CN112922899A - Axial-flow compressor rotor blade - Google Patents
Axial-flow compressor rotor blade Download PDFInfo
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- CN112922899A CN112922899A CN202110170169.1A CN202110170169A CN112922899A CN 112922899 A CN112922899 A CN 112922899A CN 202110170169 A CN202110170169 A CN 202110170169A CN 112922899 A CN112922899 A CN 112922899A
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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
- F04D29/38—Blades
- F04D29/384—Blades characterised by form
- F04D29/386—Skewed blades
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Abstract
The invention discloses an axial-flow compressor rotor blade, which is integrally a streamline profile, wherein the profile consists of an inner curved surface and an outer curved surface opposite to the inner curved surface, and comprises a lower half blade, a middle part and an upper half blade along the main axis of the blade; the inner curved surface is concave, and the concave degree of the inner curved surface is gradually reduced from bottom to top; the outer curved surface is convex, and the convex degree of the outer curved surface is gradually reduced from bottom to top; the middle section has zero cross-sectional curvature. According to the invention, the curvature continuity of the camber line in the blade profile section is ensured by reasonably optimizing and matching the air flow angle matching of the inlet and outlet of the movable blade and the second-order derivative of the camber line in the blade profile, so that the pneumatic load of the blade is recombined, the separation flow of a shock wave structure and the shock wave in the movable blade channel and the secondary flow in the angle area of the suction surface of the blade are effectively controlled, the pneumatic loss of a diffusion cascade is reduced, and the stable working range of the compressor is expanded; the obtained novel structure of the rotor blade has the practical requirements of both pneumatics and strength, and has good engineering application value.
Description
Technical Field
The invention relates to the field of aerodynamic design of impeller machinery, in particular to an axial-flow compressor rotor blade.
Background
The development requirement of the high thrust-weight ratio of the modern aviation power system puts extremely high requirements on a main pneumatic component, namely a fan/compressor. This requires that the fan/compressor not only have a large flow, a high stage pressure ratio and efficiency, but also have a wide stable operating range under all conditions. The significant improvement of the load level of the fan/compressor unit inevitably leads to the increase of the Mach number of the rotor blade tip, and simultaneously leads to the large-scale separation of the boundary layer due to the expansion pressure and the strong shearing action generated by the aggregation of the shock wave and the low-energy fluid in the angle area of the suction surface of the blade. Therefore, the pneumatic performance of the compressor under the non-working condition cannot be guaranteed, and the stall margin even at the designed rotating speed is seriously influenced. It is clear that conventional pneumatic design techniques have not been adequate for this difficult design task. An advanced pneumatic optimization design technology of the gas compressor under the condition of coupling of complex multi-factors such as gas viscosity, three-dimension, shock wave and vortex control and the like is urgently needed to be explored, and the gas compressor still has a high pneumatic efficiency level and a stall margin which can be practically utilized under the condition of high load and high flow.
Based on the idea, a plurality of advanced pneumatic design concepts and design systems are generated. The concept and design method of curved vane based on boundary layer migration theory are provided in the first 60 th century by the professor of the Wangzhongqi, the curved vane technology has been widely concerned with the end wall effect characteristic of personalized control of boundary layer low-energy fluid radial migration, and the curved vane technology has been widely applied to the design of stator and even rotor vane of advanced compression system. As early as 1974 through 1977, the NASA-Lewis research center developed a high speed silent fan QF-12 using the more advanced swept blade technology at that time. Wennerstrom first introduced swept blade technology into the aerodynamic design of a high load, high through-flow transonic fan. In order to study the aerodynamic design concept and effect of blade sweep, the Wright laboratory in the United states carries out a series of design and experimental study work of sweep-shaped rotors. The Denton et al also conducted numerical studies on aerodynamic performance and flow field structure associated with curved/swept blades, and the results of the studies show that curved/swept blades have a significant effect on the stall margin of the fan. In recent years, Beijing university of aerospace has designed and tested single stage swept-back pitch loaded fans ATS-2 and J285 with pressure ratios as high as 2.2 and 2.3, respectively. Along with the gradual increase of the load level of the fan/compressor stage, under the design requirement of approximate axial air outlet of a diffuser cascade, the stator is inevitably caused to present the personalized characteristic of a large deflection angle, so that the stronger boundary layer three-dimensional flow separation of a stator end area is caused, and the expansion of the load level and the stall margin of the fan/compressor is seriously influenced. Research results of Breugelmans and Shang show that reasonable application of the bent and swept stator blade technology can effectively reduce the stall of the angular region. The research results of Weingold et al on a three-stage compressor show that under the limiting conditions of blocking flow and stall margin, the bow-shaped stator can increase the pressure ratio and efficiency of the three-stage compressor. Gummer et al also show that forward-swept and forward-swept transonic stators can effectively reduce flow loss in the end regions and expand the effective working range of the stators through numerical values and experimental research results of the curved and swept transonic stators.
