A design method for a multi-stage axial flow compressor without adjustable stator blades

Abstract

The invention provides a design method for a multi-stage axial flow compressor without adjustable stator blades, which belongs to the field of aerodynamic design of multi-stage axial flow compressors. The design method for a multi-stage axial flow compressor without adjustable stator blades comprises: step one, selecting an average stage load factor of the multi-stage axial flow compressor; step two, determining a stage load factor of the first stage according to the selected average stage load factor of the multi-stage axial flow compressor; step three, selecting an average Mach number at an inlet of a first-stage stator blade according to the results of step one and step two, and calculating a radial Mach number of the first-stage stator according to the average Mach number at an inlet of the first-stage stator blade; step four, determining a relative chord-width ratio of the first-stage stator according to the radial Mach number of the first-stage stator in step three; and step five, determining a large negative attack angle of the first-stage stator blade according to the relative chord-width ratio of the first-stage stator.

Description

Technical Field

[0001] The invention relates to the field of aerodynamic technology of multi-stage axial flow compressors, and in particular to a design method of a multi-stage axial flow compressor without adjustable stator blades. Background Art

[0002] For the multi-stage axial flow compressor (hereinafter referred to as "compressor") currently used in engineering, when the compressor flow path and blade geometry are determined according to the design state, each stage has a unique inherent universal characteristic. When the speed decreases and the compressor deviates from the design speed, the front stage usually tends to stall and the rear stage tends to be blocked, resulting in a "front surge and rear blockage" phenomenon, which seriously affects the surge margin of the compressor. In order to alleviate this "front surge and rear blockage" phenomenon at low speeds of the compressor, the "adjustable stator" technology is usually used in engineering design, that is, by designing an adjustment mechanism, the installation angle of the stator blades of the first few stages of the compressor can change as the speed decreases, thereby preventing surge and improving the low-speed performance of the compressor. At present, this technology has been widely used in the design of aircraft engine compressors and has achieved good results.

[0003] However, with the large-scale application of new cruise missiles and drones, there is an urgent need to reduce the cost of engines. The adjustable stator adjustment mechanism of the compressor component is relatively complex and accounts for a high cost of the compressor component. Therefore, in the design of new low-cost engines, the cost is reduced by eliminating the adjustable stator adjustment mechanism of the compressor, that is, the compressor adopts a non-adjustable stator design.

[0004] For compressors with non-adjustable stator blades, in order to alleviate the "front gasping and back blocking" phenomenon at low speeds, special consideration must be given to the selection of aerodynamic design parameters, and their design parameters are significantly different from those of conventional compressors with adjustable stator blades. The present invention mainly describes the selection characteristics of the design parameters of compressors without adjustable stator blades relative to conventional compressors with adjustable stator blades. Summary of the invention

[0005] In view of this, the present invention provides a design method for a multi-stage axial flow compressor without adjustable stator blades, so as to achieve the purpose of aerodynamic parameter design of the compressor without adjustable stator blades.

[0006] The present invention provides the following technical solutions: a design method for a multi-stage axial compressor without adjustable stator blades, comprising: step one, selecting an average stage load factor of the multi-stage axial compressor; step two, determining the stage load factor of the first stage according to the selected average stage load factor of the multi-stage axial compressor; step three, selecting the average Mach number at the inlet of the first-stage stator blade according to the results of step one and step two, and calculating the radial Mach number of the first-stage stator according to the average Mach number at the inlet of the first-stage stator blade; step four, determining the relative chord-width ratio of the first-stage stator according to the radial Mach number of the first-stage stator in step three; step five, determining the large negative attack angle of the first-stage stator blade according to the relative chord-width ratio of the first-stage stator.

[0007] Furthermore, through the formula Calculate the average stage load factor for a multi-stage axial compressor, where is the i-th level of wheel merit, is the tangential speed of the blade tip at the first-stage rotor inlet, is the tangential velocity of the blade tip at the inlet of the nth stage rotor.

