阅读说明：本技术 一种用于火箭发动机的涡轮泵装置 (Turbo pump device for rocket engine ) 是由 李晓俊 林言丕 朱祖超 李林敏 于 2021-09-15 设计创作，主要内容包括：本发明公开了一种用于火箭发动机的涡轮泵装置,其包括壳体(1)、离心叶轮(2)、连接轴组件(3),轮毂的外周设有主叶片、螺旋诱导叶片,其特征在于：连接轴组件包括第一轴(31)、第二轴(32),第一轴的左端具有台阶部(37),台阶部插入轮毂的相应台阶孔内,第一轴的左端具有第一螺纹孔,第一螺纹孔内连接有第一螺杆(33),第一螺杆的左端具有肩头部(36),肩头部具有挤压锥形面(38),第二螺纹孔内连接有第二螺杆(34),第二螺杆的左端连接有导流帽(35)。本发明具有连接强度大、安装/拆卸方便的优点；其通过凹槽和/或辅助导叶的设计,能够减少扩压器的压力损失,从而提高涡轮泵的出口压力。(The invention discloses a turbo pump device for a rocket engine, which comprises a shell (1), a centrifugal impeller (2) and a connecting shaft assembly (3), wherein the periphery of a hub is provided with main blades and spiral induction blades, and the turbo pump device is characterized in that: the connecting shaft assembly comprises a first shaft (31) and a second shaft (32), a step part (37) is arranged at the left end of the first shaft, the step part is inserted into a corresponding step hole of the hub, a first threaded hole is formed in the left end of the first shaft, a first screw rod (33) is connected into the first threaded hole, a shoulder part (36) is arranged at the left end of the first screw rod, an extrusion conical surface (38) is arranged at the shoulder part, a second screw rod (34) is connected into the second threaded hole, and a diversion cap (35) is connected to the left end of the second screw rod. The invention has the advantages of high connection strength and convenient assembly/disassembly; the pressure loss of the diffuser can be reduced through the design of the grooves and/or the auxiliary guide vanes, so that the outlet pressure of the turbo pump is improved.)

1. The utility model provides a turbo pump device for rocket engine, its includes casing (1), centrifugal impeller (2), connecting axle subassembly (3), installs centrifugal impeller in the pump chamber of casing, and centrifugal impeller passes through the connecting axle subassembly to be installed on the casing, and centrifugal impeller includes wheel hub (21), and wheel hub's periphery is equipped with main blade (22), spiral induction blade (23), and spiral induction blade is located main blade's upper reaches, its characterized in that: connecting axle subassembly (3) are including first axle (31), second axle (32), the left end of second axle inserts in the step concave hole of first axle right-hand member, the right-hand member of second axle is connected with the combustion part or the injection portion of engine, the left end of first axle has step portion (37), step portion inserts in wheel hub's corresponding step hole, the left end of first axle has first screw hole, first threaded hole is connected with first screw rod (33), the left end of first screw rod has shoulder head portion (36), shoulder head portion has extrusion conical surface (38), extrusion conical surface cooperatees with wheel hub's corresponding conical surface, the right-hand member of first screw rod has the second screw hole, be connected with second screw rod (34) in the second threaded hole, the left end of second screw rod is connected with water conservancy diversion cap (35).

2. A turbo pump apparatus for a rocket engine according to claim 1, wherein the hub (21) is provided at a left end thereof with a connection cavity (24), the shoulder portion (36) and the second screw (34) are located in the connection cavity, the first shaft (31) is provided at an outer circumference thereof with a key groove and a connection key (25), the first shaft is connected to the hub through the connection key, and a tapered surface of the hub, which is engaged with the pressing tapered surface (38), is located adjacent to the stepped recess.

3. A turbo pump device for a rocket engine according to claim 2, wherein the pump chamber of the housing (1) has a front side wall (11) and a rear side wall (12) which are opposite to each other, a diffuser guide vane (13) is disposed between the front side wall and the rear side wall, the plurality of diffuser guide vanes are uniformly distributed along the circumferential direction, the diffuser guide vanes are located at the radial outer side of the centrifugal impeller (2), the front side wall and the rear side wall include a first inclined portion (14), a protruding portion (16) and a second inclined portion (15) which are sequentially connected, the first inclined portion, the protruding portion and the second inclined portion are sequentially connected along the radial direction, a groove (17) is disposed on the rear side wall (12) between adjacent diffuser guide vanes, and the groove is formed in a recessed manner to remove material.

4. A turbo pump device for a rocket engine according to claim 3, wherein one or more auxiliary vanes (18) are provided in said groove (17).

