aircraft applications, the weight associated with an engine that is not used in its mode of operation is inherently disadvantageous.

A very recent attempt to design a practical variable cycle engine involves reversing fan flow through concentric annular ducts (US Pat. No. 3,792,584).

One undesirable feature of such a concept is the complexity of the required switching mechanism (rotating one section of the engine's fan duct relative to another), and the length required for the device. It is also undesirable for it to be long.

Another attempt is to switch the fan flow into alternating fan ducts.

Each duct is switched using a valve mechanism.

One undesirable feature of this type of variable cycle engine is that during switching, flow may be restricted, potentially causing the fan to stall.

The present invention overcomes these drawbacks by providing two separate fan ducts and a means for regulating airflow therethrough without bulky redundancy or mechanical complexity. Overcome.

It is therefore an object of the present invention to provide an improved adjustable variable bypass of a simple type which can operate efficiently at subsonic and supersonic speeds and which can change between these modes of operation without intervening stall conditions. The purpose is to provide a specific turbofan engine.

The above and other objects and advantages will be more clearly understood from the following description of specific examples with reference to the drawings.

These examples are illustrative of the invention and are not intended to limit its scope in any way.

Briefly, to achieve the above objectives, an auxiliary fan with a wide flow range is provided around at least a portion of the conventional fan.

Therewith, variable guide vanes and other flow conditioning elements such as nozzles are utilized to provide a wide range of operational bypass ratios.

Although the subject matter of the invention is specifically and distinctly set forth in the claims, the invention will be more fully understood from the following drawings and a description of two embodiments.

The same reference numerals are used throughout the drawings for like elements.

FIG. 1 schematically shows an engine 10 embodying the invention.

The engine as a whole may be considered to be comprised of a core engine 12, a fan assembly 14, and a fan turbine 16, with the fan turbine connected to the fan assembly 14 by a shaft 18; Drive your body.

The core engine includes an axial compressor 20 having a rotor 22 .

The fan assembly comprises an inner first compression means or fan 24
and an external auxiliary compression means or fan 26.

The inner fan 24 is three-stage, with each stage consisting of alternating rows of rotating and stationary vanes, as is common in fans or compressors.

The number of stages or the particular arrangement required for the fans, as with the compressor or turbine, will depend on the particular cycle conditions, and no limitations are hereby placed on the type or arrangement of the components of such turbofluid machines. I have no intention of doing so.

Thus, the term core engine refers to a compressor, combustion system, and turbine arranged in series with the flow, as is well known.

The term driven fan assembly refers to a fan compressor with additional compression means.

Referring to the engine of FIG. 1, an outer auxiliary compression device (auxiliary fan) 26 is disposed outwardly from the inner fan 24 and includes at least one row of radial extensions of the rows 28 of rotor blades of the inner fan. It is shown that the rotary blade 34 includes a rotating blade 34.

To this end, at least one row of rotating vanes of the fan blades extend across both the concentric inner and outer fan annulus 36,38.

Therefore, the auxiliary fan 26 is considered to be driven by the inner fan 24.

In fighter aircraft where a high pressure ratio is required, it is conceivable to provide a radial extension to the rows of rotor blades of the inner fan to form a multi-stage outer fan.

In the embodiment of FIG. 1, air entering annulus 36 through inlet 40 is initially compressed by inner fan 24. In the embodiment of FIG.

A first portion of the air thus compressed enters the inlet 42, is further compressed by the axial compressor 20, and is then discharged to the combustion device 42 where it combusts fuel to produce high-energy combustion gases. This gas drives the core engine turbine 44.

This turbine has a variable area turbine nozzle 45.

As is common in gas turbine engines, the turbine 44 is connected to the shaft 4.
6 (gas turbine engines with two shafts, such as those shown at 18 and 46, are commonly referred to as two-spool machines).

Hot combustion gases are sent to and drive fan turbine 16.

The turbine has a variable area turbine nozzle 47 and drives the fan assembly 14.

The variable turbine nozzle 45 is used to set the core engine operating system (relationship between pressure ratio and power fluid flow rate), whereas the variable turbine nozzle 47 is used to set the core engine operating system (relationship between pressure ratio and power fluid flow rate).
A relationship is established between the speed of engine 12 and the speed of fan assembly 14.

