Dynamic differential pressure blade of gas compressor of gas turbine

Abstract

The invention relates to a dynamic differential pressure blade of a gas compressor of a gas turbine. A blade airfoil comprises an upper airfoil surface and a lower airfoil surface. The section lines of the airfoils are all composed of curves of a monotonically decreasing function from the front edge to the rear edge. The structure is simple, and when the lower wing surface does work on windward airflow, the flow speed is reduced, the static pressure intensity is increased, and upward thrust is generated; the upper airfoil does not do work, airflow flows back, static pressure is reduced, upward pulling force is generated, and therefore the double-airfoil lift airfoil profile with the upper airfoil and the lower airfoil simultaneously and consistently upward lift force is formed, and the airfoil efficiency is further improved by optimizing the lift-drag ratio. A front edge point is accurately positioned by adopting a wing section line rear corner, a spanwise line, a relative velocity line and a backward moving distance, a shock wave is weakened to be low-sound-velocity object surface static pressure, and the adverse effect of the shock wave is eliminated.

Description

Technical Field

[0001] The present invention belongs to the technical fields of industry and shipbuilding, and relates to gas turbine compressor blades, such as industrial gas turbine compressor blades, power generation gas turbine compressor blades, marine gas turbine compressor blades, gas turbine compressor blades for pumps and fans, etc. Background Art

[0002] In the existing industry and shipbuilding, axial-flow gas turbine compressor blades play a role in axially pushing the air flow, pushing the air flow backward axially to do work and achieving the effect of increasing air pressure. The planform of its airfoil is the same as that of the wing of a fixed-wing aircraft. The lower surface of the airfoil is flat, while the upper wing surface is upwardly curved. At a zero angle of attack, lift is provided by the principle of lift due to velocity difference. The flow velocity on the upper wing surface is fast, and the static pressure drops, forming a single-sided upward lift. In addition, when the lower wing has an angle of attack, the flow velocity will slow down, the static pressure will increase, and the air flow will generate a thrust on the wing surface, with the direction upward. If the lift on the upper and lower wing surfaces can be superimposed and manifested simultaneously, the efficiency of the airfoil will be improved.

[0003] In the Chinese patent "202121320870.9 A Ship-Moving Differential Pressure Propeller", the principle of lift due to hydrodynamic differential pressure of incompressible water fluid is applied to achieve an airfoil that can generate lift on both the upper and lower wing surfaces simultaneously, and the efficiency is significantly improved. Therefore, it is promoted and applied to compressible air fluid, so that the differential pressure blade airfoil of the gas turbine compressor also has a double-sided lift wing surface, which can also improve its efficiency. Summary of the Invention

[0004] The purpose of the present invention is to provide a double-sided lift airfoil and a differential pressure blade for a gas turbine compressor that can improve the propulsion efficiency of the gas turbine compressor blade, so as to solve the above problems.

[0005] In order to achieve the above purpose, the technical solution of the present invention is realized as follows:

[0006] A differential pressure blade for a gas turbine compressor, the sectional planform of which includes an upper wing surface and a lower wing surface, characterized in that the cross-section lines of the upper wing surface and the lower wing surface are both curved surfaces composed of monotonically decreasing curves from the leading edge to the trailing edge.

[0007] The main wing surface of the upper wing surface is a descending curved surface, and the curved surface is composed of a monotonically decreasing function. That is, from the leading edge to the trailing edge, the height of the cross-section line of the curved surface in the vertical direction decreases successively, and generally there is no inflection point.

[0008] The main wing surface of the lower wing surface is a descending curved surface, and the curved surface is composed of a monotonically decreasing function. That is, from the leading edge to the trailing edge, the height of the cross-section line of the curved surface in the vertical direction decreases successively, and generally there is no inflection point.

[0009] At near-sonic and supersonic speeds, the leading edge of the upper wing surface and the leading edge of the lower wing surface intersect directly.

[0010] At near-sonic and supersonic speeds, the trailing edge turning angle of the leading edge point of the lower wing surface section line is equal to or greater than the normal angle of the required leading edge point.

[0011] At near-sonic and supersonic speeds, the trailing edge turning angle of the lower wing surface section line from the leading edge to the trailing edge satisfies that the trailing edge turning angle of the latter point is equal to or greater than that of the former point.

[0012] At near-sonic and supersonic speeds, the trailing edge turning angle of the leading edge point of the spanwise line is equal to or greater than the normal angle of the required point at that location.

[0013] At near-sonic and supersonic speeds, the position of the leading edge point of the wing section line is determined by the relative velocity direction line and the value of the backward displacement distance.

[0014] Increasing the normal angle of the wing section line can increase the lift-to-drag ratio.

[0015] For a fixed-wing aircraft, the zero-degree chord of the airfoil of the dynamic pressure difference machine is collinear with the relative velocity direction line.

[0016] Compared with the prior art, the advantages of the embodiments of the present invention are as follows: The structure of the present invention is simple. When the lower wing surface does work on the oncoming air flow, the flow velocity slows down, the static pressure increases, and an upward thrust is generated on the wing surface; the air flow on the upper wing surface flows backward, the static pressure decreases, and an upward pulling force is generated on the wing surface. Thus, the lift forces on the upper and lower wing surfaces are both upward, and the sum of the two lift forces is greater than the lift force of the existing single-sided airfoil. After further optimizing the lift-to-drag ratio, the efficiency of the blade wing is further improved. At near-sonic and supersonic speeds of the air flow, the trailing edge turning angle of the wing section line is adopted, and the leading edge point of the spanwise line is accurately positioned by the relative velocity line and the backward displacement distance, weakening the shock wave into the static pressure of the low-sonic object surface and eliminating the adverse effects of the shock wave. Description of the Drawings

[0017] Figure 1 Schematic diagram of a horizontal pipe without thickness

[0018] Figure 2 Schematic diagram of a horizontal straight uniform flow

[0019] Figure 3 Schematic diagram for calculating the static pressure on the oncoming surface

[0020] Figure 4 Schematic diagram for calculating the static pressure on the backflow surface

[0021] Figure 5 Schematic diagram of the horizontal position of the impeller

[0022] Figure 6 Top view of the velocity triangle at point A

[0023] Figure 7 Projection diagram of the backward displacement distance of M

[0024] Figure 8 Schematic diagram of the backward displacement distance in the top view

[0025] Figure 9 Schematic diagram of the airfoil section of a fixed-wing aircraft

[0026] Figure 10 Horizontal symmetry diagram of the left part

[0027] Figure 11 Transition diagram after doubling the horizontal length

[0028] Figure 12 Schematic diagram of the new airfoil section at 75% length

[0029] Figure 13 Schematic diagram after optimization and improvement Specific implementation manners

[0030] The technical solutions of the embodiments of the invention will be clearly and completely described below in conjunction with the drawings. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts fall within the scope of protection of the present invention.

