CN113339325B - Inlet stage blade assembly for compressor and axial flow compressor comprising same - Google Patents
Inlet stage blade assembly for compressor and axial flow compressor comprising same Download PDFInfo
- Publication number
- CN113339325B CN113339325B CN202110905389.4A CN202110905389A CN113339325B CN 113339325 B CN113339325 B CN 113339325B CN 202110905389 A CN202110905389 A CN 202110905389A CN 113339325 B CN113339325 B CN 113339325B
- Authority
- CN
- China
- Prior art keywords
- blade
- stage
- inlet
- tip
- inlet guide
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Classifications
-
- 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
-
- 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
- F04D29/38—Blades
- F04D29/384—Blades characterised by form
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses an inlet stage blade assembly for a gas compressor and an axial flow gas compressor comprising the same, wherein the inlet stage blade assembly comprises an inlet guide blade and a first stage movable blade, the position of the inlet guide blade with the largest tail edge metal angle is a first position, and the relative blade height T at the first position meets the condition that T is more than or equal to 0.8 and less than or equal to 0.9; the blade profile section of the first-stage movable blade at the blade tip has a first chord length, the blade profile section of the first-stage movable blade at a second position, relative to the blade height T, has a second chord length, and the first chord length is 3% -10% larger than the second chord length; between the second position and the blade tip of the first stage movable blade, the forward or backward sweeping degree of the stacking shaft of the first stage movable blade is not more than 1% of the absolute blade height of the first stage movable blade. By adopting the structure, the stall margin of the blade tip of the first-stage movable blade can be improved, and the radial deformation of the leading edge and the trailing edge of the first-stage movable blade from a static state to a running state is more coordinated.
Description
Technical Field
The invention relates to an inlet stage blade assembly for a compressor and an axial flow compressor comprising the same.
Background
Inlet guide vanes are a common component in compressor parts of aircraft engines, gas turbines, and ground axial compressors. The inlet guide vane is the first row of vanes at the compressor inlet and generally functions to deflect an axial incoming flow with a tangential component velocity in the same direction as the direction of rotor rotation. Thus, the incoming flow mach number of the first stage bucket (rotor blade) is lower than that without guide vanes, so that the loss is smaller and the efficiency is higher. In addition, the inlet guide vane is also designed to be adjustable, and the installation angle of the inlet guide vane is changed along with the change of the operation condition of the compressor, so that the first-stage movable vane is kept to work at a proper inflow prerotation angle (included angle between inflow in front of a rotor blade of the compressor and the axial direction) all the time, and high efficiency and margin are realized.
In order to reduce the volume and weight of the unit and reduce the number of stages, the current axial flow compressor usually adopts a higher design rotating speed, so that the relative mach number of the incoming flow of the front stage rotor blade is close to or exceeds 1, and shock waves exist in the blade channel. Shock waves have a strong supercharging capacity but also cause significant losses. The relative incoming flow mach number of the first stage bucket is typically highest among all the blades, and increases gradually from the root to the tip. To ensure the efficiency of the first stage bucket, the inlet guide vanes are usually designed with a gradually increasing metal angle from the root to the trailing edge of the tip, as shown in fig. 1. Therefore, the incoming flow prerotation angle of the downstream rotor blade gradually increases from the blade root to the blade tip, so that the trend that the relative Mach number of the incoming flow gradually increases from the blade root to the blade tip is relieved, and the effect is shown in FIG. 2.
Furthermore, the first stage bucket tends to stall at the root and tip due to factors such as upstream boundary layers. According to the aerodynamic principle of the multistage axial flow compressor, when the rotating speed is lower than the designed rotating speed, the front stage is easy to surge and the rear stage is easy to block, the trend of the first stage of surge is most obvious, and the first stage of surge is easy to become the leading speed-losing stage. For the current aviation axial flow compressor, the root of the first stage movable blade can be designed into a form that the radius is obviously increased from the front edge to the tail edge, so the pressurizing capacity of the first stage movable blade not only comes from the expansion of a flow passage, but also comes from the centrifugal force, the mechanism of the first stage movable blade is similar to that of a mixed flow stage, and the pressurizing capacity is strong and the speed is not easy to stall. In addition, the chord length of the first-stage movable blade generally does not change greatly from the blade root to the blade tip, and the hub ratio of the first-stage movable blade is relatively small (usually about 0.5), so that the consistency of the blade root of the first-stage movable blade is very large and can generally reach more than 2, and the possibility of stalling the first-stage movable blade at the blade root is further reduced. However, the blade tip of the first stage bucket has a consistency of only half of the blade root, and the blade tip is not influenced by the clearance leakage flow without the action of the centrifugal force, so that the stall risk is high. Therefore, the improvement of the stall resistance of the blade tip of the first-stage movable blade is significant.