A second portion of the air emerging from the inner fan 24 partially penetrates the core engine 12 and its circumscribing flow path boundary 5.
0 and an inner annular fan duct (first
duct) enters 48.

Air entering the annulus 38 from the inlet 52 as a supplemental airflow is compressed by the outer fan 26 driven from the fan turbine 16 via the fan rotor 30 to the outer annular fan duct (second duct) 544. , sent by .

The outer fan has variable inlet and outlet guide vanes 5 so that it can be operated at full flow at high bypass ratios and at reduced flow at low bypass ratios.
It is shown to have an adjustment device consisting of a guiding device such as 6.58.

For supersonic operation, the outer fan duct (second duct)
It is desirable to minimize the auxiliary air flow at 54 as much as possible.

It is of interest to what extent the auxiliary air flow through the outer duct 54 can be reduced for high-speed operation.
become a target.

Complete closure is probably not practical and is also undesirable in terms of air resistance losses as well as local internal heating.

Additionally, in supersonic flight conditions, the inlet typically experiences a tip-induced distortion of the radial pressure pattern, and it is desirable to pass this region of low energy flow into the outer fan duct 54.

Therefore, during supersonic operation, the variable guide vanes 56, 58 (which act as loopers) are set to minimize the flow rate (essentially zero flow rate).

The airflow passing through the inner fan assembly 24 is transferred to the inner fan assembly 24.
It is divided into a duct (first duct) 48 and a core engine 12.

These streams are combined and mixed again in a well-known mixer 60.

A second or auxiliary combustion device 6 of the known afterburner type
2 is provided and a combustion and gas mixture is supplied therein to combust and generate an exhaust stream.

This exhaust flow passes through a variable area primary nozzle arrangement 64;
The compressed outer fan duct 54 joins the flow through a communication means, such as an opening or nozzle 66, between the variable secondary nozzle device 68 and the primary device 64.

The dimensions of opening 66 must be compatible with the flow rate and conditions of outer duct 54.

That is, in supersonic operation, the bypass ratio (the value obtained by dividing the flow rate through the outer duct 54 and the inner duct 48 by the flow rate through the core engine 12) is small, and the specific impulse is large.

In subsonic cruise operation, the second combustion device 62 is not used.

Adjusting the variable guide vanes 56,58 and the apertures 66 maximizes the flow in the outhole duct 54 and reduces the rotational speed of the inner fan 24, thus reducing the pressure ratio.

As would be expected, this lower pressure is at the mixer 60.
to the mixing plane 70 of the fan turbine 16, reducing the pressure on the discharge side of the fan turbine 16.

This allows the fan turbine 16i'i to do more work to meet the increasing demands of the outer fan 26.

By appropriately varying the areas of the variable area turbine nozzles 45, 47 and the primary exhaust nozzle 64, the fan assembly 14 and core engine 12 can be balanced for optimal operating conditions.

The variable cycle concept thus far described can also be applied to turbofan engines using separate flows for combustion in two-spool type ducts, such as the embodiment shown in FIG.

In this case, the six mixers of FIG. 1 are removed and a divider 72 is added to separate the flow exiting the turbine 16 from the flow in the inner fan duct 48.

The basic difference between the two embodiments is that in supersonic mode, the auxiliary duct combustion device 74 only increases the flow in the inner fan duct 48 and the core engine nozzle 80 and the inner fan duct nozzle, respectively. The area of the throats 76, 78 of 82 is to be separately controlled for optimal performance.

The extra energy required to drive the outer auxiliary fan 26 in a subsonic manner increases the area of the throat 76;
This was obtained by lowering the pressure on the discharge side of the fan turbine 16.

The manner in which the variable cycle concept is used in the two embodiments described above is substantially similar.

However, the embodiment of FIG. 1 has the inherent constraint that the static pressure in the inner fan duct 48 as well as the static pressure of the combustion residue exiting the core engine 12 must be equal at the mixing plane 70. There is.

Thus, whether the fan turbine 16 is in the high pressure ratio, low flow regime typical of supersonic operation, or the low pressure ratio, high flow regime typical of efficient subsonic operation; It is clear that either mode of operation acts as an energy handling device capable of supplying the required energy.

