CN213331289U - Intake fairing, gas turbine engine and aircraft - Google Patents

Abstract

The utility model relates to an admit air radome fairing, gas turbine engine and aircraft. The air inlet fairing comprises an inner skin, an outer skin and an annular cascade, wherein the outer skin surrounds the inner skin, and a hot air channel is provided between the inner skin and the outer skin; the inlet of the hot gas channel is a port of the inner skin, the annular blade grid is arranged at the front section of the hot gas channel and is adjacent to the port of the inner skin, the blade root of the blade of the annular blade grid is fixedly connected with the inner skin, and the outer skin is provided with the gas outlet part of the hot gas channel.

Description

Intake fairing, gas turbine engine and aircraft

Technical Field

The utility model relates to a gas turbine engine field especially relates to an admit air radome fairing, gas turbine engine and aircraft.

Background

Icing can occur during operation of the gas turbine engine due to low temperatures. In the case of aircraft engines, as cloud layers may contain metastable supercooled liquid water at temperatures below freezing, icing may easily occur on the surface of the aircraft's windward components as the aircraft passes through the cloud layers. For air inlet components of an aircraft engine, such as an air inlet fairing, a fan blade, an air inlet support plate, an engine splitter ring and the like, the probability of icing is higher because the airflow is accelerated and cooled when being sucked by the engine. Icing can deteriorate the starting performance of the components and cause the center of gravity of the rotating member to shift, thereby increasing vibration, which is very disadvantageous for flight safety. Therefore, anti-icing systems are commonly deployed on currently in-service aircraft and their engines.

In the prior art, a hot air anti-icing system (hot air anti-icing system) is the most mature at present and is also the most commonly used anti-icing system, hot air is mainly led out from an engine air system and is conveyed to an inner cavity of an anti-icing component through a specific pipeline and a valve, and therefore the purposes of increasing the surface temperature of the anti-icing component and preventing the surface of the component from being iced are achieved. The hot gas anti-icing system generally comprises a pipeline, a valve and an anti-icing cavity structure inside an anti-icing part, wherein the anti-icing part of the aircraft engine mainly comprises an air inlet fairing, fan blades, an air inlet support plate, an engine shunt ring, a supercharging stage inlet guide blade, a supercharging stage first-stage rotor blade, a supercharging stage second-stage rotor blade and the like. The anti-icing bleed air is typically either funneled into the other air system flow paths after exiting the anti-icing chamber or vented directly to the outside atmosphere and to the engine main flow path.

The engine intake cowling is at the forward-most end of the engine intake components and is typically an anti-icing component, the common forms of which are hot gas anti-icing and structural anti-icing. When hot air is adopted for anti-icing, the scheme of the prior art mainly comprises that the hot air is directly exhausted from a hollow or a seam on the surface of the cap cover after being converged into the cavity of the cap cover. Referring to fig. 1, a double-skin structure in which an inner skin 200 and an outer skin 100 provide a heat exchange channel 300 is adopted to limit a hot air flow area, so that a hot air flow rate is increased to achieve heat exchange between hot air and a fairing, but the increase of the flow rate also shortens the time for the hot air to stay inside the fairing per unit mass, so that the temperature of the hot air at an exhaust outlet is still maintained at a higher level, and the energy (enthalpy) of the hot air entering the fairing cannot be efficiently converted into anti-icing energy, so that in order to ensure the anti-icing effect, only the pressure and temperature of bleed air can be increased, so that the amount of bleed air is increased, and waste of the bleed air is indirectly caused.

In the prior art, there is also an improvement on the technical scheme in fig. 1 to improve the heat exchange effect, for example, the chinese utility model patent application with publication number CN203753413U, entitled "anti-icing heat transfer structure of aircraft engine inlet fairing" has a jet hole formed in the inner skin to introduce hot air into the heat exchange channel, but this structure will increase the pressure drop along the air flow path, so that the air-entraining pressure cannot be reduced; on the other hand, hot gas needs to pass through the front edge of the fairing cap and then enters the heat exchange channel, so that excessive energy loss of the hot gas occurs on the front edge of the fairing cap, partial area of the fairing can be heated only through solid heat conduction, and the heat exchange efficiency is reduced to a certain extent.

