CN113931882A - Compressor, aircraft engine and aircraft - Google Patents

Compressor, aircraft engine and aircraft Download PDF

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Publication number
CN113931882A
CN113931882A CN202111536564.3A CN202111536564A CN113931882A CN 113931882 A CN113931882 A CN 113931882A CN 202111536564 A CN202111536564 A CN 202111536564A CN 113931882 A CN113931882 A CN 113931882A
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China
Prior art keywords
blade
section
profile
root
tip
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Granted
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CN202111536564.3A
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Chinese (zh)
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CN113931882B (en
Inventor
刘天一
王进春
朱伟
李继保
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AECC Commercial Aircraft Engine Co Ltd
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AECC Commercial Aircraft Engine Co Ltd
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Priority to CN202111536564.3A priority Critical patent/CN113931882B/en
Publication of CN113931882A publication Critical patent/CN113931882A/en
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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
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • 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
    • F02C3/06—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The invention discloses a compressor, an aircraft engine and an aircraft. The compressor comprises stator blades, rotor blades, a casing and a hub, wherein the rotor blades are arranged on the axial rear sides of the stator blades at intervals in the axial direction, and an airflow channel is formed between the stator blades and the rotor blades. The casing is arranged on the radial outer sides of the stator blades and the rotor blades and is provided with a first air-entraining channel communicated with the air flow channel. The hub is arranged on the radial inner sides of the stator blade and the rotor blade and is provided with a second air entraining channel communicated with the air flow channel, wherein the rotor blade comprises a plurality of blade profile sections stacked and arranged from the blade root to the blade tip direction, and the chord lengths of the blade profile sections are gradually increased and then gradually reduced from the blade root to the blade tip. According to the invention, a part of chord length is additionally added upstream from the middle part of the front edge part of the rotor blade, so that the working potential of the middle part of the rotor blade is increased.

Description

Compressor, aircraft engine and aircraft
Technical Field
The invention relates to the technical field of aircrafts, in particular to a compressor, an aero-engine and an aircraft.
Background
The axial flow compressor is a multi-stage compression device with the airflow flowing direction consistent or nearly consistent with the rotating axial lead direction of a working wheel, is formed by correspondingly and alternately arranging a root tip flow passage and a series of stator-rotor blades, and is commonly used for an aircraft engine or a gas turbine. As shown in fig. 1, the combination of adjacent stator blades 1a and rotor blades 2a is referred to as a stage. The first row of stator blades 1a at the inlet of the compressor provides an oncoming flow with a tangential component velocity to the adjacent first row of rotor blades 2a downstream thereof, typically in an angularly adjustable manner.
In axial-flow compressors of gas turbines and aircraft engines, a part of gas is generally extracted from a certain position in the middle of the compressor to meet the requirements of turbine cooling, bearing sealing and the like. The common embodiment is that a bleed air channel is opened at the side of a casing or a hub behind a certain stage of stator blade by punching or the like, and the air in the main flow area is led out.
As shown in fig. 2, a first bleed air passage 3b is provided in the casing between the stator blade 1b and the rotor blade 2b, and a second bleed air passage 4b is provided in the hub between the stator blade 1b and the rotor blade 2 b. In order to make room for the bleed air channel, it is necessary to increase the axial distance between the stator blades 1b and the rotor blades 2b, as shown in fig. 2, compared to the way in which no bleed air channel is provided in fig. 1.
Because the boundary layer can be sucked by the bleed air, if the shape of the bleed air port is designed better, the speed and the pressure of incoming flow in front of the root and the tip of the rotor blade at the downstream of the bleed air port are higher, so that the root tip has stronger acting potential. However, the part of the leaf has no such advantage, and the work potential of the leaf is not different from the conventional situation. Therefore, if the working potential of the root tip of the rotor blade is used, the pressure ratio of the rotor blade is integrally increased, the middle part of the blade is overloaded, and if the root tip pressure ratio is only increased, the radial pressure distribution is uneven, stronger secondary flow is generated, and the overall performance benefit brought by the increased pressure ratio is offset to a certain extent. In the present case, therefore, the pressure ratio of the rotor blades behind the bleed air duct is still suitable for conventional design and should not be increased significantly. Therefore, how to improve the work capacity of the rotor blade behind the bleed air channel is the problem to be solved.