5. A turbo pump device for a rocket engine according to claim 4, wherein the radially inner end and the radially outer end of said groove (17) are both arc-shaped portions.

6. A turbo pump device for a rocket engine according to claim 5, wherein the axial depth of said grooves (17) increases gradually from the radially inner end to the radially outer end.

7. A turbo pump device for rocket engines according to claim 6 wherein two auxiliary guide vanes (18) are provided between adjacent diffuser guide vanes (13), the inner radius of the auxiliary guide vanes being equal to the inner radius of the diffuser guide vanes, the outer radius of the auxiliary guide vanes being not equal to the outer radius of the diffuser guide vanes; and the diffuser vane has an outer radius R, the auxiliary vane has an outer radius R1, R1 < R, and R1 ═ 0.70-0.97) R.

8. A turbo pump device for a rocket engine according to claim 7, wherein the diffuser guide vane (13) has a different blade profile from the auxiliary guide vane (18), and the thickness of the blade of the diffuser guide vane is gradually increased in the direction of the liquid flow, and the blade of the auxiliary guide vane is a blade having a uniform thickness.

9. A turbo pump apparatus for a rocket engine according to claim 2 or 7, wherein the inducer includes a spiral inducing blade (23) including a plurality of first spiral blades (231), a plurality of divided spiral blades (232), the divided spiral blades being disposed between the first spiral blades, a blade entrance angle of the divided spiral blades being larger than a blade entrance angle of the first spiral blades, a blade exit angle of the divided spiral blades being larger than a blade exit angle of the first spiral blades, and a magnitude relation being 1.05 to 1.3 times.

Technical Field

The invention relates to the technical field of rocket engines, in particular to a turbopump device for a rocket engine.

Background

The turbopump of the rocket engine mainly comprises an inducer, a centrifugal impeller, a mechanical seal, a bearing, a shafting supporting system, a shell and the like. However, the existing turbine pump device has the problems of insufficient impeller installation strength, inconvenient installation/disassembly and large pressure loss of the pump outlet.

Disclosure of Invention

The invention aims to overcome the defects in the prior art, and provides a turbopump device for a rocket engine, which has the advantages of high connection strength and convenience in mounting/dismounting compared with the existing impeller mounting structure through the design of a connecting shaft assembly. The design of the groove and/or the auxiliary guide vane can reduce the pressure loss of the diffuser, so that the liquid flow orderly flows to the outlet pipe of the volute, the outlet pressure of the turbo pump is improved, and the power loss of the turbo pump is reduced.

In order to achieve the purpose, the invention adopts the technical scheme that:

the utility model provides a turbo pump device for rocket engine, its includes casing (1), centrifugal impeller (2), connecting axle subassembly (3), installs centrifugal impeller in the pump chamber of casing, and centrifugal impeller passes through the connecting axle subassembly to be installed on the casing, and centrifugal impeller includes wheel hub (21), and wheel hub's periphery is equipped with main blade (22), spiral induction blade (23), and spiral induction blade is located main blade's upper reaches, its characterized in that: connecting axle subassembly (3) are including first axle (31), second axle (32), the left end of second axle inserts in the step concave hole of first axle right-hand member, the right-hand member of second axle is connected with the combustion part or the injection portion of engine, the left end of first axle has step portion (37), step portion inserts in wheel hub's corresponding step hole, the left end of first axle has first screw hole, first threaded hole is connected with first screw rod (33), the left end of first screw rod has shoulder head portion (36), shoulder head portion has extrusion conical surface (38), extrusion conical surface cooperatees with wheel hub's corresponding conical surface, the right-hand member of first screw rod has the second screw hole, be connected with second screw rod (34) in the second threaded hole, the left end of second screw rod is connected with water conservancy diversion cap (35).

Furthermore, the left end of wheel hub (21) is equipped with connection chamber (24), and shoulder head portion (36), second screw rod (34) are located connection chamber, and the periphery of primary shaft (31) is provided with keyway, connection key (25), and the primary shaft is connected with wheel hub through the connection key, and the conical surface and the step shrinkage pool of wheel hub with extrusion conical surface (38) matched with are close to be set up.

Further, the pump cavity of casing (1) has preceding lateral wall (11), back lateral wall (12) that relative setting, is provided with diffuser stator (13) between preceding lateral wall and the back lateral wall, a plurality of diffuser stator along circumference evenly distributed, the diffuser stator is located centrifugal impeller (2) radial outside, preceding lateral wall, back lateral wall are including the first rake (14), bellying (16), second rake (15) that connect gradually, first rake, bellying, second rake radially connect gradually, be provided with recess (17) on back lateral wall (12) between the adjacent diffuser stator, the recess is got rid of the material with sunken mode and is formed.