In order to improve the stall resistance of the blade tip of the first-stage movable blade, the prior art mostly adopts a local forward-swept design for the blade tip of the first-stage movable blade, and can effectively meet the aerodynamic requirement. However, improper sweep design tends to result in uneven radial deformation of the leading and trailing edges of the tip of the first stage bucket, especially for wide chord fan/compressor blades. Limited to this, the local forward sweep of the tip of the first stage bucket is generally unlikely to be too intense, and so there is a limit to how much margin can be increased by virtue thereof.
Disclosure of Invention
The invention aims to solve the technical problems of how to improve the acting capacity of the blade tip of the first-stage movable blade of the compressor, improve the stall margin of the first-stage movable blade, and simultaneously avoid the defect of strong radial deformation imbalance of the leading edge and the trailing edge, and provides an inlet-stage blade assembly for the compressor and an axial-flow compressor comprising the inlet-stage blade assembly.
The invention solves the technical problems through the following technical scheme:
the invention provides an inlet stage blade assembly for a gas compressor, which comprises an inlet guide blade and a first stage movable blade, wherein the position of the inlet guide blade with the largest trailing edge metal angle is a first position, the relative blade height T at the first position meets the condition that T is more than or equal to 0.8 and less than or equal to 0.9, and the trailing edge metal angle is an included angle between the tangential direction of a blade profile middle arc line of the inlet guide blade at the trailing edge of the inlet guide blade and the axial direction of the gas compressor;
the blade profile section of the first-stage movable blade at the blade tip has a first chord length, the blade profile section of the first-stage movable blade at a second position, where the relative blade height is T, has a second chord length, the first chord length is 3% -10% longer than the second chord length, and the relative blade height is the ratio of the distance from a certain point on the blade to the blade root of the blade to the distance from the blade tip to the blade root of the blade;
between the second position and the blade tip of the first-stage movable blade, the forward-swept or backward-swept degree of the stacking shaft of the first-stage movable blade does not exceed 1% of the absolute blade height of the first-stage movable blade, and the absolute blade height is the distance from the blade tip of the first-stage movable blade to the blade root of the first-stage movable blade.
In this scheme, the inlet stage blade subassembly of compressor adopts above-mentioned structure, reduces the trailing edge metal angle near the apex of inlet stator to for the apex of the first-stage movable vane of low reaches provides less incoming flow prewhirl angle, suitably reduce the mach number of first-stage movable vane at the apex department, increase the reaction of first-stage movable vane, in order to improve its power-doing ability. By limiting the chord length of the first-stage movable blade at the blade tip and the chord length of the blade profile section at the second position, which has the same relative blade height with the first position at the maximum tail edge metal angle of the inlet guide blade, the radial deformation of the front and tail edges of the first-stage movable blade from a static state to an operating state is more coordinated, and the blade tip part of the first-stage movable blade is lighter in load and less prone to stall on the premise of realizing the same pressure ratio and flow, so that the stall margin of the blade tip of the first-stage movable blade is improved.
Preferably, the derivative of the trailing edge metal angle of the inlet guide vane with respect to the relative blade height of the inlet guide vane is continuous;
and/or the derivative of the position of the centre of gravity of the chord length of the first stage bucket with respect to the relative blade height of the first stage bucket is continuous;
and/or the derivative of the barycentric position of the elementary sections of the first stage bucket with respect to the relative blade height of the first stage bucket is continuous.
In the scheme, the tail edge metal angle of the inlet guide vane, the gravity center position of the chord length of the first-stage movable vane and the gravity center position of the element section of the first-stage movable vane adopt the structure, so that the pneumatic performance of the compressor is improved.
Preferably, the metal angle of the trailing edge at the tip of the inlet guide vane is 0.5 to 5 degrees smaller than the maximum metal angle of the trailing edge of the inlet guide vane.
In the scheme, the metal angle of the tail edge of the blade tip of the inlet guide blade is controlled to be smaller than the maximum metal angle of the tail edge by 0.5-5 degrees, and a proper incoming flow prerotation angle can be provided for the blade tip of the downstream first-stage movable blade, so that the reaction degree of the first-stage movable blade is controlled to be within a reasonable range, and the working capacity of the first-stage movable blade is enhanced.
