JP2005226559A - Gas turbine engine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a gas turbine engine, not only capable of suppressing decline in efficiency of a high pressure turbine and preventing the high pressure turbine itself from becoming multistage but also capable of preventing a low pressure turbine arranged at a rear stage of the high pressure turbine from becoming multistage. <P>SOLUTION: The gas turbine engine (1) has the one-stage high pressure turbine (HT) and a low pressure turbine (LT) comprising an axial flow turbine. The relationship between an area S<SB>1</SB>of an inlet part of a nozzle N which comprises a plurality of stationary blades (31) arranged in rows in the circumferential direction and is arranged immediately before moving blades (32) of the high pressure turbine (HT) and an area S<SB>2</SB>of a throat part is shown by the expression: 0.18 ≤ S<SB>1</SB>/S<SB>2</SB>≤ 0.4. In addition, a flow velocity (Mach number M) of gas flowing through the nozzle N is defined as M0.05 ≤ V ≤ M0.2. The relationship between an area S<SB>3</SB>in the inlet side and an area S<SB>4</SB>in the outlet side of the moving blades (32) of the high pressure turbine (HT) adjacent to each other is shown by the expression: 0.3 ≤ S3/S4 ≤ 1.2. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

  The present invention relates to a gas turbine engine, and more particularly, to a gas turbine engine in which a single-stage high-pressure turbine and an low-pressure turbine each including an axial turbine are arranged in series.

In a gas turbine engine in which a nozzle composed of a plurality of stationary blades arranged in the circumferential direction is arranged immediately before a moving blade of a high-pressure turbine, in addition to increasing the efficiency of the high-pressure turbine, the pressure loss of the stationary blades constituting the nozzle And the resonance of the rotor blades becomes a problem. In particular, in a small gas turbine engine having a thrust of 3000 lb or less, there are large restrictions in properly determining the relationship between the moving blade and the stationary blade, and various ideas have been devised to improve efficiency (Patent Document 1). See).
JP 2003-106103 A

  In the study of the optimum setting of the aspect ratio of the stationary blade proposed in Patent Document 1, the present applicant improved the efficiency of the high-pressure turbine by appropriately setting the channel cross-sectional area between the adjacent stationary blades. It was found that a significant action can be obtained.

  The present invention is based on such knowledge, and its main purpose is to suppress the reduction in efficiency of the high-pressure turbine and to avoid the multi-stage of the high-pressure turbine itself. Another object of the present invention is to provide a gas turbine engine that can avoid the increase in the number of stages of the low-pressure turbine.

In order to solve such a problem, according to claim 1 of the present invention, gas turbine engines each having a high-pressure turbine (HT) and a low-pressure turbine (LT) each including an axial flow turbine are arranged in a circumferential direction. consists of a set by a plurality of vanes (31), relationship between the area S ₂ of the surface area S ₁ and the throat portion of the inlet portion of the nozzle N which is positioned immediately before the rotor blades of the high pressure turbine (32) 0.18 ≦ S ₁ / S ₂ ≦ 0.4.

According to a second aspect of the present invention, in addition to the above configuration, the flow velocity V (Mach number M) of the gas flowing through the nozzle is set to M0.05 ≦ V ≦ M0.2, and the high-pressure turbines are adjacent to each other. the relationship between the inlet side area S ₃ and the area S ₄ of the outlet side of between blades each other fit, and shall be characterized in that a _{0.3 ≦ S 3 / S 4 ≦} 1.2.

  According to the present invention, since the efficiency of the high-pressure turbine can be maintained within a predetermined range, it is possible to avoid the multistage of the high-pressure turbine and the low-pressure turbine, and a great effect in realizing a small and high-thrust turbine engine. Can be played.

  Hereinafter, the present invention will be described in detail with reference to the accompanying drawings.

