CN104234756B - A kind of transonic speed type film cooling holes - Google Patents

A kind of transonic speed type film cooling holes Download PDF

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Publication number
CN104234756B
CN104234756B CN201410465542.6A CN201410465542A CN104234756B CN 104234756 B CN104234756 B CN 104234756B CN 201410465542 A CN201410465542 A CN 201410465542A CN 104234756 B CN104234756 B CN 104234756B
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cooling holes
throat
film cooling
section
air
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CN201410465542.6A
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CN104234756A (en
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朱惠人
苏文超
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Northwestern Polytechnical University
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Northwestern Polytechnical University
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Abstract

The invention discloses a kind of transonic speed type film cooling holes.Between turbo blade suction surface film cooling holes and the center line of pressure face film cooling holes and blade surface, angle is 30 ° 60 °, and transonic speed type film cooling holes general structure is divided into contraction section, throat and expansion segment.Transonic speed type film cooling holes contraction section centerline length l1It is 6～10 times of throat radius, calculates the molded line of wall from film cooling holes entrance to throat with Vito Xin Siji formula；Throat is cold air from subsonic speed to ultrasonic changeover portion, and throat section is long-pending utilizes cold air flow to determine；Smoothly transitting from throat to expansion segment, use an essentially identical radius of curvature, obtain the expansion segment wall smoothed, expansion segment semiapex angle β is 4 °～6 °, tries to achieve expansion segment centerline length l by semiapex angle2.Transonic speed type film cooling holes is designed as shrinkage expansion shape, and when cold air flows through air film hole, flow velocity can be increased to supersonic speed by subsonic speed, it is ensured that Film Cooling.

