CN113958537A - Compressor and aircraft engine - Google Patents
Compressor and aircraft engine Download PDFInfo
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- CN113958537A CN113958537A CN202111536740.3A CN202111536740A CN113958537A CN 113958537 A CN113958537 A CN 113958537A CN 202111536740 A CN202111536740 A CN 202111536740A CN 113958537 A CN113958537 A CN 113958537A
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- 230000001154 acute effect Effects 0.000 claims abstract description 7
- 238000011144 upstream manufacturing Methods 0.000 claims description 10
- 230000007704 transition Effects 0.000 claims description 5
- 230000005484 gravity Effects 0.000 description 7
- 230000007423 decrease Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000010408 sweeping Methods 0.000 description 2
- 230000004323 axial length Effects 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 238000001816 cooling Methods 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 230000006866 deterioration Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
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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
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- 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
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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
- F04D29/26—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
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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
- F04D29/26—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
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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
- F04D29/522—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 relates to a compressor and an aero-engine, wherein the compressor comprises a stator blade, a rotor blade, a casing and a gas-guiding channel, the rotor blade is arranged adjacent to the stator blade, a preset distance is reserved between the stator blade and the rotor blade to form a gas flow channel, the casing is arranged at the radial outer side of the stator blade and the rotor blade, the gas-guiding channel is communicated with the gas flow channel and extends from the gas flow channel to the direction close to the casing, an included angle between a first straight line and the axis of the compressor is an acute angle, the first straight line is a connecting line between a first middle point and a second middle point, the first middle point is a middle point of a connecting line between a first leading edge point and a first trailing edge point on a first blade profile section of the rotor blade at the blade tip, the second middle point is a middle point of a connecting line between a second leading edge point and a second trailing edge point on a second blade profile section of the rotor blade at the blade root, the chord length of the rotor blade is monotonous and is reduced from the root to the tip, the rate of change of the chord length of the rotor blade in the blade height direction is variable.
Description
Technical Field
The invention relates to the technical field of aero-engines, in particular to a gas compressor and an aero-engine.
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, and as shown in figure 1, the combination of adjacent stator blades 1a and rotor blades 2a is called as a stage. The first row of stator blades at the compressor inlet provides an oncoming flow with a tangential component velocity to the adjacent first row of rotor blades downstream thereof, typically in an angularly adjustable fashion.
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. In a typical embodiment, a bleed air passage is formed by punching holes or the like in the casing or hub side after a certain stage of stationary blade to lead out the gas in the main flow region.
As shown in fig. 2, a bleed air passage 3b is provided on the casing side between the stator blade 1b and the rotor blade 2b, and in order to make room for the bleed air passage 3b, it is necessary to increase the axial distance between the blades in front and rear of the bleed air port. As shown in fig. 1 and 2, after the bleed air passage 3b is opened, the axial distance between the stator blade 1b and the rotor blade 2b is significantly greater than the axial distance between the stator blade 1a and the rotor blade 2 a.
Because the upstream boundary layer can be sucked by the bleed air at the casing side, if the shape design of the bleed air port is good, the speed and the pressure of the incoming flow in front of the blade tip part of the movable blade at the downstream of the bleed air port are higher, and the boundary layer is very thin, so that the blade tip of the movable blade has stronger work potential compared with the scheme without the bleed air port. However, the part below the blade tip of the movable blade has no such advantages, the acting potential of the movable blade is not different from the scheme without the air bleed port, the acting potential of the whole movable blade cannot meet the requirement, the radial pressure distribution of the movable blade is uneven, stronger secondary flow is generated, and the loss is increased.
It is noted that the information disclosed in this background section is only for enhancement of understanding of the general background of the invention and should not be taken as an acknowledgement or any form of suggestion that this information constitutes prior art already known to a person skilled in the art.
Disclosure of Invention
The embodiment of the invention provides a gas compressor and an aero-engine, and solves the problem that the overall acting potential of a rotor blade in the related technology cannot meet the requirement.
According to an aspect of the present invention, there is provided a compressor comprising:
a stator blade;
the rotor blades are arranged adjacent to the stator blades, and a preset distance is reserved between the stator blades and the rotor blades to form an airflow channel;
the casing is arranged on the radial outer sides of the stator blades and the rotor blades; and
the bleed air channel is communicated with the airflow channel and extends from the airflow channel to the direction close to the casing;
the included angle between the first straight line and the axial line of the compressor is an acute angle, the first straight line is a connecting line between a first middle point and a second middle point, the first middle point is the middle point of the connecting line between a first leading edge point and a first trailing edge point on a first blade profile section of the rotor blade at the blade tip, the second middle point is the middle point of the connecting line between the second leading edge point and a second trailing edge point on a second blade profile section of the rotor blade at the blade root, the chord length of the rotor blade is monotonically reduced from the root to the tip, and the change rate of the chord length of the rotor blade along the blade height direction is variable.