[0031] Embodiment

[0032] 1. Principle of lift of blades due to differential pressure in a gas turbine compressor

[0033] The principle of lift of blades due to differential pressure in a gas turbine compressor, the main inferences and conclusions can directly apply the principle of lift of blades due to differential pressure in water fluid, the main inferences and conclusions given in the Chinese patent "202121320870.9 A ship differential pressure propeller". Since water and air both belong to incompressible fluids (when the air velocity is equal to or greater than 0.2 times the speed of sound, air can be considered an incompressible gas), so replacing the water parameters with the parameters of incompressible air and making the same inferences for the concepts will do. When the air velocity is equal to or greater than 0.2 times the speed of sound, using compressible fluids, perfect gases (air is considered a perfect gas), isentropic, one-dimensional steady flow at high speed to expand the above formulas, inferences and conclusions, it can be applied to the airfoil of a low-speed fixed-wing aircraft with differential pressure, and then applied to the airfoil of the blades of a gas turbine compressor; at near-sonic and supersonic speeds, combined with the influence of shock waves, applying the trailing edge angle of the lower surface section line, the relative velocity line of the spanwise line and the position of the leading edge point after moving backward can weaken the strong static pressure of the shock wave on the surface to the static pressure of the surface at low speed of sound, thus resolving the adverse effects of the shock wave.

[0034] 2. Work done by the blades in the air flow field

[0035] Basic formula for work done in pipe fluidics

[0036] Assume that the irrotational ideal flow field is a steady flow, with only gravity in the body force, flowing along the elementary stream (streamline) and being an incompressible gas fluid. By integrating the motion differential equation along the streamline and adding the input mechanical energy, such as that from a fan, etc., the Bernoulli equation for the total flow is obtained as follows:

[0037]

[0038] For a straight, thin, smooth circular pipe, horizontally placed in the gas flow field, the position term can be neglected, and we get:

[0039]

[0040] In the formula: z is the height measured from a certain reference plane; u is the flow velocity; p is the static pressure; g is the acceleration due to gravity; ρ is the density. The subscript 1 represents the initial cross-section, 2 represents the final cross-section, H m is the mechanical energy obtained by the fluid per unit weight through the fluid machinery, which is the external energy. α is the kinetic energy correction coefficient, and in the formula, it has been taken as 1.

[0041] Multiply both sides of the equation by ρg, and let ρgH m = H, where H is the input mechanical energy, called the external pressure. After rearrangement, we get:

[0042]

[0043] The parameters in equations (1) and (2) are the parameter values of the centroid. For a circular pipe, the center line of the pipe is the centroid. The above formulas are applicable to incompressible air fluids with a velocity lower than 0.2 times the speed of sound, and can be used for the airfoils of fixed-wing aircraft and the compressor blades of gas turbines. When applying, the air parameters should be substituted.

[0044] Assume a straight pipe with a finite length, circular shape, smooth surface, and a relatively small radius R, horizontally placed in the air flow field. The pipe is stationary, and the coordinate axes are fixed on it, with the horizontal axis coinciding with the center line of the pipe. Among them, B is the inlet end, C is the external machinery, A is the outlet end, and the air flow in the pipe flows from right to left. The schematic diagram of the horizontal thin pipe is as Figure 1 shown.

[0045] When the air flow velocity in the flow field is zero and the stable flow velocity in the pipe is u, select any point in the inlet pipe BC as section 1 and any point in the outlet pipe AC as section 2.

[0046] At section 1, there is a velocity u, flowing from right to left. Its kinetic energy is transformed from the static energy of the flow field (also known as the background static energy). Therefore, the static energy of pipe BC drops by twice the kinetic energy compared to the background static energy; at section 2, there is a velocity u, but air needs to be discharged to the outlet. Therefore, the static energy must be equal to or greater than the background static energy to smoothly discharge the gas. Here, it is a constant velocity, and we take it as equal. Substituting into equation (1), we get the work done:

[0047]

[0048]

[0049]

[0050]

[0051] It shows that as long as the air velocity in the pipe is greater than zero, the external machinery will always do work. In the formula, p 背 is the static pressure of the background flow field.

[0052] If we imagine the flow field as a pipe with an infinite diameter, with a horizontal straight pipe of finite length and small diameter inside, then, Figure 1 the small pipe and the large pipe form a fluid circuit. As long as the air flow passes through point C of the external machinery, its energy will be increased.

[0053] In this way, the entire flow field consists of three parts: the outlet pipe, the large pipe (background flow field), and the inlet pipe. The flow direction is: the inlet pipe section sucks the fluid of the flow field from the inlet end. After the energy is increased by the external power, it is discharged into the outlet pipe section, flows into the background flow field through the outlet end, and then flows to the inlet end through the background flow field.

[0054] The energy of the entire flow field is divided into two regions. Among them, the energy in the outlet pipe AC is one region, and its kinetic energy is twice that of other regions.

[0055] If formula (2) is used and the effect of velocity is considered, the static pressure of the entire flow field can be divided into three regions. The static pressure of the outlet pipe section is the highest (when the fluid is set as the stationary coordinate, the velocity is the dynamic pressure, which can be converted into static pressure), the background flow field is the second, and the inlet pipe section is the lowest. The differences are one-fold dynamic pressure respectively.