[0008] Furthermore, the average load factor of the multi-stage axial compressor The value range is .

[0009] Furthermore, through the formula Calculate the stage loading factor for the first stage, where It is the correction factor of the first-level load factor relative to the average load factor.

[0010] Furthermore, the correction factor of the first-level load factor relative to the average load factor is The value range is 1.15~1.25.

[0011] Furthermore, through the formula:

[0012] Calculate the Mach number of the first-stage stator along the radial direction, where is the relative height of the blades at the first stage stator inlet, is the average Mach number at the first-stage stator inlet.

[0013] Furthermore, step four is specifically as follows:

[0014] A three-dimensional simulation is performed based on the radial Mach number of the first-stage stator to determine the increase in the chord length of the first-stage stator blade tip. ;

[0015] According to the formula , calculate the relative chord aspect ratio, where is the height of the first stage stator inlet blades.

[0016] Furthermore, step five is specifically as follows:

[0017] According to the relative chord-aspect ratio of the first-stage stator, a three-dimensional simulation is performed to determine the average inlet angle of attack of the first-stage stator. ;

[0018] By formula , calculate the large negative angle of attack of the first stage stator blades.

[0019] Compared with the prior art, the beneficial effects that can be achieved by at least one of the above technical solutions adopted by the present invention include at least:

[0020] Firstly, in the one-dimensional scheme design, by defining the average load factor of the multi-stage compressor, the selection range of the average stage load factor of the compressor without adjustable static blades is determined, and the distribution law of the first stage load factor is stipulated;

[0021] Secondly, in the S2 stream surface design stage, the matching parameter selection ranges of the first-stage stator blade inlet Mach number and the first-stage blade tip to stator chord length are specified in turn;

[0022] Finally, in the blade modeling design stage, the selection characteristics of the arc angle and attack angle in the first-stage stator blade modeling parameters are given, and finally a design method suitable for compressors without adjustable stator blades is formed. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0024] Figure 1 It is a schematic diagram of the process of the present invention;

[0025] Figure 2 This is a schematic diagram of the Mach number distribution at the first-stage stator inlet of the compressor;

[0026] Figure 3 This is a schematic diagram of the compressor meridian flow passage;

[0027] Figure 4 is the dimensionless geometric turning angle of the middle arc line of the first stage rotor tip of the compressor;

[0028] Figure 5 This is a schematic diagram of the distribution of the attack angles of the first stage stator blades of the compressor;

[0029] Figure 6 This is the characteristic diagram of a three-stage compressor. DETAILED DESCRIPTION

[0030] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0031] The following describes the implementation methods of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the following embodiments and the features in the embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0032] like Figure 1 As shown, an embodiment of the present invention provides a design method for a multi-stage axial flow compressor without adjustable stator blades, which specifically includes the following steps:

[0033] Step 1: Selection of average load factor.

[0034] The most important task in the one-dimensional design of the compressor is to select the average stage load factor, because it is an important parameter that determines the overall aerodynamic load level and the number of stages of the compressor. The average stage load factor of a multi-stage axial flow compressor is defined as:

[0035] (1)

[0036] In formula (1): is the i-th level of wheel merit, is the tangential speed of the blade tip at the first stage rotor inlet, It is the tangential speed of the blade tip at the rotor inlet of the nth stage (the nth stage is the last stage of the compressor).

[0037] A lot of engineering design experience shows that for compressors without adjustable stator blades, the average load factor of the stage is It cannot be selected too high, because too high aerodynamic load will cause flow separation to occur early, aggravate the "front surge and back blockage" at low speed, and reduce the compressor surge margin. The maximum value must be limited. This engine specifies the average load factor The maximum value is limited to 0.32. The value is less than 0.32, and the compressor without adjustable static blades often has a good low-speed surge margin. When it is greater than 0.32, compressors without adjustable static blades often have poor surge margin at low and medium speeds.

[0038] Step 2: Selection of the first-level load factor.

