JP5738555B2 - Method and structure for cooling airfoil surfaces using asymmetric chevron film holes - Google Patents
Method and structure for cooling airfoil surfaces using asymmetric chevron film holes Download PDFInfo
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Description
本発明は、総括的にはフィルム冷却式部品に関し、より具体的には、非対称シェブロンフィルム孔を使用して翼形部表面上の一般的部位をフィルム冷却する方法に関する。 The present invention relates generally to film-cooled parts, and more specifically to a method for film cooling a general site on an airfoil surface using asymmetric chevron film holes.
ガスタービンその他の高温装置は、タービンブレードのような高温ガス通路構成要素の効果的な保護のためにフィルム冷却を広く使用している。フィルム冷却というのは、部品の外部壁内における複数の小孔を通して冷却空気を吐出して、該部品の外部表面に沿って薄い低温バリヤを形成しかつ高温ガスとの直接的な接触を防止するか又は減少させる部品を冷却する方法を意味する。 Gas turbines and other high temperature devices widely use film cooling for effective protection of hot gas path components such as turbine blades. Film cooling refers to the discharge of cooling air through a plurality of small holes in the exterior wall of the part, forming a thin cold barrier along the exterior surface of the part and preventing direct contact with the hot gas. It means a method of cooling a part to be reduced or reduced.
ベーン及びブレード翼形部をフィルム冷却するのに用いる一般的部位には、とりわけ前縁、正圧側面及び負圧側面が含まれると共に、内側及び外側ベーン端部壁並びにブレードプラットフォームを含む端部壁フィルム冷却が含まれる。タービン翼形部の端部壁領域におけるフィルム冷却は、該端部壁が翼形部正圧及び負圧側面表面の両方に見られる全範囲の静圧分布を受ける点で翼形部自体のフィルム冷却とは異なる。このような全圧力場により、翼形部表面では生じないような、噴射フィルム冷却に影響を与える大きな二次流れパターンが生じる。流路にわたるフィルム冷却の大きな移動が生じ、噴射及び有効な冷却を極めて困難にする。 Common sites used to film cool vanes and blade airfoils include, among other things, leading edges, pressure and suction sides, and end walls including inner and outer vane end walls and blade platforms. Film cooling is included. Film cooling in the end wall region of the turbine airfoil is the film of the airfoil itself in that the end wall is subject to the full range of static pressure distribution found on both the airfoil positive pressure and suction side surfaces. Different from cooling. Such a total pressure field results in a large secondary flow pattern that affects spray film cooling that does not occur on the airfoil surface. A large movement of film cooling across the flow path occurs, making jetting and effective cooling extremely difficult.
噴射フィルム孔は一般的に、円形形状又はディフューザ形状のいずれかである。これらの孔は、局所表面流線の近似方向に沿って噴射して混合損失を最少にするように配向される。これは、特定の領域においてフィルム冷却の集積を生じ、またその他の領域において関連するフィルム冷却の不足を引き起こすことが多い。 The spray film holes are generally either circular or diffuser shaped. These holes are oriented to jet along the approximate direction of local surface streamlines to minimize mixing loss. This often results in the accumulation of film cooling in certain areas and the associated lack of film cooling in other areas.
上記のことに鑑みて、保護することを意図した領域から離れる方向にフィルム冷却媒体を移動させることになる強力な横方向圧力勾配の存在下で表面上にフィルム冷却を噴射する構造及び方法を得ることは有益であると言える。この構造及び方法は、主高温ガスの流れに逆らって流れを単に噴射することで生じる不適当な混合損失を発生させずに、所望の領域内にフィルム冷却媒体を維持するものでなければならない。 In view of the above, a structure and method for injecting film cooling onto a surface in the presence of a strong lateral pressure gradient that would move the film cooling medium away from the area intended to be protected is obtained. It can be said that it is beneficial. This structure and method must maintain the film cooling medium in the desired area without inadequate mixing losses caused by simply injecting the flow against the main hot gas flow.
