WO2024242589A2 - Fan blade configuration - Google Patents

Abstract

The invention relates to gas-turbine engine fans. The technical result is achieved by the use of a one-piece blade structure with a variable angle of inclination of the chord of the airfoil in relation to the direction of the airflow from the edge to the root section of the blade, in a limited range of ratios of the outside diameter of the impeller to the inside diameter thereof, D2 to D1; the design of the airfoil of the blade such as to be described by arcs of circles, the geometric cross-sectional centres of which lie in a line along the length of the blade, provides for ease of manufacture, while the angles of inclination of the chord of the airfoil to the airflow, together with the chord length, the radii of the concave part of the section, the radii of the root part and the width of the blade in the root section and the tip section, provide for high blade efficiency at speeds greater than the speed of sound, while also providing sufficient strength across a wide range of rotational speeds of the fan.

Description

Fan blade shape

The invention relates to fans of gas turbine engines, in particular to fans of turbojet fan engines. The most effective application is in low-power turbojet fan engines with a thrust of 1000 to 5000 N.

A gas turbine engine usually contains a fan that forces air flow into the engine compressor; in a dual-circuit engine, the fan also serves as a source of additional thrust by forcing air flow into the outer contour of the engine. The fan wheel consists of a base on which a plurality of blades are placed, forming a non-separable structure with the base, or blades that are fixed to the base with the possibility of removal without destroying the structure and even turning around its axis to change the engine thrust. Further development of the technical characteristics of turbofan engines leads to the need to use fans with blades or vanes operating at supersonic speeds, while having high efficiency and acceptable strength.

The blade design is known (invention certificate RU No. 2255248 C2). The blade includes a chord convexity on the leading edge between the root and the tip and has a reverse aerodynamic sweep at the tip.

The disadvantages of this design are: the blade geometry is difficult to manufacture, the split design of the fan makes it expensive and significantly difficult to use in low-power engines, and installing the blade in the confuser of the engine housing reduces the efficiency of the fan as a whole.

A design of a fan blade with a coordinated platform is known (Invention Certificate RU No. 2272180 C2). The fan blade contains an aerodynamic surface having a root and made in one piece with the platform. The platform includes a front and rear edge and arc-shaped first and second sides between them. The root of the aerodynamic surface is shifted in the transverse direction closer to the first side of the platform than to the second side.

The disadvantages of this design are: the blade geometry is difficult to manufacture, the split design of the fan makes it expensive and significantly difficult to use in low-power engines, installing the blade in the confuser relative to the outer casing of the engine reduces the efficiency of the fan as a whole, The rounded leading edge limits the maximum blade rotation speeds, creating high flow resistance at Mach speeds above 1.

A design of a turbojet engine fan blade is known (Invention Certificate RU No. 2486346C2). A turbojet engine fan blade with a double flow, comprising an internal aerodynamic part and an external aerodynamic part, combined in the radial direction and separated by a platform, wherein the internal aerodynamic part comprises one aerodynamic profile, and the external aerodynamic part comprises at least two aerodynamic profiles, wherein the attack ribs of said aerodynamic profiles of said aerodynamic part of the blade are axially aligned in a row. A reduction in the number of fan blades is achieved while maintaining satisfactory quality due to maintaining an increased relative pitch of the internal aerodynamic part of the fan blade.

The disadvantages of this design are: the complex design of the blade, which requires complex manufacturing technology and the practical difficulty of using it in low-thrust turbojet engines.

A known axial fan blade (Invention Certificate RU No. 2516739 C1) is accepted as the closest analogue based on a number of features. The blade comprises a convex approach edge, a concave trailing edge, a root portion and an upper edge. The radius of curvature of the concave trailing edge is made equal to the radius of curvature of the blade centerline, and the radius of curvature of the convex approach edge is selected based on ensuring the required blade thickness in the average cross-section along its length. The angle of entry of the air mass into the interblade channel and the angle of exit from it are related by the ratio: y = tan-l (< p / cosa), where y is the angle of exit of the air mass from the interblade channel; ( is a coefficient taking into account the loss of air mass velocity in the interblade channel, ranging from 1 to 0.9; a is the angle of entry of the air mass into the interblade channel. The blade parameters are made in accordance with the following expressions: R = a / (cosa - cosy), where R is the radius of curvature of the blade centerline; a is the blade width in the axial direction, l = (27iR (ya)) / 360, where 1 is the total length of the blade, b = R (siny - sin), where b is the width of the blade along the front. The performance, pressure and efficiency of the fan are increased while maintaining the required strength and rigidity of the blade during operation. The disadvantages of this solution are: operation at low, subsonic peripheral speeds of the blade, the geometry of the blade is designed for relatively low axial air flow speeds, the relationship between the outer diameter of the fan wheel and the inner diameter to which this solution is applicable is not taken into account.