  FIG. 1 is a schematic view of a multi-shaft bypass jet engine to which the present invention is applied. The engine 1 includes an outer casing 3 and an inner casing 4 that are connected to each other by a rectifying plate 2 and are coaxially arranged, and each has a cylindrical shape. The outer shaft 7 and the inner shaft 8 are formed of concentrically combined hollow shafts and are supported at the center of the inner casing 4 by bearings 5f, 5r, 6f, and 6r that are independent of each other. .

  The outer shaft 7 is integrally coupled with the impeller wheel 9 of the high-pressure centrifugal compressor HC on the front side and the turbine wheel 11 of the high-pressure turbine HT disposed adjacent to the nozzle N of the backflow combustion chamber 10 on the rear side. ing.

  The inner shaft 8 has a front fan 12 at its front end, a compressor wheel 13 constituting a moving blade of a low-pressure axial compressor LC behind the front fan 12, and a low-pressure turbine in the combustion gas injection duct 14 at its rear end. Turbine wheels 15 on which LT blades are placed are integrally coupled to each other.

  A nose cone 16 is provided at the center of the front fan 12, and a stationary blade 17 having an outer end coupled to the inner peripheral surface of the outer casing 3 is disposed behind the front fan 12.

  A stationary blade 18 of a low-pressure axial compressor LC is disposed on the inner periphery of the front end portion of the inner casing 4. Behind that, a suction duct 19 for feeding the air sucked by the front fan 12 and pre-compressed by the low-pressure axial compressor LC to the high-pressure centrifugal compressor HC, and the impeller casing 20 of the high-pressure centrifugal compressor HC continuous therewith. And are formed. Further, the bearing box 21 of the bearings 5f and 6f that supports the front end side of the outer shaft 7 and the inner shaft 8 is coupled to the inner peripheral side of the suction duct 19.

  Part of the air sucked by the front fan 12 is sent to the high-pressure centrifugal compressor HC via the low-pressure axial compressor LC as described above. The remaining relatively low speed and a large amount of air is jetted backward from the bypass duct 22 formed between the outer casing 3 and the inner casing 4 and becomes the main thrust in the low speed range.

  A diffuser 23 is coupled to the outer periphery of the high-pressure centrifugal compressor HC, and high-pressure air is fed into the backflow combustion chamber 10 provided immediately after that.

  In the reverse flow combustion chamber 10, the fuel injected from the fuel injection nozzle 24 provided on the rear end face and the high-pressure air sent from the diffuser 23 are mixed and burned. The main thrust in the high speed range is obtained by the combustion gas injected from the nozzle N facing backward to the atmosphere through the injection duct 14.

  A bearing box 25 of bearings 5r and 6r for supporting the rear end side of the outer shaft 7 and the inner shaft 8 is coupled to the inner peripheral side of the injection duct 14.

  An output shaft of a starter motor 26 is connected to the outer shaft 7 of the engine 1 via a gear mechanism (not shown). When this starter motor 26 is driven, the impeller wheel 9 of the high-pressure centrifugal compressor HC is driven together with the outer shaft 7, and high-pressure air is sent into the reverse flow combustion chamber 10. When this high pressure air and fuel are mixed and burned, the turbine wheel 11 of the high pressure turbine HT and the turbine wheel 15 of the low pressure turbine LT are driven by the injection pressure of the combustion gas. The impeller wheel 9 of the high-pressure centrifugal compressor HC is driven by the rotational force of the high-pressure turbine wheel 11, and the front fan 12 and the compressor wheel 13 of the low-pressure axial compressor LC are driven by the rotational force of the low-pressure turbine wheel 15, respectively. When the high pressure turbine wheel 11 and the low pressure turbine wheel 15 are driven by the combustion gas injection pressure, the engine 1 continues to rotate in a state determined according to a self-feedback balance between the fuel supply amount and the intake air amount. Become.

  In such an engine 1, in order to reduce the loss of the high-pressure turbine HT that drives the high-pressure compressor HC, the flow passage area of the stationary blades constituting the nozzle N that feeds the combustion gas to the high-pressure turbine HT is an important factor. ing.