However, due to the limitations of the current technical conditions, many research results only partially break through some aspects and even a single technology, and even are limited to the mechanism research of the technology. However, it is well established that the importance of blade geometry on aerodynamic performance is well recognized. At present, the requirement of the pneumatic design is diversified, a set of pneumatic optimization design and blade three-dimensional geometric pneumatic forming system which can solve the practical engineering problem, is efficient and has strong fault-tolerant capability is urgently needed so as to meet the requirement of the development of the current impeller mechanical design technology. For this reason, many research institutes and scholars have been struggling with them. Benini and JANG, etc. respectively carry out pneumatic optimization design research on an NASARotor37 rotor, take the complicated three-dimensional viscous flow characteristics inside a cascade into consideration, achieve the purposes of controlling the shock wave structure of the suction surface of the blade and the size of a separation area through the change of the blade geometry, and obtain good effects. Lian et al take the NASAROTOR67 rotor as a blue book, and carry out pneumatic optimization design research on the transonic blades by adopting a plurality of optimization algorithms including a multi-objective genetic algorithm. The research result shows that the pressure ratio and the heat insulation efficiency of the rotor are improved to different degrees. In addition, research results also show that the boundary layer flow of the transonic rotor surface is very sensitive to the change of the geometric parameters of the blade, and especially in the blade tip section with a high incoming flow Mach number, the small change of the section geometry can bring about the huge improvement of the near-wall pressure distribution.
Aiming at the current high-load fan/compressor pneumatic design technology, under the limiting conditions of higher load level and through-flow capacity, how to meet the requirement of the lowest stall margin (SM > 10% -15%) becomes a bottleneck of the technology. Ellbrant et al developed an optimization design study that trades aerodynamic efficiency against stall margin. The total pressure loss coefficient (2D)/polytropic efficiency (3D) and the static pressure rise coefficient are taken as objective functions, and a pareto front optimization method for solving quasi-three-dimensional and full-three-dimensional NS equations is combined. Research results show that aerodynamic efficiency and stall margin are a pair of contradictions. If the aerodynamic efficiency requirement is too high, the desired stall margin must not be achieved. Therefore, in the engineering design process of the actual compressor, the two technical indexes of the aerodynamic efficiency and the stall margin must be considered in a compromise mode. And an Oyama and the like coupled three-dimensional NS equation solver establishes a high-precision optimal design system, and the system is used for carrying out optimal design research on the NASCATOR 67 rotor. The results of the study show that the adiabatic efficiency increases by 1.783% at a constant flow to pressure ratio. Finally, it is worth mentioning that in recent years, the MarkTurner-led topic group of the university of cincinnati establishes a three-dimensional CFD coupled turbomachinery aerodynamic optimization design system based on blade geometric curvature control. The obvious characteristic of the optimization design system is that the curvature of the key geometric parameter of the blade, which affects the channel flow field, is associated with the aerodynamic performance of the blade cascade, and the association between the geometric characteristic of the blade and the flow field parameter is explored and constructed. The above-mentioned optimization design system has three important features, one of which is to couple a CFD solution system. Secondly, the engineering practicability is emphasized, and not only the breakthrough of a single technology. Thirdly, the importance of the blade channel geometry is more and more emphasized.
In summary, the method is based on the optimization design idea of flow matching between blade rows and controlling the camber line curvature of the blade profile, the front 1.5 stages of a high-pressure compressor of an E3 engine of GE company are used as blueprints, and the optimization design is carried out on rotor blades of the engine under the stage environment under the condition of ensuring the strength of the blades so as to obtain the overall improvement of comprehensive aerodynamic performance indexes such as through-flow capacity, aerodynamic efficiency, pressure ratio, stall margin and the like.