The purpose of this invention is to develop a simple design of a fan blade for a gas turbine engine, preferably a turbofan type, operating in a small thrust range from 1000 to 5000 N with the blade edge moving at a speed, including above the speed of sound, Mach number>1, with a relatively simple manufacturing technology and a relatively low cost.

The set objective is achieved by: using simple geometric shapes when constructing the blade profile, optimizing the blade inclination angles to the air flow and optimizing the number of blades on the fan impeller for maximum performance and maximum efficiency of the fan impeller, finding a dependence for the blade thickness to ensure its sufficient strength when operating at peripheral blade edge speeds exceeding the speed of sound. The geometric dimensions of the blade are related as follows: the fan blade contains a convex profile edge determined by geometric construction described by the radius of the circle Ria for the root part of the profile 3 of the blade and R16 for the end part of the blade 4; the leading edge of the blade 1 and the trailing edge of the blade 2 are made straight; The chord inclination angles to the air flow are taken for the root section A = 53...55°, for the tip section of the blade profile B = 77...79°, where the ratio of the outer diameter of the impeller to the inner D2/D1 lies in the range of 2.4...2.6; the chord length of the root section profile Ll = kl n Dl/N; where kl is the coefficient taking into account the properties of the blade material from 2 to 3.1; л is a mathematical constant -3.1415; N is the number of impeller blades, an even number in the range from 14 to 24; for the tip section of the blade, the chord length L2 = k2 - 7i - D2/N; where k2 is the coefficient of the material properties lies in the range from 1.3 to 1.5; the radii of curvature of the concave part of the profile are determined by the ratio R2a = pi LI; pi is the coefficient of flow properties from 7.0 to 7.3 for the root section, R26= i L1; for the end section, where the greatest thickness of the blade - I in the root section is determined from the expression: tl=Ll/y; where y is the coefficient taking into account the blade material from 15.5 to 15.7; accordingly, for the end section t2=L2/y. The essence of the invention is explained by the drawings: Fig. 1 is an isometric view of a fan blade wheel; Fig. 2 is a main view of a blade wheel with the main dimensions of the blade end profile indicated; Fig. 3 is a view of a blade wheel from the right with the diameter designations and the location of the root section indicated; Fig. 4 is a root section of a blade with the main dimensions indicated.

Thus, the proposed technical solution for the profile of the fan blade of a turbojet engine allows the blade wheel to be made in one piece using, for example, a CNC milling machine with four degrees of freedom or more, which is currently a widespread technological process, reducing the cost of producing a fan wheel to a minimum, while ensuring high rigidity of the fan wheel and, as a result, ensuring the possibility of its operation at high, even supersonic at the edge, peripheral speeds. The rigidity of the design also ensures the possibility of operating the fan at significant angular accelerations of the rotor during a rapid change in the operating mode of the turbojet engine. The choice of blade geometry within the boundaries of the given parameters allows obtaining the highest possible efficiency of the fan wheel.

Claims

5 Invention formula A turbofan engine fan blade comprising a convex approach edge described by a radius of a circle Ria for the root section of the profile and R16 for the tip section of the blade, a concave trailing edge described by a radius of a circle R2a for the root section of the profile and R26 for the tip section of the blade, a leading edge made along a straight line and a trailing edge made along a straight line, the angle of inclination of the profile chord at the root section relative to the direction of the air flow - A lies in the range of 53...55°, the angle of inclination of the profile chord at the tip section relative to the direction of the air flow - B lies in the range of 77...79°, wherein the ratio of the outer diameter of the impeller with blades D2 to the inner diameter D1 is from 2.4 to 2.6; wherein in the root section the chord length L1 is determined by the relationship: LI = kl л D1/N; where kl is the coefficient taking into account the properties of the blade material from 2 to 3.1; л is a mathematical constant -3.1415; N is the number of blades of the impeller, an even number in the range from 14 to 24; accordingly, for the tip of the blade, the chord length L2 = k2 л В2/М; where k2 is the coefficient of the material properties lies in the range from 1.3 to 1.5; and the radius of curvature of the concave part is determined from the expression: R2a = pi Ll; pi is the coefficient of the flow properties from 7.0 to 7.3 for the root part, R26 = pi L1 ; for the tip part, and the greatest thickness of the blade - И in the root section is determined from the expression: tl ^ Ll / y; where y is the coefficient taking into account the blade material from 15.5 to 15.7; accordingly, for the tip part t2 = L2/y; the radii of the convex surface of the blade profile in the root and end parts are determined by geometric construction.