Further, one or more auxiliary guide vanes (18) are arranged in the groove (17).

Further, the radially inner end and the radially outer end of the groove (17) are both arc-shaped portions.

Further, the axial depth of the groove (17) gradually increases from the radially inner end to the radially outer end.

Furthermore, two auxiliary guide vanes (18) are arranged between adjacent diffuser guide vanes (13), the inner radius of each auxiliary guide vane is equal to that of each diffuser guide vane, and the outer radius of each auxiliary guide vane is not equal to that of each diffuser guide vane; and the diffuser vane has an outer radius R, the auxiliary vane has an outer radius R1, R1 < R, and R1 ═ 0.70-0.97) R.

Furthermore, the blade profile of the diffuser guide blade (13) is different from the blade profile of the auxiliary guide blade (18), the thickness of the blade of the diffuser guide blade is gradually increased along the flow direction, and the blade of the auxiliary guide blade is an equal-thickness blade.

Further, the inducer includes spiral induction blade (23), and spiral induction blade includes a plurality of first helical blade (231), a plurality of reposition of redundant personnel helical blade (232), and reposition of redundant personnel helical blade sets up between first helical blade, and reposition of redundant personnel helical blade's blade entrance angle is greater than first helical blade's blade entrance angle, and reposition of redundant personnel helical blade's blade exit angle is greater than first helical blade's blade exit angle, and the size relation is 1.05-1.3 times.

Compared with the existing impeller mounting structure, the turbopump device for the rocket engine has the advantages of high connection strength and convenience in mounting/dismounting through the design of the connecting shaft assembly. The design of the groove and/or the auxiliary guide vane can reduce the pressure loss of the diffuser, so that the liquid flow orderly flows to the outlet pipe of the volute, the outlet pressure of the turbo pump is improved, and the power loss of the turbo pump is reduced. According to the inducer, the diversion helical blades are designed in an offset independent mode, the inlet and outlet angles of the diversion blades are slightly larger than the inlet and outlet angles of the first helical blades, the effect of preventing the inlet blades of the inducer from being blocked can be achieved, and meanwhile the effect of improving the total lift of the inducer can be improved.

Drawings

FIG. 1 is a schematic view showing the construction of a turbopump apparatus for a rocket engine according to the present invention (a partial structural view of a turbopump);

FIG. 2 is a schematic view of a partial enlarged structure of a turbopump apparatus for a rocket engine according to the present invention;

FIG. 3 is a schematic side view of a turbopump apparatus for a rocket engine according to the present invention;

fig. 4 is a schematic structural view of an inducer of the turbopump apparatus for a rocket engine according to the present invention.

In the figure: the centrifugal impeller comprises a shell 1, a centrifugal impeller 2, a connecting shaft assembly 3, a front side wall 11, a rear side wall 12, diffuser guide vanes 13, a first inclined part 14, a second inclined part 15, a convex part 16, a groove 17, auxiliary guide vanes 18, a hub 21, main vanes 22, spiral induction vanes 23, a connecting cavity 24, a connecting key 25, a first shaft 31, a second shaft 32, a first screw 33, a second screw 34, a deflector cap 35, a shoulder part 36, a step part 37, an extrusion conical surface 38, first spiral vanes 231 and split spiral vanes 232.

Detailed Description

In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.

The present invention will be described in further detail with reference to the accompanying drawings.

As shown in fig. 1 to 4, a turbo pump device for a rocket engine includes a housing 1, a centrifugal impeller 2, and a connecting shaft assembly 3, the centrifugal impeller 2 is installed in a pump cavity of the housing 1, the centrifugal impeller 2 is installed on the housing 1 through the connecting shaft assembly 3, the centrifugal impeller 2 includes a hub 21, a main blade 22 and a spiral inducing blade 23 are provided at the periphery of the hub 21, the spiral inducing blade 23 is located at the upstream of the main blade 22, and the turbo pump device is characterized in that: the connecting shaft assembly 3 comprises a first shaft 31 and a second shaft 32, the left end of the second shaft 32 is inserted into a stepped concave hole at the right end of the first shaft 31, the right end of the second shaft 32 is connected with a combustion part or an injection part of an engine, the left end of the first shaft 31 is provided with a stepped part 37, the stepped part 37 is inserted into a corresponding stepped hole of the wheel hub 21, the left end of the first shaft 31 is provided with a first threaded hole, a first screw 33 is connected into the first threaded hole, the left end of the first screw 33 is provided with a shoulder head part 36, the shoulder head part 36 is provided with an extrusion conical surface 38, the extrusion conical surface 38 is matched with a corresponding conical surface of the wheel hub 21, the right end of the first screw 33 is provided with a second threaded hole, a second screw 34 is connected into the second threaded hole, the left end of the second screw 34 is connected with a diversion cap 35, the left end of the wheel hub 21 is provided with a connecting cavity 24, the shoulder head part 36 and the second screw 34 are positioned in the connecting cavity 24, the periphery of the first shaft 31 is provided with a key groove, The connecting key 25, the first shaft 31 is connected with the hub 21 through the connecting key 25, and the conical surface of the hub 21 matched with the extruding conical surface 38 is arranged adjacent to the step concave hole.