簡潔に言えば、１つの実施形態によると、フィルム冷却式翼形部又は翼形部領域は、１以上の非対称シェブロンフィルム冷却孔を有するように構成される。 Briefly, according to one embodiment, the film cooled airfoil or airfoil region is configured to have one or more asymmetric chevron film cooling holes.
別の実施形態では、フィルム冷却式タービン構造は、少なくとも１つの非対称シェブロンフィルム冷却孔を含み、シェブロンの一方の側部は、該フィルム冷却式タービン構造の表面上に噴射冷却媒体の一部分を導くことに関して該シェブロンの他方の側部よりも支配的（「優勢」ともいう。）である。
In another embodiment, the film cooled turbine structure includes at least one asymmetric chevron film cooling hole, and one side of the chevron directs a portion of the jet cooling medium onto the surface of the film cooled turbine structure. Is more dominant (also referred to as “dominant”) than the other side of the chevron.
本発明の上記その他の特徴、態様及び利点は、図面全体を通して同じ参照符号が同様な部品を表している添付図面を参照しながら以下の詳細な記述を読むことにより、一層良好に理解されるであろう。 These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like reference characters represent like parts throughout the drawings, wherein: I will.
上記の図面の図は、幾つかの代替的な実施形態を表しているが、以下の説明において述べるように本発明のその他の実施形態もまた考えられる。全てのケースにおいて、本開示は、本発明の例示した実施形態を限定としてではなく説明として提示している。当業者には、多くのその他の変更形態及び実施形態を案出することができるが、それらもまた、本発明の原理の技術的範囲及び技術思想の範囲内に属する。 While the above drawing figures represent several alternative embodiments, other embodiments of the invention are also contemplated, as will be described in the following description. In all cases, this disclosure presents illustrated embodiments of the present invention by way of illustration and not limitation. Many other modifications and embodiments can be devised by those skilled in the art, but they also fall within the scope and spirit of the principles of the present invention.
シェブロンフィルム孔は、翼形部表面上でのフィルム有効性を高めるのに有益であることが分かっている。現在のシェブロンフィルム孔は常に、フィルム孔中心線の周りでの対称設計に基づいている。 Chevron film holes have been found to be beneficial in increasing film effectiveness on the airfoil surface. Current chevron film holes are always based on a symmetrical design around the film hole centerline.
本明細書で前述したように、タービン翼形部の端部壁領域におけるフィルム冷却は、該端部壁が翼形部正圧及び負圧側面表面の両方に見られる全範囲の静圧分布を受ける点で翼形部自体のフィルム冷却とは異なる。このような全圧力場により、翼形部表面では生じないような、噴射フィルム冷却に影響を与える大きな二次流れパターンが生じる。しかしながら、翼形部表面は、翼形部が端部壁領域と交わる領域を除いて、全体的に非常により少ない程度までではあるがそのような二次流れ作用を受ける。流路にわたるフィルム冷却の大きな移動が生じ、噴射及び有効な冷却を極めて困難にする。 As previously described herein, film cooling in the end wall region of the turbine airfoil results in a full range of static pressure distributions where the end wall is found on both the airfoil positive and suction side surfaces. It differs from film cooling of the airfoil itself in that it receives. Such a total pressure field results in a large secondary flow pattern that affects spray film cooling that does not occur on the airfoil surface. However, the airfoil surface is subject to such secondary flow effects to a much lesser extent overall, except in the region where the airfoil intersects the end wall region. A large movement of film cooling across the flow path occurs, making jetting and effective cooling extremely difficult.
噴射フィルム孔は一般的に、円形形状又はディフューザ形状のいずれかである。これらの孔は通常、局所表面流線の近似方向に沿って噴射して混合損失を最少にするように配向される。これは、特定の領域においてフィルム冷却の集積を生じ、またその他の領域において関連するフィルム冷却の不足を引き起こすことが多い。 The spray film holes are generally either circular or diffuser shaped. These holes are usually oriented so that they are jetted along the approximate direction of local surface streamlines to minimize mixing losses. This often results in the accumulation of film cooling in certain areas and the associated lack of film cooling in other areas.