Preferably, the derivative of the magnitude of the trailing edge metal angle of the inlet guide vane with respect to the relative blade height of the inlet guide vane is positive and continuously increasing from the blade root to the first position of the inlet guide vane.
In the scheme, by adopting the structure, the incoming flow prerotation angle of the first-stage movable blade at the downstream of the inlet guide vane can be gradually increased from the blade root to the blade tip so as to relieve the trend that the relative Mach number of the incoming flow is gradually increased from the blade root to the blade tip.
Preferably, from the blade root to the first position of the inlet guide vane, the derivative of the size of the trailing edge metal angle of the inlet guide vane with respect to the relative blade height of the inlet guide vane is a fixed value greater than zero.
In the scheme, by adopting the structure, the incoming flow prerotation angle of the first-stage movable blade at the downstream of the inlet guide vane can be gradually increased from the blade root to the blade tip so as to relieve the trend that the relative Mach number of the incoming flow is gradually increased from the blade root to the blade tip.
Preferably, the axial position of the leading edge point of the profiled section of the first stage bucket at the tip is located upstream of the axial position of the leading edge point of the profiled section at the second location.
In this scheme, adopt above-mentioned structure, can make the apex leading edge of first order movable vane have the effect of sweepforward, further strengthen the anti stall ability of the apex of first order movable vane.
Preferably, from the second position to the tip of the first stage bucket, the maximum absolute thickness of the profile section of the first stage bucket is constant or monotonically decreasing, and the maximum absolute thickness is the diameter of the largest inscribed circle in the profile section of the first stage bucket.
In this scheme, adopt above-mentioned structural design for the centrifugal force increase that the apex chord length increase of first order movable vane leads to is offset to a certain extent, further strengthens the anti stall ability of the apex of first order movable vane.
Preferably, a ratio of a maximum absolute thickness of a profile section of the first stage bucket at a tip to a maximum absolute thickness of a profile section of the first stage bucket at the second position is M, a ratio of a chord length of the profile section of the first stage bucket at the second position to a chord length of the profile section of the first stage bucket at the tip is N, M = N × t, wherein: t is more than or equal to 0.8 and less than or equal to 0.95.
In this scheme, adopt above-mentioned structural design, the centrifugal force increase that further makes the apex chord length increase of first order movable vane lead to is offset to a certain extent, and then strengthens the anti stall ability of the apex of first order movable vane, can also make the radial deformation of first order movable vane at the leading edge of apex and trailing edge more harmonious.
Preferably, the first chord length is 5% -8% longer than the second chord length.
In this scheme, adopt above-mentioned structural design, avoid first order movable vane to sweep forward too obviously in apex department for the radial deformation of first order movable vane at the leading edge of apex and trailing edge is more harmonious.
The invention also provides an axial flow compressor which comprises the inlet stage blade assembly.
The positive progress effects of the invention are as follows: by adopting the structure, the metal angle of the tail edge is reduced near the blade tip of the inlet guide blade, so that a smaller inflow prerotation angle is provided for the blade tip of the downstream first-stage movable blade, the Mach number of the first-stage movable blade at the blade tip is properly reduced, and the reaction degree of the first-stage movable blade is increased, so that the power-applying capacity of the first-stage movable blade is improved. By limiting the chord length of the first-stage movable blade at the blade tip and the chord length of the blade profile section at the second position, which has the same relative blade height with the first position at the maximum tail edge metal angle of the inlet guide blade, the radial deformation of the front and tail edges of the first-stage movable blade from a static state to an operating state is more coordinated, and the blade tip part of the first-stage movable blade is lighter in load and less prone to stall on the premise of realizing the same pressure ratio and flow, so that the stall margin of the blade tip of the first-stage movable blade is improved.
Drawings
Fig. 1 is a top view of an inlet guide vane in a compressor.
FIG. 2 is a schematic diagram illustrating the change of the relative Mach number of the incoming flow of the prior first-stage bucket from the root to the tip.
Fig. 3 is a definition reference diagram of a metal corner.
FIG. 4 is a schematic diagram of the variation of the metal angle of the inlet guide vane of the present invention from the root to the trailing edge of the blade tip of the inlet guide vane of the prior art.
FIG. 5 is a schematic view of the change of the center of gravity of the blade profile cross section from the root to the tip of the first stage moving blade of the present invention and the first stage moving blade of the prior art.