In some embodiments, the rotor blade includes a first portion and a second portion located above the first portion in the blade height direction, a rate of change of a chord length of the first portion in the blade height direction is from small to large, and a rate of change of a chord length of the second portion in the blade height direction is from large to small.
In some embodiments, the airfoil section of the rotor blade at the tip thereof has a first chord length, the airfoil section of the rotor blade at the root thereof has a second chord length, the first chord length is greater than the second chord length by X1%, and X1% ranges from 5% to 10%.
In some embodiments, the rotor blade has a projection onto the meridian plane having a first leading edge line, a line connecting the first leading edge point and the second leading edge point being a second line, the first leading edge line being located downstream of the second line near the tip of the rotor blade; the first leading edge line is located upstream of the second line near the root of the rotor blade.
In some embodiments, the projection of the rotor blade onto the meridian plane has a first leading edge line, a line connecting the first leading edge point and the second leading edge point is a second straight line, and the first leading edge line and the second straight line have a unique intersection point between the first profile section and the second profile section.
In some embodiments, the chordal position of maximum camber CmaxThe ratio of a first distance to the length of the chord line of the preset blade profile, the first distance being the distance between a foot perpendicular to the chord line of the preset blade profile from the position of the camber line of the preset blade profile farthest relative to the chord line and the leading edge point of the preset blade profile, and the chord direction position C of the maximum camber of the section of the first blade profile1maxSatisfies the following conditions: c1max>X2, wherein the value range of X2 is 0.5-0.55; chord-wise position C of maximum camber of the second-leaf section1maxSatisfies the following conditions: c1max<The value range of X3 and X3 is 0.4-0.45.
In some embodiments, the chordwise location B of maximum thicknessmaxThe ratio of the second distance to the length of the chord line of the preset profile, the second distance being the distance between the foot perpendicular to the chord line of the preset profile from the position with the maximum radius of the inscribed circle of the preset profile and the leading edge point of the preset profile, and the chord position B with the maximum thickness of the section of the first profile1maxSatisfies the following conditions: b is1max<Taking X4, X4The value range is 0.4-0.45; chord position B of maximum thickness of second leaf-shaped cross section1maxSatisfies the following conditions: b is1max>The value range of X5 and X5 is 0.5-0.55.
In some embodiments, the thickness TmaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A midpoint of a connecting line between a foot obtained by drawing a perpendicular line from a position farthest to a chord line on a mean camber line of the preset blade profile and the chord line of the preset blade profile and a leading edge point of the preset blade profile is a first position E on the section of the first blade profile1Thickness T ofE1Satisfies the following conditions: t isE1>TmaxThe value range of X6 and X6 is 0.7-0.8; in the second leaf-shaped cross section, at a first position E1Thickness T ofE1Satisfies the following conditions: t isE1<TmaxAnd X7, wherein the value range of X7 is 0.6-0.7.
In some embodiments, the thickness TmaxIn a second position E which is the maximum diameter of the inscribed circle of the predetermined blade profile2A midpoint of a connecting line between a foot perpendicular to the chord line of the preset blade profile and a tail edge point of the preset blade profile, wherein the foot perpendicular to the chord line of the preset blade profile is taken from a position farthest relative to the chord line on a mean camber line of the preset blade profile, and a second position E is arranged on the section of the first blade profile2Thickness T ofE2Satisfies the following conditions: t isE2<TmaxX8, wherein the value range of X8 is 0.6-0.7; in a second leaf-shaped cross-section, at a second position E2Thickness T ofE2Satisfies the following conditions: t isE2>TmaxAnd X9, wherein the value range of X9 is 0.6-0.7.