[0056] Use a rigid flat thin sheet with no thickness, with a radius equal to the inner radius R of the pipe, perpendicular to the paper surface, and drive the air flow to move leftward along the pipe center line in the pipe at a constant speed u to replace the work done by the external machinery. The work done can be calculated using the pipe formula (1). At this time, for the air flow on the right surface of the sheet, due to the leftward movement of the sheet, the length of the inlet pipe section becomes longer. The incompressible air flow has not flowed in from the inlet end yet, and the air flow separates from the sheet. The separation velocity is u, and this space is a vacuum. At the inlet end, the air flow is the static pressure of the flow field. On the right side of the sheet is a vacuum. Under the action of the static pressure difference between the two, the air flow of the flow field flows into the inlet pipe section to fill the vacuum area. When the flow velocity reaches u, which is equal to the separation velocity of the sheet, the flow reaches equilibrium and the vacuum disappears. According to formula (2), the static pressure of the inlet pipe section is reduced by one-fold dynamic pressure compared to the static pressure of the flow field.

[0057] When the pipe is extremely short, formula (1) still applies; when it tends to be infinitely short, it still applies. When it is equal to zero, it is equivalent to the independent movement of the slice without a pipe segment. For the slice with a pipe and the slice without a pipe, assuming that both are moving horizontally from right to left with the same speed u, the static pressure of the fluid field on the left side of the two slices is the same and equal to the static pressure of the background fluid field. On the right side and the back of the two slices, due to the movement of the slices, the air separates from the slices, creating a vacuum on the back of the slices. The gas on the right side of the slices is the static pressure of the background fluid field. The air between the static pressure of the fluid field and the vacuum moves towards the vacuum area under the action of the static pressure of the fluid field. When the flow reaches equilibrium, the vacuum area disappears, the flow velocity is u, and the static pressure on the right side of the slice is the static pressure of the fluid field minus half of the dynamic pressure. Therefore, the forces on the two slices are the same, and the work done by both is equal. The amount of work done can be calculated using formula (1) for the case with a pipe. The movement of the slice without a pipe can be regarded as the movement of the slice with a pipe. Both are making the same horizontal movement in the fluid field, and the amount of work done is equal.

[0058] When a rigid flat thin slice without thickness moves alone in a stationary fluid field, the energy of the gas in front of it is relatively high, and the static pressure on the back is the lowest.

[0059] When the compressor blade of a gas turbine moves in an air fluid, it can be considered a rigid body. The surface of the blade is divided into countless small slices. Formula (1) can be applied to each small slice to obtain the micro work amount. Finally, by summing them up, the work done by the blade can be obtained.

[0060] In the calculation of the flow rate of a pipe fluid, when the air flow velocity forms an angle with the center line of the pipe, the flow velocity needs to be decomposed into two components parallel and perpendicular to the center line. Only the parallel component forms the pipe flow rate and constitutes the work - done flow rate, while the perpendicular flow rate does not do work. Therefore, we only need to find the component velocity parallel to the center line and substitute it into the formula for calculation.

[0061] After changing the conditions of formula (1) to a compressible gas and adding perfect gas, isentropic, one - dimensional high - speed steady flow, the pipe work formula is:

[0062]

[0063] In the formula, γ is the specific heat ratio (the specific heat ratio of air is equal to 1.4), and the subscripts 1 and 2 are the corresponding parameters of pipe cross - sections 1 and 2 respectively, and the usage is the same as that of formula (1).

[0064] Similarly, when the compressor blade of a gas turbine moves in a compressible air fluid field, the surface of the blade can be divided into countless small slices. Formula (3) can be applied to each small slice to obtain the micro work (amount). Finally, by summing them up, the work done by the blade can be obtained.

[0065] 3. Static pressure of the airfoil section of the compressor blade of a gas turbine

[0066] It is divided into the static pressure of the incompressible airfoil section and the static pressure of the compressible airfoil section.

[0067] The airfoil section of incompressibility is the airfoil section of the fixed-wing aircraft. Under incompressibility, they are the same as the airfoil section of the underwater wing. Therefore, the fluid formulas and results of the underwater wing can be directly applied, but the parameters are air parameters. For compressibility, the theory of incompressibility needs to be extended. At low speeds, the post-corner angle and the post-shift distance do not need to be calculated; but at near-sonic and supersonic speeds, the post-corner angle of the lower wing section, the included angle of the spanwise line relative velocity, and the post-shift distance need to be calculated to determine the position of the leading edge point of the wing, so as to weaken the shock wave.

[0068] Static pressure of the incompressible airfoil section.

[0069] Static pressure difference between the left and right (front and back) of a rigid flat thin plate without thickness.

[0070] Fix the origin of coordinates to the center point of the plate, place the X-axis horizontally to the right, and the Y-axis upward. The air flow in the flow field becomes a horizontal straight and uniform flow, moving from left to right. Schematic diagram of the horizontal straight and uniform flow, as Figure 2 shown.

[0071] According to the incompressible Bernoulli horizontal formula in the flow field:

[0072]

[0073] Find the static pressure p on the left side of the plate 片左 . Select the far left end of the flow field as section 1 and the left side of the plate as section 2, and substitute into formula (4) to get:

[0074]

[0075]

[0076] That is, the left side of the plate converts the dynamic pressure of the far-end gas into static pressure. The static pressure p on the left side of the plate 片左 is increased by twice the dynamic pressure compared to the static pressure of the flow field.

[0077] Find the static pressure p on the right side of the plate 片右 . The air in the flow field moves to the right at a speed of u. The gas separates from the right side of the plate at a speed of u, and the separation zone is a vacuum. Under the action of this static pressure difference between the static pressure of the flow field and the vacuum, the air quickly moves along the horizontal line to the right back surface of the plate. When dynamic equilibrium is reached, the speed of the air moving to the back surface is u, and the static pressure on the back surface of the plate is the static pressure of the flow field decreased by twice the dynamic pressure, that is:

[0078]

[0079]

[0080] The static pressure difference between the left and right sides of the plate is twice the dynamic pressure:

[0081] p 片左右差 = p 片左 - p 片右 = ρu 2 。

[0082] The airfoil section of a gas turbine compressor blade is a closed smooth transition curve, which we represent with a circular curve. By finding the static pressure at any point on the curve, we can apply it to obtain the static pressure at any point on the airfoil section. When calculating the area, we take 1 unit length in the spanwise direction.

