WO2005068782A1 - Stabilisation de la zone de transition laminaire/turbulente sur une pale de turbine - Google Patents

Stabilisation de la zone de transition laminaire/turbulente sur une pale de turbine Download PDF

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Publication number
WO2005068782A1
WO2005068782A1 PCT/US2004/009897 US2004009897W WO2005068782A1 WO 2005068782 A1 WO2005068782 A1 WO 2005068782A1 US 2004009897 W US2004009897 W US 2004009897W WO 2005068782 A1 WO2005068782 A1 WO 2005068782A1
Authority
WO
WIPO (PCT)
Prior art keywords
turbine fan
fan blade
recited
boundary layer
blade
Prior art date
Application number
PCT/US2004/009897
Other languages
English (en)
Inventor
Preston Henne
Robert Mills
Original Assignee
Gulfstream Aerospace Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Gulfstream Aerospace Corporation filed Critical Gulfstream Aerospace Corporation
Priority to US10/710,723 priority Critical patent/US7878759B2/en
Priority to EP04821310A priority patent/EP1711688A4/fr
Priority to PCT/US2004/042885 priority patent/WO2005074469A2/fr
Publication of WO2005068782A1 publication Critical patent/WO2005068782A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • 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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/31—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor with roughened surfaces
    • 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

Definitions

  • the present invention relates to Turbo Fan Engine Design and Modification and Fan
  • Blade Aerodynamics and more specifically to systems and methods for Improving Operating
  • the invention is easy and cost effective to implement; the invention(s) is an extremely effective and efficient method of reducing fan disc stresses; the adaptations and/or devices are passive in nature, and once installed, require substantially no attention or monitoring.
  • Figures 1 and 2 illustrate an untreated turbofan fan blade and Figures 3-14 illustrate various embodiments of the presently disclosed invention(s) in which boundary layer control is exemplarily exerted over aerodynamic induced shock forces, which otherwise may also resonate.
  • boundary layer control is exemplarily exerted over aerodynamic induced shock forces, which otherwise may also resonate.
  • boundary layer control devices attached or integrated with the fan blade geometry to either fix the point of transition between laminar and turbulent flow, or to control flow separation over the fan blade.
  • This invention is the application of proven boundary layer control methods for the containment of flow disturbances on fan blades for turbo fan engines.
  • the purpose of this invention is the stabilization of the flow field adjacent to the inlet fan blade and the subsequent reduction of the structural stresses in the fan disc, together with any associated improvements in the operating efficiency of the engine, in particular specific fuel consumption.
  • the boundary layer devices which apply in the development of this invention, comprise 1.
  • grit placement those that are mounted externally on the fan blade, and include but are not limited to grit placement, vortex generators, beads, turbulators, trip strips, or2.any geometry or surface treatment that is used to generate a recessed shape in the air passage surface of the fan blade that is intended to stabilize the boundary layer adjacent to the fan blade surface either by transitioning it from laminar to turbulent flow, or by reattaching the separated boundary layer.
  • Figure 1 A typical implementation of this invention is shown in Figure 1.
  • grit has been located on the high-speed side of an engine fan blade.
  • the grit strip is attached to the leading edge of the fan blade and extends from the tip of the blade to approximately half the span. In this application the grit was bonded to the blade using a high strength epoxy.
  • the invention takes the form of a turbine fan blade 40 adapted to trip and control a boundary layer transition at a side surface of the blade 40 during operation as a component in a turbine fan assembly 35.
  • the turbine fan blade 40 includes a leading edge 55, a trailing edge 58, and two side surfaces including a high-pressure side surface 49 and a low- pressure side surface 52.
  • At least one of the two side surfaces has an essentially smooth surface portion 61 located between the leading and trailing edges, and the essentially smooth surface portion is interrupted by a surface deviation 64.
  • the surface deviation is configured to trip a positionally stabilized boundary layer transition 24 at a location toward the trailing edge from the surface deviation during operation of the turbine fan blade in the turbine fan assembly. In this manner, fatigue inducing aerodynamic forces experienced upon the blade during operation are controlled.
  • the surface deviation 64 exemplifies a flutter suppression feature, or in functional terms, a flutter suppression means 90 configured for tripping a positionally stabilized boundary layer transition at a location toward the trailing edge from the flutter suppression means during operation of the turbine fan blade in a turbine fan assembly thereby controlling fatigue inducing aerodynamic forces 25 experienced upon the blade during operation.
  • a turbine fan blade assembly also constitutes an invention.
  • an invention is also constituted by a turbine 32
  • the invention takes the form of a turbine-powered aircraft 20 having an aircraft body to which at least one turbine fan based power unit 30 is mounted.
  • the invention may be alternatively described as a turbine fan blade adapted to trip and control a boundary layer transition at a side surface of the blade during operation as a component in a turbine fan assembly.
  • the turbine fan blade is produced by a method that includes obtaining a turbine fan blade having a leading edge, a trailing edge, and two side surfaces including a high- pressure side surface and a low-pressure side surface.
  • a flutter suppression means is provided on the turbine fan blade that trips a positionally stabilized boundary layer transition at a location toward the trailing edge from the flutter suppression means during operation of the turbine fan blade in a turbine fan assembly thereby controlling fatigue inducing aerodynamic forces experienced upon the blade during operation.
  • a turbine fan blade adapted to reattach a separated boundary layer thereby controlling a boundary layer transition at a side surface of the blade during operation as a component in a turbine fan assembly is also disclosed.
  • the turbine fan blade being produced by a method including: obtaining a turbine fan blade having a leading edge, a trailing edge, and two side surfaces including a high-pressure side surface and a low-pressure side surface; and providing a separated boundary layer reattachment device on the turbine fan blade that reattaches a separated boundary layer to the turbine fan blade thereby controlling fatigue inducing aerodynamic forces experienced upon the blade during operation.