It will be apparent to those skilled in the art that various modifications can be made to the above-described mechanism without departing from the scope of the invention.

For example, it is contemplated that the auxiliary airflow of the outer fan duct 54 could be discharged into high drag areas of the aircraft installation to fill the areas and thus improve the overall performance of the system.

In another embodiment, supplemental airflow can be passed over or through the wing flaps to increase lift in a manner known in the art.

In either case, aperture 66 may not be necessary.

Furthermore, in some configurations, it may not be necessary to make some or all of the turbine and exhaust nozzle variable.

The claims are intended to cover all similar modifications falling within the scope of the invention.

[Brief explanation of the drawing]

FIG. 1 is a schematic diagram of an engine using one embodiment of the present invention, and FIG.
The figure is a schematic representation of an engine similar to FIG. 1 but using an alternative embodiment. Explanation of main symbols, 12: Core engine, 16: Fan turbine, 24: Inner fan, 26: Outside (JIIJ fan (auxiliary compressor), 44: Core engine turbine, 45: Variable area turbine Nozzle, 4T: variable area turbine nozzle, 48: inner annular fan duct),
54: Outer annular fan duct, 56, 58: Variable guide vane, 62: Auxiliary combustion device, so: Core engine
nozzle.

Claims (1)

Claims: 1. (a) a core engine including a compressor, a combustion device, and a core turbine; and (b) a first fluid portion entering the core engine and exiting the core engine as a hot gas stream. (c) a fan rotor vane pressurizing a power fluid comprising: a second fluid portion passing as bypass flow through a substantially annular first duct generally surrounding the core engine; (d) adjusting the flow rate of the auxiliary power fluid stream between a high flow rate in the subsonic mode of operation and essentially zero flow rate in the supersonic mode of operation; (e) a fan turbine device driven by the hot gas flow to rotate and drive both fan rotating blades and an auxiliary compression device in both subsonic and supersonic modes of operation; (f) A variable cycle gas turbofan engine comprising: an auxiliary combustion device for receiving and combusting fuel and a bypass flow in a supersonic mode of operation; and (g) a variable area primary nozzle device for discharging said hot gas flow. 2. said auxiliary power fluid flow is directed into a substantially annular second duct partially surrounding the annular first duct, a guiding device for guiding the auxiliary power fluid flow to the auxiliary compression device; Claim 1 provided on the upstream side
Variable cycle gas turbofan engine as described in Section. 3. The variable cycle gas turbofan engine of claim 2, wherein the fan turbine device is of a variable area type. 4. The variable cycle gas turbofan engine according to claim 2, wherein the guide device comprises a non-rotating first stage variable area guide vane. 5. The auxiliary power fluid flow is further guided axially downstream of the auxiliary compression device by a non-rotating second stage variable area guide vane. A variable cycle gas turbofan engine according to claim 4. 6. The variable cycle gas turbofan engine of claim 2, wherein the auxiliary combustion device is located substantially within the annular first duct. 7. The variable cycle gas turbofan engine of claim 6, wherein the post-combustion bypass flow is discharged by a variable area secondary nozzle arrangement. 8. Claim 7, further comprising a communication device for interacting said auxiliary power fluid flow with said post-combustion bypass flow prior to exiting said variable cycle gas turbofan engine.
Variable cycle gas turbofan engine as described in Section. 9. The variable cycle gas turbofan engine of claim 8, wherein the communication device is of the variable area aperture type. 10. The variable cycle gas turbofan engine of claim 2, wherein the bypass flow is mixed with the hot gas flow by a mixer in a supersonic mode of operation, characterized by a mixed flow. 11. the auxiliary combustion device is located downstream of the mixer and receives the mixed flow in a supersonic mode of operation;
A variable cycle gas turbofan engine according to claim 10. 12 the post-combustion mixed stream is discharged through the auxiliary combustion device and subsequently through the variable area primary nozzle device;
A variable cycle gas turbofan engine according to claim 11. 13 A device for adjusting the flow rate of the auxiliary power fluid flow is connected to the annular second
13. The variable cycle gas turbofan engine of claim 12, comprising a variable area opening between the downstream end of the duct and the variable area primary nozzle.