FIG. 6 is a schematic view of the chord length variation of the blade profile cross-section from the root to the tip of the first stage moving blade of the present invention and the first stage moving blade of the prior art.
FIG. 7 is a schematic diagram of two different distribution laws of the metal angles of the trailing edge of the inlet guide vane from the blade root to the blade tip according to the present invention.
FIG. 8 is a meridional view of one embodiment of a first stage bucket of the present invention.
Description of reference numerals:
inlet guide vane 100
Guide vane pressure surface 101
Guide vane suction surface 102
Guide vane leading edge 103
Guide vane trailing edge 104
Detailed Description
The invention will be more clearly and completely described below by way of examples and with reference to the accompanying drawings, without thereby limiting the scope of the invention to these examples.
It should be understood that the terms "first", "second", etc. are used for limiting technical features only for the convenience of distinguishing corresponding technical features, and the terms have no special meanings if not stated otherwise, and therefore, the scope of the present invention should not be construed as being limited. Reference to "one embodiment" or "another embodiment" means that a feature, structure, or characteristic described in connection with at least one embodiment of the application. Furthermore, some features, structures, or characteristics of one or more embodiments of the present application may be combined as appropriate.
It is noted that these and other figures which follow are merely exemplary and not drawn to scale and should not be considered as limiting the scope of the invention as it is actually claimed. Further, the conversion methods in the different embodiments may be appropriately combined.
For ease of understanding, one or more of the terms herein are to be interpreted as follows:
axial-flow compressor: the multistage compression equipment with the airflow flowing direction consistent or nearly consistent with the rotating axis direction of the 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 aeroengine or a gas turbine; the combination of adjacent stator and rotor blades is referred to as a stage.
Relative leaf height: the ratio of the distance from a certain point on a blade in the compressor to the blade root to the distance from the blade tip to the blade root of the blade.
Absolute leaf height: the distance from a certain point on the blade in the compressor to the blade root.
Pre-rotation angle: the angle between the incoming flow in front of the rotor blade of the compressor and the axial direction is formed.
The pneumatic performance is as follows: the pneumatic performance of the compressor (or a compressor stage, a compressor rotor blade, the same below) mainly comprises four indexes, namely inlet conversion flow (air flow converted from inlet conditions to standard atmospheric conditions in kg/s), pressure ratio (ratio of total outlet pressure to total inlet pressure, dimensionless), efficiency (degree ratio of mechanical function converted into gas pressure by the compressor, calculated by ideological parameters of total inlet temperature and total pressure, and total outlet temperature and total pressure, dimensionless), surge margin (the size of a range in which the compressor can stably work is measured, and calculated by the conversion flow, the pressure ratio of the compressor at a design point, and the flow and the pressure ratio of the compressor at a near surge point, dimensionless).
Metal angle: included angles between tangential directions of blade profile camber lines of the blades at the front edges and the tail edges of the blades and the axial direction are respectively called as a front edge metal angle and a tail edge metal angle. The trailing edge metal angle of the inlet guide vane is defined as positive and negative if the trailing edge of the inlet guide vane is oriented such that the tangential velocity of the outlet flow of the inlet guide vane is the same as the direction of rotation of the downstream rotor. And (3) making a ray upstream from the leading edge point of the inlet guide vane, wherein the ray is tangent to a local mean camber line, and if the tangential component of the direction of the ray is the same as the rotating direction of the downstream rotor, the leading edge metal angle of the inlet guide vane is negative, otherwise, the leading edge metal angle of the inlet guide vane is positive. Such as the case shown in fig. 3, where the leading edge metal angle of the inlet guide vane is negative and the trailing edge metal angle is positive.
Blade tip clearance: in the operation process of the compressor, the inlet guide vane needs to be adjusted frequently, so that a gap is inevitably formed between the vane tip and the casing. This causes gas to flow along the gap from the pressure side to the suction side during compressor operation. The presence of tip clearances results in increased inlet guide vane losses and a reduction in the incoming flow pre-swirl angle of the rotor tip as compared to the ideal case without tip clearances.
Stacking shafts: the centers of gravity of all the blade profile sections (elementary sections) of the movable blade are connected to form a space curve.