[0083] Let the section curve of the airfoil be a circular curve with a radius of R. The circular curve is parallel to the plane of the paper, and an infinitely distant horizontal parallel air flow with a velocity u flows from left to right. The origin of the coordinate axis is fixedly connected to the center of the circle. The horizontal axis X is placed horizontally and points to the right, and the vertical axis Y is vertically upward. The curve is subjected to the static pressure of the air with the inner normal direction as positive. Point A is the intersection of the straight uniform air flow and the circumference. Line B is the normal line at point A, line C is the tangent line at point A, and α is the angle between the straight uniform air flow and the normal line. According to the meanings of formulas (2) and (4), on the upstream surface, in the second and third quadrants, the fluid impacts directly, which is the normal component of the force. The normal velocity becomes zero, and the dynamic pressure will be completely converted into static pressure, and the increased static pressure direction points to the center of the circle; the tangential component of the force has no impact and remains unchanged. The schematic diagram of the static pressure calculation on the upstream surface is as Figure 3 shown.

[0084] According to the meanings of formulas (2) and (4), the normal velocity = u × cosα, and the sum of the background static pressure and the newly added static pressure is the static pressure value at point A on the upstream surface:

[0085]

[0086] The second and third quadrants are symmetric figures with the horizontal axis as the axis of symmetry. The static pressure values of their symmetric points are equal in magnitude and the direction is along the inner normal direction. The newly formed static pressure is the thrust, which is equal in magnitude and the direction is also along the inner normal direction; the newly formed static pressure component force in the horizontal direction is the resistance, which is equal in magnitude and the direction is the same. The newly formed static pressure component force in the vertical direction is the lift, which is equal in magnitude. The component force in the second quadrant is the negative lift, and in the third quadrant, it is the positive lift.

[0087] The downstream surface consists of the curves in the first and fourth quadrants of the circular curve. Point A is the intersection of the reverse straight uniform air flow (with the same magnitude as the straight uniform flow velocity but in the opposite direction) and the circumference. Line B is the normal line at point A, line C is the tangent line at point A, and α is the angle between the reverse straight uniform air flow line and the normal line. The schematic diagram of the static pressure calculation on the downstream surface is as Figure 4 shown.

[0088] According to the meanings of formulas (2) and (4), the normal velocity = u×cosα. The normal velocity on the backflow surface is transformed from the static pressure of the background flow field, causing the static pressure to decrease. The difference between the background static pressure and the newly added static pressure is the static pressure value at point A on the backflow surface:

[0089]

[0090] The direction of the static pressure is the same as the inner normal direction, pointing to the center of the circle, but it is decreasing.

[0091] The first and fourth quadrants are symmetric figures with respect to the horizontal axis. The static pressure values of their symmetric points are equal in magnitude, and the direction is along the inner normal direction. The newly formed static pressure is a tensile force (suction force), equal in magnitude and along the outer normal direction; the newly formed component of the static pressure in the horizontal direction is a resistance force, equal in magnitude and the same in direction. The newly formed component of the static pressure in the vertical direction is a lift force, equal in magnitude. The component force in the first quadrant is a positive lift force, and in the fourth quadrant is a negative lift force.

[0092] In summary, the static pressure value at a certain point A on the wing section is:

[0093]

[0094] In the formula, the plus sign represents the oncoming flow surface, and the minus sign represents the backflow surface. The formula shows that for the upper wing surface, the intercept line in the first quadrant should be used. The intercept line is a monotonically decreasing function curve without inflection points; for the lower wing surface, the intercept line in the third quadrant should be used. The intercept line is a monotonically decreasing function curve without inflection points; from the perspective of height, their height coordinates decrease in sequence to form a consistent double-sided positive lift airfoil, that is, a dynamic pressure difference airfoil.

[0095] For the convenience of discussion, we limit it to the first and third quadrants. The lift-drag ratio of a certain point on the wing section is only considered to be the ratio of the absolute value of the lift increased by the dynamic pressure at this point to the absolute value of the resistance increased by the dynamic pressure. Then, the static pressure increment in formula (5) is multiplied by the static pressure of this extremely small 1 unit area, and then projected onto the vertical and horizontal directions respectively to find the ratio, which is the lift-drag ratio of this point:

[0096]

[0097] It means that the larger the included angle, the higher the lift-drag ratio. Under the condition of meeting the lift requirement, the included angle should be increased as much as possible to improve the efficiency of the wing; the larger the included angle, the larger the rear turning angle, and the lower the static pressure on the object surface, which can better avoid the influence of shock waves. The infinitesimal line segment of the intercept line of the point can be considered as a straight line. If the straight line is only extended in the horizontal direction, the normal included angle becomes larger, the tangent value increases, and the lift-drag ratio increases. Using the horizontal extension of the intercept line to increase the lift-drag ratio is an effective method. When using a horizontal uniform flow design, the chord of the wing is horizontal. For those that are relatively not horizontal, the normal included angle of their intercept line is large, the tangent value is large, and the lift-drag ratio is relatively high.

[0098] The above is the reasoning and conclusion of the dynamic pressure difference lift principle in an incompressible fluid when the air flow is below 0.2 times the speed of sound.

[0099] Static pressure of a compressible airfoil section.

[0100] In the flow field, the parameter relationship between section 1 and section 2 of a compressible, perfect gas, isentropic, one-dimensional high-speed steady flow is:

[0101]

[0102]

[0103] The static pressure value p on the left side of the slice 片左 . When the air flow moves from the left far end to the left side of the slice, it is an isentropic compression process. Select the left far end of the flow field as section 1, and the parameters are the flow field parameters; the left side of the slice is section 2, and the velocity parameter is zero. Substitute into formula (6) to get:

[0104]

[0105]

[0106] When the air flow at point A of the section line has a normal angle, the static pressure p α Substitute with the normal velocity u cosα to get:

[0107]

[0108] The static pressure value p on the right side of the slice 片右 . When the gas in contact with the right side of the slice moves to the right together with the flow field velocity, it separates from the slice, creating a vacuum space. The gas in the flow field static pressure and the vacuum space is under the action of these two forces, and the gas fills the vacuum space from right to left. Its flow is an isentropic expansion process. When it reaches equilibrium, from formula (6), we get:

[0109]

[0110] When the air flow at point A of the section line has a normal angle, the static pressure p α Substitute with the normal velocity u cosα to get:

[0111]

[0112] Formula (7) is used for the upstream surface, and formula (8) is used for the downstream surface.