In some embodiments, the chordwise location B of maximum thicknessmaxThe ratio of the second distance to the chord length of the preset profile, the second distance being the distance between the foot perpendicular to the chord of the preset profile and the leading edge point of the preset profile, and the thickness T being the distance between the foot perpendicular to the chord of the preset profile from the position with the maximum radius of the inscribed circle of the preset profile and the leading edge point of the preset profilemaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A second position E is the midpoint of a connecting line between a foot, which is obtained by drawing a perpendicular line from the farthest position relative to the chord line on the mean camber line of the preset blade profile to the chord line of the preset blade profile, and the leading edge point of the preset blade profile2Is arranged from the mean camber line of the preset blade profile relative to the chordThe middle point of a connecting line between a foot obtained by drawing a perpendicular line from the farthest position of the line to the chord line of the preset blade profile and the tail edge point of the preset blade profile is the chord position B with the maximum thickness on the blade profile section with the blade height of more than X10 percentmaxFirst position E1Thickness T ofE1And thickness TmaxAnd a second position E2Thickness TE2And thickness TmaxThe ratio values are all equal in size; chord-wise position B of maximum thickness on the blade profile cross section below X10% blade heightmaxFirst position E1Thickness T ofE1And thickness TmaxA ratio of (A), a second position E2Thickness TE2And thickness TmaxThe values of the ratios are respectively in smooth transition to the value of the position of the blade root, the derivative of the value of the smooth transition section relative to the blade height of the rotor blade is continuous, and the value range of X10% is 15% -25%.
According to another aspect of the invention, an aircraft engine is provided, comprising the compressor.
Based on the technical scheme, an included angle between a first straight line of the rotor blade and an axis of the compressor is an acute angle, the chord length of the rotor blade is monotonically reduced from the root to the tip, and the change rate of the chord length of the rotor blade along the blade height direction is variable, so that the chord length of the rotor blade can be increased upstream by using the axial distance brought by the air bleed channel, the acting potential of the middle part and the root of the rotor blade is enhanced, the pressure ratio of the rotor blade behind air bleed is improved, the strength risk of the tail edge of the blade root can be reduced by setting the change of the chord length of the rotor blade, and the stress condition of the rotor blade is optimized.
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 view of a partial structure of a compressor in the related art.
Fig. 2 is a partial structural view of another compressor in the related art.
Fig. 3 is a schematic view of a partial structure of another compressor in the related art.
Fig. 4 is a partial structural schematic diagram of an embodiment of the compressor of the present invention.
Fig. 5 is a schematic structural diagram of a preset blade profile of an embodiment of the compressor of the present invention.
Fig. 6 is a schematic view of a projection of a rotor blade onto a meridian plane in an embodiment of the compressor according to the invention.
Fig. 7 is a schematic view of the variation of the thickness of the blade profile section in the axial direction at the blade tip and the blade root of the rotor blade in an embodiment of the compressor of the present invention.
FIG. 8 is a thickness T of a rotor blade at a first location in an embodiment of a compressor according to the present inventionE1And thickness TmaxAnd the thickness T at the second positionE2And thickness TmaxThe ratio of (a) to (b) is shown schematically in the direction of the leaf height.
FIG. 9 is a schematic representation of the variation in the height of the chord-wise location of the maximum thickness of the rotor blade in one embodiment of the compressor of the present invention.
In the figure: 1a, stator blades; 2a, rotor blades; 1b, stator blades; 2b, rotor blades; 3b, a gas-guiding channel; 1c, stator blades; 2c, rotor blades; 3c, a gas-guiding channel; 1. a stator blade; 2. a rotor blade; 3. a bleed air passage; 4. a first leading edge line; 5. a first trailing edge line; 6. and (4) a projection surface.
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. It is to be understood that the described embodiments are merely a few embodiments of the invention, and not all embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments of the present invention without making any creative effort, shall fall within the protection scope of the present invention.
In the description of the present invention, it is to be understood that the terms "central," "lateral," "longitudinal," "front," "rear," "left," "right," "upper," "lower," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like are used in the orientation or positional relationship indicated in the drawings for convenience in describing the invention and for simplicity in description, and are not intended to indicate or imply that the referenced device or element must have a particular orientation, be constructed and operated in a particular orientation, and are therefore not to be considered limiting of the scope of the invention.
For ease of understanding, reference is made to fig. 5, where the terms referred to herein are explained as follows:
chord position: any point on the element blade-shaped camber line is perpendicular to the chord line, and the distance between the foot and the front edge is divided by the chord length.
Chordal position of maximum thickness: the chord-wise position of the maximum radius of the inscribed circle of the primitive leaf profile is recorded as Bmax，BmaxAnd the T point 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 basic leaf profile.
Chordal position of maximum camber: the chord-wise position of the point on the camber line of the elementary leaf-shaped camber line which is farthest from the chord line is denoted as Cmax，CmaxAnd the T point 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 elementary leaf profile.