The pneumatic performance is as follows: the pneumatic performance of the compressor (or a compressor stage, a compressor rotor blade, the same below) mainly comprises four indexes, namely inlet conversion flow (air flow converted from inlet conditions to standard atmospheric conditions in kg/s), pressure ratio (ratio of total outlet pressure to total inlet pressure, dimensionless), efficiency (degree ratio of mechanical function converted into gas pressure by the compressor, calculated by ideological parameters of total inlet temperature and total pressure, and total outlet temperature and total pressure, dimensionless), surge margin (the size of a range in which the compressor can stably work is measured, and calculated by the conversion flow, the pressure ratio of the compressor at a design point, and the flow and the pressure ratio of the compressor at a near surge point, dimensionless).
As shown in fig. 1, 4-8, the present embodiment is an inlet stage vane assembly for a compressor, which is used in an axial flow compressor. The inlet stage bucket assembly includes an inlet guide vane 100 and a first stage bucket 200. Referring to fig. 1, the inlet guide vane 100 has a guide vane pressure surface 101, a guide vane suction surface 102, a guide vane leading edge 103, a guide vane trailing edge 104, a guide vane tip 105 and a guide vane root (not shown), and referring to fig. 8, the first stage bucket 200 has a bucket pressure surface 201, a bucket suction surface (not shown), a bucket root 202, a bucket leading edge 203, a bucket trailing edge 204 and a bucket tip 205.
Referring to FIG. 4, in the present embodiment, the position of the inlet guide vane 100 having the largest metal corner at the trailing edge is the first position, and the relative blade height T at the first position satisfies 0.8 ≦ T ≦ 0.9. Referring to FIG. 5, the blade profile section of the first stage movable blade 200 at the blade tip has a first chord length, the blade profile section of the first stage movable blade 200 at a second position relative to the blade height T has a second chord length, and the first chord length is 3% -10% larger than the second chord length. Referring to FIG. 6, between the second position and the tip of the first stage bucket 200, the stacking axis of the first stage bucket 200 is swept forward or swept backward by no more than 1% of the absolute blade height of the first stage bucket 200.
The inlet stage blade assembly of the compressor adopts the structure, and the metal angle of the tail edge is reduced near the blade tip of the inlet guide vane 100, so that a smaller inflow prewhirl angle is provided for the blade tip of the downstream first stage movable blade 200, the Mach number of the first stage movable blade 200 at the blade tip is properly reduced, and the reaction degree of the first stage movable blade 200 is increased, so that the working capacity of the first stage movable blade is improved. By limiting the chord length of the first-stage movable blade 200 at the blade tip and the chord length of the blade profile section at the second position with the same relative blade height as the first position at the maximum tail edge metal angle of the inlet guide blade, the radial deformation of the leading edge and the tail edge of the first-stage movable blade 200 from a static state to an operating state is more coordinated, and the blade tip part of the first-stage movable blade 200 is lighter in load and less prone to stall on the premise of realizing the same pressure ratio and flow rate, so that the stall margin of the blade tip of the first-stage movable blade 200 is improved.
As the chord length of the blade tip of the first stage movable blade 200 is increased, even if the stacking axis at the blade tip is not swept forward, the front edge at the blade tip has the effect of sweeping forward.
In the present embodiment, the value of the relative blade height T is not particularly limited, and it should be understood that, assuming that the relative blade height T at the first position of the inlet guide vane 100 is 0.85, the relative blade height T at the second position on the first-stage movable blade 200 is also 0.85. Of course, in actual manufacturing, due to machining errors, certain deviations may occur, and reasonable deviations may be within an acceptable range.
Preferably, the first chord length is 5% -8% larger than the second chord length. The forward sweepback of the first-stage movable blade 200 at the blade tip is avoided from being too obvious, so that the radial deformation of the leading edge and the trailing edge of the first-stage movable blade 200 is more coordinated when the blade tip is in a static state to a running state.
In the present embodiment, the derivative of the trailing edge metal angle of the inlet guide vane 100 with respect to the blade height of the inlet guide vane 100 is continuous. The derivative of the position of the center of gravity of the chord length of the first stage bucket 200 with respect to the blade height of the first stage bucket 200 is continuous. The derivative of the barycentric position of the elementary sections of the first stage bucket 200 with respect to the blade height of the first stage bucket 200 is continuous. The structure is adopted for the tail edge metal angle of the inlet guide vane 100, the gravity center position of the chord length of the first-stage movable vane 200 and the gravity center position of the elementary section of the first-stage movable vane 200, and the pneumatic performance of the compressor is favorably improved.
In other embodiments, the metal angle of the trailing edge of the inlet guide vane 100, the position of the center of gravity of the chord length of the first stage moving blade 200, and the position of the center of gravity of the section of the element of the first stage moving blade 200 may also satisfy only one or two of the cases, and are not described herein again.