It is important to note here that the statements in this background section merely provide background information related to the present disclosure and may not necessarily constitute prior art.
Disclosure of Invention
The invention provides a gas compressor, an aircraft engine and an aircraft, which are used for improving the working capacity of rotor blades on the axial rear side of a gas-guiding channel.
The invention provides a compressor, which comprises stator blades, rotor blades, a casing and a hub, wherein the rotor blades are arranged on the axial rear sides of the stator blades at intervals in the axial direction, and an airflow channel is formed between the stator blades and the rotor blades. The casing is arranged on the radial outer sides of the stator blades and the rotor blades and is provided with a first air-entraining channel communicated with the air flow channel. The hub is arranged on the radial inner sides of the stator blade and the rotor blade and is provided with a second air entraining channel communicated with the air flow channel, wherein the rotor blade comprises a plurality of blade profile sections stacked and arranged from the blade root to the blade tip direction, and the chord lengths of the blade profile sections are gradually increased and then gradually reduced from the blade root to the blade tip.
In some embodiments, the meridional projection of the rotor blade includes a leading edge line that is a smooth curve.
In some embodiments, the plurality of profiled sections includes a root profiled section and a tip profiled section, an average of a chord length of the root profiled section and a chord length of the tip profiled section is an average chord length, and a maximum chord length of the chord lengths of the plurality of profiled sections is at least 5% greater than the average chord length.
In some embodiments, a maximum chord length of the chord lengths of the plurality of airfoil sections is at least 10% greater than the average chord length.
In some embodiments, the plurality of profile sections includes a root profile section and a tip profile section, the root profile section having a maximum thickness chordwise location less than 0.45 and the tip profile section having a maximum thickness chordwise location greater than 0.5.
In some embodiments, the plurality of profiled sections includes a tip profiled section, the profiled section having a leading edge point and a trailing edge point, a point of maximum distance between a camber line and a chord line of the profiled section being taken along the chord line as a reference point for a foot formed by the tip profiled section, a midpoint between the leading edge point and the reference point on the chord line being a first location, the profile thickness at the first location of the tip profiled section being less than 0.7 the maximum thickness of the tip profiled section, the profile thickness at the first location of the root profiled section being greater than 0.7 the maximum thickness of the root profiled section.
In some embodiments, a midpoint between the trailing edge point and the reference point is a second location on the chord line, the profile thickness at the second location of the tip profile cross section is greater than 0.6 the maximum thickness of the tip profile cross section, and the profile thickness at the second location of the root profile cross section is less than 0.7 the maximum thickness of the root profile cross section.
In some embodiments, the plurality of profile sections includes a root profile section, the pressure face of the root profile section having a local convex region convex to the outboard side.
In some embodiments, the root profile section has a maximum thickness to chord ratio greater than 0.07.
In some embodiments, the thickness of the pressure side is greater than the thickness of the suction side at the maximum thickness of the blade root profile cross-section.
In some embodiments, the localized protrusion region is located between a first point having a distance to chord ratio between the foot and the leading edge point of less than 0.3 and a second point having a distance to chord ratio between the foot and the trailing edge point of less than 0.2.
In some embodiments, the local convex surface transitions smoothly in a direction from the root profile section to the tip profile section.
In some embodiments, the local convex surface tapers from the root to within 25% of the blade height of the rotor blade.
In some embodiments, in a direction extending from the blade root to the blade tip of the rotor blade, a maximum thickness position of the plurality of profile sections between 25% of the blade height of the rotor blade and the blade tip, a ratio of the profile thickness to the maximum thickness at the first position, and a ratio of the profile thickness to the maximum thickness at the second position are the same, and the smooth transition is made from 25% of the blade height of the rotor blade to the maximum thickness position of the plurality of profile sections between the blade root, the ratio of the profile thickness to the maximum thickness at the first position, and the ratio of the profile thickness to the maximum thickness at the second position.
