US3836283A - Construction of axial-flow turbine blades - Google Patents
Construction of axial-flow turbine blades Download PDFInfo
- Publication number
- US3836283A US3836283A US00313843A US31384372A US3836283A US 3836283 A US3836283 A US 3836283A US 00313843 A US00313843 A US 00313843A US 31384372 A US31384372 A US 31384372A US 3836283 A US3836283 A US 3836283A
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- United States
- Prior art keywords
- temperature
- blade
- alpha
- delta
- fluid
- 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.)
- Expired - Lifetime
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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
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
- F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
- F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
-
- 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
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- 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
- FIG. l is a cross-sectional, end view of an axial-flow turbine blade constructed in accordance with the present invention, and the cooling thereof comprises impingement cooling at the leading edge, convection cooling in the mid-chord region and film cooling at the trailing edge;
- FIG. 2 is a graph illustrating the distribution of the effective local heat transfer coefficients
- FIG. 3 is a schematic diagram used for the calculation of the non-steady state, one-dimensional temperature
- FIG. 4 is a graph illustrating the non-steady state temperature distribution at the leading edge of a blade with elapsed time
- FIG. 5 is a graph of the blade thickness distribution in the chordwise direction, i.e., leading edge to trailing edge direction, of the outer shell of the blade;
- FIG. 6 is a cross-sectional, end view of another example of an axial-flow turbine blade constructed in accor dance with the present invention, in which film cooling is used at the leading edge region and the trailing edge region;
- FIG. 7 is a graph illustrating the distribution, of nonsteady state thermal stresses without the use of the invention.
- FIG. 8 is a graph illustrating the distribution of nonsteady state thermal stresses with the use of the present invention.
- the present invention provides a method for making the shell thickness distribution on the pressure side and the suction side of the hollow blade correspond to the distribution of the effective local heat transfer coefficients along the blade surface in the chordwise direction, i.e., the direction from the leading edge to the trailing edge.
- the temperature at each part of the blade changes almost uniformly even in the case of the transient operation such as starting, stopping, acceleration and deceleration. Therefore, no excessive thermal stresses occur in the blade, and consequently, the durability of the blade constructed by such method is remarkably increased compared with that of the conventional hollow blade which is constructed without taking into consideration the transient operation.
- the said effective local heat transfer coefficient a is a constant of proportionality, defined by the following equation,
- the blade illustrated in cross-section in FIG. 1 has an outer shell 1 having a lower pressure or suction surface wall and a high pressure or pressure surface wall, the suction side being designated by the numeral 3 and the pressure side being designated by the numeral 4, and a cooling fluid insert or duct 2 is' within the shell 1 and has its outer wall spaced from the inner wall of the shell 1.
- the insert 2 has an opening 2a for directing cooling fluid against the leading edge portion of the shell 1, and the fluid flows rearwardly of the blade between the outer wall of the insert 2 and the inner wall of the shell 1 and is exhausted through the channel 1a.
- Reference numeral 5 designates one of the small elements or portions of the outer shell 1 which is used for the application of numerical calculations.
- Reference numerals 6 and 7 designate main air flow side and cooling air flow side of the hollow blade respectively.
- the intersections of the extensions of the wall of the'impingement hole 2a of said insert 2 and the inner surface of the said outer shell 1 is designated by the letter P.
- Extensions which are 6 50 on both sides of the impingement hole center line and which go through the center of the circle which contains the blade leading edge will intersect the inner surface of the said outer shell 1 at Q.
- the main flow is divided into two parts, suction side surface flow and pressure side surface flow, at the outer surface stagnation point R.
- the cooling air impinges on the inner surface stagnation point S which is located at the inner side of the shell 1, and opposite to the point R.
- the local heat transfer coefficients a in the main air flow side 6 along the outer surface of the shell 1 and a in the cooling air flow side 7 along the inner surface of the shell 1, in the chordwise direction can be calculated from the empirical equations explained below.
- the empirical equation on the convective heat transfer is univers'ally described with some dimensionless numbers as follows,
- Nut R m P a proper textbook of Heat Transfer, e.g., Heat & Mass Transfer by Eckert, Drake, McGraw-Hill, or Heat Transmission by McAdams', McGraw-Hill, and c, m and n are numerical constants.
- the values of the heat transfer coefficients are calculated from each empirical equation applied to the blade portion identified hereinafter.
- the leading edge region can be considered as a circular cylinder in the'range from the leading edge stagnation point R to 0 Therefore, the following empirical equation by Schmidt and Wenner (see Anlagen, 12, (1941)) is used for the heat transfer coefficients along the circumference of a circular cylinder,
- Cooling air flow side (01) i. the leading edge stagnation point S and the adjacent region:
- the heat transfer coefficient a at the leading edge stagnation point S in the cooling air flow side is obwhere S and 5, represent surface heat transfer area in the main air flow side and in the cooling air flow side, respectively.