Therefore, there is a need in the art for an inlet cowl, a gas turbine engine, and an aircraft to improve the heat exchange effect in the hot gas inlet cowl, improve the anti-icing performance thereof, improve the hot gas utilization efficiency, reduce the induced flow, improve the performance of the gas turbine engine, and further improve the flight safety performance of the aircraft in the flight process.

SUMMERY OF THE UTILITY MODEL

An object of the utility model is to provide an inlet air fairing.

It is another object of the present invention to provide a gas turbine engine.

It is yet another object of the present invention to provide an aircraft.

An air intake fairing according to one aspect of the present invention includes an inner skin, an outer skin, and an annular cascade; wherein the outer skin surrounds the inner skin, a hot gas path being provided between the inner skin and the outer skin; the inlet of the hot gas channel is a port of the inner skin, the annular blade grid is arranged at the front section of the hot gas channel and is adjacent to the port of the inner skin, the blade root of the blade of the annular blade grid is fixedly connected with the inner skin, and the outer skin is provided with the outlet of the hot gas channel.

In one or more embodiments of the intake fairing, the inner skin and the outer skin are tapered, and the port projects axially a first length from a small end of the inner skin.

In one or more embodiments of the intake fairing, the outlet of the hot gas path is an outlet hole penetrating through the outer skin, the outlet hole is located near the large end of the outer skin, and the extension direction of the outlet hole is the radial direction of the intake fairing; the large ends of the outer skin and the inner skin close the hot gas channel.

In one or more embodiments of the air intake fairing, a connection region between a blade root of the blade of the annular blade cascade and the inner skin is an aerodynamic smooth structure.

In one or more embodiments of the air intake fairing, the inlet metal angle of the airfoil of the annular cascade coincides with the direction of extension of the hot gas path.

In one or more embodiments of the intake fairing, the exit metal angles of the airfoil of the annular cascade are circumferentially offset.

In one or more embodiments of the intake fairing, the airfoil of the annular cascade is an airfoil of a turbine blade.

In one or more embodiments of the inlet cowl, the hot gas path has a plurality of annular cascades therein, one and the annular cascade is disposed at a forward section of the hot gas path, and the remaining annular cascades are disposed at a mid-section and/or a rear section of the hot gas path.

A gas turbine engine in accordance with an aspect of the present invention includes the inlet cowling of any one of the above.

An aircraft according to an aspect of the present invention comprises a fuselage, wings and a gas turbine engine.

The utility model has the advantages of but not limited to:

through setting up annular cascade at hot gas passage's anterior segment, the back is got into from the entry to the hot gas air current, and through annular cascade, the effect of annular cascade is to change the hot gas flow direction, makes its axial velocity reduce and circumferential velocity increase, and the air current gets into double-skinned structure and the radome fairing carries out the heat transfer behind the cascade. Because the axial speed of the airflow is reduced and the position of the axial outlet is not changed, the time for the hot air of unit mass to flow out of the fairing is prolonged, so that the heat exchange efficiency between the hot air and the fairing is improved, and the using amount of the anti-icing air entraining can be saved or the temperature of the anti-icing air entraining can be reduced. The reduction of the air entraining amount can not only reduce the influence of the work of the anti-icing system on the performance of the whole engine, but also reduce the sizes of pipelines and valves of the anti-icing system, realize the weight reduction of the engine and make positive contribution to the improvement of the economic index of the engine; meanwhile, the anti-icing performance is improved, and the flight safety performance of the airplane in the flight process can also be improved.

