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Publication numberUS3678802 A
Publication typeGrant
Publication dateJul 25, 1972
Filing dateAug 21, 1969
Priority dateAug 21, 1968
Also published asDE1751938B1
Publication numberUS 3678802 A, US 3678802A, US-A-3678802, US3678802 A, US3678802A
InventorsButter Karl
Original AssigneeMesserschmitt Boelkow Blohm
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Arrangement of cooling chambers for rocket engine combustion chambers
US 3678802 A
Abstract
Longitudinally extending coolant channels are formed by means of a cutting tool in a monolithic tubular wall section used as the convergent-divergent thrust nozzle for a rocket engine combustion chamber. The width of the channels at particular locations along their length is established in inverse relationship to the amount of heat to be removed from the combustion chamber at that location. In forming the coolant channels their side wall planes are established and radial planes of the tubular wall section are formed in parallel relationship with the side wall planes. The cutting tool is aligned in the radial planes and then is displaced laterally into the side wall planes for cutting the channels. The number of passes required for the cutting operation depends on the width of the tool and the width of the channel which varies over the length of the combustion chamber.
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[151 3,678,802 [451 July 25,1972

[54] ARRANGEMENT OF COOLING CHAMBERS FOR ROCKET ENGINE COMBUSTION CHAMBERS [72] Inventor: Karl Butter, Munich, Germany [73] Assignee: Messerschmitt-Bolkow-Blohm Gesellschaft mit beschrankter Ilaftung, Ottobrunn, near Munich, Germany [22] Filed: Aug. 21, 1969 [21] Appl.No.: 852,121

[58] Field of Search ..60/267, 39.66, 260; 90/15. 1 90/15, 11 R, 11 C, 9, 9.4; 29/157 C [5 6] References Cited UNITED STATES PATENTS Brady ..90/l 1.3

FOREIGN PATENTS OR APPLICATIONS 129,351 12/1928 Switzerland .90/l 5.l

Schenk ..90/l 1.3

Primary Examiner-Gil Weidenfeld Attorney-McGlew and Toren I ABSTRACT Longitudinally extending coolant channels are formed by means of a cutting tool in a monolithic tubular wall section used as the convergent-divergent thrust nozzle for a rocket engine combustion chamber. The width of the channels at particular locations along their length is established in inverse relationship to the amount of heat to be removed from the combustion chamber at that location. In forming the coolant channels their side wall planes are established and radial planes of the tubular wall section are formed in parallel relationship with the side wall planes. The cutting tool is aligned in the radial planes and then is displaced laterally into the side wall planes for cutting the channels. The number of passes required for the cutting operation depends on the width of the tool and the width of the channel which varies over the length of the combustion chamber.

1 Claim, 11 Drawing; Figures PATENTEDamzs 1912 SHEET 2 [IF 4 Fig. 5

Fig. 3a

INVENTOR Karl Butter ARRANGEMENT OF COOLING CHAMBERS FOR ROCKET ENGINE COMBUSTION CHAMBERS SUMMARY OF THE INVENTION The present invention is directed to the formation of cooling channels over the variable diameter length of a rocket engine combustion chamber having a convergent-divergent thrust nozzle and, more particularly, it is concerned with the formation of the cooling channels by cutting the channels in the outer surface of a monolithic tubular wall section shaped in the form of the convergent-divergent thrust nozzle. To complete the cooling channels an outer wall covering is placed about the tubular wall section to form individual channels.

In the operation of a rocket engine, high pressure conditions are required in order to achieve the requisite efficiency and the operation takes place under extremely high temperatures. Accordingly, in liquid fuel rocket engines, it is customary to cool the highly thermally stressed combustion chamber walls by flowing at least one propellant component of the combustion process through the cooling channels, entering at the rear end of the thrust nozzle through a feed ring, passing through the cooling channels in the longitudinal direction of the combustion chamber, collecting the propellant component in another ring at the opposite end of the nozzle wall from which it is fed through an injection head into the combustion chamber. As mentioned, the cooling channels extend in the longitudinal direction of the combustion chamberthrust nozzle body. The formation of the outer wall of the cooling channels can be effected in a number of different ways known in the art.