- T represents the main air flow inlet temperature and T represents the cooling air inlet temperature.
- U represents the local velocity of the cooling air flow
- 1:,- represents the distance in the chordwise direction from the leading edge stagnation point S in the cooling air flow side along the inner surface of the outer shell ll
- U and X are values of U and X at the point P, respectively.
- FIG. 2 is a graph which shows the distribution of the blade surface local heat transfer coefficients a a using the methods of calculation just described.
- the ordinate is the heat transfer coefficient
- the abscissa is the distance along the outer blade surface and the origin corresponds to the leading edge stagnation points R and S.
- FIG. 3 is a schematic diagram referred to for the calculation of the non-steady state temperature in a small element of the blade, such as the small element 5.
- the blade shell thickness be I, and assume that the y axis is oriented in the blade shell thickness direction with its origin located at the blade surface in the main air flow side.
- T represents the main air flow temperature
- T represents cooling air flow temperature.
- T, and T represents the blade surface temperatures at y 0 and y 1 respectively, under steady state conditions.
- To represents the temperature of the entire region kept in an equilibrium state that is realized before heating or after cooling.
- Temperature T(y,t) at an arbitrary position and arbitrary time in the small element can be obtained from the following fundamental equation:
- the response of temperature T(y,t) is not exactly the same as the first-order response to the step input used in the linear dynamic system, but its trend is very similar.
- the time constant 1' of the blade temperature T(y,t) is defined by the same method as is used in said first-order response, namely is the elapsed time when 63.2 percent of the value at the steady state, T(y,), is reached.
- the shell thickness 1 along the blade surface so that the time constant 7 may be considered much the same in every part of the blade.
- the blade thickness distribution in the chordwise direction calculated by the said procedure is shown by the graph of FIG. 5.
- the ordinate is the blade thickness 1
- the abscissa is the distance along the blade surface and the origin corresponds to the leading edge.
- the representative time constant rm is equal to 2.391 sec.
- FIG. 7 and FIG. 8 are graphs illustrating the distribution of non-steady state thermal stresses obtained from non-steady state blade temperature distribution calculated by equations (13) (17).
- the or dinate is the thermal stress aKg/mm
- the abscissa is the distance along the blade surface
- the elapsed time t is taken as the parameter.
- FIG. 7 is the result obtained in the case of constant blade thickness that does not take into consideration the desirability of equal time constants.
- FIG. 8 is a graph of the results obtained by the methods of the present invention which considers the time constants and makes them substantially equal. From these two figures, it is apparent that if the transient response at every part of the blade is taken into consideration, thermal stresses can be remarkably reduced. Then, according to the present invention, crack initiation on the blade surface can be avoided for far longer times than have heretofore been accomplished, and consequently, the blade can sufficiently withstand the frequent starts and stops of the engines including the blades.
- the blade thickness calculated by the methods of the present invention conflict with the blade profile designed on the basis of the aerodynamic performance, especially in the trailing edge region, it is sufficient to make the effective local heat transfer coefficient correspond to the profile desired from the aerodynamic performance and then introducing a film cooling or a transpiration cooling to the relevant region.
- FIG. 6 is another embodiment of the turbine blades to which the present invention is applied.
- the cooling thereof comprises impingement cooling and film cooling at the leading edge, convection cooling in the midchord region and film cooling in the trailing edge region.
- the outer shell 1 encloses a pair of inserts 2b and 2c.
- the holes 8 and 9 are made at the leading edge region for film cooling.
- the equality of transient response of the various portions of the blade is easily realized within the required blade profile because of the application of the film cooling through the channels or holes 10 and 11 at the trailing edge.
- axial-flow turbine blades constructed in accordance with the present invention are very strong and resistant to frequent heat variations, such as by reason of starts and stops. In other words, the durability of the blade is remarkably increased.
- the turbine inlet temperature of the motive fluid can be higher, resulting in improvement of the thermal efficiency of a gas turbine or a steam turbine.
- axial-flow turbine blades in accordance with the present invention is useful not only in aircraft engines, but also in marine turbines, steam turbines, automobile engines, etc. Accordingly, the present invention is extremely useful for industrial purposes.
- a hollow turbine part for use in a hot fluid medium said part having a pressure surface wall and a suction surface wall and having a leading edge and a trailing edge, said walls having a thickness distribution in the direction from said leading edge to said trailing edge such that, with changes of the temperature of'said fluid, the temperature response at each portion of said walls in substantially the same as the temperature re sponse at the other portions of said walls, whereby the temperature distribution in said walls changes substantially uniformly in response to changes in temperature of said fluid.