本明細書では、表面流体流線湾曲（曲率）が大きい領域及び用途において同様の流体流れの利点を達成する非対称シェブロンフィルム孔の実施形態について説明する。これらの実施形態は、依然として単一の円形貫通孔を使用しているが、シェブロンフットプリントの２つの半部分をトラフの異なる寸法又は配向を有するように変更している。この非対称性は、冷却しようとする表面上に噴射冷却媒体の一部分を導くことに関して、シェブロンの一方の側部を他方側よりも支配的にする利点がある。シェブロンの支配的又は主要側部は、高温ガスによって与えられる流線湾曲を打消すように方向付ける／配向する必要がある。 Described herein are embodiments of asymmetric chevron film holes that achieve similar fluid flow benefits in areas and applications where the surface fluid streamline curvature (curvature) is large. These embodiments still use a single circular through hole, but modify the two halves of the chevron footprint to have different trough dimensions or orientations. This asymmetry has the advantage of making one side of the chevron more dominant than the other with respect to directing a portion of the jet cooling medium onto the surface to be cooled. The dominant or main side of the chevron needs to be oriented / oriented to counteract the streamline curvature imparted by the hot gas.
本明細書では、最初に図１〜図３を参照して対称シェブロンフィルム冷却孔の説明を行なって、その後に説明する非対称フィルム冷却孔の原理及び実施形態の良好な理解が得られるようにしている。図１は、当技術分野で公知の対称シェブロンフィルム冷却孔を示す上面図である。リッジ部１２は、２つのウィングトラフ１４間でその深さが横方向に外向き凸面形である。凸面形リッジ部１２は、アーチ形でありかつその輪郭がほぼ三角形であり、また入口ボア１６と該凸面形リッジ部の下流端部の外表面１８との接合部との間で下流方向に発散している。リッジ部１２の後端縁は、シェブロン出口の横方向アーチ形下流端部に沿って外表面１８と同一平面になるように滑らかに連続しており、凸面形後端縁は、入口孔１６に向かって上流方向に弓形になっている。複合シェブロンフィルム冷却孔１０の湾曲形態は、高温ガス２０の流れに対してシェブロンフィルム冷却孔１０の側面図を示す図２にさらに詳細に示すような複合傾斜角度Ａ，Ｂの利点をもたらし、ここでは、シェブロン出口は、入口孔１６から後方に傾斜角度Ａで傾斜して異なった状態に発散している。より具体的には、傾斜角度Ａ及びＢは、２つの限定角度、すなわち一方が中心線に沿った角度であり、また他方が各トラフ内における角度である。 In this specification, the symmetrical chevron film cooling holes are first described with reference to FIGS. 1-3 so that a better understanding of the principles and embodiments of the asymmetric film cooling holes described below can be obtained. Yes. FIG. 1 is a top view of a symmetrical chevron film cooling hole known in the art. The ridge portion 12 is convex outwardly in the lateral direction between two wing troughs 14. The convex ridge 12 is arcuate and has a substantially triangular outline, and diverges in the downstream direction between the inlet bore 16 and the junction of the outer surface 18 at the downstream end of the convex ridge. doing. The rear edge of the ridge 12 is smoothly continuous to be flush with the outer surface 18 along the lateral arcuate downstream end of the chevron outlet, and the convex rear edge is in the inlet hole 16. It is bowed in the upstream direction. The curved form of the composite chevron film cooling holes 10 provides the advantage of composite tilt angles A and B as shown in more detail in FIG. 2 which shows a side view of the chevron film cooling holes 10 for the flow of hot gas 20. In this case, the chevron outlet diverges in a different state by being inclined at an inclination angle A rearward from the inlet hole 16. More specifically, the tilt angles A and B are two limiting angles, one angle along the center line and the other angle within each trough.