Maximum thickness: the diameter of the inscribed circle of the leaf profile at the maximum thickness position is designated as Tmax。
For the convenience of the subsequent description, point E is defined on the basis of this1And E2Point E1Is located at a first position, E2The position is a second position which respectively satisfies the relation CE1= CE/2 and DE2And (d) = DE/2. Will be at the first position E1And a second position E2The thickness of the profile at each position is denoted TE1And TE2。
Aiming at the problem that the overall acting potential of the rotor blade cannot meet the requirement, after a great deal of research, the inventor finds that if the acting potential of the blade tip of the rotor blade is used, the pressure ratio of the rotor blade is integrally increased, and the parts in the blade root and the blade are overloaded; if only the blade tip pressure ratio is increased, the radial pressure distribution is uneven, stronger secondary flow is generated, and the loss is increased. Therefore, the pressure ratio of the rotor blade behind the bleed air is not increased particularly.
On the basis of the related art shown in fig. 2, a condition that an axial gap between a moving blade and a stationary blade at a bleed air passage is large is utilized, a part of chord length is additionally increased from the front edge of the rotor blade to the upstream, so that the rotor blade is in a structural form shown in fig. 3, the bleed air passage 3c is arranged on the side of a casing between the stator blade 1c and the rotor blade 2c, and the front edge line of the rotor blade 2c inclines towards the direction close to the stator blade 1c, so that the work potential of the middle part and the root part of the rotor blade 2c can be enhanced, and the overall pressure ratio of the rotor blade 2c can be designed to be higher. However, this may cause the rotor blade 2c to experience excessive stresses during operation. Because such rotor blades 2c are generally swept back, centrifugal forces tend to cause the rotor blades 2c to pitch forward during operation; moreover, since the rotor blade 2c pressurizes the gas at this time, the pressure is higher downstream than upstream of the rotor blade 2c, so the aerodynamic load also tends to tilt the rotor blade 2c forward. The superposition of the two results in a high stress at the trailing edge of the blade root of the rotor blade 2c, which brings a risk of strength. For this reason, it is necessary to make the trailing edge of the root of the rotor blade 2c extra thick, which in turn deteriorates aerodynamic performance, so that the benefits of increasing the chord length of the middle and lower portions of the blade are offset.
The centers of gravity of all elementary sections of the rotor blade are connected to form a spatial curve, which is referred to as the center-of-gravity stacking axis. Considering the sensitive influence of the stacked shaft shape on the strength performance of the rotor blade, the rotor blade is not suitable to be designed into a pure backswept form.
Based on the above research, the inventors further modified the structure of the rotor blade.
As shown in fig. 4 to 6, in some embodiments of the compressor provided by the present invention, the compressor includes a stator blade 1, a rotor blade 2, a casing, and a bleed air passage 3, the rotor blade 2 is disposed adjacent to the stator blade 1, and the stator blade 1 and the rotor blade 2 have a predetermined distance therebetween to form an air flow passage, the casing is disposed radially outside the stator blade 1 and the rotor blade 2, and the bleed air passage 3 is communicated with the air flow passage and extends from the air flow passage in a direction close to the casing.
An included angle between a first straight line and an axis of the compressor is an acute angle, the first straight line is a connecting line between a first middle point and a second middle point, the first middle point is a middle point of a connecting line between a first leading edge point and a first trailing edge point on a first blade profile section of the rotor blade 2 at a blade tip, the second middle point is a middle point of a connecting line between a second leading edge point and a second trailing edge point on a second blade profile section of the rotor blade 2 at a blade root, the chord length of the rotor blade 2 is monotonically reduced from the root to the tip, and the change rate of the chord length of the rotor blade 2 along the blade height direction is variable.
As shown in fig. 4, in the meridian projection plane, the connecting line of the leading edge point and the trailing edge point of the blade tip is a midpoint a, the connecting line midpoint of the leading edge point and the trailing edge point of the blade root is a midpoint B, an included angle ABZ between the straight line AB and the axial direction of the compressor is an acute angle, and the included angle ABZ is greater than 0 ° and less than 90 °.
In the above embodiment, the included angle between the first straight line of the rotor blade 2 and the axis of the compressor is an acute angle, and the chord length of the rotor blade 2 is monotonically decreased from the root portion to the tip portion, and the chord length of the rotor blade 2 is increased upstream by using the axial distance brought by the air-entraining channel, so that the pressure ratio of the rotor blade 2 after air-entraining is improved, which is helpful for improving the overall acting potential of the rotor blade 2, rather than only improving the acting potential at the tip portion, and can meet more demands.