Drawings

The above and other features, properties and advantages of the present invention will become more apparent from the following description of the embodiments with reference to the accompanying drawings, in which:

FIG. 1 is a schematic view of a prior art air intake fairing.

FIG. 2 is a schematic view of an intake fairing and hot gas flow of an embodiment.

Fig. 3 is a view from a-a of fig. 2.

FIG. 4 is a block diagram of an embodiment of an inner skin of an intake fairing provided with an annular cascade.

Reference numerals:

10-inlet fairing

20-hot gas path

201-front section

202-middle section

203-rear section

30-air intake channel

40-hot gas flow

1-outer skin

2-inner skin

3-inlet of hot gas channel

4-outlet of hot gas channel

5-ring blade cascade

51-blade

511-blade root

Detailed Description

The present invention will be further described with reference to the following embodiments and drawings, and more details will be set forth in the following description in order to provide a thorough understanding of the present invention, but it is obvious that the present invention can be implemented in various other ways different from those described herein, and those skilled in the art can make similar generalizations and deductions according to the actual application without departing from the spirit of the present invention, and therefore, the scope of the present invention should not be limited by the contents of the embodiments.

Also, the present application uses specific words to describe embodiments of the application, such as "one embodiment," "an embodiment," and/or "some embodiments" to mean that a particular feature, structure, or characteristic described in connection with at least one embodiment of the application. Therefore, it is emphasized and should be appreciated that two or more references to "an embodiment" or "one embodiment" in various places throughout this specification are not necessarily all referring to the same embodiment. Furthermore, some features, structures, or characteristics of one or more embodiments of the present application may be combined as appropriate.

It should be noted that, in the following embodiments, the axial direction, the circumferential direction, and the radial direction are axial directions, circumferential directions, and radial directions of the intake cowling; the front and the rear are the axial relative positions of the air inlet fairing.

The following embodiments describe an inlet cowl for a gas turbine engine in general, and specifically for example, a turbo fan aero-engine (turbo fan-engine) of an aircraft, the aircraft including a fuselage and a wing, the aero-engine being disposed on the wing of the aircraft.

Referring to fig. 2-4, in some embodiments, the intake fairing 10 includes an outer skin 1, an inner skin 2, and an annular cascade 5. The inner skin 2 and the outer skin 1 can be both in a conical structure. The outer skin 1 radially surrounds the inner skin 2, and a hot gas channel 20 is provided between the conical surface of the inner skin 2 and the conical surface of the outer skin 1. The inlet 3 of the hot air channel 20 is a port of the inner skin 2 in the axial direction, for example, as shown in fig. 2 and 4, the position of the inlet 3 is located at a position axially protruding from the small end of the conical surface of the inner skin 2 by a first length L1, but not limited to this, the inlet 3 may also be a structure directly opening at the small end of the conical surface of the inner skin 2, and the like. However, the inlet 3 axially protrudes from the small end by a first length, and an inlet passage 30 of the first length L is formed therebetween, so that the flow of the hot gas can be concentrated and guided, and the flow of the hot gas entering the hot gas passage 20 is stable. It will be appreciated by those skilled in the art that in addition to fulfilling the basic function of anti-icing, the amount of hot gas bleed should be as small as possible to improve the thermal efficiency of the engine, so that the smaller the pore size of the inlet 3, the better, but the minimum size of the inlet 3 should not be smaller than that required to achieve the required critical anti-icing bleed flow locally to ensure that there is sufficient hot gas to act as anti-icing.