図３は、図１に示す対称フィルム冷却孔１０の正面図である。この正面図は、図２に示す高温ガス２０の方向におけるものである。シェブロンフィルム孔１０は、図１に示すフィルム孔１０中心線の周りでの対称設計に基づいており、かつリッジ部１２及びウィングトラフ１４の更なる詳細を示している。 FIG. 3 is a front view of the symmetric film cooling hole 10 shown in FIG. This front view is in the direction of the hot gas 20 shown in FIG. The chevron film hole 10 is based on a symmetrical design about the film hole 10 centerline shown in FIG. 1 and shows further details of the ridge 12 and wing trough 14.
図４は、タービンベーン２４の端部壁領域２２内に配置されたフィルム冷却孔を示す斜視図であり、一方、図５は、タービンブレード２８の端部壁領域内に配置されたフィルム冷却孔２６を示す斜視図である。上述のように、タービン翼形部の端部壁領域におけるフィルム冷却は、本明細書で前述したように、該端部壁が翼形部正圧及び負圧側面表面の両方に見られる全範囲の静圧分布を受ける点で翼形部自体のフィルム冷却とは異なる。しかしながら、翼形部表面は、これもまた上述のように、翼形部が端部壁領域と交わる領域を除いて、全体的に非常により少ない程度までではあるがそのような二次流れ作用を受ける。このような全圧力場により、翼形部表面では生じないような、噴射フィルム冷却に影響を与える大きな二次流れパターンが生じて、流路にわたるフィルム冷却の大きな移動を引き起こし、噴射及び有効な冷却を極めて困難にする。 FIG. 4 is a perspective view showing film cooling holes disposed in the end wall region 22 of the turbine vane 24, while FIG. 5 shows film cooling holes disposed in the end wall region of the turbine blade 28. FIG. As noted above, film cooling in the end wall region of the turbine airfoil is the full range where the end wall is found on both the airfoil positive and suction side surfaces, as previously described herein. It differs from film cooling of the airfoil itself in that it receives a static pressure distribution. However, the airfoil surface also has such secondary flow effects to a much lesser extent overall, except where the airfoil intersects the end wall region, also as described above. receive. Such a total pressure field results in a large secondary flow pattern that affects spray film cooling, which does not occur on the airfoil surface, causing a large movement of film cooling across the flow path, spraying and effective cooling. Makes it extremely difficult.
図６は、タービンベーン２９の端部壁領域における局所表面流線の近似方向に沿って噴射するように配向された円形噴射フィルム孔を示している。図６には、円形噴射フィルム孔を示しているが、噴射フィルム孔は一般的に、円形形状又はディフューザ形状のいずれかである。これらの孔は、局所表面流線の近似方向に沿って噴射して混合損失を最少にするように配向される。これは、上述したように、特定の領域においてフィルム冷却の集積を生じ、またその他の領域において関連するフィルム冷却の不足を引き起こすことが多い。 FIG. 6 shows circular spray film holes oriented to spray along the approximate direction of local surface streamlines in the end wall region of the turbine vane 29. Although FIG. 6 shows circular spray film holes, the spray film holes are generally either circular or diffuser shaped. These holes are oriented to jet along the approximate direction of local surface streamlines to minimize mixing loss. This, as mentioned above, often results in film cooling accumulation in certain areas and the associated lack of film cooling in other areas.