[0113] The formula shows that the upper wing surface should use the section line in the first quadrant, and the height of the section line monotonically decreases from the leading edge to the trailing edge without inflection points; the lower wing surface should use the section line in the third quadrant, and the height of the section line monotonically decreases from the leading edge to the trailing edge without inflection points; only in this way can a consistent positive lift double-sided wing be formed simultaneously to constitute a dynamic pressure difference airfoil.

[0114] At low subsonic speeds, the above formula can complete the lift design and improvement of the cross-section line.

[0115] However, at near-sonic and supersonic speeds, the influence of shock waves also needs to be considered. It can weaken the strong static pressure on the object surface of the normal shock wave to the static pressure on the object surface of low subsonic speed by the angle between the normal of the oncoming flow surface.

[0116] When the normal angle rotates from 0 degree to the α angle, its tangent also rotates by the α angle correspondingly, which is called the post-rotation angle of the tangent line, abbreviated as the post-rotation angle. The word "post-" means rotation along the velocity direction; the post-rotation angle of the object surface is that when the normal angle rotates from 0 degree to the α angle, the object surface also rotates by the α angle correspondingly, which is called the post-rotation angle of the object surface here. The angle values of the post-rotation angle, the post-rotation angle of the object surface, and the normal angle are equal. At near-sonic and supersonic speeds, increasing the post-rotation angle, the normal velocity decreases, the work done decreases, the static pressure on the object surface also decreases accordingly, and the resistance also decreases; for the post-rotation angle of the cross-section line and the post-rotation angle in the spanwise direction, the points with static pressure exceeding the requirement are weakened to points with static pressure equal to or less than the static pressure requirement of low subsonic speed.

[0117] The straight and uniform air flow velocity u at point A is the relative velocity of the velocity triangle of the point during the rotation of the blade. In the range of 5 times supersonic speed, the calculation results of the above formula are generally considered to be within the acceptable error range.

[0118] At near-sonic and supersonic speeds, shock waves will be generated on the windward surface. It is necessary to find out the angle between the normal of the windward surface to reduce the oncoming flow velocity to the normal velocity and the required static pressure on the object surface as specified. There is no shock wave on the leeward surface, so there is no need to solve it. From formula (7), the angle between the normal of the windward surface can be obtained:

[0119]

[0120] We limit the value range of the α value to 0 to 90 degrees. The larger the value, the smaller the normal velocity and the lower the static pressure on the object surface. We often recommend the highest specified value of the low subsonic speed value according to practice (using the highest specified value of the low subsonic speed value, the highest specified value of the static pressure on the object surface of the point can be calculated). If it is equal to or less than this value, it is considered to reach the static pressure value of the low subsonic object surface, and the adverse influence of the shock wave is eliminated. Here we only give an example of the specified value lower than the speed range because when the value is low, by increasing the rotation speed of the rotating shaft, the force on the cross-section line increases, and the compression ratio can also be increased, and the projection angle in the axial direction is smaller, with higher efficiency.

[0121] At near-sonic and supersonic speeds, the relative velocity of the leading-edge point of the airfoil section increases with the increase of the blade radius, and the leading-edge point angle also increases. Substitute the highest low-speed velocity specified value and the cosine zero-degree value into formula (7) for the leading-edge point angle of the lower surface section line, calculate the highest low-speed specified static pressure value, and then substitute this static pressure value and the relative velocity value of this point into formula (9) to calculate α i value. When the post-deflection angle of each point behind the section line is equal to or greater than this angle, the normal velocity and static pressure value of each point are not higher than those of the leading-edge point; if the post-deflection angle of the latter point is equal to or greater than that of the former point, no secondary shock wave will be generated. In this way, the static pressure value of each point on the airfoil section does not exceed the highest specified value of the static pressure on the object surface at low speed.

[0122] The spanwise line is a line perpendicular to the relative velocity. At near-sonic and supersonic speeds, it will also encounter the strong static pressure generated by the normal shock wave same as the section line, and the static pressure value on the object surface needs to be reduced. Using the same above-mentioned derivation process of finding the angle of the airfoil section, the same formula (9) for finding the angle of the spanwise line can be deduced.

[0123] As the radius increases, the angle of the points on the spanwise line also increases, and the angle needs to be calculated for each point. Substitute the highest low-speed velocity specified value and the cosine zero-degree value into formula (7) for the spanwise line angle, calculate the highest low-speed specified static pressure value, and then substitute this static pressure value and the relative velocity value of this point into formula (9) to calculate α i value. When the post-deflection angle of the spanwise line is equal to or greater than this angle, the static pressure value on the object surface of this point does not exceed the highest specified value of the static pressure on the object surface at low speed. Here we only calculate the spanwise line where the leading-edge point is located, because other spanwise lines generally meet the requirements. All calculations are required only in special cases.

[0124] Different airfoil materials have different highest limiting values of the specified low-speed. Here we take a metal airfoil as an example. The static pressure value of the points on the airfoil section is specified not to be higher than the static pressure value p 0.95 at 0.95 (to 0.97, here we take 0.95) times the speed of sound u 0.95 . We can substitute u 0.95 and the cosine value equal to 1 into formula (7) to find the p 0.95 of this point; then substitute the true relative velocity u i of the leading-edge point at the radius r i and p 0.95 into formula (9) to find the normal angle α i . The static pressure value of this point at this angle is equal to p 0.95 , and when it is greater than this angle, the static pressure value is less than p 0.95 . If applied to all the airfoil sections and spanwise lines of the blade, then the blade surface will all meet the requirement of p 0.95 , achieving the purpose of weakening the shock wave to the static pressure on the object surface at low speed.

[0125] The blade shall use the maximum outer diameter R 2 Perform the relative velocity value calculation. For the calculated velocity, if it is equal to or less than u 0.95 , it is considered a low-speed sound impeller; if it is greater than u 0.95 , it is considered a near-sonic and supersonic impeller, and it is necessary to calculate the radius r 0.95 at u 0 , those less than the radius r 0 shall be treated as low-speed sound; those equal to the radius r 0 shall be marked as point A 0 , and it is necessary to calculate the post-turning angle of the cut-off line for near-sonic and supersonic impellers from point A 1 to point A n , as well as the calculation of the post-turning angle and the post-shift value of the spanwise line.