- a hollow turbine part as claimed in claim ll wherein said part is a hollow blade and wherein said thickness distribution is such that each portion of said walls has a mean temperature time constant which is substantially equal to a predetermined time constant, said mean time constant at each portion of said walls being the mean of the temperature time constants at the outer surface thereof, at the inner surface thereof and at an intermediate point between the surfaces thereof, each of said outer surface, inner surface and intermediate point time constants being determined by replacing the temperature response at said outer surface, said inner surface and said point to a step change of said fluid temperature with approximately a first order response thereto, and said predetermined time constant being substantially equal to the mean time constant at said leading edge of said blade.
- a hollow turbine blade for use in a fluid medium, said blade comprising a pressure surface wall and a suction surface wall and having a leading and trailing edge, said walls having a thickness distribution in the direction from the leading edge to the trailing edge of said blade such that the temperature response at each portion of said walls in substantially the same as the other portions of said walls with changes of the temperature of said fluid and such that the mean temperature time constant is substantially equal to the mean temperature time constant at said leading edge of said blade, said mean time constant at each portion of said walls being the mean of the temperature time constants at the outer surface thereof, at the inner surface thereof and at an intermediate point between the surfaces thereof, each of said outer surface, inner surface and intermediate point time constants being determined by replacing the temperature response at said each point to a step change of said fluid temperature with approximately a first order response thereto, said temperature response at each point being calculated from the following equa trons:
- T(y,t) represents a temperature at an arbitrary position and an arbitrary time in a small element of said wall
- 1 represents elapsed time after a sudden temperature change of said motive fluid
- a represents thermal diffusivity of the blade material
- .and y represents the axis oriented in the blade wall thickness direction with its origin located at the blade surface in the main air flow side;
- the boundary conditions and initial conditions are:
- T represents the recovery temperature of the fluid
- T represents the cooling air flow temperature
- To represents the temperature of the entire region kept in an equilibrium state that is realized before heating or after cooling
- 1 represents the wall thickness and represents the conductivity of the blade material
- T and T are:
- T represents the non-steady state term of the 1 o TA temperagtcure, and is expressed as follows: 5 C2 (TA TB) 2 2 K ne /A, and
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP47044663A JPS527482B2 (es) | 1972-05-08 | 1972-05-08 |
Publications (1)
Publication Number | Publication Date |
---|---|
US3836283A true US3836283A (en) | 1974-09-17 |
Family
ID=12697673
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US00313843A Expired - Lifetime US3836283A (en) | 1972-05-08 | 1972-12-11 | Construction of axial-flow turbine blades |
Country Status (3)
Country | Link |
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US (1) | US3836283A (es) |
JP (1) | JPS527482B2 (es) |
FR (1) | FR2184261A5 (es) |
Cited By (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4183716A (en) * | 1977-01-20 | 1980-01-15 | The Director of National Aerospace Laboratory of Science and Technology Agency, Toshio Kawasaki | Air-cooled turbine blade |
US4456428A (en) * | 1979-10-26 | 1984-06-26 | S.N.E.C.M.A. | Apparatus for cooling turbine blades |
US6050777A (en) * | 1997-12-17 | 2000-04-18 | United Technologies Corporation | Apparatus and method for cooling an airfoil for a gas turbine engine |
US6206637B1 (en) * | 1998-07-07 | 2001-03-27 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
EP1101901A1 (de) * | 1999-11-16 | 2001-05-23 | Siemens Aktiengesellschaft | Turbinenschaufel sowie Verfahren zur Herstellung einer Turbinenschaufel |
EP1132574A2 (en) * | 2000-03-08 | 2001-09-12 | Mitsubishi Heavy Industries, Ltd. | Gas turbine cooled stationary blade |
US20100068069A1 (en) * | 2006-10-30 | 2010-03-18 | Fathi Ahmad | Turbine Blade |