図７は、本発明の１つの実施形態による、図４及び図５に示す端部壁領域内で使用するのに好適な非対称シェブロンフィルム孔３０を示す上面図である。平坦なリッジ部３２が、２つのウィングトラフ３４、３６間でその幅が横方向に増大している。リッジ部３２は、平坦でありかつその輪郭がほぼ三角形であり、また入口ボア３８と該リッジ部の下流端部の外表面４０との接合部との間で下流方向に発散している。リッジ部３２の後端縁は、シェブロン出口の横方向平坦下流端部に沿って外表面４０と同一平面になるように滑らかに連続している。ウィングトラフ３４の寸法は、ウィングトラフ３６の寸法とは異なっていて、ウィングトラフ３４、３６が各々、互いに対して異なった状態で入口ボア３８の周囲部分内に滑らかに連続する。 FIG. 7 is a top view of an asymmetric chevron film hole 30 suitable for use in the end wall region shown in FIGS. 4 and 5 according to one embodiment of the present invention. The flat ridge 32 has a width that increases laterally between the two wing troughs 34, 36. The ridge portion 32 is flat and has a substantially triangular outline, and diverges in the downstream direction between the inlet bore 38 and the junction between the outer surface 40 at the downstream end of the ridge portion. The rear edge of the ridge 32 is smoothly continuous along the laterally flat downstream end of the chevron outlet so as to be flush with the outer surface 40. The dimensions of the wing trough 34 are different from the dimensions of the wing trough 36, and the wing troughs 34, 36 each continue smoothly into the peripheral portion of the inlet bore 38 in a different state relative to each other.
シェブロン出口が入口ボア３８から後方に発散するような複合傾斜角度を有する非対称シェブロンフィルム冷却孔３０（図示せず）の湾曲形態は、対称シェブロンフィルム冷却孔１０について図２に示した複合傾斜角度Ｂ、Ｃにおける上述した利点と同様の利点をもたらすことになる。 The curved form of the asymmetric chevron film cooling holes 30 (not shown) having a compound inclination angle such that the chevron outlet diverges back from the inlet bore 38 is the compound inclination angle B shown in FIG. , C will provide the same advantages as described above.
非対称シェブロンフィルム冷却孔３０は、特定の非対称フィルム冷却孔３０の実施形態が、図８に示すような異なるトラフ深さ、異なるトラフ幅、異なるトラフ拡散角度、異なるトラフ形状、等々を含むことができる点で、対称フィルム冷却孔１０とは大幅に異なっている。図８は、そこでは各シェブロン領域が、互いに対して同様でないジオメトリを有する一対のウィングトラフを含む２つの非対称シェブロン領域を示している。 Asymmetric chevron film cooling holes 30 may include specific asymmetric film cooling hole 30 embodiments having different trough depths, different trough widths, different trough diffusion angles, different trough shapes, etc. as shown in FIG. In this respect, it is significantly different from the symmetrical film cooling hole 10. FIG. 8 shows two asymmetric chevron regions, where each chevron region includes a pair of wing troughs that have dissimilar geometries relative to each other.
１つの実施形態によると、例えば、ウィングトラフ３４は、ウィングトラフ３６の深さとは異なる深さを有する。別の実施形態によると、ウィングトラフ３４は、ウィングトラフ３６の幅とは異なる幅を有する。さらに別の実施形態によると、ウィングトラフ３４は、拡散角度Ｃとは異なる拡散角度Ｂを有する。なおさらに別の実施形態によると、ウィングトラフ３４は、ウィングトラフ３６の形状とは異なる形状を有する。 According to one embodiment, for example, the wing trough 34 has a depth that is different from the depth of the wing trough 36. According to another embodiment, the wing trough 34 has a width that is different from the width of the wing trough 36. According to yet another embodiment, the wing trough 34 has a diffusion angle B that is different from the diffusion angle C. According to yet another embodiment, the wing trough 34 has a shape that is different from the shape of the wing trough 36.
ウィングトラフ３４、３６間の前述の非対称性は次に、対称シェブロンの場合と同様な過渡領域となり、特定の実施形態では、その形状が平坦であるか又は多平面であるか又は非対称アーチ形であるかの状態になる。１つの実施形態によると、非対称シェブロンフィルム冷却孔３０は、シェブロン領域に冷却媒体を供給する単一の円形貫通孔を用いている。 The aforementioned asymmetry between the wing troughs 34, 36 then becomes a transition region similar to that of a symmetric chevron, and in certain embodiments, its shape is flat, multi-planar or asymmetric arched. It becomes a certain state. According to one embodiment, the asymmetric chevron film cooling hole 30 uses a single circular through hole that supplies a cooling medium to the chevron region.