[0126] The post-shift distance value is the straight-line length between the position of the leading edge point of the airfoil and the position of the initial leading edge point. When the normal angle of a certain point on the spanwise line rotates from 0 degrees to the α angle, the intersection point with the relative velocity direction line also moves a certain straight-line distance accordingly. This straight-line moving distance (value) is called the post-shift distance (value), abbreviated as post-shift (value). Therefore, the initial intersection point with a post-turning angle of 0 degrees is the starting point of the post-shift distance, and the final intersection point is the ending point of the post-shift distance. The straight-line distance between the two points is the post-shift distance.

[0127] The impellers of a gas turbine compressor include a fan impeller and a compressor impeller, which can be regarded as impellers with different lengths of blades. They all push the air flow behind the wheel shaft, forming an axial acceleration flow and compressing the gas. It can be imagined that the impeller is placed in a horizontal position and rotates in place. The air flow flows in from the front of the shaft and flows behind the wheel blade after passing through the wheel blade for discussion. The coordinate axes coincide horizontally with the rotation axis, the coordinate origin is represented by O as the in-situ position of the impeller, the Z axis is to the left, the Y axis is vertically upward, and the X axis is horizontally perpendicular to the book facing us. r is an arbitrary radius, A represents the leading edge point of the wing. When the speed of all leading edge points is zero, they are all on the spanwise (radial) line of the origin. Among them, we assume that the leading edge point of the impeller blade being studied is exactly on the positive Y axis, and R 2 represents the maximum radius of the blade. At near-sonic and supersonic speeds, R 1 represents the starting radius r 0 from which the blade starts to require a post-turning angle, A 0 represents the leading edge point at the radius r 0 (which is also the highest limit value point of the speed specified by low-speed sound); it is assumed that there are n equally spaced points between R 1 and R 2 . Starting from the zero-mark point A 0 , the corresponding radius is r 0 (R 1 is equal to r 0 ); the nth punctuation mark ends, corresponding to the punctuation mark An , the radius r n (R 2 is equal to r n ); the i-th punctuation point in the middle is represented by point A i , corresponding to the radius r i ; the punctuation point before the i-th point is represented by A i-1 , corresponding to the radius r i-1 . The rotation direction of the impeller is set as rotating from the inside to the outside above the axis. The schematic diagram of the horizontal position of the impeller is as Figure 5 shown.

[0128] The velocity triangle of point A on the blade of the impeller consists of absolute velocity, circumferential velocity, and relative velocity to form a right triangle on the horizontal cutting plane at point A. Among them, the horizontal axial absolute air flow velocity u 气 of point A is one side of the right angle; the circumferential velocity u 牵 (u 牵 = 2πrn s , where n s is the number of revolutions per second) is the other side of the right angle; the relative velocity u 相 (equal to the square root of the sum of their squares) is the hypotenuse. At Figure 5 make a top view of point A. The included angle between the relative velocity and the circumferential velocity is β, called the relative velocity angle. The M direction is the projection direction perpendicular to the relative velocity. The top view of the velocity triangle of point A is as Figure 6 shown.

[0129] Make a horizontal rotating cylinder with a radius of r. On the 360-degree cylindrical surface, on the generatrix at any angle, the horizontal absolute air flow velocities u 气 , circumferential velocities u 牵 , and relative velocities u 相They are all of equal size. The three sides of the velocity triangle at each point are equal, and the figures are also exactly the same. The velocity triangles at each point are all on the tangent plane at that point. The generatrix at each point is a line on the tangent plane and also a line on the cylindrical surface. It is the common line of the tangent plane and the cylinder. The velocity triangles overlap and share. The generatrices on the cylindrical surface are all parallel to the axis. The angles between the relative velocity direction lines at each point and the generatrix and the axis are equal. Starting from point A (assumed to be zero degree, but in actual applications, the midpoint is often used), the cylindrical surface rolls horizontally by 360 degrees on the horizontal tangent plane at point A, forming a developed view on the horizontal tangent plane at point A with a radius of r. The velocity triangles at each point on the developed plane are all imprinted from the generatrix on the cylindrical surface, which are congruent triangles. Among them, the relative velocity magnitudes are equal, and the relative velocity directions are the same, constituting a plane relative velocity uniform flow. If the spanwise thickness is 1 unit (i.e., the radius increment is 1), a three-dimensional relative velocity uniform flow is formed. If the radius increment approaches zero, the thickness of the developed view approaches zero, and the gas parameters of the developed view will be very consistent and accurate. If the designed airfoil section line of the plane developed view is on the developed plane, and the cylinder rolls back from the end point of rolling (360 degrees) to the starting point A (0 degree), then the plane section line is imprinted on the cylindrical surface, becoming a cylindrical surface section line. Adding a small spanwise thickness is a small section of the blade. Here, the sectional view of the blade we consider is a cylindrical surface formed by a generatrix parallel to the impeller axis with a radius of r rotating around the axis for one week and developed into a plane to form a blade sectional view. We only take one blade for discussion. The sectional view intercept line of the blade is designed under a horizontal relative velocity uniform flow. The chord line is consistent with the airflow direction, and the efficiency is relatively high (when the airflow direction is parallel to the chord line, for the same lift, the normal angle is larger than that when they are not parallel, the tangent value is higher, and the efficiency is relatively high); if the angle between the relative velocity direction line and the generatrix is marked on the horizontal line and a line is drawn, this line is the axial line, then the projection of the force generated by the uniform flow on the airfoil section line on the axial line is the axial thrust; reflected on the cylindrical surface, the force of the airflow with the relative velocity on this section line in the axial projection is the axial thrust, and the magnitudes of the two axial thrusts are equal. The airfoil of the dynamic pressure difference wing of a fixed-wing aircraft is designed in a horizontal uniform flow. If its horizontal uniform flow velocity is equal to the relative velocity, and then the chord line is designed to be at a zero angle (the chord length of the dynamic pressure difference airfoil in this article refers to the distance between the projections of the leading and trailing edges on the chord line, and the chord is the horizontal line), the designed airfoil is the airfoil of the blade sectional view and can be directly used as the airfoil of the blade sectional view. If the axial thrust needs to be calculated, the axial line can be marked, and after finding the sum of the forces at each point, its projection is the axial thrust. Placing the designed airfoil of the fixed-wing dynamic pressure difference wing on the developed view of the horizontal tangent plane at point A, with the leading edge point of the airfoil section placed at point A on the spanwise line of the blade and the chord aligned with the relative velocity direction (collinear), it becomes the airfoil of the dynamic pressure difference blade. Therefore, directly using a plane relative velocity uniform flow and a zero-degree chord to design the sectional view of the dynamic pressure difference fixed-wing wing, a well-matched sectional view of the dynamic pressure difference blade airfoil can be designed; and for those not equal to the relative velocity, it is also a dynamic pressure difference airfoil, but the degree of matching is not that high.In this way, to design the dynamic pressure difference blade airfoil, we first design the dynamic pressure difference wing airfoil of the fixed-wing aircraft, make the airfoil and the blade cross-section diagram in the same plane, and then place the zero-degree chord in the same direction as the relative velocity and in the correct position with the leading edge point, and it becomes the dynamic pressure difference blade airfoil.