In use, a rocket engine combustion chamber-thrust nozzle is exposed over its entire length to mechanical and, especially, to thermal stresses which vary over the extent of the combustion chamber. To obtain a uniform heat balance for the combustion chamber and to maintain a mean wall temperature over its length, it is necessary to remove greater amounts of heat from certain locations due to the higher temperature at those locations and this heat removal is achieved by a liquid cooling medium passing through the cooling channels. One means of achieving heat removal is by varying the cross sectional area of the cooling channels over the length of the combustion chamber for increasing the velocity of the liquid coolin g medium and thereby effecting variable heat removal. Such variable heat removal is easier to accomplish in combustion chamber-thrust nozzle bodies which are formed of a multiplicity of individual members within which the cooling channels are formed rather than in combustion chambers formed of a monolithic tubular wall section. With regard to monolithic tubular wall sections it has been the experience in the past that the necessary dimensioningof the cooling channels can be achieved only in cast combustion chamber-thrust nozzles. However, such units made of cast iron are, for various reasons, unsuitable for heavy duty combustion chambers. Accordingly, the formation of cooling channels which are adequate for removing the heat generated in the combustion chamber thrust noules still provides considerable difficulties apart from the fact that the manufacturing methods used up to the present time are expensive.

Accordingly, it is the primary object of the present invention to provide a method and arrangement of cooling channels in a monolithic tubular wall section for a combustion chamber-thrust nozzles of a rocket engine in which the cooling channels are easily and economically formed.

Therefore, in accordance with the present invention, the individual cooling channels are formed for their entire length or for individual sections of the combustion chamber between cutting planes which extend longitudinally through the monolithic wall section and intersect in a vertex line which may be offset from the central axis of the combustion chamber. The cutting planes are determined by the mean channel width of the cooling channels at specific locations, such as at the opposite ends of the combustion chamber and at the transition plane between the converging and diverging sections of its thrust nozzle.

After establishing the cutting planes for the cooling channel side walls radial planes of the combustion chamber are formed disposed in parallel relationship with the cutting planes. Initially, a tool reference plane is established along the plane of symmetry of the cooling channel and then the tool is displaced into a working plane parallel to the radial plane and congruent with the channel side wall to be machined. With the tool located in position to cut or otherwise form the channel side wall the tool is placed in operation for cutting the required depth of the channel and by means of its longitudinal feed it produces one side wall of the channel in a single pass. The formation of the opposite side wall of they same channel is effected in a similar manner with the displacement of the tool from the original tool reference plane corresponding to the plane of symmetry of the channel into the cutting plane disposed in parallel relationship with the adjacent radial plane of the combustion chamber.

By means of the present invention, the cooling channels can be easily formed to accommodate the various thermal stresses which exist within a rocket engine combustion chamber-thrust nozzle and .it affords a simple production method within the geometric and constructional limits of the variable dimensioning required of the channels.

The various features of novelty which characterize the invention are pointed out with particularity in the claims annexed to and forming a part of this specification. For a better understanding of the invention, its operating advantages and specific objects attained by its use, reference should be had to the accompanying drawings and descriptive matter in which there is illustrated and described a preferred embodiment of the invention.

In the drawing:

FIG. 1 is a partial longitudinal sectional view of a rocket engine combustion chamber-thrust nozzle illustrating the method of forming the cooling channels in accordance with the present invention;

FIGS. 2, 2a, and 2b aretransverse sections taken along the lines II II in FIG. 1 indicating the operation of forming the cooling channels in the wall section of the combustion chamberthrust nozzle;

FIGS. 3, 3a, and 3b are transverse sections similar to FIGS. 2, 2a,,and 2b taken along the line III-III in FIG. 1;

FIGS. 4 and 5 are top views of the cooling channels illustrating the positions of the cutting tool in the formation of the channels;

FIG. 6 is a perspective view of apparatus including a cutting tool for forming cooling channels in the wall section of the combustion chamber-thrust nozzle; and

FIG. 7 is a side view of the apparatus shown in FIG. 6.