Disclosure of Invention
The invention provides a rotor blade structure of an axial-flow compressor, aiming at solving the problem that the aerodynamic performance and the strength of the blade are difficult to simultaneously consider in the prior related art.
The specific technical scheme of the invention is as follows:
the axial-flow compressor rotor blade is integrally a streamline profile which is composed of an inner curved surface and an outer curved surface opposite to the inner curved surface, and comprises a lower half blade, an intermediate part and an upper half blade along the main axis; the inner curved surface is concave, and the concave degree of the inner curved surface is gradually reduced from bottom to top; the outer curved surface is convex, and the convex degree of the outer curved surface is gradually reduced from bottom to top; the intermediate portion has zero cross-sectional curvature.
Preferably, the height of the lower half blade accounts for 40-90% of the height of the blade.
Preferably, the section barycenter of the blade along the length direction is overlapped to obtain a three-dimensional blade.
According to the invention, through reasonably optimizing matching of the angle matching of the air flow of the inlet and the outlet of the movable blade and the second-order derivative of the camber line of the blade profile, the curvature continuity of the camber line of the blade profile section is ensured, thereby recombining the pneumatic load of the blade, effectively controlling the shock wave structure in the movable blade channel, the separated flow after the shock wave and the secondary flow of the blade suction surface angle region, reducing the pneumatic loss of a diffusion cascade and expanding the stable working range of the compressor; the obtained novel structure of the rotor blade gives consideration to the actual requirements of both pneumatics and strength, reflects the actual design requirement that the engine blade has no additional function, and has good practical application value.
Drawings
FIG. 1 is a front view of a blade of the present invention;
FIG. 2 is a cross-sectional comparison of the present invention and the prior art;
FIG. 3 is a cross-sectional centroid stacking diagram of the present invention;
FIG. 4 is a graph of the mean camber line and first and second derivative characteristics of a representative cross-section of the present invention;
FIG. 5 is a plot of compressor performance versus lobe height for the present invention in comparison to the prior art, wherein (a) efficiency; (b) the total pressure ratio.
Detailed Description
In order to make the aforementioned objects, features and advantages of the present invention comprehensible, embodiments accompanied with figures are described in detail below.
In the description of the present invention, it should be noted that, in the description of the embodiments of the present application, the description of the term "some specific embodiments" means that a specific feature, structure, material, or characteristic described in connection with the embodiments or examples is included in at least one embodiment or example of the present invention. Throughout this specification, the schematic representations of the terms used above do not necessarily refer to the same implementation or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments or examples.
Referring to fig. 1, the present invention provides an axial compressor rotor blade structure, which is a streamline profile as a whole, the profile is composed of an inner curved surface 1a and an outer curved surface 1b opposite to the inner curved surface 1a, and includes a lower half blade 11, an intermediate portion 12, and an upper half blade 13 along a main axis; the inner curved surface 1a is concave, and the concave degree is gradually reduced from bottom to top; the outer curved surface 1b is convex, and the convex degree is gradually reduced from bottom to top; the intermediate portion 12 has zero cross-sectional curvature.
In the invention, the inner curved surface 1a faces the air outlet, the outer curved surface 1b is positioned on one side of the air inlet, and the height H of the lower half blade 11 accounts for 40-90% of the height H of the blade. Further, the section of the blade along the length direction adopts a gravity center stacking structure to form a three-dimensional blade.
Specifically, the maximum deflection position of the camber line of the blade profile is located after 50% of the chord length of the blade; the mean camber line of the upper half blade section at the inlet portion is a partial negative curvature. The slope and the curvature of the camber line of the blade profile are smoothly distributed along the chord direction.
The combination of aerodynamic and strength properties specifically obtained by the present invention is illustrated below by aerodynamic inspection, strength analysis and verification.
One, pneumatic inspection
In the prior art, the main design parameters of the first 1.5 stages of the E3 high-pressure compressor are shown in a table 1, and the performances of the blade structure can be shown in a table 2.
TABLE 1 Main design parameters
TABLE 2 comparison of maximum efficiency point characteristics of compressor before and after optimization
Table 2 shows the compressor characteristics before and after the control curvature optimization. Research results show that the gradient of diffusion flow field parameters on the surface of the blade is effectively controlled by controlling the curvature of the geometric curve of the section of the blade, and the overall aerodynamic performance of the compressor is remarkably improved. Compared with the prototype scheme, under the optimized design working condition, the adiabatic efficiency and the pressure ratio of the gas compressor are respectively increased by 1.67 percent and 0.223 percent. The clogging flow rate was increased by 0.537%. The stall flow is reduced by 1.54%, so that the stall margin of the compressor is increased, and the stable working range of the compressor is effectively expanded. In addition, compared with the prior inlet and outlet geometric angle matching scheme, the blocking flow is reduced, but the pneumatic efficiency, the pressure ratio and the stall margin are greatly improved instead.