[0039] For a compressor without adjustable static blades, the rear stage always works in a blocked state with a low pressure ratio when working at low speed. At this time, to ensure that the compressor has sufficient boosting capacity, the front stage of the compressor needs to have a sufficient pressure ratio. Therefore, the first stage load factor of the compressor needs to be higher than the average stage load factor to ensure that it has sufficient boosting capacity at low speed. Studies have shown that: the first stage load factor Take 1.15-1.25 times the average load factor , compressors without adjustable stator blades often have higher low-speed boost capability and higher surge margin.

[0040] Step 3: Selection of the inlet Mach of the first-stage stator blades.

[0041] Usually, when a compressor without adjustable stator blades is running at low speed, the first-stage stator blades tend to stall first. Therefore, in order to improve the low-speed surge margin, the primary goal is to delay the occurrence of stall of the first-stage stator blades to ensure that there is sufficient margin. According to the aerodynamics of the impeller, the anti-stall ability of the blades is closely related to the available angle of attack range. The lower the blade inlet Mach, the larger the available angle of attack range and the stronger the blade's anti-stall ability. Based on the above theory, in the S2 flow surface design link, the Mach of the first-stage stator blades must be smaller than that of conventional adjustable stator blades to ensure that they have sufficient anti-stall ability. Studies have shown that for compressors with non-adjustable stator blades, the Mach of the first-stage stator blades should be less than 0.7. For example, Figure 2 shown.

[0042] Step 4: Matching the chord lengths of the first-stage rotor and stator blade tips.

[0043] The results of simulation and experimental research show that the specific area where the first stator blade tip occurs stall in a low-speed compressor with non-adjustable stator blades is the tip of the first-stage stator blade. In order to increase the anti-stall capability of this area, increasing the chord length of the tip of the first-stage stator blade is an effective means. Figure 3 As shown in the figure, during the design, the overall anti-stall capability of the first stage can be improved, the occurrence of stall can be delayed, and the low-speed surge margin of the compressor can be improved by increasing the chord length of the first-stage static blade tip (i.e., the stator forward sweep design) and appropriately reducing the chord length of the first-stage rotor tip. The specific increase in the chord length of the first-stage stator tip can be determined by defining the relative chord-width ratio To quantify, it is specifically defined as:

[0044] (2)

[0045] In formula (2), if Figure 3 As shown, is the increase in the chord length of the tip of the first-stage stator blade, is the height of the first-stage stator inlet blade. Studies have shown that for compressors with non-adjustable stator blades, The selection range is usually between 0.13-0.17.

[0046] Step 5: Design a large forward curved blade shape at the tip of the first-stage rotor.

[0047] In the aerodynamic design of compressors, the blade curvature distribution is usually characterized by defining the dimensionless geometric turning angle of the mid-arc line. The dimensionless geometric turning angle of the mid-arc line is the ratio of the geometric turning angle of a point on the mid-arc line of the blade relative to the leading edge of the blade to the total turning angle. Since the Mach number at the tip of the first-stage rotor blade is high (usually supersonic), in conventional adjustable compressor design, the first 20% chord length of the blade ( Figure 4 The bending angle of the blade tip (the area of ​​dimensionless chord length 0-0.2) is designed to be smaller to obtain higher blade efficiency. However, such a design will result in a small range of available angle of attack of the blade. For compressors with adjustable stator blades, this problem can be solved by adjusting the angle of the 0th stage stator so that the tip of the first-stage rotor always works within the optimal angle of attack range. However, for compressors with non-adjustable stator blades, when working at low speed, the angle of attack of the tip of the first-stage rotor will definitely deviate greatly. Therefore, the available angle of attack range of the tip of the first-stage rotor must be designed to be large enough to ensure that it is still within the available angle of attack range under low speed conditions. Studies have shown that increasing the curvature within the first 20% chord length of the blade can significantly increase the available angle of attack range of the blade shape, which can be quantified as follows: the first 20% chord length gives 20% of the total curvature, that is, the dimensionless geometric turning angle of the arc line in the first 20% chord length is about 0.2.

