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Publication numberUS3585800 A
Publication typeGrant
Publication dateJun 22, 1971
Filing dateJul 27, 1967
Priority dateJul 27, 1967
Publication numberUS 3585800 A, US 3585800A, US-A-3585800, US3585800 A, US3585800A
InventorsRobert J Kuntz, Albert L Blubaugh, Richard J Labotz, Charles B Mcgough
Original AssigneeAerojet General Co
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Transpiration-cooled devices
US 3585800 A
Abstract  available in
Images(2)
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Claims  available in
Description  (OCR text may contain errors)

United States Patent [72] Inventors Robert J. Kuntz Rancho Cordova; Albert L. Blubaugh, Sacramento; Richard J. Labotz, Fair Oaks; Charles B. McGough, Fair Oaks, all of, Ala. [21] Appl. No. 656,522 [22] Filed July 27, 1967 [45] Patented June 22, 1971 [73] Assignee Aerojet-General Corporation El Monte, Calif.

[54] TRANSPIRATION-COOLED DEVICES 10 Claims, 3 Drawing Figs.

[52] US. Cl 60/265, 60/39.66, 239/127.1 [51] lmt. Cl ..F02k 11/02, 864d 33/08 [50] Field of Search 60/258, 265, 260, 39.66; 2-39/127.3

[56] References Cited UNITED STATES PATENTS 2,705,399 4/1955 Allen... 60/39.66 2,926,490 3/1960 Eaton 60/267 3,115,746 12/1963 Hsia 60/265 3,153,320 10/1964 Prosser... 239/1273 3,289,405 12/1966 Eng 60/39.74 3,309,026 3/1967 Loprete 60/266 I IIII'IIIIII'IIIIIII'III'Ie (\""""'mummmti........-.-...............

Primary Examiner-Douglas Hart Attorneys-Edward O. Ansell and D. Gordon Angus ABSTRACT: This invention relates to transpiration cooling and particularly to transpiration cooling of devices subjected to heat flux above the capabilities'of the material forming the device, such as thrust chambers for rockets.

According to the present invention, a device to be cooled by transpiration cooling has a wall constructed of a plurality of discrete wafers, each having planar surfaces and edge surfaces. A controlled pattern of flow passages is fonned on one or both planar surfaces of each wafer, each flow passage terminating at an end surface of the wafer. The wafers are joined together in a stack to form a unitary structure. The passage terminating edge surfaces together form the wall surface of the device, and the wall surface resembles a porous wall. The wafers are thin enough to give excellent wafer-to-coolant heat transfer area.

Transpiration cooling is accomplished by permitting coolant to flow through the passages in such a manner that heat input into the wall of the device is efi'lciently transferred to the coolant within the wall. This enables the coolant to reach the maximum limiting temperature of the wall material before it is introduced onto the wall surface, thus fully exploiting the cooling capacity of the coolant.

According to an optional but desirable feature of this invention, the flow passages in the wafers are divided into discrete groups, thereby allowing isolation of areas where coolant boiling might occur. Restrictor passages are provided to each group of passages so that the flow resistance into each group of passages is independent of the heat transfer on the surface of the device.

PATENTED JUN22 I9?! SHEET 2 OF 2 DIVERCENT SECTION DISTANCE FROM THE INJECTOR IIN.)

INVENTORS F l G 5 ROBERT J. KUNTZ ALBERT L. OLUBAUCII RICHARD J. LA BOTZ Z CNZRLES B. N CH ATTORN TRANSPIRATION-COOLED DEVICES This invention relates to transpiration cooling and particularly to transpiration cooling of devices subjected to heat flux above the capabilities of the material forming the device, such as rocket thrust chambers.

Devices, such as thrust chambers, cooled with conventional techniques, are susceptible to wall failures due to erosion in regions where the heat transfer conditions are unusually severe. Local melting or erosion of a small area of the surface of the chamber can cause a complete failure of the chamber wall since high-pressure coolant within a coolant tube will erupt through the failure opening.