  • the separated boundary layer reattachment device may take the form of a vortex generator or a turbulator.
  • the invention is also constituted by a method for providing a turbine fan blade configured to trip and control a boundary layer transition at a side surface of the blade during operation as a component in a turbine fan assembly.
  • the method includes providing a turbine fan blade having a leading edge, a trailing edge, and two side surfaces including a high-pressure side surface and a low-pressure side surface; and providing a flutter suppression means on the turbine fan blade that trips a positionally stabilized boundary layer transition at a location toward the trailing edge from the flutter suppression feature during operation of the turbine fan blade in a turbine fan assembly thereby controlling fatigue inducing aerodynamic forces experienced upon the blade during operation.
  • the invention takes the form of amethod for retrofitting an earlier manufactured turbine fan blade with an adaptation configured to trip and control a boundary layer transition at a side surface of the turbine fan blade during operation as a component in a turbine fan assembly.
  • This method includes obtaining an earlier manufactured turbine fan blade having a leading edge, a trailing edge, and two side surfaces including a high- pressure side surface and a low-pressure side surface; and providing a flutter suppression feature on the turbine fan blade that trips a positionally stabilized boundary layer transition at a location toward the trailing edge from the flutter suppression feature during operation of the turbine fan blade in a turbine fan assembly thereby controlling fatigue inducing aerodynamic forces experienced upon the blade during operation.
  • the invention takes the form of a method for producing a turbine fan blade configured to trip and control a boundary layer transition at a side surface of the turbine fan blade during operation as a component in a turbine fan assembly, the method including adapting a turbine fan blade having a leading edge, a trailing edge, and two side surfaces including a high-pressure side surface and a low-pressure side surface with a flutter suppression feature in association with at least one of the side surfaces, the flutter suppression feature configured to trip a positionally stabilized boundary layer transition at a location toward the trailing edge from the flutter suppression feature during operation of the turbine fan blade in a turbine fan assembly thereby controlling fatigue inducing aerodynamic forces experienced upon the blade during operation.
  • Another embodiment is a method for achieving flutter suppression in a turbine fan based power unit of a flying jet aircraft and thereby controlling operationally induced stress levels to within an acceptable degradation range.
  • the method includes utilizing, on a flying jet aircraft, a turbine fan based power unit including a rotating fan blade assembly including multiple fan blades, mounted on a hub, and each of the fan blades having a leading edge, a trailing edge, and two side surfaces that include a high-pressure side surface and a low-pressure side surface; the jet aircraft flying under conditions that cause at least laminar and turbulent boundary layers to form at at least one of the side surfaces of each of the blades with a boundary layer transition occurring therebetween; and a plurality of the turbine fan blades each including a boundary layer transition trip feature at the side surface at which the boundary layer transition occurs which trips a positionally stabilized boundary layer transition at a location toward the trailing edge of the respective blade from the trip feature during flight-operation of the turbine fan blade assembly and thereby controlling operationally induced stress levels within the hub to
  • the flutter suppression means is located on the low- pressure side surface and/or the high-pressure side surface.
  • the flutter suppression means can be located on at least one of the two side surfaces and is at least partially constituted by a rough surface portion adjacently located to an essentially smooth surface portion positioned on a trailing edge side of the rough surface portion.
  • the rough surface portion includes an area of the at least one of the two side surfaces having grit adhered thereto.
  • the flutter suppression means is located on at least one of the two side surfaces having an essentially smooth surface portion located between the leading and trailing edges and the flutter suppression means being characterized by a surface deviation constituting a departure from the essentially smooth surface portion.
  • the departure from the essentially smooth surface portion can be constituted by a reduced-elevation surficial portion, compared to the essentially smooth surface portion and/or can be constituted by a raised-elevation surficial portion, compared to the essentially smooth surface portion.
  • the raised-elevation surficial portion being provided by applying an adhesive to the at least one of the two side surfaces, a thickness of the adhesive constituting a thickness of the raised-elevation surficial portion.
  • the raised-elevation surficial portion is provided by applying grit fixed to the at least one of the two side surfaces using an adhesive, the grit establishing peak elevations of the raised-elevation surficial portion.
  • the raised-elevation surficial portion is provided by plasma spray upon the at least one of the two side surfaces.
  • the departure from the essentially smooth surface portion is approximately six inches long and is positioned approximately five percent of a chord length of the respective turbine fan blade toward the trailing edge from the leading edge.
  • the departure from the essentially smooth surface portion is an elongate strip-shaped area of raised elevation having a length less than three inches.
  • the width is preferably less than one-half inch, more preferably less than one-hundredth of an inch; and most preferably, approximately one one-thousandth of an inch.
  • the departure from the essentially smooth surface portion is an elongate strip-shaped area of raised elevation having a tip-end distanced approximately one-half inch from a tip of the turbine fan blade.
  • the distance is at least one-half inch from a tip of the turbine fan blade.
  • the departure from the essentially smooth surface portion is an elongate strip-shaped area of raised elevation having a length of approximately two and one-half inches.
  • the elongate strip-shaped area of raised elevation is positioned approximately one-third of a chord length of the respective turbine fan blade toward the trailing edge from the leading edge. It is preferred that the elongate strip-shaped area of raised elevation is positioned no more than one-third of a chord length of the respective turbine fan blade toward

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

La présente invention se rapporte à une pale de turbine (40) conçue pour déclencher et réguler une transition de couche limite au niveau d'une surface latérale de la pale (40) lors du fonctionnement en tant que composant dans un ensemble turbine (35). La pale de turbine (40) comprend un bord avant (55), un bord arrière (58) et deux surfaces latérales incluant une surface latérale haute pression (49) et une surface latérale basse pression (52). L'une au moins des deux surfaces latérales possède une partie superficielle sensiblement lisse (61) disposée entre le bord avant et le bord arrière, et cette partie superficielle sensiblement lisse est interrompue par une déviation de surface (64). Cette déviation de surface est conçue pour déclencher une transition de couche limite stabilisée en position (24) au niveau d'un emplacement en direction du bord avant à partir de la déviation superficielle en cours de fonctionnement de la pale de turbine dans l'ensemble turbine. De cette manière, il est possible de réguler les forces aérodynamiques induisant une fatigue qui s'appliquent sur la pale pendant le fonctionnement.
PCT/US2004/009897 2003-12-20 2004-03-26 Stabilisation de la zone de transition laminaire/turbulente sur une pale de turbine WO2005068782A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US10/710,723 US7878759B2 (en) 2003-12-20 2004-07-30 Mitigation of unsteady peak fan blade and disc stresses in turbofan engines through the use of flow control devices to stabilize boundary layer characteristics
EP04821310A EP1711688A4 (fr) 2003-12-20 2004-12-20 Attenuation des contraintes sur les pales et disques de soufflante a crete instable dans des turboreacteurs double flux au moyen de dispositifs de regulation de debit pour stabiliser les caracteristiques de couches limites
PCT/US2004/042885 WO2005074469A2 (fr) 2003-12-20 2004-12-20 Attenuation des contraintes sur les pales et disques de soufflante a crete instable dans des turboreacteurs double flux au moyen de dispositifs de regulation de debit pour stabiliser les caracteristiques de couches limites

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
US48182003P 2003-12-20 2003-12-20
US60/481,820 2003-12-20
US48188904P 2004-01-13 2004-01-13
US60/481,889 2004-01-13

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US10/710,723 Continuation US7878759B2 (en) 2003-12-20 2004-07-30 Mitigation of unsteady peak fan blade and disc stresses in turbofan engines through the use of flow control devices to stabilize boundary layer characteristics

Publications (1)

Publication Number Publication Date
WO2005068782A1 true WO2005068782A1 (fr) 2005-07-28

Family

ID=34798700

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2004/009897 WO2005068782A1 (fr) 2003-12-20 2004-03-26 Stabilisation de la zone de transition laminaire/turbulente sur une pale de turbine

Country Status (1)

Country Link
WO (1) WO2005068782A1 (fr)

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5169290A (en) * 1991-11-07 1992-12-08 Carrier Corporation Blade for centrifugal flow fan

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5169290A (en) * 1991-11-07 1992-12-08 Carrier Corporation Blade for centrifugal flow fan

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