Moreover, the change rate of the chord length of the rotor blade 2 along the blade height direction is set as a variable, so that the strength risk of the trailing edge of the blade root is reduced, and the stress condition of the rotor blade 2 is optimized; the radial pressure distribution of the rotor blade 2 is more uniform, stronger secondary flow is avoided, and aerodynamic loss is reduced.
According to the embodiment of the invention, by setting the chord length of the rotor blade 2, the acting potential of the middle part and the root part of the rotor blade 2 can be enhanced by increasing the chord length, the strength risk of the tail edge of the blade root can be avoided, the thickness of the tail edge does not need to be increased in order to overcome the strength problem of the tail edge of the blade root, and the deterioration of the aerodynamic performance is avoided.
In the exemplary embodiment provided by the invention, the stator blades 1 and the rotor blades 2 are not provided with bleed air ports on the side close to the hub, which has a lower influence on aerodynamic losses.
In the embodiment provided by the invention, the stator blade 1 and the rotor blade 2 can be any one of a plurality of sets of adjacently arranged stator blades 1 and rotor blades 2 in the compressor, and the stator blade 1 is positioned at the upstream of the rotor blade 2 and is arranged adjacent to the rotor blade 2.
In some embodiments, the chord length of the rotor blade 2 decreases monotonically as a preset linear function from its root to its tip. In other embodiments, the chord length of the rotor blade 2 decreases monotonically from its root to its tip as a predetermined nonlinear function.
In some embodiments, the rotor blade 2 includes a first portion and a second portion located above the first portion in the blade height direction, the rate of change of the chord length of the first portion in the blade height direction is from small to large, and the rate of change of the chord length of the second portion in the blade height direction is from large to small. The setting can improve the working capacity and reduce the intensity risk caused by the sweepback of the blade profile through the change of the chord length.
In some embodiments, the airfoil section of the rotor blade 2 at the tip thereof has a first chord length, and the airfoil section of the rotor blade 2 at the root thereof has a second chord length, and the first chord length is greater than the second chord length by X1%, and X1% has a value in the range of 5% to 10%, such as 5%, 6%, 7%, 8%, 9%, or 10%.
The first chord length is set to be larger than the second chord length by X1%, and the value range of X1% is controlled within 5% -10%, so that the working potential of the middle part and the root part of the rotor blade 2 can be improved, the overall working capacity of the rotor blade 2 is improved, the radial pressure distribution of the rotor blade 2 from the blade root to the blade tip tends to be uniform, and the pneumatic performance is improved.
As shown in fig. 6, in some embodiments, a projection plane 6 of the rotor blade 2 onto the meridian plane has a first leading edge line 4 and a first trailing edge line 5, a line connecting a first leading edge point on a first blade profile section of the blade tip and a second leading edge point on a second blade profile section of the blade root is a second straight line, and the first leading edge line 4 is located downstream of the second straight line near the blade tip of the rotor blade 2; near the root of the rotor blade 2, the first leading edge line 4 is located upstream of the second straight line. The arrangement can lead the whole rotor blade 2 to be in a structural form of sweeping forward and then sweeping backward from the blade root to the blade tip, can weaken the sweeping-backward degree of the gravity center stacking shaft of the rotor blade 2, is beneficial to improving the balance of the whole stress, and reduces the intensity risk caused by the fact that the blade type is wholly swept backward.
In some embodiments, the projection of the rotor blade 2 onto the meridian plane has a first leading edge line 4, a line connecting the first leading edge point and the second leading edge point is a second straight line, and the first leading edge line 4 and the second straight line have a unique intersection point between the first profile cross-section and the second profile cross-section. The arrangement can ensure that the overall change trend of the rotor blade 2 is forward sweep before backward sweep, avoid the influence on the stability of the airflow caused by too much change, reduce the strength risk of the rotor blade 2 and reduce the aerodynamic loss as much as possible.