要約して説明すると、本明細書では、特に表面流体流線湾曲が大きい領域及び用途において多様な翼形部表面に対するフィルム冷却を高める非対称シェブロンフィルム冷却孔について説明している。この非対称フィルム冷却孔の使用は、保護することを意図した領域から離れる方向にフィルム冷却媒体を移動させることになる強力な横方向圧力勾配の存在下での表面上へのより効率的なフィルム冷却の噴射を可能にし、かつ不適当な混合損失を発生させずに所望の１つ又は複数の領域内にフィルム冷却媒体を維持する。単に主高温ガスの流れに逆らって流れを噴射して圧力勾配を打消そうとしている従来型のフィルム孔では、一層大きな混合損失が生じる。より効率的な冷却媒体の使用により、より長寿命を有する産業用エンジンのようなより高効率のエンジンが得られる。 In summary, this specification describes asymmetric chevron film cooling holes that enhance film cooling for a variety of airfoil surfaces, particularly in regions and applications where the surface fluid streamline curvature is high. The use of this asymmetric film cooling hole allows more efficient film cooling onto the surface in the presence of a strong lateral pressure gradient that will move the film cooling medium away from the area intended to be protected. The film cooling medium in the desired region or regions without inadequate mixing loss. In conventional film holes that simply jet the flow against the main hot gas flow to try to counteract the pressure gradient, even greater mixing losses occur. The use of a more efficient cooling medium results in a more efficient engine such as an industrial engine with a longer life.
非対称シェブロンフィルム冷却孔は、冷却の適切さ及び損失の妥協が見出されるまで試行錯誤する公知のフィルム孔の配置を使用して達成可能である利点に優る、或いは円形フィルム孔に対して単にディフューザの形状を付加することによってまた場合によっては所望の方向に冷却媒体を導くのに役立つ複合角度を該ディフューザ上に付加することによって達成可能である利点に優る利点をもたらす。非対称シェブロンフィルム冷却孔はさらに、フィルム孔を変更するのではなくて二次流れ及び圧力勾配を軽減するのに役立つように単に端部壁自体の形状を変更することによって達成可能である利点に優る利点をもたらす。 Asymmetric chevron film cooling holes outperform the advantages that can be achieved using known film hole arrangements that are trial and error until a compromise of proper cooling and loss is found, or simply the diffuser's relative to circular film holes. Adding a shape and possibly providing an advantage over that which can be achieved by adding a compound angle on the diffuser that helps guide the cooling medium in the desired direction. Asymmetric chevron film cooling holes further outweigh the advantages that can be achieved by simply changing the shape of the end walls themselves to help mitigate secondary flow and pressure gradients rather than changing the film holes. Bring benefits.
本明細書では、本発明の一部の特徴のみを例示しかつ説明してきたが、当業者には多くの修正及び変更が想起されるであろう。従って、特許請求の範囲は、全てのそのような修正及び変更を本発明の技術思想の範囲内に属するものとして保護することを意図していることを理解されたい。 Although only certain features of the invention have been illustrated and described herein, many modifications and changes will occur to those skilled in the art. Accordingly, it is to be understood that the appended claims are intended to protect all such modifications and changes as fall within the scope of the spirit of the invention.