Transpiration cooling has been recognized to be an effective method of cooling chambers under severe heat transfer operating conditions; that is, chambers operating under high pressures and high heat fluxes. Transpiration cooling is basically a process whereby a surface device, vessel, wall, container or the like in contact with heat energy at a temperature above the capabilities of the material forming the surface, is kept cool by the flow of a coolant fluid through the material toward the heat source. 1

Prior to the present invention, practical application of transpiration cooling has not been accomplished with any degree of success. Difficulties in fabricating complex shapes, coolant flow control problems, and porosity nonuniformities, have been major stumbling blocks to the successful development of transpiration-cooled devices. In addition, those devices constructed with conventional porous materials have been subject to "hot spot instability." Hot spot instability is a phenomenon in which an overheated region is deprived of coolant and the overheated region grows in size and intensity. Hot spot instability occurs because the hydraulic resistance of a conventional porous wall increases as the wall temperature rises due to the decreasing coolant density. Thus, when a region overheats, the coolant circumvents it due to its higher hydraulic resistance and the hot spot, now partially deprived of coolant, partially becomes hotter. Furthermore, conventional porous materials, being permeable in all directions, have the inherent limitation of being susceptible to the flow of hot combustion gases throughthe porous material, especially when utilized in a high-pressure gradient region of a throat.

The present invention overcomes the foregoing disadvantages of prior known techniques in that wafers or plates having passages to permit coolant flow are used; the wafers being thin enough to provide intimate thermal contact between the metal and the coolant within the wall. This enables the coolant to approach the maximum limiting temperatures of the wall material before it is introduced onto the chamber surface. In this manner, the cooling capability of the fluid is completely exploited and the coolant flow may be minimized. Furthermore, because with this invention the coolant passages are not random but discrete, they can be designed such that the coolant is brought to the heated surface in a great multiplicity of hydraulically isolated parallel passages. By designing these passages such that their area of major hydraulic resistance is some distance removed from the heated surface, their hydraulic resistance is made essentially independent of the surface heat flux. This uncouples the coolant flow from the surface temperature and substantially eliminates hot spot instability.

It is an object of the present invention to provide a transpiration cooling system for a device, such as a thrust chamber for a rocket.

Another object of the present invention is to provide a wafer design which makes possible the use of thin wafers having passages for coolant flow for transpiration cooling of devices.

According to the present invention, a device to be cooled by transpiration cooling has a wall constructed of a plurality of discrete wafers, each having planar surfaces and edge surfaces. A controlled pattern of flow passages is formed in one planar surface of each wafer, each flow passage terminating at an end surface of the wafer. The wafers are joined together in a stack to form a unitary structure. The passage terminating edge surfaces together form the wall surface of the device. Each wafer is sufiiciently thin so that the resulting large number of wafers in the structure result in large coolant-towafer contact areas and provide excellent thermal communication between the coolant and wafers.

According to an optional but desirable feature of the present invention, the flow passages in the wafer are divided into groups, each group being in an individual flow compartment.

.The present invention offers significant advantages over conventional cooling techniques in the following respects:

1. Precise flow control.-Control of the flow of the coolant in the lateral direction can be accurately maintained by varying the depth or length of the passage through the wafer, and providing internal manifold sizes confonning to the flow desired. If desired, the flow can easily be varied circumferentially over the surface of the chamber to accommodate nonuniform combustion conditions. I 2. Failure protection.The wafertype transpiration-cooled design is not susceptible to failure caused by high-pressure coolant erupting through a failure opening in the chamber wall, since no high-pressure coolant is necessary in the hightemperature area of the chamber wall. Furthermore, localized overheatingwill not result in hot spot instability because the local coolant flow is uncoupled from the surface heat flux through the use of hydraulically isolated parallel coolant flow passages which have their hydraulic resistance significantly removed from the heated surface.