In some embodiments, the plurality of profiled sections includes a tip profiled section having a chord-wise location of maximum camber greater than 0.5.
In a second aspect, the invention provides an aircraft engine comprising the compressor.
A third aspect of the invention provides an aircraft comprising an aircraft engine as described above.
Based on the technical scheme provided by the invention, the gas compressor comprises stator blades, rotor blades, a casing and a hub, wherein the rotor blades are arranged on the axial rear sides of the stator blades at intervals in the axial direction, and an airflow channel is formed between the stator blades and the rotor blades. The casing is arranged on the radial outer sides of the stator blades and the rotor blades and is provided with a first air-entraining channel communicated with the air flow channel. The hub is arranged on the radial inner sides of the stator blade and the rotor blade and is provided with a second air entraining channel communicated with the air flow channel, wherein the rotor blade comprises a plurality of blade profile sections stacked and arranged from the blade root to the blade tip direction, and the chord lengths of the blade profile sections are gradually increased and then gradually reduced from the blade root to the blade tip. According to the invention, a part of chord length is additionally added upstream from the middle part of the front edge part of the rotor blade, so that the working potential of the middle part of the rotor blade is increased.
Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention.
Fig. 1 is a schematic partial structure diagram of a compressor in the prior art.
Fig. 2 is a partial structural schematic diagram of a compressor provided with a bleed air passage in the prior art.
Fig. 3 is a schematic structural view of a blade profile section in the related art.
Fig. 4 is a partial structural view of a compressor according to some embodiments of the present invention.
FIG. 5 is a meridional projection of a rotor blade according to some embodiments of the invention.
FIG. 6 is a graph illustrating the distribution of thickness from the leading edge to the trailing edge of a rotor blade according to some embodiments of the present invention.
FIG. 7 is a graph illustrating the variation in thickness distribution in the blade height direction of a rotor blade according to some embodiments of the present invention.
FIG. 8 is a schematic root profile cross-section of a rotor blade according to some embodiments of the present invention.
Fig. 9 is an enlarged view of a portion of the blade root profile cross-section shown in fig. 8.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The relative arrangement of the components and steps, the numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail, but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.
Spatially relative terms, such as "above … …," "above … …," "above … …," "above," and the like, may be used herein for ease of description to describe one device or feature's spatial relationship to another device or feature as illustrated in the figures. It will be understood that the spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientation depicted in the figures. For example, if a device in the figures is turned over, devices described as "above" or "on" other devices or configurations would then be oriented "below" or "under" the other devices or configurations. Thus, the exemplary term "above … …" can include both an orientation of "above … …" and "below … …". The device may be otherwise variously positioned and the spatially relative descriptors used herein interpreted accordingly.
For ease of understanding, with reference to fig. 3, the terms referred to herein are explained as follows:
chord position: any point on the camber line of the blade-shaped section makes a perpendicular line towards the chord line, and the ratio of the distance between the foot and the front edge point to the chord length is determined.
Maximum thickness: the diameter of the inscribed circle of the profile at the location of maximum thickness of the profile cross section.
Maximum thickness ratio: the ratio of the diameter to the chord length of the inscribed circle of the profile at the maximum thickness position of the profile section is designated as Tmax。
Chordal position of maximum thickness: the chord-wise position of the maximum radius of the inscribed circle of the profile of the blade-shaped cross section is marked as Bmax，BmaxAnd the point T is a foot obtained by drawing a perpendicular line to the chord line at the position with the maximum radius of the inscribed circle of the profile of the blade section.
Chordal position of maximum camber: the chord-wise position of the point of the camber line of the blade-shaped cross-section which is farthest from the chord line is denoted as Cmax，CmaxAnd the point E is the foot obtained by drawing a perpendicular line to the chord line from the point which is farthest relative to the chord line on the camber line of the blade-shaped section.