The centrifugal impeller 2 is a semi-open impeller, and the turbo pump is used for pumping liquid fuels such as low-temperature methane or low-temperature liquid oxygen.

Compared with the existing impeller mounting structure, the turbopump device for the rocket engine has the advantages of high connection strength and convenience in mounting/dismounting through the design of the connecting shaft assembly 3.

The first screw 33 and the deflector cap 35 are provided with wrench grooves for operating a wrench tool.

Further, as shown in fig. 2-3, the pump chamber of the casing 1 has a front side wall 11 and a rear side wall 12 which are arranged opposite to each other, diffuser guide vanes 13 are arranged between the front side wall 11 and the rear side wall 12, the diffuser guide vanes 13 are uniformly distributed along the circumferential direction, the diffuser guide vanes 13 are located on the radial outer side of the centrifugal impeller 2, the front side wall 11 and the rear side wall 12 include a first inclined portion 14, a protruding portion 16 and a second inclined portion 15 which are connected in sequence, the first inclined portion 14, the protruding portion 16 and the second inclined portion 15 are connected in sequence along the radial direction, a groove 17 is arranged on the rear side wall 12 between adjacent diffuser guide vanes 13, the groove 17 is formed in a recessed manner to remove material, and one or more auxiliary guide vanes 18 are arranged in the groove 17.

The turbo pump device for the rocket engine can reduce the pressure loss of a diffuser through the design of the groove 17 and/or the auxiliary guide vane 18, so that liquid flow orderly flows to the outlet pipe of the volute, the outlet pressure of the turbo pump is increased, and the power loss of the turbo pump is reduced.

Further, the radial inner end and the radial outer end of the groove 17 are arc-shaped parts, so that a better flow guide effect is achieved.

Further, the axial depth of the groove 17 gradually increases from the radially inner end to the radially outer end, enabling the pressure loss of the diffuser to be further reduced.

Further, two auxiliary guide vanes 18 are arranged between adjacent diffuser guide vanes 13, the inner radius of each auxiliary guide vane 18 is equal to that of each diffuser guide vane 13, and the outer radius of each auxiliary guide vane 18 is not equal to that of each diffuser guide vane 13; the diffuser vane 13 has an outer radius R and the auxiliary vane 18 has an outer radius R1, R1 < R, R1 ═ (0.80-0.95) R, preferably R1 ═ 0.85-0.90R.

The blade profile of the diffuser guide vane 13 is different from the blade profile of the auxiliary guide vane 18, the blade thickness of the diffuser guide vane 13 is gradually increased along the liquid flow, and the blade of the auxiliary guide vane 18 is an equal-thickness blade.

As shown in fig. 4, the inducer includes a spiral inducer blade 23, the spiral inducer blade 23 includes a plurality of first spiral blades 231 and a plurality of split spiral blades 232, and the split spiral blades 232 are disposed between the first spiral blades 231; the blade entrance angle of the spiral splitter blade 232 is slightly larger than the blade entrance angle of the first spiral blade 231, and the blade exit angle of the spiral splitter blade 232 is slightly larger than the blade exit angle of the first spiral blade 231, for example, 1.05 to 1.3 times. According to the inducer, the shunting spiral blades 232 are designed in an offset independent mode, the inlet and outlet angles of the shunting blades 232 are slightly larger than the inlet and outlet angles of the first spiral blades 231, the effect of preventing the inlet blades of the inducer from being blocked can be achieved, and meanwhile the effect of improving the total lift of the inducer can be improved.

Compared with the existing impeller mounting structure, the turbopump device for the rocket engine has the advantages of high connection strength and convenience in mounting/dismounting through the design of the connecting shaft assembly 3. The design of the groove 17 and/or the auxiliary guide vane 18 can reduce the pressure loss of the diffuser, so that the liquid flow orderly flows to the outlet pipe of the volute, the outlet pressure of the turbo pump is increased, and the power loss of the turbo pump is reduced.

The above-described embodiments are illustrative of the present invention and not restrictive, it being understood that various changes, modifications, substitutions and alterations can be made herein without departing from the principles and spirit of the invention, the scope of which is defined by the appended claims and their equivalents.