In some embodiments, the chordal position of maximum camber CmaxThe ratio of a first distance to the length of the chord line of the preset blade profile, the first distance being the distance between a foot perpendicular to the chord line of the preset blade profile from the position of the camber line of the preset blade profile farthest relative to the chord line and the leading edge point of the preset blade profile, and the chord direction position C of the maximum camber of the section of the first blade profile1maxSatisfies the following conditions: c1max>X2, wherein the value range of X2 is 0.5-0.55, such as 0.5, 0.51, 0.52, 0.53, 0.54 and 0.55; chord-wise position C of maximum camber of the second-leaf section2maxSatisfies the following conditions: c2max<The value range of X3 and X3 is 0.4-0.45, such as 0.4, 0.41, 0.42, 0.43, 0.44 and 0.45. The chord-wise positions of the maximum camber of the blade tip and blade root blade profile sections are controlled within the range, so that the blade profile sweepback degree of the root tip position can be effectively controlled.
As shown in FIG. 7, in some embodiments, the chordwise location B of maximum thicknessmaxThe ratio of the second distance to the length of the chord line of the preset profile, the second distance being the distance between the foot perpendicular to the chord line of the preset profile from the position with the maximum radius of the inscribed circle of the preset profile and the leading edge point of the preset profile, and the chord position B with the maximum thickness of the section of the first profile1maxSatisfies the following conditions:B1max<x4, wherein the value range of X4 is 0.4-0.45, such as 0.4, 0.41, 0.42, 0.43, 0.44 and 0.45; chord position B of maximum thickness of second leaf-shaped cross section2maxSatisfies the following conditions: b is2max>The value of X5 and X5 is 0.5-0.55, such as 0.5, 0.51, 0.52, 0.53, 0.54 and 0.55. By controlling the chord-wise position of the maximum thickness, the control of the blade profile sweepback degree of the root tip position is facilitated.
In some embodiments, the thickness TmaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A midpoint of a connecting line between a foot obtained by drawing a perpendicular line from a position farthest to a chord line on a mean camber line of the preset blade profile and the chord line of the preset blade profile and a leading edge point of the preset blade profile is a first position E on the section of the first blade profile1Thickness T ofE1Satisfies the following conditions: t isE1>TmaxX6, X6 ranges from 0.7 to 0.8, such as 0.7, 0.71, 0.72, 0.73, 0.74, 0.75, 0.76, 0.77, 0.78, 0.79 and 0.8; in the second leaf-shaped cross section, at a first position E1Thickness T ofE1Satisfies the following conditions: t isE1<TmaxX7, X7 ranges from 0.6 to 0.7, such as 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69 and 0.7.
In some embodiments, the thickness TmaxIn a second position E which is the maximum diameter of the inscribed circle of the predetermined blade profile2A midpoint of a connecting line between a foot perpendicular to the chord line of the preset blade profile and a tail edge point of the preset blade profile, wherein the foot perpendicular to the chord line of the preset blade profile is taken from a position farthest relative to the chord line on a mean camber line of the preset blade profile, and a second position E is arranged on the section of the first blade profile2Thickness T ofE2Satisfies the following conditions: t isE2<TmaxX8, wherein X8 has a value in the range of 0.6 to 0.7, such as 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69 and 0.7; in a second leaf-shaped cross-section, at a second position E2Thickness T ofE2Satisfies the following conditions: t isE2>TmaxX9, X9 ranges from 0.6 to 0.7, such as 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69 and 0.7.
By aligning a first profile section of the blade tip and a second blade rootChordal position of maximum thickness and maximum camber of the profiled section, first position E1And a second position E2The blade root adopts a relative forward loading form, the blade tip adopts a relative backward loading form, the shape of the blade profile can be effectively controlled, the gravity center of the blade root is axially moved backward, the gravity center of the blade tip is moved forward, and the backswept degree of the gravity center stacking axis is reduced on the premise that the meridian projection shape of the rotor blade 2 is unchanged.
As shown in FIGS. 8 and 9, in some embodiments, the chordwise location B of maximum thicknessmaxThe ratio of the second distance to the chord length of the preset profile, the second distance being the distance between the foot perpendicular to the chord of the preset profile and the leading edge point of the preset profile, and the thickness T being the distance between the foot perpendicular to the chord of the preset profile from the position with the maximum radius of the inscribed circle of the preset profile and the leading edge point of the preset profilemaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A second position E is the midpoint of a connecting line between a foot, which is obtained by drawing a perpendicular line from the farthest position relative to the chord line on the mean camber line of the preset blade profile to the chord line of the preset blade profile, and the leading edge point of the preset blade profile2The middle point of a connecting line between a foot obtained by drawing a perpendicular line from the farthest position relative to a chord line on a mean camber line of the preset blade profile to the chord line of the preset blade profile and the tail edge point of the preset blade profile is the chord direction position B with the maximum thickness on a blade profile section with the blade height of more than X10 percentmaxFirst position E1Thickness T ofE1And thickness TmaxAnd a second position E2Thickness TE2And thickness TmaxThe ratio values are all equal in size; chord-wise position B of maximum thickness on the blade profile cross section below X10% blade heightmaxFirst position E1Thickness T ofE1And thickness TmaxA ratio of (A), a second position E2Thickness TE2And thickness TmaxThe values of the ratios are respectively and smoothly transited to the values of the parameters corresponding to the blade root position of the rotor blade 2, and the derivative of the numerical value of the smooth transition section relative to the blade height of the rotor blade 2 is continuous, wherein the value range of X10% is 15% -25%.