For the convenience of the following description, on the basis of this, on a chord line, a first position E is defined1And a second position E2Respectively satisfy the relation CE1= CE/2 and DE2And (d) = DE/2. Will be at the first position E1And a second position E2The ratio of the thickness of the blade profile to the chord length is respectively recorded as a first thickness ratio TE1And a second thickness ratio TE2。
The problem to be solved is how to improve the working capacity of the rotor blade behind the air guide channel. As shown in fig. 4, by utilizing the condition that the axial gap between the stator blade 1 and the rotor blade 2 at the bleed air passage is large, a part of chord length is additionally increased from the front edge part of the rotor blade 2 to the upstream, so that the work potential of the middle part of the rotor blade 2 can be increased, and the pressure ratio in the root tip and the blade of the rotor blade can be simultaneously designed to be higher.
As shown in fig. 4, an embodiment of the present invention provides a compressor. The compressor comprises stator blades 1, rotor blades 2, a casing and a hub, wherein the rotor blades 2 are arranged on the axial rear side of the stator blades 1 at intervals in the axial direction, and an airflow channel is formed between the stator blades 1 and the rotor blades 2. The casing is arranged radially outside the stator blades 1 and the rotor blades 2 and has a first bleed air channel 3 communicating with the air flow channel. The hub is arranged radially inside the stator blades 1 and the rotor blades 2 and has a second bleed air channel 4 communicating with the air flow channel. The rotor blade 2 includes a plurality of blade profile sections stacked in a direction from a blade root to a blade tip, and chord lengths of the blade profile sections gradually increase and then gradually decrease in the direction from the blade root to the blade tip.
In the above solution of the embodiment of the present invention, in the direction from the blade root to the blade tip, the chord lengths of the blade-shaped cross sections gradually increase and then gradually decrease, that is, a part of the chord length is additionally added upstream from the middle of the leading edge of the rotor blade 2, so as to increase the work potential in the middle of the rotor blade 2, and on this basis, the pressure ratio between the blade tip and the blade tip of the rotor blade 2 may also be designed to be higher.
In some embodiments, referring to fig. 5, a meridional projection of the rotor blade 2 includes a leading edge line 21, the leading edge line 21 being a smooth curve. That is, the chord length of the blade profile section changes smoothly from the blade root to the blade tip.
Referring to fig. 5, X is an axial direction, and a meridional projection of the rotor blade 2 includes a leading edge line 21 and a trailing edge line 22. The midpoint of the connecting line of the leading edge point and the trailing edge point of the blade tip is marked as A; the midpoint of the line connecting the leading edge point and the trailing edge point of the blade root is marked as B; the midpoint of the line between the leading edge point and the trailing edge point at 50% of the leaf height is denoted as M, where M is located upstream of AB.
Specifically, as shown in fig. 5, in the meridional projection of the rotor blade 2, the line connecting the midpoint a of the line connecting the leading edge point and the trailing edge point of the blade tip and the midpoint B of the line connecting the leading edge point and the trailing edge point of the blade root is approximately perpendicular to the line connecting the leading edge point and the trailing edge point at 50% of the blade height.
In some embodiments, the plurality of airfoil sections includes a root airfoil section and a tip airfoil section. The average of the chord lengths of the blade root profile sections and the blade tip profile sections is an average chord length, and the maximum chord length of the chord lengths of the blade root profile sections is at least 5% longer than the average chord length.
Specifically, for example, in the middle of the blade, the blade height = hmAt a maximum value that is at least 5% greater than the average chord length of the root and tip of the blade.
In some embodiments, a maximum chord length of the chord lengths of the plurality of airfoil sections is at least 10% greater than the average chord length.
Also in some embodiments, from blade height =0, i.e. the position of the blade root, to blade height = hmThe chord length is monotonously increased; from leaf height = hmBy leaf height =1 tip position, the chord length monotonically decreases. The chord length distribution law here may be linear or any nonlinear law.
The inventor finds that the working potential of the rotor blade 2 can be increased by increasing the chord length of the middle part of the rotor blade 2, but because the sweepforward degree of the rotor blade is high, the rotor blade has a strong tendency of inclining backwards during operation due to centrifugal force, so that the tensile stress of the front edge of the root part of the rotor blade is very large, and the arrangement may cause the rotor blade 2 to be subjected to overlarge stress during operation. And the increase of the chord length in the middle of the blade also leads to the increase of the weight, and further leads to the increase of the integral centrifugal stress of the root part of the blade.