[0048] Design of large negative attack angle of first stage stator blades

[0049] like Figure 5 As shown in the figure, the angle of attack of the first-stage stator blades of a conventional compressor with adjustable stator blades is generally between 0° and -5°, showing an inverted C-shaped distribution, with a small angle of attack at the root tip and a large angle of attack at the middle section of the blade. The angle of attack of the tip section is generally between -2° and -5°. However, for compressors with non-adjustable stator blades, the inlet airflow angle of the first-stage stator blades will seriously deviate from the conventional design range at low speeds, so the angle of attack of the first-stage stator blades needs to be designed to be smaller. Specifically, it can be quantified as follows: the overall angle of attack range is between -3° and -9°, and the blade tip angle of attack is more reasonably selected between -6° and -9°.

[0050] The following is a further detailed description using a 3-stage non-adjustable compressor as an example:

[0051] A three-stage compressor adopts a non-adjustable stator blade design. In order to alleviate the "front gasping and back blocking" phenomenon at low speed and improve the margin for medium and low speeds, the key design parameters are selected according to the five steps described in the present invention.

[0052] Table 1--Selection of key design parameters for 3-stage stationary blade non-adjustable compressor

[0053]

[0054] According to the five steps in Table 1, the key parameters of the aerodynamic design of the three-stage fixed vane non-adjustable compressor are selected to complete the design of the modified compressor. The final compressor characteristics are as follows: Figure 6 As shown in the figure, it can be seen that between the low speed of 0.7-0.8, the compressor surge boundary is much higher than the engine working line, indicating that the compressor has sufficient surge margin in this speed range. It also proves the effectiveness of the aerodynamic design method of a compressor without adjustable static blades described in the present invention.

[0055] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A design method for a multi-stage axial flow compressor without adjustable stator blades, characterized in that: include: Step 1, selecting the average stage load factor of the multi-stage axial flow compressor; Step 2: determining the stage load factor of the first stage according to the average stage load factor of the selected multi-stage axial flow compressor; Step 3: According to the results of step 1 and step 2, the average Mach number at the inlet of the first-stage stator blades is selected, and the Mach number of the first-stage stator in the radial direction is calculated according to the average Mach number at the inlet of the first-stage stator blades; Step 4: Perform a three-dimensional simulation based on the radial Mach number of the first-stage stator to determine the increase in the chord length of the first-stage stator blade tip. According to the formula , calculate the relative chord aspect ratio, where is the height of the first stage stator inlet blade; Step 5: Perform a three-dimensional simulation based on the relative chord-aspect ratio of the first-stage stator to determine the average inlet angle of attack of the first-stage stator. ; Through the formula , calculate the large negative angle of attack of the first-stage stator blades.

2. The design method of a multi-stage axial flow compressor without adjustable stator blades according to claim 1, characterized in that: By formula Calculate the average stage load factor for a multi-stage axial compressor, where is the i-th level of wheel merit, is the tangential speed of the blade tip at the first stage rotor inlet, is the tangential velocity of the blade tip at the inlet of the nth stage rotor.

3. The design method of a multi-stage axial flow compressor without adjustable stator blades according to claim 2, characterized in that: Average load factor of multi-stage axial flow compressor The value range is .

4. The design method of a multi-stage axial flow compressor without adjustable stator blades according to claim 3 is characterized in that: By formula Calculate the stage loading factor for the first stage, where It is the correction factor of the first-level load factor relative to the average load factor.

5. The design method of a multi-stage axial flow compressor without adjustable stator blades according to claim 4, characterized in that: Correction factor of the first level load factor relative to the average load factor The value range is 1.15~1.

25.

6. The design method of a multi-stage axial flow compressor without adjustable stator blades according to claim 5, characterized in that: By formula: Calculate the Mach number of the first-stage stator along the radial direction, where is the relative height of the blades at the first stage stator inlet, is the average Mach number at the first-stage stator inlet.

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