[0130] The blade span line, at the specified static pressure intensity p of 0.95 times the supersonic speed 0.95 When 0 (The corresponding radius is R 1 ). The radius is increased, the leading edge point A i The relative speed direction line and the backward distance value must be placed appropriately to ensure that the static pressure strength p is not exceeded. 0.95 .

[0131] In A 0 To A n The relative velocity angle β between the points is β 0 To β n ,u 牵 As the radius increases, the angle β gradually decreases. The angle β value is between 0 and 90 degrees. i Through the velocity triangle we can obtain:

[0132] tanβ=u 气 ÷u 牵

[0133] β=arctan(u 气 ÷2πr n s )………………(10)

[0134] The above formula determines the direction of the relative velocity at point A on the horizontal tangent plane. 1 Start, β = β i =arctan(u 气 ÷2πr i n s ), r=r i , A uses A i Mark.

[0135] After determining the relative velocity direction, determine the i The distance that a point moves in a straight line in the direction of relative velocity. i Indicates spanwise line A i The back angle value of the point; the tangent parallel line is A i The parallel line of the tangent line of point A i-1 The point intersects and intersects with A i The horizontal line intersects at point K; l i kA is the straight line segment i Distance, A i The backward distance of the point, M is the projection diagram of the backward distance, such asFigure 7 as shown

[0136] The difference in radius length Δr between adjacent leading edge points (this point and the previous point) is: Δr = (R 2 - R 1 ) / n

[0137] A 1 The backward movement value of this point of point A is l 1 = Δr × tanα 1 ; accordingly, the backward movement value of this point of point A is l i = Δr × tanα i = Δr × tanα i .

[0138] Let S i represent the cumulative backward movement value of point A. It is the sum of the projections of l i to l 1 along the β i direction respectively. Make i the top view, establish a new coordinate system. The origin still uses the old coordinate origin, which is represented by O. The Y-axis is upward and the X-axis is to the right. The leading edge A point coincides with the origin, and point A is above. The schematic diagram of the backward movement distance in the top view is as Figure 5 shown Figure 8 as shown

[0139] The cumulative backward movement value S i of the i-th point is

[0140] S i = l 1 cos(β 1 - β i ) + l 2 cos(β 2 - β i ) + … + l i-1 cos(β i-1 - β i ) + l i cos(β i - β i ) …… (11)

[0141] The above is the principle of the dynamic pressure difference blade lift of a gas turbine compressor, and it is also the principle of the dynamic pressure difference lift of a gas turbine compressor blade. Now apply it

[0142] 4. The airfoil of the dynamic pressure difference wing of a fixed-wing aircraft

[0143] In the design of dynamic pressure difference airfoils, the relative velocity is regarded as a horizontal uniform flow, and the chord elevation angle is zero degrees. According to formulas (7) and (8), for the upper airfoil surface, the intercept line in the first quadrant is selected, and the intercept line is a monotonically decreasing function curve; for the lower airfoil surface, the intercept line in the third quadrant is selected, and the intercept line is a monotonically decreasing function curve. When the speed is near the speed of sound or supersonic, the radius r needs to be calculated at the time of u 0.95 at that time 0 , those less than the radius r 0 are processed according to the low speed of sound; those equal to the radius r 0 are marked as point A 0 , and for the leading edge point of the lower airfoil from point A 1 to point A n , using p 0.95 and the relative velocity at this point, substitute into formula (9) to calculate the normal angle between points A 1 to A n . When the post-rotation angle at the leading edge point of the lower airfoil is equal to or greater than this angle, the lower airfoil and the upper airfoil are directly connected. The post-rotation angle of the intercept line from the leading edge point to the trailing edge point should satisfy that the post-rotation angle of the latter point is equal to or greater than that of the former point to avoid the influence of shock waves. In this way, positive lift will be generated on both the upper and lower airfoil surfaces at the same time. Then, according to the point lift-drag ratio judgment, select the intercept line with higher efficiency to further improve the lift-drag ratio.

[0144] Modify the airfoil with the current lift principle of velocity difference to an airfoil with dynamic pressure difference

[0145] Suppose it is a fixed-wing airfoil in a relative velocity uniform flow field. The upper airfoil surface is an upward arc, point A is the leading edge, point B is the trailing edge, the lower airfoil surface is horizontal, AB is the chord length, C is the highest point, D is the foot of the perpendicular of C, and the distance from point D to point A is 25% of the AB length. If the uniform air flow flows horizontally from left to right, the curve AC is the oncoming surface, and the curve CB is the backflow surface. Therefore, the curve ACB on the upper airfoil surface should remove the oncoming surface and leave the backflow surface. A schematic diagram of the cross-section of a fixed-wing aircraft airfoil is as Figure 9 shown

[0146] Take the left part ACD and make a symmetric image ADE part with the horizontal line as the axis of symmetry. The lift magnitudes of the two curves of the symmetric image are equal, but the directions are opposite; the horizontal resistances are equal. The horizontal symmetric image of the left part is as Figure 10 shown

[0147] Make a drawing of the ADE part. The length of AD is 25% of AB. Horizontally extend AD by two times to obtain the horizontal line AF. The length of AF is 75% of AB. Pass through point F and draw a perpendicular line. Take point G such that GF is equal to the length of DE. Make the AG curve (shown as a dotted line) of the AFG part, and its curve type is the same as that of the AE curve. Due to the horizontal extension, the normal angle of the AG curve is larger, the value of tanα is higher, and the lift-drag ratio is higher. The transition drawing after the horizontal length is extended by two times is as Figure 11 shown.