In order to alleviate the tendency of backward tilting of the rotor blade during operation, as shown in fig. 6, the thicknesses of the sections of the blade root and the blade tip of the rotor blade are designed to be distributed along the chord direction, and such a thickness distribution rule will cause the center of gravity of the blade root to move forward and the center of gravity of the blade tip to move backward.
Specifically, in some embodiments, the plurality of profile sections includes a root profile section and a tip profile section, the root profile section having a maximum thickness chordwise location less than 0.45 and the tip profile section having a maximum thickness chordwise location greater than 0.5.
The chord-wise position of the maximum thickness of the blade root profile section is closer to the front edge, and the chord-wise position of the maximum thickness of the blade tip profile section is closer to the rear edge, so that the center of gravity of the blade root moves forwards, the center of gravity of the blade tip moves backwards, and the tendency that the rotor blade tilts backwards in operation is relieved.
Still further, in some embodiments, the maximum thickness chordwise location of the root profile section is less than 0.45 and the maximum thickness chordwise location of the tip profile section is greater than 0.5.
In some embodiments, the plurality of profiled sections includes a tip profiled section having a leading edge point and a trailing edge point D, and a distance between a mean camber line and a chord line CD of the profiled sectionThe foot formed by the blade root profile section with the largest point perpendicular to the chord line CD is the reference point E, and the midpoint between the leading edge point C and the reference point E on the chord line CD is the first position E1First position E of blade tip profile section1The blade profile thickness of (a) is less than 0.7 of the maximum thickness of the blade tip profile cross section, and the first position E of the blade root profile cross section1The profile thickness of (a) is greater than 0.7 of the maximum thickness of the blade root profile cross section.
In some embodiments, the midpoint between the trailing point and the reference point on the chord line is the second location E2Second position E of blade tip profile section2The blade profile thickness is greater than 0.6 of the maximum thickness of the blade tip profile section and the second position E of the blade root profile section2The profile thickness of (a) is less than 0.7 of the maximum thickness of the blade root profile cross section.
The blade tip is arranged in such a way that the thickness of the blade tip is smaller than that of the blade root in the area close to the front edge, but the thickness of the blade tip is larger than that of the blade root in the area close to the rear edge, namely, the chordwise position of the maximum thickness of the blade-root profile section is closer to the front edge, and the chordwise position of the maximum thickness of the blade-tip profile section is closer to the rear edge, so that the gravity center of the blade root moves forwards, the gravity center of the blade tip moves backwards, and the tendency that the rotor blade tilts backwards in operation is relieved.
Further, in other embodiments, the first location E of the tip profile cross-section1The blade profile thickness of (a) is less than 0.6 of the maximum thickness of the blade tip profile cross section, and the first position E of the blade root profile cross section1The profile thickness of (a) is greater than 0.8 of the maximum thickness of the blade root profile cross section.
In other embodiments, the second position E of the blade tip profile section2The blade profile thickness is greater than 0.7 of the maximum thickness of the blade tip profile section and the second position E of the blade root profile section2The profile thickness of (a) is less than 0.6 of the maximum thickness of the blade root profile cross section.
In the above embodiments, only the thicknesses of the tip profile section and the root profile section are described, and the thicknesses of the plurality of blade profile sections between the tip profile section and the root profile section are not described, the thicknesses of the blade profile sections are described below,
as shown in FIG. 7, in some embodiments, the maximum thickness position, the first position E, of the plurality of blade profile sections between the blade height 25% of the rotor blade 2 and the blade tip in the direction extending from the blade root to the blade tip of the rotor blade 21And a second position E2Has the same ratio of profile thickness to maximum thickness, from 25% of the blade height of the rotor blade 2 to the position of maximum thickness of the profile sections between the blade root, the first position E1And a second position E2The ratio of the thickness of the blade profile to the maximum thickness of (a) is smoothly transited. That is, the maximum thickness, first position E, below 25% of the leaf height1And a second position E2The ratio of the blade profile thickness to the maximum thickness is smoothly transited to be consistent with the position of the blade root profile section. A smooth transition means that the derivative of this parameter with respect to the leaf height is continuous.