Here, the area of the nozzle N, that is, the area between two stationary blades adjacent to each other in the circumferential direction is given by the product of the dimension between the stationary blades 31 and the radial dimension of the stationary blade 31 as shown in FIG. Lal is, minimum distance dimension to the area S ₁ between the combustion gas inlet-side edges to each other, and the curved surface of the outlet-side edge and the other stationary blade 31 of one vane 31 of the vane 31 adjacent to each other The relationship between the ratio (S ₂ / S ₁ ) of the area S ₂ of the portion (throat portion) to be obtained and the efficiency of the high-pressure turbine HT is as shown in FIG.

As can be seen from FIG. 3, the efficiency is maximized when the value of S ₂ / S ₁ is 0.18, and the efficiency is remarkably lowered in the region of 0.18 or less. If this is the flow rate of the nozzle outlet increases the area S ₂ of the throat portion is in accordance with a relatively small flow rate of the nozzle outlet V (Mach number M) is more than M0.2, surface of the blade 32 in the high pressure turbine HT This is because a shock wave is generated. When the value of S ₂ / S ₁ exceeds 0.18, that is, when the flow velocity V at the nozzle outlet decreases as the area S _{2 of the} throat portion relatively increases, the efficiency gradually decreases, and S ₂ / S _When the value of ₁ exceeds 0.4, the flow velocity V (Mach number M) at the nozzle outlet is less than M0.05, and the efficiency is halved from the peak value.

  This means that in order not to halve the work load of the high-pressure turbine HT, the high-pressure turbines HT must be arranged in a plurality of stages, or even if the work load of the high-pressure turbine HT is halved, the same work load as the low-pressure turbine LT. This means that a plurality of low-pressure turbines LT must be connected in series. This is not preferable because it leads to an increase in the axial length of the engine 1.

  In view of such knowledge, the present invention addresses this problem by setting the area ratio of the inlet / outlet of the nozzle N that ejects combustion gas to the high-pressure turbine HT in the range of 0.18 to 0.4.

In addition, regarding the area given by the product of the dimension between adjacent blades 32 of the high-pressure turbine HT and the radial dimension of the blade 32, the combustion gas inlet side edge of one blade 32 and the other blade to the area S ₃ of the portion spacing between the curved surface of the blade 32 is minimized, the site of the area S ₄ the distance between the curved surface is minimized on the outlet side edge and the other blades 32 of one of the blades 32 By setting the ratio (S ₄ / S ₃ ) in the range of 0.3 to 1.2, the efficiency of the high-pressure turbine HT can be set in a range that does not halve from the peak value. As a result, the high pressure turbine HT and the low pressure turbine LT need not be multistaged.

1 is a schematic cross-sectional view of a jet engine to which the present invention is applied. It is a relationship diagram of the area ratio of the entrance and exit of the nozzle by this invention, and the efficiency of a high pressure turbine. It is explanatory drawing which shows the relationship between the stationary blade which comprises a nozzle, and the moving blade of a high pressure turbine.

Explanation of symbols

HT High-pressure turbine LT Low-pressure turbine N Nozzle 1 Engine 31 Stator blade 32 Rotor blade

Claims (2)

A gas turbine engine in which a single-stage high-pressure turbine and a low-pressure turbine each consisting of an axial turbine are arranged in series,
Is composed of a plurality of vanes which are arrayed in the circumferential direction, the relationship between the area S ₂ of the surface area S ₁ and the throat portion of the inlet portion of the nozzle which is arranged immediately before the rotor blades of the high pressure turbine,
A gas turbine engine characterized by 0.18 ≦ S ₁ / S ₂ ≦ 0.4. The flow velocity V of gas flowing through the nozzle (Mach number M), and M0.05 ≦ V ≦ M0.2, and the high-pressure inlet side area S ₃ and the outlet side of the inter-rotor blades adjacent to each other of the turbine The gas turbine engine according to claim 1, wherein a relationship with the area S ₄ is 0.3 ≦ S ₃ / S ₄ ≦ 1.2.

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