Description

A kind of transonic speed type film cooling holes
Technical field
The invention belongs to aero engine turbine blades technical field, specifically, relate to a kind of transonic speed type Film cooling holes.
Background technology
Along with the fast development of aero-engine, turbine inlet temperature (TIT) improves constantly, and the one-level of thrust-weight ratio 10 is sent out The turbine inlet temperature (TIT) of motivation has reached 1900K～2000K, this heat-resisting far beyond turbo blade material The limit.At present, in aero-engine, turbo blade is cooled down by widely used air film cooling technology, i.e. from Compressor extraction section gas introduces blade interior as cold air, effectively cools down the inner surface of blade, The most a part of cold air is sprayed by the air film hole of blade wall, makes cold air under the effect of blade exterior main flow Form one layer of cold air film at blade surface, blade surface is isolated with high-temperature fuel gas, and then to turbo blade It is effectively protected.
Present stage, in advanced aero engine, turbo blade pressure face and suction surface subregion combustion gas speed are Reach supersonic speed, in order to make gaseous film control have preferable cooling effect, need to keep certain air blowing ratio, Require that corresponding cooling rates is the highest, owing to cold air temperature is more much lower than fuel gas temperature, therefore cold air Local velocity of sound is less than the local velocity of sound of combustion gas, and cold air more easily reachs velocity of sound even supersonic speed.
At present the most cylindrical air film hole on turbo blade, expansion shape air film hole, slit air film hole and double Fan-shaped air film hole, when cold air flows through air film hole, it is impossible to make cooling rates reach supersonic speed from subsonic speed.Send out Bright patent 200910121310.8 discloses one side and becomes kidney-shaped slit film cooling holes, owing to air film hole cuts The change of face shape, cold air gradually extends to both sides after flowing into air film hole, and thickness is the most thinning, this hole The hole row of type composition is easily formed continuous print air film, can obtain being better than the cooling effect of cylindrical hole row, but this The cross sectional shape planting pass is become kidney-shaped from rectangle, and cold air can not be made to reach supersonic speed from subsonic speed continuously. In continuous print flows, due to the impact of air-flow compressibility, air-flow to be allowed just accelerates to supersonic speed from subsonic speed Must first go through one and shrink shape hole, then by an expansion shape hole.Therefore, in pressure face and the suction of blade On the subregion in power face, in order to cold air to be made accelerates to supersonic speed from subsonic speed in air film hole, it should will Air film hole Cross section Design is the shape first shrinking further expansion.
Summary of the invention
The deficiency existed in order to avoid prior art, overcomes air film hole type can not flow through the cold air in hole from Asia Velocity of sound accelerates to ultrasonic problem, and the present invention proposes a kind of transonic speed type film cooling holes.Use and starting Being respectively provided with transonic speed type film cooling holes on the pressure face of machine turbo blade and suction surface, film cooling holes cuts Area is first to shrink to expand afterwards, when cold air flows through transonic speed type film cooling holes, can accelerate to surpass from subsonic speed Velocity of sound.
The technical solution adopted for the present invention to solve the technical problems is: include turbine blade internal cooling passage, suction Power face film cooling holes, pressure face film cooling holes, is characterized in that, described film cooling holes is shrinkage expansion Shape structure, it is oval for being divided into contraction section, throat, expansion segment, film cooling holes entrance section and outlet Shape, throat section is circular, and film cooling holes is symmetricly set on turbo blade two sides, suction surface air film Cooling Holes is identical with pressure face gaseous film control hole shape, and is connected with turbine blade internal cooling passage, suction surface The center line of film cooling holes and pressure face film cooling holes respectively and between turbine blade surface angle α be 30 °～60 °, film cooling holes contraction section centerline length is 6～10 times of throat radius, film cooling holes entrance Molded line to throat's wall is carried out by Vito Xin Siji formula:
( r 2 r ) 2 = 1 - ( 1 - 1 ( r 1 / r 2 ) 2 ) [ 1 - ( x / l 1 ) 2 ] 2 [ 1 + 1 3 ( x / l 1 ) 2 ] 3
In formula, r is the radius in cross section, any perpendicular flow direction, r in contraction section1For an o1To line a1-b1Vertical Distance, x is contraction section center line o1-o2Upper any point-to-point o1Distance；Film cooling holes throat is cold air From subsonic speed to ultrasonic changeover portion, the long-pending calculating in throat section uses formula:Wherein m For by the cold air flow of air film hole, ρ is the density of throat's air-flow, and T is the temperature of throat's air-flow；Air film is cold But hole smoothly transits as expansion segment wall from throat to expansion segment, and expansion segment semiapex angle β is 4 °～6 °, expansion segment Center line is:R in formula3For an o3To line b2-c2Vertical range, r2For throat's vertical current The radius in dynamic cross section, direction.