3. Cooling effectiveness-By permitting excellent thermal communication between the metal of the wall and the coolant at all points within the wall, the cooling capability of the ecolant is completely exploited and the cooling efficiency is maximized.

4. Fabrication simplicity.--Due to the basic simplicity and repetitious nature of the design, the same technology can be applied to rocket engines of all sizes. Building of micromotors can be accomplished with the same techniques employed in construction of huge multimillion pound boosters. In addition, larger chambers can be fabricated by using a number of smaller easily constructed modules.

The above and other features of this invention will be more fully understood from the following detailed description and the accompanying drawings, in which:

FIG. 1 is a perspective view, partly in cutaway cross section, showing a thrust chamber according to a presently preferred embodiment of the present invention;

FIG. 2 is a perspective view, partly cut away, of a plurality of wafers stacked together to form a portion of the thrust chamber illustrated in FIG. 1; and

FIG. 3 is a graphical illustration of the coolant flow rate distribution in the thrust chamber illustrated in FIG. 1.

Referring now to the drawings, and particularly to FIGS. I and 2, there is illustrated a thrust chamber 10 having a cylindrical section 11 and an exhaust nozzle comprising a convergent section 12, a throat 13 and a divergent section 14. The inside face of the thrust chamber resembles a porous wall, containing many thousands of small openings or holes 15 which are in the order of several thousandths of an inch across and arranged in a precise, predetermined pattern. Holes 15 have hydraulic diameters within the range of 0.0005 inch to 0.020 inch; a typical size being 0.004 inch. Each hole is in fluid communication with a manifold 17 by means of a precisely controlled metering passage 16 which accurately meters the coolant flow to the holes. The term hydraulic diameter as used herein means four times the cross-sectional areadivided by the wetted perimeter. The term wetted perimeter as used herein means the size of the perimeter of the passage actually wetted by fluid, which for purposes of this invention is the actual perimeter size.

The thrust chamber is made of a stack of a large number of thin wafers 19 whose inner and outer peripheries are cylinders which define the inside and outside diameters of the chamber. The wafers are housed within housing 110a which forms an outer shell for the thrust chamber and defines the outer diameter of the chamber.

Holes are formed between islands 18 on each wafer 19; the islands also serving as separators between adjacent wafers and being an integral part of the flow pattern on the wafer surface. Each wafer 19 is between 0.001 and 0.050 inches in thickness, 0.007 inch being typical although other thicknesses may be employed dependent upon the application.

Flow restrictors 20 having shapes approximating T-shaped islands, are formed as an integral part of each wafer 18, and form passages 21 between the broadest section of the restrictors. Preferably separation between adjacent flow restrictors is such that passages 21 have a hydraulic diameter of the order between 0.0005 inch and 0.010 inch, a typical hydraulic diameter being 0.002 inch. As a result of the small hydraulic diameter of these passages, the flow through them will be substantially laminar.

The narrow portion between adjacent restrictors 20 forms a flared passage or flow channel 28 which provides a low-resistance flow path in the region where the flow will be subjected to influence of heat transfer conditions on the face of the thrust chamber. This flared passage makes the total flow resistance of each path relatively insensitive to heat transfer on the face of the thrust chamber.

A distribution plenum 22 is formed in each wafer 19 so that the flow of coolant from manifold header passage 23 is permitted to flow through metering passage 16, plenum 22, passages 21 and to holes 15. Header passages 23 are in fluid communication with passage 24 of manifold 17.

The individual wafers 19 may be formed any one of several different ways. One method known to those skilled in the art is photoetching. With photoetching, a thin metal sheet is imprinted with an acid resistant ink which outlines all the flow passages. The sheet is then immersed in an acid bath which etches out the flow paths at a precisely known, predetermined rate. Several depths of etch can be obtained on a single sheet of stock by repeating the above process with different patterns. In this way, it is possible to make distribution plenum 22 considerably deeper than passages 21 or metering passages 16 to obtain the proper distribution and pressure drop characteristics. It should be pointed out however, that it may be possible to form the individual wafers by embossing them or by electroplating areas to form raised areas rather than etching out the depressed areas.