With continued reference to fig. 2 and 4, the annular cascade 5 is positioned at the forward section 201 of the hot gas path 20, adjacent to the port of the inner skin 2, i.e. the inlet 3 of the hot gas path 20, the root 511 of the blades 51 of the annular cascade 5 being fixedly connected to the inner skin 2, while the outer skin 1 has the outlet 4 of said hot gas path. As shown in fig. 2, after the hot gas flow 40 enters the hot gas channel 20 from the inlet 3, the hot gas flow first flows along the extending direction of the hot gas channel 20, and after flowing through the annular blade row 5, the direction of the hot gas flow 40 is deflected under the influence of the annular blade row 5, so that the axial velocity of the hot gas flow 40 is reduced, and the circumferential velocity thereof is increased, as shown in fig. 2, because the axial velocity of the hot gas flow 40 is reduced and the position of the axial outlet is not changed, the time for the unit mass of hot gas to flow out of the inlet fairing 10 is prolonged, so that the heat exchange between the unit mass of hot gas and the cap is more sufficient, and thus the heat exchange efficiency between the hot gas and the inlet fairing 10 is improved, so that the amount of the anti-icing bleed air can be saved or the temperature of the anti-. The reduction of the air entraining amount can not only reduce the influence of the work of the anti-icing system on the performance of the whole engine, but also reduce the sizes of pipelines and valves of the anti-icing system, realize the weight reduction of the engine and make positive contribution to the improvement of the economic index of the engine.

With continued reference to fig. 3 and 4, in some embodiments, the fixed connection between the blade root 511 of the blade 51 of the annular blade row 5 and the inner skin 2 may be a tongue-and-groove connection structure similar to the connection between a rotor blade and a rotor disk in a gas turbine engine, a welded structure, an integrally formed fixed connection structure formed by 3D printing, or other common fixed connection structures, which are not limited to this. The connecting area of the blade root 511 of the blade 51 of the annular blade cascade 5 and the inner skin 2 is an aerodynamic smooth structure, i.e. the influence of the connecting area on the smoothness of the flow channel is reduced as much as possible, so that the flow loss of the hot gas flow 40 is small. The method for realizing the pneumatic smooth structure may be a conventional leaf-type design method, for example, a spline curve design method, and a second derivative continuity of a curve at a junction is realized, for example, a bezier curve or the like, and the spline curve design method is a means well known to those skilled in the art and will not be described herein again.

With continued reference to fig. 2-4, in one or more embodiments, the inlet metal angles of the airfoils of the annular cascade 5 are aligned with the direction of extension of the hot gas path 20 to reduce the flow losses of the hot gas stream 40, it being understood that the "alignment" described above is not strictly aligned and may also allow for tolerances within certain positive and negative angles of attack, but is preferably strictly aligned. The outlet metal angle of the blade profile can be circumferentially deflected by a certain angle, so that the circumferential flow velocity of the hot gas flow 40 can be increased as much as possible under the condition of not causing gas flow separation, and the axial flow velocity of the hot gas flow 40 is reduced, so that the flowing time of the hot gas flow in the hot gas channel 20 is longer, and further, the heat exchange is fully carried out. The blade profile design of the annular blade cascade 5 can refer to the blade profile of a turbine blade similar to a gas turbine engine, the turbine blade has a large bending angle, more axial speed can be converted into circumferential speed, so that the flowing time of the hot gas flow 40 in the hot gas channel 20 is prolonged, and the design difficulty of the annular blade cascade 5 is reduced by referring to the blade profile design of the existing turbine blade.

With continued reference to fig. 2, the specific structure of the outlet 4 of the hot gas channel 20 may be to include an outlet hole 41 penetrating through the outer skin 1, the outlet hole 41 being located near the large end of the outer skin 1, and the extending direction of the outlet hole 41 being the radial direction of the inlet fairing 10, while the large ends of the outer skin 1 and the inner skin 2 axially close the hot gas channel 20, so that the hot gas flow 40 is difficult to axially exhaust and partially stagnates near the large end of the outer skin 1 and the large end of the inner skin 2 of the hot gas channel 20, and the flow speed of the hot gas in the hot gas channel 20 is slowed down, and the time that the hot gas per unit mass can stagnate inside the inlet fairing 10 is further prolonged, thereby further improving the heat exchange efficiency between the hot gas and the inlet fairing 10.