US20100250155A1 (en) * | 2009-03-30 | 2010-09-30 | General Electric Company | Method for quantifying hole flow rates in film cooled parts |
EP3808939A1 (en) * | 2019-10-14 | 2021-04-21 | Raytheon Technologies Corporation | Airfoil vane, baffle for an airfoil vane assembly and method of assembling a ceramic matrix composite airfoil vane |
Families Citing this family (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE2537027C2 (de) * | 1975-08-20 | 1982-06-03 | Bayer Ag, 5090 Leverkusen | Verfahren zur Herstellung von 4,4'-Dihydroxy-3,3',5,5'-tetraalkyl- diphenylalkanen |
DE2935316A1 (de) * | 1979-08-31 | 1981-03-26 | Bayer Ag, 51373 Leverkusen | Verfahren zur herstellung reiner 4,4'-dihydroxydiphenylalkane bzw. -cycloalkane |
JPH0161908U (es) * | 1987-10-14 | 1989-04-20 | ||
JP7196607B2 (ja) | 2017-07-19 | 2022-12-27 | 東洋紡株式会社 | グルコース測定方法およびグルコースセンサ |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR924012A (fr) * | 1946-02-18 | 1947-07-24 | Const Aeronautiques Du Ct Soc | Perfectionnement apporté aux turbines à combustion |
DE892698C (de) * | 1943-05-21 | 1953-10-08 | Messerschmitt Boelkow Blohm | Luftgekuehlte Hohlschaufel, insbesondere fuer Gas- und Abgasturbinen |
GB910400A (en) * | 1960-11-23 | 1962-11-14 | Entwicklungsbau Pirna Veb | Improvements in or relating to blades for axial flow rotary machines and the like |
US3420502A (en) * | 1962-09-04 | 1969-01-07 | Gen Electric | Fluid-cooled airfoil |
US3650635A (en) * | 1970-03-09 | 1972-03-21 | Chromalloy American Corp | Turbine vanes |
-
1972
- 1972-05-08 JP JP47044663A patent/JPS527482B2/ja not_active Expired
- 1972-12-11 US US00313843A patent/US3836283A/en not_active Expired - Lifetime
- 1972-12-14 FR FR7244599A patent/FR2184261A5/fr not_active Expired
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE892698C (de) * | 1943-05-21 | 1953-10-08 | Messerschmitt Boelkow Blohm | Luftgekuehlte Hohlschaufel, insbesondere fuer Gas- und Abgasturbinen |
FR924012A (fr) * | 1946-02-18 | 1947-07-24 | Const Aeronautiques Du Ct Soc | Perfectionnement apporté aux turbines à combustion |
GB910400A (en) * | 1960-11-23 | 1962-11-14 | Entwicklungsbau Pirna Veb | Improvements in or relating to blades for axial flow rotary machines and the like |
US3420502A (en) * | 1962-09-04 | 1969-01-07 | Gen Electric | Fluid-cooled airfoil |
US3650635A (en) * | 1970-03-09 | 1972-03-21 | Chromalloy American Corp | Turbine vanes |
Cited By (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4183716A (en) * | 1977-01-20 | 1980-01-15 | The Director of National Aerospace Laboratory of Science and Technology Agency, Toshio Kawasaki | Air-cooled turbine blade |
US4456428A (en) * | 1979-10-26 | 1984-06-26 | S.N.E.C.M.A. | Apparatus for cooling turbine blades |
US6050777A (en) * | 1997-12-17 | 2000-04-18 | United Technologies Corporation | Apparatus and method for cooling an airfoil for a gas turbine engine |
US6210112B1 (en) * | 1997-12-17 | 2001-04-03 | United Technologies Corporation | Apparatus for cooling an airfoil for a gas turbine engine |
US6206637B1 (en) * | 1998-07-07 | 2001-03-27 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
WO2001036791A1 (de) * | 1999-11-16 | 2001-05-25 | Siemens Aktiengesellschaft | Turbinenschaufel sowie verfahren zur herstellung einer turbinenschaufel |
EP1101901A1 (de) * | 1999-11-16 | 2001-05-23 | Siemens Aktiengesellschaft | Turbinenschaufel sowie Verfahren zur Herstellung einer Turbinenschaufel |
EP1132574A2 (en) * | 2000-03-08 | 2001-09-12 | Mitsubishi Heavy Industries, Ltd. | Gas turbine cooled stationary blade |
EP1132574A3 (en) * | 2000-03-08 | 2003-07-16 | Mitsubishi Heavy Industries, Ltd. | Gas turbine cooled stationary blade |
US20100068069A1 (en) * | 2006-10-30 | 2010-03-18 | Fathi Ahmad | Turbine Blade |
US20100250155A1 (en) * | 2009-03-30 | 2010-09-30 | General Electric Company | Method for quantifying hole flow rates in film cooled parts |
US7890274B2 (en) * | 2009-03-30 | 2011-02-15 | General Electric Company | Method for quantifying hole flow rates in film cooled parts |
EP3808939A1 (en) * | 2019-10-14 | 2021-04-21 | Raytheon Technologies Corporation | Airfoil vane, baffle for an airfoil vane assembly and method of assembling a ceramic matrix composite airfoil vane |
US11280201B2 (en) | 2019-10-14 | 2022-03-22 | Raytheon Technologies Corporation | Baffle with tail |
Also Published As
Publication number | Publication date |
---|---|
FR2184261A5 (es) | 1973-12-21 |
JPS496310A (es) | 1974-01-21 |
JPS527482B2 (es) | 1977-03-02 |
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