Beneficial effect
The transonic speed type film cooling holes that the present invention proposes is shrinkage expansion shape structure, and film cooling holes is divided into receipts Contracting section, throat, expansion segment, film cooling holes entrance section and outlet are oval, and throat section is Circle, film cooling holes is symmetricly set on turbo blade two sides；Turbo blade suction surface film cooling holes And the center line of pressure face film cooling holes respectively and between blade surface angle be 30 °～60 °, transonic speed type gas Film Cooling Holes contraction section centerline length l1It is 6～10 times of throat radius, from film cooling holes entrance to throat, Vito Xin Siji formula is used to calculate the molded line of wall；Throat be cold air from subsonic speed to ultrasonic transition Section, throat section is long-pending utilizes cold air flow to determine；Smoothly transit from throat to expansion segment, use one Individual essentially identical radius of curvature, obtains the expansion segment wall smoothed, and expansion segment semiapex angle β is 4 °～6 °, logical Cross semiapex angle and try to achieve expansion segment centerline length l2。
Transonic speed type film cooling holes of the present invention is compared with the Cooling Holes pass of prior art, due to transonic speed type Film cooling holes Cross section Design is shrinkage expansion shape, when cold air is when by air film hole, and flow velocity can be by subsonic speed Increase to supersonic speed, make cold air spray from air film hole with the speed more than 1Ma, hence with transonic speed type gas Film Cooling Holes can reach the air blowing ratio of anticipation when cooling down, thus ensures that gaseous film control has preferably cooling Effect.
Accompanying drawing explanation
With embodiment, one transonic speed type film cooling holes of the present invention is made the most below in conjunction with the accompanying drawings Describe in detail bright.
Fig. 1 is transonic speed type film cooling holes structural representation of the present invention.
Fig. 2 is transonic speed type film cooling holes front view of the present invention.
Fig. 3 is transonic speed type film cooling holes axonometric drawing of the present invention.
In figure:
1. suction surface film cooling holes 2. pressure face film cooling holes 3. turbine blade internal cooling passage
Detailed description of the invention
Embodiment one
The present embodiment is the transonic speed type gaseous film control pore structure on certain h type engine h moving turbine blade.
Refering to Fig. 1, Fig. 2, Fig. 3, the present embodiment is close trailing edge position on the suction surface of engine turbine blade The place of putting is provided with transonic speed type film cooling holes 1, and suction surface film cooling holes 1 leads to turbine blade internal cooling Road 3 is connected.
In the present embodiment, suction surface transonic speed type film cooling holes general structure is divided into contraction section, throat and expansion Open section.Angle α, angle α is had to take between center line and the blade surface of suction surface transonic speed type film cooling holes Being 45 °, suction surface transonic speed type film cooling holes entrance section is oval, a1With a0Between distance for entering Length s of mouth cross section ellipse semi-major axis1, a3With a0Between the length that distance is entrance section ellipse semi-minor axis s2.Suction surface transonic speed type film cooling holes throat takes perpendicular flow direction cross section b1-b2-b3-b4, this cross section is Circle, b1With o2Between the radius r that distance is cross section, throat's perpendicular flow direction2.Suction surface transonic speed type Film cooling holes outlet is also oval, c1With c0Between distance be outlet ellipse semi-major axis Length s3, c3With c0Between length s that distance is outlet ellipse semi-minor axis4。o1With o2Between away from From for streamwise contraction section center line o1-o2Length l1, o2With o3Between distance be streamwise Expansion segment center line o2-o3Length l2。
The contraction section Main Function of suction surface transonic speed type film cooling holes is to accelerate cold air, and ensures to enter Throat's air-flow is uniform and stable, and the performance of contraction section is had a major impact by the shape of contraction section curve.Suction surface across Velocity of sound type film cooling holes contraction section centerline length l16 times of throat radius r can be taken2.From contraction section entrance Using Vito Xin Siji formula to calculate the molded line of wall to throat portion, its formula is:
( r 2 r ) 2 = 1 - ( 1 - 1 ( r 1 / r 2 ) 2 ) [ 1 - ( x / l 1 ) 2 ] 2 [ 1 + 1 3 ( x / l 1 ) 2 ] 3
The radius in cross section, any perpendicular flow direction, r during in formula, r is contraction section1For an o1To line a1-b1Vertical away from From, x is contraction section center line o1-o2Upper any point-to-point o1Distance.Utilize the suction surface that this formula obtains Transonic speed type film cooling holes contraction section wall molded line smooths, and makes air-flow gradually expand in air film hole, it is thus achieved that Uniform flow field.