[0148] Take Figure 9 the CB curve as the upper wing surface and connect it with Figure 11 the AG curve as the lower wing surface. That is, point C is connected to point A, and point B is connected to point G. Their horizontal lengths are equal, and their vertical heights are equal, forming a new dynamic pressure difference airfoil. Among them, the solid line represents the upper wing surface, and the dotted line represents the lower wing surface. The schematic diagram of the cross-section of the new airfoil with a length of 75% is as Figure 12 shown.

[0149] For the new dynamic pressure difference airfoil, positive lift is generated on both the upper and lower wing surfaces, forming a double-sided lift airfoil. Since there is one more lift surface than before, the lift-drag ratio is significantly improved.

[0150] Increasing the normal angle can further improve the lift-drag ratio of the airfoil.

[0151] According to this characteristic, we horizontally extend the C(A) leading edge point in Figure 12 to the left by 25% of the length of AB. For the upper wing section line CB (solid line) and the lower wing section line AG (dotted line), and then extend the solid line and the dotted line respectively according to their same type of curve. The CB (solid line) and AG (dotted line) after the extension result form a smoother and more efficient new airfoil, further improving the lift-drag ratio of the airfoil. This airfoil is the dynamic pressure difference airfoil of the fixed-wing aircraft after optimization and improvement. The schematic diagram after optimization and improvement is as Figure 13 shown.

[0152] 5. Dynamic pressure difference blades for gas turbine compressors.

[0153] The dynamic pressure difference airfoil of the fixed-wing aircraft obtained above can be directly used as the airfoil of the dynamic pressure difference blade for the gas turbine compressor. Make the plane expansion diagram of the airfoil of the aircraft wing and the leading edge point A of the blade in the same plane, with the chord line and the relative velocity collinear, and then overlap the leading edge point with point A, and it becomes the airfoil of the dynamic pressure difference blade for the gas turbine compressor.

[0154] In Figure 6 , select a certain radius r from the root to the tip of the impeller blade, the position of the leading edge point A, and make the horizontal plane expansion diagram of point A. Figure 13The leading edge point of the airfoil coincides with point A on the plane; calculate the β angle according to formula (10) to determine the relative velocity angle β, so that the zero chord line of the fixed airfoil is collinear with the relative velocity motion direction line, and the airfoil also overlaps with the plane (in the same plane); this airfoil section is the required airfoil section of the gas turbine compressor differential pressure blade. For low subsonic speeds, a flat airfoil development drawing of the blade has been completed. However, at near-sonic and supersonic speeds, it is necessary to calculate the radius r 0.95 at u 0 , those less than the radius r 0 are processed according to low subsonic speed; those equal to the radius r 0 are marked as point A 0 , and it is necessary to use p 1 for the lower leading edge point from point A n to point A 0.95 , substitute the relative velocity at this point into formula (9) to calculate the normal angle from point A 1 to point A n . When the trailing angle at the lower leading edge point is equal to or greater than this angle, the trailing angle of the section line from the leading edge point to the trailing edge point should satisfy that the trailing angle of the latter point is equal to or greater than that of the former point; the spanwise line then needs to calculate the relative velocity angle β i according to formula (10) to determine the relative velocity direction passing through point A i ; calculate the cumulative backward movement value S i according to formula (11), and then Figure 13 start from point A i for the leading edge point of the airfoil, move backward by S i along the relative velocity direction line to locate the appropriate position of the leading edge point. Only then is a flat airfoil development drawing of the gas turbine compressor differential pressure blade completed. Take different r values between the blade root and the blade tip, repeat the above process, and different flat development drawings of different radii can be obtained until the production of 1 blade is completed. If several blades are required for the entire circumference of the impeller, they are evenly distributed on the same circumference, repeat the above process several times until the production of several blades is completed.

[0155] The above-completed impeller can be arranged as coaxial same-direction impellers and coaxial opposite-direction impellers.

[0156] For coaxial same-direction impellers, the production process of the impeller can be repeated once using the above method to complete two impellers, which are arranged in series, and are formed with static guide impellers usually used for spacing in between.

[0157] For coaxial opposite-direction impellers, after completing the above single wheel, the entrainment velocity of the second impeller is in the opposite direction, and then use the above method to repeat the production of an impeller, and arrange the two completed single wheels in series to form, and there can be no static guide impeller spacing in between.

Claims

1. A gas turbine compressor dynamic pressure difference blade, the cross-sectional development diagram of which includes an upper wing surface and a lower wing surface, characterized in that: The upper wing surface and the lower wing surface section lines are both curved surfaces consisting of monotonically descending curves from the leading edge to the trailing edge.

2. A gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: The main wing surface of the upper wing is composed of a monotonic descending function surface from the leading edge to the trailing edge, and has no inflection point.

3. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: The main wing surface of the lower wing is composed of a monotonic descending function surface from the leading edge to the trailing edge, and has no inflection point.

4. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: At near-sonic and supersonic speeds, the leading edge of the upper wing and the leading edge of the lower wing directly intersect.

5. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: At near-sonic and supersonic speeds, the rearward turning angle of the leading edge point of the lower wing section line is equal to or greater than the required normal angle of the leading edge point.

6. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: At near-sonic and supersonic speeds, the section line of the lower wing surface is the rearward turning angle from the leading edge to the trailing edge, and the rearward turning angle of the latter point is equal to or greater than the rearward turning angle of the former point.

7. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: At near-sonic and supersonic speeds, the rearward turning angle of the leading edge point of the spanwise line is equal to or greater than the required normal angle of the point.

8. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: At near-sonic and supersonic speeds, the position of the leading edge point of the wing section line is determined by the relative velocity direction line and the backward distance value.

9. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: Increasing the normal angle of the wing section can improve the lift-to-drag ratio.

10. The gas turbine compressor dynamic pressure difference blade according to claim 1, characterized in that: The zero-degree chord of the dynamic pressure difference wing airfoil of a fixed-wing aircraft is collinear with the relative velocity direction line.

Images