With continued reference to FIG. 2, in some embodiments, in addition to providing the annular cascade 5 at the forward section of the hot gas path 20, the intermediate section 202 and/or the aft section 203 of the hot gas path 20 may also be provided with the annular cascade 5, which may further enhance the circumferential flow effect of the hot gas flow 40 of the hot gas path 20, preventing insufficient circumferential rotation of the hot gas at the aft section of the hot gas path 20 in the case of a longer length of the hot gas path 20, thereby affecting the heat exchange efficiency.

In summary, the intake fairing and the gas turbine engine introduced by the above embodiments have the beneficial effects that, but not limited to, the annular blade cascade is arranged at the front section of the hot gas channel, the hot gas flow enters from the inlet and passes through the annular blade cascade, the annular blade cascade has the function of changing the flow direction of the hot gas, so that the axial speed of the hot gas is reduced, the circumferential speed of the hot gas is increased, and the hot gas flows through the blade cascade and enters the double-skin structure to exchange heat with the fairing. Because the axial speed of the airflow is reduced and the position of the axial outlet is not changed, the time for the hot air of unit mass to flow out of the fairing is prolonged, so that the heat exchange efficiency between the hot air and the fairing is improved, and the using amount of the anti-icing air entraining can be saved or the temperature of the anti-icing air entraining can be reduced. The reduction of the air entraining amount can not only reduce the influence of the work of the anti-icing system on the performance of the whole engine, but also reduce the sizes of pipelines and valves of the anti-icing system, realize the weight reduction of the engine and make positive contribution to the improvement of the economic index of the engine; meanwhile, the anti-icing performance is improved, and the flight safety performance of the airplane in the flight process can also be improved.

Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make various changes and modifications without departing from the spirit and scope of the present invention. Therefore, any modification, equivalent changes and modifications made to the above embodiments according to the technical spirit of the present invention, all without departing from the content of the technical solution of the present invention, fall within the scope of protection defined by the claims of the present invention.

Claims (10)

1. An intake fairing, comprising:

an inner skin;

outer skin;

an annular cascade;

wherein the outer skin surrounds the inner skin, a hot gas path being provided between the inner skin and the outer skin; the inlet of the hot gas channel is a port of the inner skin, the annular blade grid is arranged at the front section of the hot gas channel and is adjacent to the port of the inner skin, the blade root of the blade of the annular blade grid is fixedly connected with the inner skin, and the outer skin is provided with the outlet of the hot gas channel.

2. The air intake fairing as recited in claim 1, wherein said inner skin and said outer skin are tapered, said port projecting axially a first length from a small end of said inner skin.

3. The intake fairing as recited in claim 2, wherein the outlet of said hot gas path is an outlet aperture through said outer skin, said outlet aperture being located adjacent a large end of said outer skin, said outlet aperture extending in a direction radially of said intake fairing; the large ends of the outer skin and the inner skin close the hot gas channel.

4. The air intake fairing as recited in claim 1, wherein a connection area of a blade root of said annular cascade blade to said inner skin is aerodynamically smooth.

5. The air intake fairing as recited in claim 1, wherein an inlet metal angle of said airfoil of said annular cascade coincides with a direction of extent of said hot gas path.

6. The air intake fairing as recited in claim 5, wherein exit metal angles of the airfoils of said annular cascade are circumferentially offset.

7. The air intake fairing as recited in claim 1, wherein the airfoil of said annular cascade is an airfoil of a turbine blade.

8. The intake fairing of claim 1, wherein said hot gas path has a plurality of annular cascades therein, one of said cascades being disposed in a forward section of said hot gas path and the remaining cascades being disposed in a mid-section and/or an aft section of said hot gas path.

9. A gas turbine engine comprising an air intake fairing as claimed in any one of claims 1 to 8.

10. An aircraft comprising a fuselage, a wing and a gas turbine engine as claimed in claim 9.

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