The backswept degree of the gravity center stacking shaft of the rotor blade 2 can be further weakened by setting different thickness distribution characteristics for the blade profile elements at different blade heights, and the strength risk is reduced.
Through the description of the multiple embodiments of the compressor, the embodiments of the compressor provided by the invention can be seen that the rotor blades are designed to be swept forward and then swept backward from the root, and the chord length of the rotor blades is increased upstream by using the additional axial distance brought by the air bleed channel, so that the working capacity of the first row of rotor blades behind the air bleed port on the casing side of the compressor can be improved, the pressure ratio of the rotor blades behind the air bleed is improved, the axial length of the compressor is favorably shortened or the total pressure ratio of the compressor is favorably improved, the aim of improving the working capacity is fulfilled, meanwhile, additional potential intensity hazard is not brought, and the efficiency of the rotor blades cannot be obviously reduced; the backswept degree of the gravity center stacking shaft of the rotor blade is weakened by setting different thickness distribution characteristics for the blade profile elements at different blade heights; the front edge of the rotor blade is designed into an S shape, so that the sweepback degree of the blade profile at the root tip position is weakened; the position of the blade tip of the rotor blade is designed to be post-loaded, so that the strength characteristic is improved, and the leakage flow of the blade tip clearance can be weakened.
Based on the compressor, the invention further provides an aircraft engine which comprises the compressor.
The positive technical effects of the compressor in the above embodiments are also applicable to an aircraft engine, and are not described herein again.
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 principles of the invention, and these modifications and equivalents are intended to be included within the scope of the claims.
Claims (11)
1. A compressor, comprising:
a stator blade (1);
a rotor blade (2) arranged adjacent to the stator blade (1), and the stator blade (1) and the rotor blade (2) have a preset distance therebetween to form an air flow channel;
a casing provided radially outside the stator blade (1) and the rotor blade (2); and
a bleed air channel (3) which is communicated with the airflow channel and extends from the airflow channel to the direction close to the casing;
the included angle between a first straight line and the axial line of the gas compressor is an acute angle, the first straight line is a connecting line between a first middle point and a second middle point, the first middle point is the middle point of the connecting line between a first leading edge point and a first trailing edge point on a first blade profile section of the rotor blade (2) at the blade tip, the second middle point is the middle point of the connecting line between the second leading edge point and a second trailing edge point on a second blade profile section of the rotor blade (2) at the blade root, the chord length of the rotor blade (2) is monotonically reduced from the root to the tip, and the change rate of the chord length of the rotor blade (2) along the blade height direction is variable.
2. The compressor according to claim 1, characterized in that the rotor blade (2) comprises a first portion and a second portion located above the first portion in the blade height direction, the rate of change of the chord length of the first portion in the blade height direction is from small to large, and the rate of change of the chord length of the second portion in the blade height direction is from large to small.
3. The compressor as claimed in claim 1, characterised in that the profile cross section of the rotor blade (2) at its tip has a first chord length and the profile cross section of the rotor blade (2) at its root has a second chord length, the first chord length being greater than the second chord length by X1%, the X1% having a value in the range 5% to 10%.
4. Compressor according to claim 1, characterised in that the projection of the rotor blade (2) on a meridian plane has a first leading edge line (4), the line connecting the first leading edge point and the second leading edge point being a second straight line, the first leading edge line (4) being located downstream of the second straight line close to the tip of the rotor blade (2); -the first leading edge line (4) is located upstream of the second straight line near a root of the rotor blade (2).
5. Compressor according to claim 1, characterized in that the projection of the rotor blade (2) on a meridian plane has a first leading edge line (4), the line connecting the first leading edge point and the second leading edge point being a second straight line, the first leading edge line (4) and the second straight line having a unique intersection point between the first profile section and the second profile section.