１０ 対称シェブロンフィルム冷却孔
１２ 対称シェブロンフィルム冷却孔リッジ部
１４ 対称シェブロンフィルム冷却孔ウィングトラフ
１６ 対称シェブロンフィルム冷却孔入口ボア
１８ フィルム冷却式部品の外表面
２０ 高温ガス流
２２ タービンベーンの端部壁領域
２４ タービンベーン
２６ フィルム冷却孔
２８ タービンブレード
２９ タービンベーンの端部壁領域
３０ 非対称シェブロンフィルム孔
３２ 非対称シェブロンフィルム孔平坦リッジ部
３４ 非対称シェブロンフィルム孔ウィングトラフ
３６ 非対称シェブロンフィルム孔ウィングトラフ
３８ 非対称シェブロンフィルム孔入口ボア
４０ フィルム冷却式部品の外表面
５０ 一対の非対称フィルム孔領域
10 Symmetric Chevron Film Cooling Hole 12 Symmetric Chevron Film Cooling Hole Ridge 14 Symmetric Chevron Film Cooling Hole Wing Trough 16 Symmetric Chevron Film Cooling Hole Inlet Bore 18 Film Cooling Part Outer Surface 20 Hot Gas Flow 22 Turbine Vane End Wall Region 24 Turbine vane 26 Film cooling hole 28 Turbine blade 29 Turbine vane end wall region 30 Asymmetric chevron film hole 32 Asymmetric chevron film hole flat ridge 34 Asymmetric chevron film hole wing trough 36 Asymmetric chevron film hole wing trough 38 Asymmetric chevron film hole Inlet bore 40 Outer surface 50 of film cooled part Pair of asymmetric film hole area
Claims (9)
少なくとも１つの非対称シェブロンフィルム冷却孔（３０）であって、各シェブロン（３０）の一方の側部（３４）が、当該フィルム冷却式タービン構造（４０）の表面上に噴射冷却媒体の一部分を導くことに関して該シェブロン（３０）の他方の側部（３６）よりも優勢である、少なくとも１つの非対称シェブロンフィルム冷却孔（３０）を含んでおり、各シェブロン（３０）の優勢な側部が、当該フィルム冷却式タービン構造（４０）の表面上を流れる高温ガスによって生じる流線湾曲を打消すように配向している、フィルム冷却式タービン構造（４０）。 A film-cooled turbine structure (40) comprising:
At least one asymmetric chevron film cooling hole (30), wherein one side (34) of each chevron (30) directs a portion of the jet cooling medium onto the surface of the film cooled turbine structure (40). Including at least one asymmetric chevron film cooling hole (30) that is superior to the other side (36) of the chevron (30), wherein the dominant side of each chevron (30) A film cooled turbine structure (40) oriented to counteract streamline curvature caused by hot gases flowing over the surface of the film cooled turbine structure (40).
The at least one asymmetric chevron film cooling hole (30) includes a first trough region (34) and a second trough region (36), wherein a first trough region (34) and a second trough region ( The film-cooled turbine structure (40) according to any one of claims 1 to 8, wherein 36) have different geometries.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US12/607,586 | 2009-10-28 | ||
| US12/607,586 US20110097191A1 (en) | 2009-10-28 | 2009-10-28 | Method and structure for cooling airfoil surfaces using asymmetric chevron film holes |
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| Publication Number | Publication Date |
|---|---|
| JP2011094609A JP2011094609A (en) | 2011-05-12 |
| JP5738555B2 true JP5738555B2 (en) | 2015-06-24 |
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| JP2010184516A Expired - Fee Related JP5738555B2 (en) | 2009-10-28 | 2010-08-20 | Method and structure for cooling airfoil surfaces using asymmetric chevron film holes |
Country Status (5)
| Country | Link |
|---|---|
| US (1) | US20110097191A1 (en) |
| JP (1) | JP5738555B2 (en) |
| CN (1) | CN102052092B (en) |
| CH (1) | CH702107B1 (en) |
| DE (1) | DE102010037050A1 (en) |
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-
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- 2010-08-20 JP JP2010184516A patent/JP5738555B2/en not_active Expired - Fee Related
- 2010-08-23 CH CH01344/10A patent/CH702107B1/en not_active IP Right Cessation
- 2010-08-27 CN CN201010272834.XA patent/CN102052092B/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| US20110097191A1 (en) | 2011-04-28 |
| CN102052092A (en) | 2011-05-11 |
| CH702107A2 (en) | 2011-04-29 |
| CH702107B1 (en) | 2015-07-15 |
| CN102052092B (en) | 2016-01-20 |
| JP2011094609A (en) | 2011-05-12 |
| DE102010037050A1 (en) | 2011-05-05 |
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