FIG. 2 shows a cross-sectional profile comparison analysis of the blade profiles before and after optimization of the controlled curvature. Research results show that the optimized blade geometry shows that the blade profile turning angle below 40% of the blade span is reduced, which is beneficial to reducing the aerodynamic load of the lower half blade span of the blade and improving the flow of the angular area of the suction surface of the movable blade, thereby reducing the flow loss and improving the aerodynamic efficiency. In the 40% -90% blade span area, the blade profile folding angle is increased, the characteristics of high airflow tangential speed and strong airflow working capacity of the blade in the transonic flow area are fully utilized, the pneumatic load of the upper half blade span is increased, and the loss of the pneumatic load of the lower half blade span is made up to a certain extent. In the area, the fluid dynamics mechanism of the curvature control optimization design is that the inverse pressure gradient of transonic flow is adjusted through reasonably controlling the curvature of the geometric profile of the section, so that the pressure distribution on the surface of the blade group is reduced, and the shock wave intensity is reduced. In the blade tip part, in order to reduce the shock wave loss, the blade profile presents a pre-compression characteristic, and the maximum deflection position of the camber line of the blade profile is positioned behind 50% of the chord length of the blade (referring to fig. 3, the section barycenter of the blade along the length direction is overlapped to obtain a three-dimensional blade), which is consistent with the core idea of the existing transonic high-load movable blade design theory.
For further analysis of the geometric characteristics of the optimized forward and aft airfoil sections and the advantages of the aerodynamic optimization technique herein, FIG. 4 shows the variation characteristics of the mean camber line and its first and second derivatives in the fundamental airfoil plane for a typical section. Research results show that the slope and the curvature of the camber line of the blade profile are smoothly distributed along the chord direction, and are effectively controlled, so that the expected controllable curvature target is realized. Because the Mach number of the root is low, a large work adding amount is required, and therefore the blade profile turning angle of the root of the movable blade is large. And the Mach number of the blade tip part is higher, the supercharging effect of partial airflow mainly comes from shock waves, and meanwhile, in order to control flow separation possibly caused after the shock waves, the blade profile turning angle of the blade tip part of the movable blade is smaller. Even for effective reduction of wavefront mach number and thus shock intensity, reverse curvature (pre-compressed blade profile) designs have been developed, i.e. local negative curvature of the mean camber line near the inlet. Meanwhile, near the transonic section, the slope and curvature of the camber line in the blade profile vary greatly along the chord direction, such as 75% spanwise section, due to qualitative changes in the flow physical characteristics.
FIG. 5 shows the adiabatic efficiency and total pressure ratio distribution along the spanwise direction. Research results show that the pressure ratio characteristic of the curvature control optimization scheme is represented as follows: the 40% span is reduced while the remaining span is increased. The reason is that after the optimized design, the deflection angle of the airflow below 40% of the blade span is reduced, and the deflection angle of the airflow is increased at other span positions, and the shock wave intensity and the flow field quality after the shock wave are reasonably improved. In addition, through the adjustment of the geometric curvature of the molded surface, the gradient of the diffusion flow field parameters is effectively controlled, and the control on the characteristics of the shock wave expansion pressure flow field in the blade tip area is ideal. Therefore, the adiabatic efficiency of the optimized scheme for controlling the curvature is improved in a substantially full-span range, and the average amplification is larger than 1%.
Second, intensity analysis and check
The design principle of the rotor blade can be summarized as follows: within the limits of mechanical structure, strength, material properties, etc., a minimum loss aerodynamic profile of the desired aerodynamic load distribution is constructed. Therefore, the blade strength analysis and verification are indispensable important links in the design of the engine blade. In order to verify the reasonability and practicability of the novel rotor blade structure obtained by the invention, the CSD technology is adopted to carry out strength analysis on the prototype and the novel rotor blade structure. Table 3 gives the main material properties of the blade.