Preferably, in other embodiments, the maximum thickness position, first position E, is above 15% of the leaf height1And a second position E2The ratio of the thickness of the blade profile to the maximum thickness is the same; at the maximum thickness position and the first position E below 15% of the leaf height1And a second position E2The ratio of the thickness of the blade profile to the maximum thickness of (a) is smoothly transited.
In the above embodiments, the forward sweep degree of the gravity stacking axis of the rotor blade is weakened by setting different thickness distribution characteristics for the blade profile sections at different blade heights.
In some embodiments, referring to fig. 8 and 9, the plurality of profile sections includes a root profile section. The pressure surface of the blade root profile section has a local bulge area bulging outward. The blades are locally thickened in a mode that the pressure surface is locally convex, so that the shape of the gravity center stacking shaft is adjusted, the shape of the suction surface is kept unchanged, and better pneumatic performance is maintained.
In some embodiments, the ratio of the maximum thickness to the chord length of the blade root profile sectionTmaxGreater than 0.07. The local bulge of the blade root profile cross-section forms the maximum thickness of the blade root profile cross-section. As shown in fig. 8, the blade root profile cross-section is bounded by the solid line profile of the suction surface and the dashed line profile of the pressure surface therein.
As shown in FIG. 9, in some embodiments, at the maximum thickness of the root profile cross-section, the thickness of the pressure side is greater than the thickness of the suction side. Specifically, at the maximum thickness position, the airfoil thickness of the suction side =0.035 chord length, and the airfoil thickness of the pressure side = Tmax-0.035 chord length, due to TmaxGreater than 0.07, the pressure side profile thickness is greater than the suction side profile thickness.
As shown in fig. 8, in some embodiments, the localized salient region is located between a first point P1 and a second point P2, the distance to chord ratio between the foot and the leading edge point of the first point P1 is less than 0.3, and the distance to chord ratio between the foot and the trailing edge point of the second point P2 is less than 0.2.
In some embodiments, the local convex surface transitions smoothly in a direction from the root profile section to the tip profile section.
In some embodiments, the local convex surface tapers from the root to within 20% of the blade height of the rotor blade 2. Specifically, the blade root profile section and the profile sections of other parts of the blade are in smooth transition, and the transition area is limited to be below 20% of the blade height. By smooth is meant that the blade profile is smooth with each first derivative.
In some embodiments, the first position E, the maximum thickness position of the profile cross-section in the direction extending from the root to the tip of the rotor blade 2, is a plurality of positions between 25% of the blade height of the rotor blade 2 and the tip1And a second position E2Has the same ratio of profile thickness to maximum thickness, from 25% of the blade height of the rotor blade 2 to the position of maximum thickness of the profile sections between the blade root, the first position E1And a second position E2The ratio of the thickness of the blade profile to the maximum thickness of (a) is smoothly transited.
In some embodiments, a plurality of profilesThe section comprises a blade tip profile section, and the chord-wise position C of the maximum camber of the blade tip profile sectionmaxGreater than 0.5. Therefore, the blade tip is loaded backwards, the strength characteristic is improved, and the leakage flow of the blade tip clearance can be weakened.
The embodiment of the invention also provides an aircraft engine which comprises the air compressor.
The embodiment of the invention also provides an aircraft comprising the aero-engine.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention and not to limit it; although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art will understand that: modifications to the specific embodiments of the invention or equivalent substitutions for parts of the technical features may be made; without departing from the spirit of the present invention, it is intended to cover all aspects of the invention as defined by the appended claims.

Claims (17)

1. A compressor, comprising:
a stator blade (1);
the rotor blades (2) are arranged on the axial rear side of the stator blade (1) at intervals in the axial direction, and an airflow channel is formed between the stator blade (1) and the rotor blades (2);
the casing is arranged on the radial outer sides of the stator blade (1) and the rotor blade (2) and is provided with a first air-bleed channel (3) communicated with the airflow channel; and
the hub is arranged on the radial inner sides of the stator blades (1) and the rotor blades (2) and is provided with a second air-entraining channel (4) communicated with the airflow channel;
the rotor blade (2) comprises a plurality of blade profile sections which are stacked in the direction from the blade root to the blade tip, and the chord lengths of the blade profile sections are gradually increased and then gradually reduced in the direction from the blade root to the blade tip.