Other means include, but are not limited to, the utilization of the crude etching process in forming all the passages with the exception of the metering passage 16 and flow restrictor passage 21, which are later provided by a scribing process similar to that used in preparing defraction gratings. Grooving of channels can also be obtained by conventional indentation processes or forming rolls. These methods are likewise well known to those skilled in the art.

The completed wafers 19 are then stacked together to form the entire thrust chamber 10. For some applications it may be desirable to bond the wafers together in forming the thrust chamber. Bonding of the wafers 19 can be accomplished in several ways. One method is to electroplate to the thin sheet stock a thin flash of brazed material prior to applying acid-resistant ink. The wafers will then end up with a coating of brazing alloy on all surfaces which contact adjoining wafers, but with none in any passageway through which coolant flows. The entire thrust chamber 10 may then be placed in a furnace and brazed together.

Other methods which might be used for joining the wafers together are diffusion bonding, resistant welding, or simply applying some advanced bonding agent to the contacting surfaces.

When fully assembled, faces 25, 25a, 25b, etc., form the inside surface of the thrust chamber 10; thereby resembling a porous wall having many thousands of holes 15 on the inside face of the chamber. The face of each wafer is sized so as to provide proper dimensions for the thrust chamber upon assembly.

Manifold 17 is attached to the uppermost portion of the thrust chamber and is provided with passage 24 for distribution of coolant to the respective header passages 23. Ports 26 are provided in manifold 17 for the introduction of coolant. By way of example, a coolant may be introduced into the injector (not shown) by means of fittings (not shown) and transmitted from the injector to the ports 26 of the thrust chamber. The injector is mounted to the thrust chamber at the end of cylindrical section 11 of the chamber and over manifold 17 to close that end of the chamber. By way of example, the injector may be attached to the chamber by mating bolts to suitable threaded receptacles 27 or the like.

FIG. 3 illustrates a typical coolant flow along the length of the thrust chamber from the injector to the nozzle opening at full thrust. ln the graph of FIG. 3, the coolant flow,

in pounds per second per unit length is plotted against the distance from the injector along the length of the thrust chamber. The flow rate ideally should be substantially constant throughout the cylindrical section and into the conver gent section of the thrust chamber. The flow rate should increase in the vicinity of the throat because of the high heat flux there, and then decrease through the divergent section. Since coolant flow rate drops as the coolant flows along passage 23 from manifold 17 to the various wafers (due to the release of coolant through the wafer passages), passage 23 is preferably divided into successive increments 29, 30, 31 and 32, each having successively smaller diameters so as to maintain a sufficient flow rate through the passage 23. Increment 32 is of the smallest diameter and continues along the length of the thrust chamber through the vicinity of the convergent section, the throat, and the divergent section. By providing incrementally decreasing sizing of passage 23, the volume of coolant in the chamber walls and hence the amount of coolant lost on shutdown is minimized.

As examples of suitable coolant for use with the present invention, a propellant such as N 0, ClF or gaseous or liquid hydrogen may be used. lnert coolant such as water may likewise be used.

In operation, coolant is introduced into passage 23 by way of manifold 17 and is permitted to flow through metering passages 16 to each distribution plenum 22. As the coolant enters passages 21, the coolant develops a laminar flow due to the small size of the hydraulic diameter of each passage 21. The coolant then flows into the flared portion 28 and finally through holes 15 to the inside surface of the thrust chamber. Inside the flared portion 28 the coolant is heated and may vaporize before reaching the inside surface of the thrust chamber. Once the coolant has left the wall it is driven out the exhaust nozzle with other exhaust gases. If the coolant is a propellant it will react with the combustion gas and contribute to the performance of the thrust motor.