Suction surface transonic speed type film cooling holes throat be cold air from subsonic speed to ultrasonic changeover portion, throat Determination whole film cooling holes design in important.Flow by transonic speed type film cooling holes is big Little being amassed by throat section is limited, and throat opening area calculates employing formula and is:
A = m ρ κ R g T
Wherein, m is the cold air flow by air film hole, and ρ is the density of throat's air-flow, and T is throat's air-flow Temperature.
Suction surface transonic speed type film cooling holes expansion segment continues to make cold air be accelerated, and now cold air reaches Supersonic speed, can produce friction loss and eddy current loss in flow process, expansion segment can not long can not Too short, if expansion segment is long, then can produce the biggest friction loss, and expansion segment is too short, this can cause Air-flow separates with tube wall, produces eddy current loss.Should smoothly transit from throat to expansion segment, use one Roughly the same radius of curvature, obtains the expansion segment wall smoothed；Owing to suction surface gas flow rate is very fast, for The air blowing ratio that guarantee is certain, the cold air flow velocity of needs is also very fast, and expansion segment semiapex angle β is taken as 6 °, expansion Section center line o2-o3The calculating formula of length is:
l 2 = r 3 - r 2 tan β
R in formula3For an o3To line b2-c2Vertical range, r2Radius for cross section, throat's perpendicular flow direction.
Embodiment two
The present embodiment is the transonic speed type gaseous film control pore structure on certain h type engine h moving turbine blade.
Refering to Fig. 1, Fig. 2, Fig. 3, the present embodiment is close trailing edge position on the pressure face of engine turbine blade The place of putting is provided with pressure face transonic speed type film cooling holes 2, and pressure face film cooling holes 2 and turbo blade Inner cooling path 3 is connected.
In the present embodiment, pressure face transonic speed type film cooling holes general structure is divided into contraction section, throat and expansion Open section.Angle α, angle α is had to take between center line and the blade surface of pressure face transonic speed type film cooling holes Being 60 °, pressure face transonic speed type film cooling holes entrance section is oval, a1With a0Between distance for entering Length s of mouth cross section ellipse semi-major axis1, a3With a0Between the length that distance is entrance section ellipse semi-minor axis s2.Pressure face transonic speed type film cooling holes throat takes perpendicular flow direction cross section b1-b2-b3-b4, this cross section is Circle, b1With o2Between the radius r that distance is cross section, throat's perpendicular flow direction2.Pressure face transonic speed type Film cooling holes outlet is also oval, c1With c0Between distance be outlet ellipse semi-major axis Length s3, c3With c0Between length s that distance is outlet ellipse semi-minor axis4。o1With o2Between away from From for streamwise contraction section center line o1-o2Length l1, o2With o3Between distance be streamwise Expansion segment center line o2-o3Length l2。
The contraction section Main Function of pressure face transonic speed type film cooling holes is to accelerate cold air, and ensures to enter Throat's air-flow is uniform and stable, and the performance of contraction section is had a major impact by the shape of contraction section curve.Pressure face across Velocity of sound type film cooling holes contraction section centerline length l1Desirable 10 times of throat radius r2.From contraction section entrance to Throat portion Vito Xin Siji formula calculates the molded line of wall, and its formula is:
( r 2 r ) 2 = 1 - ( 1 - 1 ( r 1 / r 2 ) 2 ) [ 1 - ( x / l 1 ) 2 ] 2 [ 1 + 1 3 ( x / l 1 ) 2 ] 3
In formula, r is the radius in cross section, any perpendicular flow direction, r in contraction section1For an o1To line a1-b1Hang down Straight distance, x is contraction section center line o1-o2Upper any point-to-point o1Distance.Utilize the pressure that this formula obtains Power face transonic speed type film cooling holes contraction section wall molded line smooths, and makes air-flow gradually expand in air film hole, Obtain uniform flow field.
Pressure face transonic speed type film cooling holes throat be cold air from subsonic speed to ultrasonic changeover portion, throat Determination important in whole Nozzle Design.Flow by pressure face transonic speed type film cooling holes is big Little being amassed by throat section is limited, and throat opening area calculates employing formula and is:
A = m ρ κ R g T
Wherein, m is the cold air flow by air film hole, and ρ is the density of throat's air-flow, and T is throat's air-flow Temperature.
Pressure face transonic speed type film cooling holes expansion segment continues to make cold air be accelerated, and now cold air reaches Supersonic speed, should smoothly transit from throat to expansion segment, uses a roughly the same radius of curvature, Obtain the expansion segment wall smoothed.The combustion gas speed of pressure face hole location is than the combustion gas of suction surface relevant position Speed is smaller, it is ensured that under the conditions of certain air blowing ratio, the cooling rates that this position of pressure face needs is the least by one A bit, expansion segment semiapex angle β is taken as 4 °, expansion segment center line o2-o3The calculating formula of length is:
l 2 = r 3 - r 2 tan β
In formula, r3For an o3To line b2-c2Vertical range, r2Radius for cross section, throat's perpendicular flow direction.