6. An air compressor as set forth in claim 1, characterized in that the chord-wise position C of maximum cambermaxThe first distance is the distance between a foot obtained by drawing a perpendicular line from the position farthest to the chord line on the mean camber line of the preset blade profile to the chord line of the preset blade profile and the front edge point of the preset blade profile, and the chord direction position C of the maximum camber of the section of the first blade profile is the ratio of the first distance to the length of the chord line of the preset blade profile1maxSatisfies the following conditions: c1max>X2, wherein the value range of the X2 is 0.5-0.55; chord-wise position C of maximum camber of the second leaf-shaped cross-section2maxSatisfies the following conditions: c2max<And X3, wherein the value range of X3 is 0.4-0.45.
7. Compressor according to claim 1, characterised in that the chord-wise position B of maximum thicknessmaxThe ratio of a second distance to the length of the chord line of the preset blade profile, wherein the second distance is the distance between a foot perpendicular to the chord line of the preset blade profile from the position with the maximum radius of the inscribed circle of the preset blade profile and the front edge point of the preset blade profile, and the chord direction position B with the maximum thickness of the section of the first blade profile1maxSatisfies the following conditions: b is1max<X4, wherein the value range of the X4 is 0.4-0.45; a chord-wise position of maximum thickness of the second leaf-shaped cross sectionB2maxSatisfies the following conditions: b is2max>And X5, wherein the value range of X5 is 0.5-0.55.
8. Compressor according to claim 1, characterised in that the thickness T ismaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A midpoint of a connecting line between a foot perpendicular to the chord line of the preset blade profile from a position farthest from the chord line on the mean camber line of the preset blade profile and the leading edge point of the preset blade profile, wherein on the first blade profile section, the first position E is1Thickness T ofE1Satisfies the following conditions: t isE1>TmaxX6, wherein the value range of X6 is 0.7-0.8; in the second leaf-shaped cross-section, the first position E1Thickness T ofE1Satisfies the following conditions: t isE1<TmaxAnd X7, wherein the value range of X7 is 0.6-0.7.
9. Compressor according to claim 1, characterised in that the thickness T ismaxIn a second position E which is the maximum diameter of the inscribed circle of the predetermined blade profile2A midpoint of a connecting line between a foot perpendicular to the chord line of the preset blade profile from a position farthest from the chord line on the mean camber line of the preset blade profile and the chord line of the preset blade profile, and a second position E on the section of the first blade profile2Thickness T ofE2Satisfies the following conditions: t isE2<TmaxX8, wherein the value range of X8 is 0.6-0.7; in the second leaf-shaped cross-section, the second position E2Thickness T ofE2Satisfies the following conditions: t isE2>TmaxAnd X9, wherein the value range of X9 is 0.6-0.7.
10. Compressor according to claim 1, characterised in that the chord-wise position B of maximum thicknessmaxThe second distance is the distance between a foot obtained by drawing a perpendicular line from the position with the maximum radius of the inscribed circle of the profile of the preset blade profile to the chord line of the preset blade profile and the front edge point of the preset blade profile, and the thickness is the ratio of the second distance to the length of the chord line of the preset blade profileDegree TmaxIn a first position E which is the maximum diameter of the inscribed circle of the predetermined blade profile1A second position E which is the midpoint of a connecting line between a foot made of a perpendicular line from the farthest position relative to a chord line on a mean camber line of the preset blade profile to the chord line of the preset blade profile and a leading edge point of the preset blade profile2The middle point of a connecting line between a foot perpendicular to the chord line of the preset blade profile from the position farthest to the chord line on the mean camber line of the preset blade profile and the tail edge point of the preset blade profile is the chord direction position B of the maximum thickness on the blade profile section with the blade height of more than X10 percentmaxThe first position E1Thickness T ofE1And thickness TmaxAnd said second position E2Thickness TE2And thickness TmaxThe ratio values are all equal in size; the chord-wise position B of the maximum thickness on the profile cross section below X10% of the blade heightmaxThe first position E1Thickness T ofE1And thickness TmaxThe second position E2Thickness TE2And thickness TmaxThe values of the ratios are respectively in smooth transition to the value of the blade root position, the derivative of the value of the smooth transition section relative to the blade height of the rotor blade (2) is continuous, and the value range of X10% is 15% -25%.
11. An aircraft engine, characterized in that it comprises a compressor as claimed in any one of claims 1 to 10.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN115655729A (en) * | 2022-11-09 | 2023-01-31 | 北京航空航天大学 | Low-speed simulation method and device for coupling of S-shaped transition section and stator of gas compressor |
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