2. Compressor according to claim 1, characterized in that the meridional projection of the rotor blade (2) comprises a leading edge line (21), the leading edge line (21) being a smooth curve.
3. The compressor of claim 1, wherein the plurality of airfoil sections includes a root airfoil section and a tip airfoil section, wherein an average of a chord length of the root airfoil section and a chord length of the tip airfoil section is an average chord length, and wherein a maximum chord length of the chord lengths of the plurality of airfoil sections is at least 5% greater than the average chord length.
4. The compressor of claim 3, wherein a maximum chord length of the chord lengths of the plurality of airfoil sections is at least 10% greater than the average chord length.
5. The compressor of claim 1 wherein the plurality of airfoil sections include a root airfoil section and a tip airfoil section, the root airfoil section having a maximum thickness chordal position less than 0.45 and the tip airfoil section having a maximum thickness chordal position greater than 0.5.
6. The compressor of claim 1, wherein the plurality of airfoil sections includes a tip airfoil section having a leading edge point (C) and a trailing edge point (D), wherein a point of maximum distance between a mean camber line and a chord line (CD) of the airfoil section is perpendicular to the chord line (CD) and a foot formed by the tip airfoil section is a reference point (E), and wherein a midpoint between the leading edge point (C) and the reference point (E) on the chord line (CD) is a first position (E)1) First position (E) of said blade tip profile section1) Is less than 0.7 of the maximum thickness of the blade tip profile section, and a first position of the blade root profile section (E)1) Is greater than 0.7 of the maximum thickness of the blade root profile cross section.
7. An air compressor according to claim 6,on the chord line, the midpoint between the trailing edge point and the reference point is a second position (E)2) Second position (E) of said blade tip profile section2) Is greater than 0.6 of the maximum thickness of the blade tip profile section, and a second position of the blade root profile section (E)2) Is less than 0.7 of the maximum thickness of the blade root profile cross section.
8. The compressor of claim 1 wherein the plurality of airfoil sections includes a root airfoil section, the pressure face of the root airfoil section having a local convex region projecting outboard.
9. The compressor of claim 8 wherein the root profile section has a maximum thickness to chord ratio of greater than 0.07.
10. The compressor of claim 9 wherein the thickness of the pressure side is greater than the thickness of the suction side at the maximum thickness of the blade root profile section.
11. Compressor according to claim 8, characterised in that said local convex zone is located at a first point (P)1) And a second point (P)2) In said first point (P)1) Is less than 0.3, the second point (P) is located at a distance from the foot to the leading edge point2) The ratio of the distance between the foot and the tail edge point to the chord length is less than 0.2.
12. The compressor as set forth in claim 8 wherein the local convex surface transitions smoothly in a direction from the root profile section to the tip profile section.
13. Compressor according to claim 12, characterized in that the local convex surface tapers from the root to within 25% of the blade height of the rotor blade (2).
14. Compressor according to any one of claims 1 to 13, characterised in that the first position (E), the maximum thickness position of the profile cross-section in a plurality of sections between the blade tip and the height of the rotor blade (2), in the direction extending from the blade root to the blade tip of the rotor blade (2)1) And a second position (E)2) Has the same ratio of profile thickness to maximum thickness, from 25% of the blade height of the rotor blade (2) to the first position (E), the maximum thickness position of the profile sections1) And a second position (E)2) The ratio of the thickness of the blade profile to the maximum thickness of (a) is smoothly transited.
15. The compressor of claim 1 to wherein the plurality of airfoil sections includes a tip airfoil section having a chord wise location of maximum camber greater than 0.5.
16. An aircraft engine, characterised in that it comprises a compressor as claimed in any one of claims 1 to 15.
17. An aircraft comprising an aircraft engine according to claim 16.
CN202111536564.3A 2021-12-16 2021-12-16 Compressor, aircraft engine and aircraft Active CN113931882B (en)

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