In one embodiment, the trailing edge metal angle at the tip of the inlet guide vane 100 is 0.5 ° -5 ° less than the maximum trailing edge metal angle of the inlet guide vane 100. The metal angle of the tail edge of the blade tip of the inlet guide vane 100 is controlled to be smaller than the maximum metal angle of the tail edge by 0.5-5 degrees, and a proper inflow prerotation angle can be provided for the blade tip of the downstream first-stage movable blade 200, so that the reaction degree of the first-stage movable blade 200 is controlled to be within a reasonable range, and the working capacity of the first-stage movable blade 200 is enhanced. The trailing edge metal angle is positive at any relative leaf height.
In another embodiment, the metal angle of the trailing edge at the tip of the inlet guide vane 100 may be less than the maximum metal angle of the inlet guide vane 100 by more than 5 °, but not too great, preferably controlled within 8 °.
Preferably, in the present embodiment, the trailing edge metal angle at the tip of the inlet guide vane 100 is 2 ° smaller than the maximum trailing edge metal angle of the inlet guide vane 100.
In one embodiment, referring to the distribution law 2 of the trailing edge metal angles in FIG. 7, the derivative of the magnitude of the trailing edge metal angle of the inlet guide vane 100 with respect to the blade height of the inlet guide vane 100 is positive and continuously increasing from the blade root of the inlet guide vane 100 to the first position. By adopting the structure, the incoming flow prerotation angle of the first-stage movable blade 200 at the downstream of the inlet guide vane 100 can be gradually increased from the blade root to the blade tip, so that the trend that the relative Mach number of the incoming flow is gradually increased from the blade root to the blade tip is relieved.
In another embodiment, referring to distribution law 1 of the trailing edge metal angle in fig. 7, the derivative of the magnitude of the trailing edge metal angle of the inlet guide vane 100 with respect to the blade height of the inlet guide vane 100 from the blade root to the first position of the inlet guide vane 100 is a fixed value greater than zero. By adopting the structure, the incoming flow prerotation angle of the first-stage movable blade 200 at the downstream of the inlet guide vane 100 can be gradually increased from the blade root to the blade tip, so that the trend that the relative Mach number of the incoming flow is gradually increased from the blade root to the blade tip is relieved.
Referring to FIG. 8, the axial position of the leading edge point of the profiled section of the first stage bucket 200 at tip A is located upstream of the axial position of the leading edge point of the profiled section at the second position B. The leading edge of the blade tip of the first stage bucket 200 has a sweepforward effect, and the anti-stall capability of the blade tip of the first stage bucket 200 is further enhanced.
During design of the first-stage movable blade 200, the maximum absolute thickness of the blade profile section of the first-stage movable blade 200 is unchanged or monotonically decreases from the point B at the second position to the point a at the tip of the first-stage movable blade 200. By adopting the structure design, the centrifugal force increase caused by the increase of the chord length of the blade tip of the first-stage movable blade 200 is offset to a certain extent, and the stall resistance of the blade tip of the first-stage movable blade 200 is further enhanced. The maximum absolute thickness is the diameter of the largest inscribed circle in the blade profile cross section of the first-stage movable blade 200.
In the present embodiment, the maximum absolute thickness of the profile section of the first stage bucket 200 is constant. Preferably, the following relationship may be satisfied:
the ratio of the maximum absolute thickness of the profile section of the first stage movable blade 200 at the blade tip to the maximum absolute thickness of the profile section of the first stage movable blade 200 at the second position is M, the ratio of the chord length of the profile section of the first stage movable blade 200 at the second position to the chord length of the profile section of the first stage movable blade 200 at the blade tip is N, M = N × t, wherein: t is more than or equal to 0.8 and less than or equal to 0.95.
By adopting the structural design, the centrifugal force increase caused by the increase of the chord length of the blade tip of the first-stage movable blade 200 is further counteracted to a certain extent, so that the stall resistance of the blade tip of the first-stage movable blade 200 is enhanced, and the radial deformation of the first-stage movable blade 200 at the front edge and the tail edge of the blade tip can be more coordinated.
In another embodiment, the maximum absolute thickness of the blade profile section of the first stage bucket 200 decreases monotonically, which may enable the increase in centrifugal force of the first stage bucket 200 due to increased chord length of the blade tip to be offset to some extent, further enhancing the stall resistance of the blade tip of the first stage bucket 200.