Claims (1)

1. a transonic speed type film cooling holes, including turbine blade internal cooling passage, suction surface film cooling holes, Pressure face film cooling holes, it is characterised in that: described film cooling holes is shrinkage expansion shape structure, is divided into receipts Contracting section, throat, expansion segment, film cooling holes entrance section and outlet are oval, and throat section is Circle, film cooling holes is symmetricly set on turbo blade two sides, suction surface film cooling holes and pressure face Gaseous film control hole shape is identical, and is connected with turbine blade internal cooling passage, suction surface film cooling holes and pressure The center line of power face film cooling holes respectively and between turbine blade surface angle α be 30 °～60 °, gaseous film control Hole contraction section centerline length is 6～10 times of throat radius, the molded line of film cooling holes entrance to throat's wall Carry out by Vito Xin Siji formula:
( r 2 r ) 2 = 1 - ( 1 - 1 ( r 1 / r 2 ) 2 ) [ 1 - ( x / l 1 ) 2 ] 2 [ 1 + 1 3 ( x / l 1 ) 2 ] 3
In formula, r is the radius in cross section, any perpendicular flow direction, r in contraction section1For an O1To line a1-b1Vertical Distance, r2For being perpendicular to the radius of the throat section of flow direction, l1For contraction section length, x is in contraction section Heart line O1-O2Upper any point-to-point O1Distance；Point O1Intersection point for air film hole entrance Yu center line；Point O2 For air film hole center line midpoint；Line a1-b1Molded line for air film hole contraction section；Film cooling holes throat is cold air From subsonic speed to ultrasonic changeover portion, the long-pending calculating in throat section uses formula:Wherein A For being perpendicular to the throat section area of center line, m is the cold air flow by air film hole, and ρ is throat's air-flow Density, T is the temperature of throat's air-flow；Film cooling holes smoothly transits as expansion segment wall from throat to expansion segment, Expansion segment semiapex angle β is 4 °～6 °, and expansion segment center line is:R in formula3For an O3To line b2-c2 Vertical range, r2Radius for cross section, throat's perpendicular flow direction；Point O3For air film hole outlet and center line Intersection point；Line b2-c2Molded line for air film hole expansion segment.
CN201410465542.6A 2014-09-15 2014-09-15 A kind of transonic speed type film cooling holes Expired - Fee Related CN104234756B (en)

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CN104895620B (en) * 2015-04-20 2016-08-10 西北工业大学 A kind of arrowhead-shaped diplopore cellular construction for gaseous film control
JP6550000B2 (en) 2016-02-26 2019-07-24 三菱日立パワーシステムズ株式会社 Turbine blade
CN109139127A (en) * 2018-09-17 2019-01-04 北京航空航天大学 A kind of pre- rotating gaseous film control structure of turbine guide vane
CN109595124B (en) * 2018-10-31 2020-08-14 华中科技大学 Dislocation negative pressure suction type wind gathering device
CN111578310A (en) * 2020-04-30 2020-08-25 南京理工大学 Air film cooling hole structure for turboshaft engine
CN114087027B (en) * 2021-11-23 2024-02-27 浙江燃创透平机械有限公司 Gas turbine stationary blade with honeycomb duct
CN115163559A (en) * 2022-06-24 2022-10-11 中国船舶重工集团公司第七0三研究所 Low-loss gas compressor transition section structure

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US4684323A (en) * 1985-12-23 1987-08-04 United Technologies Corporation Film cooling passages with curved corners
DE59808269D1 (en) * 1998-03-23 2003-06-12 Alstom Switzerland Ltd Film cooling hole
CN102151619B (en) * 2010-12-20 2012-06-27 北京航空航天大学 Porous wall supersonic cyclone separator and separation method thereof
JP2012202280A (en) * 2011-03-25 2012-10-22 Mitsubishi Heavy Ind Ltd Gas turbine cooling structure
CN203616135U (en) * 2013-09-24 2014-05-28 中国航天科技集团公司第六研究院第十一研究所 Jet nozzle

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