In the flared portion 28 of each wafer, the coolant remains in intimate thermal contact with the metal of the wall at all points within the wall. Thus, by the time that the coolant reaches holes 15, and surface 25 of each wafer, the coolant temperature approaches the maximum limiting temperature of the wall material. The cooling capacity of the coolant is therefore effectively exploited and the coolant flow may be minimized.

Provision of the individual flow compartments between each flow restrictor 20 allows isolation of areas of coolant boiling, thus preventing hot spots from spreading and creating large areas on the surface where overheating may occur.

Control of the flow of the coolant into the chamber can be accurately maintained by accurately sizing the depth or length of the metering passages 16 through the wafer. Thus, the flow of coolant to each plenum 22 and thereby to the holes is accurately controlled.

Flow passage 21 is set back from the edge surface 25 so that the flow resistance through passages 2l is independent of influence caused by local heat transfer at the edge surface.

As is well known, the thrust chamber operates by injecting propellant into the chamber from an injector, attached to the end of cylindrical section 1 1 over manifold 17. The propellant is ignited by suitable ignition means, well known in the art. The heat generated by this process is dissipated by the transpiration cooling techniques as described hereinbefore.

The present invention thus provides transpiration cooling of devices subjected to heat flux above the capabilities of the material forming the device, such as thrust chambers for rockets. By permitting intimate thermal contact to be maintained between the coolant and all points within the wall, the cooling capacity of the coolant is effectively exploited and the cooling efficiency is held at a high level.

Due to the simplicity and repetitious nature of the design, devices of all sizes may be fabricated by using an appropriate number of easily constructed modules.

Although the transpiration-cooling techniques of this invention have been described in connection with rocket thrust chambers, it is to be understood that these techniques are fully applicable to other devices as well. For example, cooling of rocket injectors and afterbumers may be accomplished in accordance with these techniques. Furthermore, the principles of this invention are fully applicable to fields of technology outside of the rocket field, for example, to furnaces and the like.

This invention is not to be limited by the embodiment shown in the drawings, and described in the description, which is given by way of example rather than of limitation but only in accordance with the scope of the appended claims.

What we claim is:

l. A device having a wall adapted to contact heat energy, said wall comprising a stack of a plurality of discrete wafers, each wafer having planar surfaces and edge surfaces; a controlled pattern of flow passages on a planar surface of each wafer, said flow passages terminating at an edge surface thereof, a plurality of flow separators in each flow passage forming a porous surface of said wall having a plurality of edge surface holes arranged in a pattern; delivery means for delivering coolant to said flow passages, said delivery means including metering passages on each wafer for metering coolant flow to edge surface holes; said wafers being sufficient thin as to permit thermal equilibrium between said wafers and the coolant within the flow passages, whereby the device is transpiration cooled by the coolant in the flow passages.

2. A device according to claim 1 wherein said edge surface holes are of a hydraulic diameter of substantially between 0.0005 inch and 0.020 inch.

3. A device according to claim 1 wherein the thickness of each wafer is substantially between 0.001 inch and 0.050 inch.

4. A device according to claim 1 further including flow restrictor passages set back from said edge surface, whereby coolant flow resistance through said restrictor passages will be independent of local heat transfer conditions at the edge surface.

5. A device according to claim 4 wherein said restrictor passages are of a hydraulic diameter of substantially between 0.0005 inch and 0.010 inch.

6. A device according to claim 4 wherein said restrictors further form flared passages between them, each flared passage containing a plurality of flow passages.

7. A rocket motor thrust chamber, comprising:

a combustion chamber adapted to receive propellant injection means at one end thereof,

exhaust nozzle means associated with said combustion chamber at the other end thereof,

said chamber having a wall comprising a stack of a plurality of discrete wafers, each wafer having planar surfaces and edge surfaces; a controlled attern of flow passages on a planar surface of each w er, said flow passages being adapted to receive coolant and terminating at an edge surface thereof;

a plurality of flow separators in each flow passage forming a porous surface of said wall having a plurality of edge surface holes arranged in a pattern; said wafers being sufficiently thin as to permit thermal equilibrium between said wafers and coolant within the flow passages, whereby the combustion chamber is transpiration cooled by coolant in the flow passages; and

delivery means for delivering coolant to said flow passages, said delivery means including metering passages on each wafer for metering coolant flow to the edge surface holes.