The embodiment of the invention also provides an axial flow compressor which comprises the inlet stage blade assembly.
While specific embodiments of the invention have been described above, it will be appreciated by those skilled in the art that this is by way of example only, and that the scope of the invention is defined by the appended claims. Various changes and modifications to these embodiments may be made by those skilled in the art without departing from the spirit and scope of the invention, and these changes and modifications are within the scope of the invention.
Claims (10)
1. An inlet stage blade assembly for a compressor comprises an inlet guide blade and a first stage movable blade, and is characterized in that the position of the inlet guide blade with the largest trailing edge metal angle is a first position, the relative blade height T at the first position satisfies the condition that T is more than or equal to 0.8 and less than or equal to 0.9, and the trailing edge metal angle is an included angle between the tangential direction of a blade profile middle arc line of the inlet guide blade at the trailing edge of the inlet guide blade and the axial direction of the compressor;
the blade profile section of the first-stage movable blade at the blade tip has a first chord length, the blade profile section of the first-stage movable blade at a second position, where the relative blade height is T, has a second chord length, the first chord length is 3% -10% longer than the second chord length, and the relative blade height is the ratio of the distance from a certain point on the blade to the blade root of the blade to the distance from the blade tip to the blade root of the blade;
between the second position and the blade tip of the first-stage movable blade, the forward-swept or backward-swept degree of the stacking shaft of the first-stage movable blade does not exceed 1% of the absolute blade height of the first-stage movable blade, and the absolute blade height is the distance from the blade tip of the first-stage movable blade to the blade root of the first-stage movable blade.
2. The inlet stage vane assembly for a compressor of claim 1, wherein a derivative of a trailing edge metal angle of the inlet guide vane with respect to a relative leaf height of the inlet guide vane is continuous;
and/or the derivative of the position of the centre of gravity of the chord length of the first stage bucket with respect to the relative blade height of the first stage bucket is continuous;
and/or the derivative of the barycentric position of the elementary sections of the first stage bucket with respect to the relative blade height of the first stage bucket is continuous.
3. The inlet stage vane assembly for an air compressor of claim 1, wherein a trailing edge metal angle at a tip of the inlet guide vane is 0.5 ° -5 ° smaller than a maximum trailing edge metal angle of the inlet guide vane.
4. The inlet stage vane assembly for an air compressor of claim 1, wherein a derivative of a magnitude of a trailing edge metal angle of the inlet guide vane with respect to a relative blade height of the inlet guide vane is positive and continuously increasing from a blade root of the inlet guide vane to the first position.
5. The inlet stage vane assembly for an air compressor of claim 1, wherein a derivative of a magnitude of a trailing edge metal angle of the inlet guide vane with respect to a relative blade height of the inlet guide vane is a fixed value greater than zero from a blade root of the inlet guide vane to the first position.
6. The inlet stage blade assembly for an air compressor as claimed in claim 1, wherein an axial position of a leading edge point of the profiled section of the first stage bucket at the tip is located upstream of an axial position of the leading edge point of the profiled section at the second location.
7. The inlet stage blade assembly for a compressor as claimed in claim 1, wherein a maximum absolute thickness of the profile section of the first stage bucket, from the second position to the tip of the first stage bucket, is constant or monotonically decreasing, the maximum absolute thickness being a diameter of a maximum inscribed circle within the profile section of the first stage bucket.
8. The inlet stage blade assembly for an air compressor as claimed in claim 7, wherein the ratio of the maximum absolute thickness of the profile section of the first stage bucket at the tip to the maximum absolute thickness of the profile section of the first stage bucket at the second location is M, the ratio of the chord length of the profile section of the first stage bucket at the second location to the chord length of the profile section of the first stage bucket at the tip is N, M = nt, wherein: t is more than or equal to 0.8 and less than or equal to 0.95.
9. The inlet stage vane assembly for an air compressor of claim 1, wherein the first chord length is between 5% and 8% greater than the second chord length.