TABLE 3 blade Material Properties
The stress distribution test of the rotor blade shows that the maximum stress of the root of the prototype rotor blade is 7.3251 multiplied by 108pa, the maximum stress of the root of the optimized rotor blade is 5.985 multiplied by 108pa which is far smaller than the prototype rotor scheme, and the stress distribution of the surface of the blade also tends to be reasonable. The result shows that the novel rotor blade structural layout obtained by the invention has better pneumatic effect and the strength of the blade root is superior to that of the original structure.
The above description is only for the preferred embodiment of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art should be able to cover the technical solutions and the inventive concepts of the present invention within the technical scope of the present invention.
Although the present disclosure has been described above, the scope of the present disclosure is not limited thereto. Various changes and modifications may be effected therein by one of ordinary skill in the pertinent art without departing from the spirit and scope of the present disclosure, and these changes and modifications are intended to be within the scope of the present disclosure.
Claims (3)
1. The axial-flow compressor rotor blade is integrally a streamline profile which is composed of an inner curved surface and an outer curved surface opposite to the inner curved surface, and is characterized by comprising a lower half blade, an intermediate part and an upper half blade along the main axis of the axial-flow compressor rotor blade; the inner curved surface is concave, and the concave degree of the inner curved surface is gradually reduced from bottom to top; the outer curved surface is convex, and the convex degree of the outer curved surface is gradually reduced from bottom to top; the intermediate portion has zero cross-sectional curvature.
2. The axial compressor rotor blade of claim 1, wherein the lower half blade height is 40-90% of the blade height.
3. The axial compressor rotor blade according to claim 1 or 2, characterised in that the cross section of the blade in the length direction is superimposed according to the centre of gravity to obtain a three-dimensional blade.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202110170169.1A CN112922899A (en) | 2021-02-05 | 2021-02-05 | Axial-flow compressor rotor blade |
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| CN202110170169.1A CN112922899A (en) | 2021-02-05 | 2021-02-05 | Axial-flow compressor rotor blade |
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Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB913620A (en) * | 1960-02-03 | 1962-12-19 | Szydlowski Joseph | Improvements in or relating to axial-flow compressors |
| CN101310112A (en) * | 2005-11-16 | 2008-11-19 | 西门子公司 | Impeller of radial compressor |
| CN102536327A (en) * | 2011-12-07 | 2012-07-04 | 北京航空航天大学 | Pneumatic and structural feature considered three-dimensional geometric structure of fan blade of aircraft engine |
| CN103256248A (en) * | 2012-02-21 | 2013-08-21 | 珠海格力电器股份有限公司 | Impeller and centrifugal compressor comprising same |
| DE102014226689A1 (en) * | 2014-12-19 | 2016-06-23 | Rolls-Royce Deutschland Ltd & Co Kg | Rotor blade of an axial flow machine |
| CN106250652A (en) * | 2016-08-16 | 2016-12-21 | 深圳友铂科技有限公司 | The construction method of a kind of compressor blade blade profile and compressor blade |
| CN106351872A (en) * | 2016-09-12 | 2017-01-25 | 深圳友铂科技有限公司 | Compressor rotor blade meeting both pneumatic and strength requirements |
-
2021
- 2021-02-05 CN CN202110170169.1A patent/CN112922899A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB913620A (en) * | 1960-02-03 | 1962-12-19 | Szydlowski Joseph | Improvements in or relating to axial-flow compressors |
| CN101310112A (en) * | 2005-11-16 | 2008-11-19 | 西门子公司 | Impeller of radial compressor |
| CN102536327A (en) * | 2011-12-07 | 2012-07-04 | 北京航空航天大学 | Pneumatic and structural feature considered three-dimensional geometric structure of fan blade of aircraft engine |
| CN103256248A (en) * | 2012-02-21 | 2013-08-21 | 珠海格力电器股份有限公司 | Impeller and centrifugal compressor comprising same |
| DE102014226689A1 (en) * | 2014-12-19 | 2016-06-23 | Rolls-Royce Deutschland Ltd & Co Kg | Rotor blade of an axial flow machine |
| CN106250652A (en) * | 2016-08-16 | 2016-12-21 | 深圳友铂科技有限公司 | The construction method of a kind of compressor blade blade profile and compressor blade |
| CN106351872A (en) * | 2016-09-12 | 2017-01-25 | 深圳友铂科技有限公司 | Compressor rotor blade meeting both pneumatic and strength requirements |
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