8. A device according to claim 7 wherein said delivery means comprises a coolant manifold, and a manifold header passage passing through each of said wafers, said header passage being a fluid communication with said manifold and said flow passages.

9. A device according to claim 7 wherein said nozzle means comprises a convergent section, a throat, and a divergent section.

10. A device according to claim 7 further including flow restrictor passages set back from said edge surface, whereby coolant flow resistance through said restrictor passages will be independent of local heat transfer conditions at the edge surface.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2705399 *Nov 7, 1952Apr 5, 1955Armstrong Siddeley Motors LtdCombustion chambers
US2926490 *Mar 19, 1957Mar 1, 1960Thiokol Chemical CorpLaminated fluid-jacketed thrust chamber structure
US3115746 *Jul 18, 1960Dec 31, 1963Lockheed Aircraft CorpHydrogen transpiration cooling of a high temperature surface using a metal hydride asthe coolant material
US3153320 *Nov 8, 1962Oct 20, 1964Gen Motors CorpCooled rocket nozzle design
US3289405 *Jan 16, 1964Dec 6, 1966Harvard EngPropellant oxidizer mixing system for rocket motors
US3309026 *May 15, 1964Mar 14, 1967Curtiss Wright CorpGas cooled rocket structures
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3798902 *Jun 13, 1973Mar 26, 1974Messerschmitt Boelkow BlohmArrangement of cooling channels for rocket engine combustion chambers
US3906725 *Jul 2, 1973Sep 23, 1975Aerojet General CoExhaust emissions control device
US3914348 *Feb 25, 1974Oct 21, 1975Aerojet General CoPlatelet-venturi carburetor for combustion engine
US3925983 *Apr 17, 1974Dec 16, 1975Us Air ForceTranspiration cooling washer assembly
US5127221 *May 3, 1990Jul 7, 1992General Electric CompanyTranspiration cooled throat section for low nox combustor and related process
US5363645 *May 14, 1993Nov 15, 1994Societe Europeenne De PropulsionEnclosure containing hot gases cooled by transpiration, in particular the thrust chamber of a rocket engine
US5501011 *Jul 25, 1994Mar 26, 1996Societe Europeenne De PropulsionMethod of manufacture of an enclosure containing hot gases cooled by transportation, in particular the thrust chamber of a rocket engine
US5732883 *Apr 25, 1996Mar 31, 1998Societe Europeene De PropulsionCombustion enclosure with cooling by transpiration
US5903976 *Dec 22, 1997May 18, 1999Societe Europeenne De PropulsionMethod of manufacturing a combustion enclosure with cooling by transpiration
DE102011000383A1 *Jan 28, 2011Aug 2, 2012Deutsches Zentrum für Luft- und Raumfahrt e.V.Thruster device for driving e.g. spacecraft, has walls porously formed in injection area, which forms injector for injecting fuel into thrust chamber unit space, so that required fuel is injected during operation of device
DE102011000383B4 *Jan 28, 2011Jan 22, 2015Deutsches Zentrum für Luft- und Raumfahrt e.V.Triebwerkvorrichtung und Verfahren zum Betreiben einer Triebwerkvorrichtung
EP1748253A2 *Jul 25, 2006Jan 31, 2007Deutsche Forschungsanstalt für Luft- und Raumfahrt e.V.Combustion chamber and method for producing a combustion chamber
WO2001087021A2 *May 7, 2001Nov 15, 2001Essox Res And Dev IncPlasma processing method and apparatus
Classifications
U.S. Classification60/265, 239/127.1
International ClassificationF02K9/64
Cooperative ClassificationF02K9/64, F05D2260/203
European ClassificationF02K9/64