10. An axial compressor comprising an inlet stage vane assembly according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110905389.4A CN113339325B (en) | 2021-08-09 | 2021-08-09 | Inlet stage blade assembly for compressor and axial flow compressor comprising same |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110905389.4A CN113339325B (en) | 2021-08-09 | 2021-08-09 | Inlet stage blade assembly for compressor and axial flow compressor comprising same |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113339325A CN113339325A (en) | 2021-09-03 |
| CN113339325B true CN113339325B (en) | 2022-01-07 |
Family
ID=77480904
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110905389.4A Active CN113339325B (en) | 2021-08-09 | 2021-08-09 | Inlet stage blade assembly for compressor and axial flow compressor comprising same |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113339325B (en) |
Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113605990A (en) * | 2021-09-14 | 2021-11-05 | 西安陕鼓动力股份有限公司 | Two-stage tail gas turbine, stationary blade and movable blade for high-pressure nitric acid three-in-one device |
| CN113982994B (en) * | 2021-10-28 | 2024-05-28 | 西安热工研究院有限公司 | Novel capacity-increasing transformation method for movable blade adjustable axial flow fan of power station |
| CN113958537B (en) * | 2021-12-16 | 2022-03-15 | 中国航发上海商用航空发动机制造有限责任公司 | Compressor and aircraft engine |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6354798B1 (en) * | 1997-09-08 | 2002-03-12 | Siemens Aktiengesellschaft | Blade for a fluid-flow machine, and steam turbine |
| EP3051142A1 (en) * | 2015-01-28 | 2016-08-03 | MTU Aero Engines GmbH | Gas turbine axial compressor |
| CN112283160A (en) * | 2020-12-24 | 2021-01-29 | 中国航发上海商用航空发动机制造有限责任公司 | Compressor rotor blade and design method thereof |
Family Cites Families (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9309769B2 (en) * | 2010-12-28 | 2016-04-12 | Rolls-Royce Corporation | Gas turbine engine airfoil shaped component |
| JP6468414B2 (en) * | 2014-08-12 | 2019-02-13 | 株式会社Ihi | Compressor vane, axial compressor, and gas turbine |
| EP3296573A1 (en) * | 2016-09-20 | 2018-03-21 | Siemens Aktiengesellschaft | A technique for controlling rotating stall in compressor for a gas turbine engine |
-
2021
- 2021-08-09 CN CN202110905389.4A patent/CN113339325B/en active Active
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6354798B1 (en) * | 1997-09-08 | 2002-03-12 | Siemens Aktiengesellschaft | Blade for a fluid-flow machine, and steam turbine |
| EP3051142A1 (en) * | 2015-01-28 | 2016-08-03 | MTU Aero Engines GmbH | Gas turbine axial compressor |
| CN112283160A (en) * | 2020-12-24 | 2021-01-29 | 中国航发上海商用航空发动机制造有限责任公司 | Compressor rotor blade and design method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113339325A (en) | 2021-09-03 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN113339325B (en) | Inlet stage blade assembly for compressor and axial flow compressor comprising same | |
| US8235658B2 (en) | Fluid flow machine including rotors with small rotor exit angles | |
| US7967571B2 (en) | Advanced booster rotor blade | |
| US8517677B2 (en) | Advanced booster system | |
| US8087884B2 (en) | Advanced booster stator vane | |
| EP2198167B2 (en) | Airfoil diffuser for a centrifugal compressor | |
| EP0083199B1 (en) | Surge control of a fluid compressor | |
| CN109505790B (en) | High-load high-through-flow-capacity axial flow fan | |
| US8038409B2 (en) | Turbomachine with rotors of high specific energy transfer | |
| CN216589292U (en) | Centrifugal compressor stage serial diffuser | |
| CN112283160B (en) | Compressor rotor blade and design method thereof | |
| CN113719459A (en) | Mixed-flow compressor for hundred thousand-twenty thousand cubic meter grade air separation device | |
| CN114607641A (en) | Axial fan's stator structure and axial fan | |
| CN109630469B (en) | Blade tandem diffuser and control method thereof | |
| CN220452231U (en) | Nine-stage axial flow compressor for medium-sized blast furnace | |
| CN113958537B (en) | Compressor and aircraft engine | |
| CN216429979U (en) | Mixed-flow compressor for hundred thousand-twenty thousand cubic meter grade air separation device | |
| CN113803274B (en) | Axial compressor and turbofan engine | |
| CN220101610U (en) | Blade structure of axial flow compressor for medium-sized blast furnace | |
| CN112989500B (en) | Inlet flow-dividing stability-expanding design method suitable for contra-rotating lift fan | |
| CN116753190B (en) | Tandem centrifugal compressor impeller with middle static blade grid | |
| CN112049818B (en) | Compressor and compressor blade | |
| CN116857225A (en) | Centrifugal blade with spanwise parabolic thickness distribution and design method | |
| CN110005644B (en) | Axial flow compressor stator with middle casing | |
| KR20220015700A (en) | Blade, compressor and gas turbine having the same |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |