EP2441919A1 - Aube chargée à l'arrière - Google Patents

Aube chargée à l'arrière Download PDF

Info

Publication number
EP2441919A1
EP2441919A1 EP11181843A EP11181843A EP2441919A1 EP 2441919 A1 EP2441919 A1 EP 2441919A1 EP 11181843 A EP11181843 A EP 11181843A EP 11181843 A EP11181843 A EP 11181843A EP 2441919 A1 EP2441919 A1 EP 2441919A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
turbomachine
fan
array
distance
Prior art date
Legal status (The legal status 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 status listed.)
Withdrawn
Application number
EP11181843A
Other languages
German (de)
English (en)
Inventor
Matthew B. Estes
Renee J. Jurek
Shankar S. Magge
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MTU Aero Engines AG
RTX Corp
Original Assignee
United Technologies Corp
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 United Technologies Corp filed Critical United Technologies Corp
Publication of EP2441919A1 publication Critical patent/EP2441919A1/fr
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • 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
    • 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

Definitions

  • This disclosure relates generally to a gas turbine engine and, more particularly, to loading an aft end of a low pressure turbine airfoil.
  • gas turbine engines include multiple sections, such as a fan section, a compression section, a combustor section, a turbine section, and an exhaust nozzle section.
  • the compression section and the turbine section include airfoil arrays mounted for rotation about an engine axis.
  • the airfoil arrays include multiple individual airfoils (or blades) that extend radially from a mounting platform to a tip.
  • Air moves into the engine though the fan section. Rotating the combustion section's airfoil arrays compresses the air. The compressed air is then mixed with fuel and combusted in the combustor section. The products of combustion expand to rotatably drive the airfoil arrays in the turbine section. Rotating the airfoil arrays in the turbine section drives rotation of the fan section.
  • the turbine section often includes a low pressure turbine and a high pressure turbine.
  • Each airfoil within the low pressure turbine's airfoil array is designed with the goal to position the laminar to turbulent boundary layer transition to reduce the likelihood of boundary layer separation and severe performance degradation. This typically results in a design with more forward loading, moving the transition location forward and a reducing laminar flow length.
  • reducing the length of the laminar flow increases the length of the turbulent boundary layer.
  • the turbulent boundary layer performs more poorly than laminar flow provided the boundary layer has not separated.
  • An example airfoil array includes an airfoil array configured to rotate within a turbomachine at a different rotational speed than a fan within the turbomachine.
  • the airfoils in the array are arranged such that the pitch-to-chord ratio measured at midspan is less than 1.3, wherein said pitch is a distance between the trailing edges of adjacent airfoils and said chord is the axial chord length of said airfoils.
  • An uncovered turning angle of the airfoil is greater than 15 degrees.
  • An example turbomachine arrangement includes a fan rotatable about an axis, and a turbine section that rotates at a first speed.
  • the turbine section is configured to rotatably drive the fan at second, different speed.
  • the turbine section includes an airfoil array having a first airfoil that is circumferentially spaced a pitch from an adjacent second airfoil. A ratio of the pitch to an axial chord length of the first airfoil is less than 1.3. An uncovered turning angle of the first airfoil is greater than 15 degrees.
  • An example method of establishing a transition location for a turbomachine airfoil includes positioning a first airfoil relative to a second airfoil within a turbomachine such that a ratio of a circumferential distance between the first airfoil and the second airfoil to an axial chord length of the first airfoil is less than 1.3.
  • the first airfoil has an uncovered turning angle that is greater than 15 degrees and is configured to be rotated about an axis of the turbomachine at a different rotational speed than a fan of the turbomachine.
  • Figure 1 schematically illustrates an example geared gas turbine engine 10 including (in serial flow communication) a fan 14, a low pressure compressor 18, a high pressure compressor 22, a combustor 26, a high pressure turbine 30, and a low pressure turbine 34.
  • the gas turbine engine 10 is circumferentially disposed about an engine centerline X.
  • the turbines 30 and 34 extract energy from the hot combustion gases flowing from the combustor 26.
  • the high pressure turbine 30 utilizes the extracted energy from the hot combustion gases to power the high pressure compressor 22 through a high speed shaft 38.
  • the low pressure turbine 34 utilizes the extracted energy from the hot combustion gases to power the low pressure compressor 18 and the fan 14 through a low speed shaft 42.
  • the examples described in this disclosure are not limited to the two-spool engine architecture described and may be used in other architectures, such as a single spool axial design, a three-spool axial design, and still other architectures. That is, there are various types of engines that can benefit from the examples disclosed herein, which are not limited to the design shown.
  • the low pressure turbine 34 drives the fan 14 through a gear system 36.
  • the gear system 36 can be any known suitable gear system, such as a planetary gear system with orbiting planet gears, planetary system with non-orbiting planet gears, or another type of gear system.
  • the gear system 36 has a constant gear ratio, and enables the low pressure turbine 34 to rotate at a higher speed than the fan 14. In the prior art, the rotational speed of the low pressure turbine is the same as the fan.
  • an airfoil assembly 50 from the low pressure turbine 34 includes an attachment 54 and an airfoil 58.
  • the airfoil 58 extends from a leading edge 62 to a trailing edge 66 relative to flow through the low pressure turbine 34.
  • the airfoil 58 has a pressure surface 70, a suction surface 74, and a chord length 78.
  • the chord length 78 represents the distance between the leading edge 62 and the trailing edge 66.
  • the airfoil 58 also has an axial chord 82, which is a projection of the chord length 78 onto a plane containing the axis X of the gas turbine engine 10 ( Figure 1 ).
  • fluid exiting the high pressure turbine 30 moves through passages established between adjacent ones of the airfoil 58.
  • the fluid has a laminar flow portion 86 and a turbulent flow portion 90 relative to the suction surface 74 of the airfoil 58.
  • the laminar flow portion 86 transitions to the turbulent flow portion 90 at a transition location 94.
  • transition location 94 is closer to the trailing edge 66 of the airfoil 58 than in prior art designs. Accordingly, laminar flow along a suction surface 74 of the airfoil 58 extends further than in the prior art.
  • the example airfoil 58 is positioned in a specific location relative to an adjacent airfoil 58a in the low pressure turbine 34.
  • the airfoil 58 is positioned such a ratio of a pitch distance 98 to the axial chord length 82 is less than 1.3.
  • the pitch distance 98 represents the distance, or circumferential spacing, between the trailing edge 66 of the airfoil 58 and the airfoil 58a.
  • the narrowest area of the passage 106 between the airfoil 58 and the airfoil 58a is referred to as a throat 102.
  • a throat location 110 of the airfoil 58 represents a position of the airfoil 58 corresponding to the narrowest area of a passage 106 between the airfoil 58 and the airfoil 58a.
  • the airfoil 58 has an uncovered turning angle aft of the throat location 110 that is greater than 15 degrees. This amount of uncovered turning facilitates locating the transition location 94 for the airfoil 58.
  • the uncovered turning angle in this example represents the difference between a surface angle of the airfoil 58 at the throat location 110 and a surface angle of the airfoil 58 at the trailing edge 66.
  • a person having skill in this art and the benefit of this disclosure would understand how to determine uncovered turning angle for an airfoil.
  • an axial chord length is 1.27 inches (32.36 mm) and a pitch distance is 1.29 inches (32.77 mm) for a pitch-to-chord ratio of 1.02.
  • the uncovered turning angle is 17 degrees in this example.
  • the ratio of 1.02 and angle of 17 degrees performs particularly well in the engine 10.
  • Other combinations of ratios and angles may perform well in other engines and turbomachines. That is, the specific ratio and angle can be adjusted to obtain the desired performance characteristics. The adjustments may be required due to the operating environment and other factors, for example.
  • inventions include an airfoil curved and positioned such that the transition location is located closer to the aft end of the airfoil than in the prior art.
  • the example low pressure turbine is able to rotate at a different speed than a fan of the fan section, which facilitates designing the airfoil to position the transition location toward the aft end of the airfoil portion.
  • the gearing between the fan and the low pressure turbine allows the low pressure turbine to rotate at a higher speed than the fan.

Landscapes

  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP11181843A 2010-09-21 2011-09-19 Aube chargée à l'arrière Withdrawn EP2441919A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/886,693 US20120070297A1 (en) 2010-09-21 2010-09-21 Aft loaded airfoil

Publications (1)

Publication Number Publication Date
EP2441919A1 true EP2441919A1 (fr) 2012-04-18

Family

ID=44785425

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11181843A Withdrawn EP2441919A1 (fr) 2010-09-21 2011-09-19 Aube chargée à l'arrière

Country Status (2)

Country Link
US (1) US20120070297A1 (fr)
EP (1) EP2441919A1 (fr)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2226599A (en) * 1989-01-03 1990-07-04 Gen Electric Turbofan engine
US5221181A (en) * 1990-10-24 1993-06-22 Westinghouse Electric Corp. Stationary turbine blade having diaphragm construction
US5352092A (en) * 1993-11-24 1994-10-04 Westinghouse Electric Corporation Light weight steam turbine blade
EP0937862A2 (fr) * 1998-02-20 1999-08-25 BMW Rolls-Royce GmbH Forme pour les aubes des turbines axiales

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5035578A (en) * 1989-10-16 1991-07-30 Westinghouse Electric Corp. Blading for reaction turbine blade row
US5277549A (en) * 1992-03-16 1994-01-11 Westinghouse Electric Corp. Controlled reaction L-2R steam turbine blade
US6279242B1 (en) * 1999-09-13 2001-08-28 Swagelok Company Intrinsic gauging for tube fittings
US7694505B2 (en) * 2006-07-31 2010-04-13 General Electric Company Gas turbine engine assembly and method of assembling same

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2226599A (en) * 1989-01-03 1990-07-04 Gen Electric Turbofan engine
US5221181A (en) * 1990-10-24 1993-06-22 Westinghouse Electric Corp. Stationary turbine blade having diaphragm construction
US5352092A (en) * 1993-11-24 1994-10-04 Westinghouse Electric Corporation Light weight steam turbine blade
EP0937862A2 (fr) * 1998-02-20 1999-08-25 BMW Rolls-Royce GmbH Forme pour les aubes des turbines axiales

Also Published As

Publication number Publication date
US20120070297A1 (en) 2012-03-22

Similar Documents

Publication Publication Date Title
US11193496B2 (en) Gas turbine engine airfoil
US11767856B2 (en) Gas turbine engine airfoil
US10352331B2 (en) Gas turbine engine airfoil
US10590775B2 (en) Gas turbine engine airfoil
US20150354362A1 (en) Gas turbine engine airfoil
EP3108118A2 (fr) Surface portante de moteur à turbine à gaz
US11193497B2 (en) Gas turbine engine airfoil
US11473434B2 (en) Gas turbine engine airfoil
US20220235792A1 (en) Gas turbine engine airfoil
US10502229B2 (en) Gas turbine engine airfoil
EP3108101B1 (fr) Profil aérodynamique de moteur à turbine à gaz
EP3034805A1 (fr) Joint à languette ayant une partie radiale conique et section d'un moteur à turbine à gaz ayant un tel joint à languette
US11391294B2 (en) Gas turbine engine airfoil
US20170175760A1 (en) Gas turbine engine airfoil
EP2441919A1 (fr) Aube chargée à l'arrière
EP3108103B1 (fr) Aube de soufflante pour moteur à turbine à gaz
US20200173463A1 (en) Gas turbine engine airfoil
US20170167267A1 (en) Gas turbine engine airfoil

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

AX Request for extension of the european patent

Extension state: BA ME

17P Request for examination filed

Effective date: 20121018

RIN1 Information on inventor provided before grant (corrected)

Inventor name: JUREK, RENEE J.

Inventor name: HUEBNER, NORBERT

Inventor name: ESTES, MATTHEW B.

Inventor name: MAGGE, SHANKAR S.

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: MTU AERO ENGINES GMBH

Owner name: UNITED TECHNOLOGIES CORPORATION

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: MTU AERO ENGINES AG

Owner name: UNITED TECHNOLOGIES CORPORATION

17Q First examination report despatched

Effective date: 20141210

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: MTU AERO ENGINES AG

Owner name: UNITED TECHNOLOGIES CORPORATION

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION HAS BEEN WITHDRAWN

18W Application withdrawn

Effective date: 20170724