DETAILED DESCRIPTION OF THE INVENTION In FIG. 1 a combustion chamber-thrust nozzle body 1 is shown containing a longitudinally extending cooling channel 2, though not shown the body I contains a plurality of such cooling channels and the completed combustion chamberthrust nozzle member is provided with an exterior covering about the body 1 which forms the outer wall of the cooling channels 2. The body I is formed of a monolithic tubular wall section of variable diameter along its length. One end of the body I is fon'ned by a cylindrically shaped section 1a, the opposite end is formed by a diverging section 1b and a converging section 1c extends between the cylindrical section 1a and the diverging section lb tapering inwardly from the cylindrical section to the transition plane between the converging section 1c and the diverging section 1b at which the minimum diameter of the combustion chamber exists.

The width of the cooling channels 2 is determined, in addition to geometrical and constructional requirements, by the stresses within the combustion chamber, particularly the thermal stresses. Within the combustion chamber the highest thermal stresses occur in the range of the converging section of the thrust nozzle and decrease in the diverging section. Accordingly, to increase the velocity of the liquid cooling medium flowing through the channels 2 the cross sectional area and, accordingly, the width is reduced in the sections which are more highly stressed for attaining the desired heat removal characteristics. As a result, the width of the cooling channels is less in the cylindrical and converging sections of the combustion chamber as compared to the diverging section, note FIGS. 4 and 5. In particular for geometric reasons, the cooling channels 2 are most narrow at the neck or transition point between the converging and diverging sections.

In the formation of the cooling channels, cutting planes V and V see FIGS. 2, 2a, and 2b, extend through the side walls having the width x at the transition point between the converging and diverging sections 1c, 1b and the width y for the cylindrical section 1a. Since the diameter at the location of the transition point having the width x is considerably less than for the cylindrical section having the width y, the cutting planes V,, V intersect at a line S1, 2 offset from the central axis ZL of the monolithic tubular wall section body 1. The cutting planes V V are established for a mean duct width measured at half channel height. Bi-secting the angle formed between the cutting planes V V is a plane of symmetry KT of the cooling channel. A radial plane R,, R is formed adjacent each of the cutting planes V V extending through the central axis ZL of the body 1 and disposed in parallel relationship with the adjacent cutting planes.

In the step of forming the channels, note the apparatus shown in FIGS. 6 and 7, a tool W, such as a side milling cutter is initially established in plane WB which corresponds with the plane of symmetry KT. As can be seen in FIG. 2a, the body 1 is rotated about its central axis for displacing the side of the tool W into the radial plane R Next, from the radial plane R the tool W is displaced laterally in the direction of the arrow 11 into the cutting plane V arranged in parallel relationship with the radial plane. With the side of the tool W corresponding to the cutting plane V,, the cutting or formation of the channel side wall is commenced by permitting the tool W to move into the monolithic tubular section body I for the depth of the channel and, as indicated in FIG. 4, the tool progresses from the position a in the cylindrical section la of the combustion chamber to the position I; located at the transition plane between the converging section It and diverging section lb. In this manner one side wall corresponding to the cutting plane V is formed. Similarly, in a mirror image fashion, as shown in FIG. 2b, the opposite side wall of the channel 2 is formed following the same procedure as set forth above. In the formation of the side wall V the tool W is moved in the direction of the arrow 1' from the radial plane R into the cutting plane V As is apparent, the width of the tool W must be equal to or smaller than the smallest channel width x. As a result, if the width y is greater than twice the width x of the channel 2, it is necessary that an additional cutting operation be performed to remove the portion of the body 1 remaining within the channel cross section between its side walls. Such an additional cutting operation may require one or more individual longitudinally extending cutting steps.

As indicated previously and as shown in FIGS. 3, 4 and 5, the mean channel width 1 at the end of the diverging section lb remote from the converging section 16 is greater than the width y at the opposite end of the combustion chamber. As a result of this characteristic, the cutting planes for the cooling channels extending from the end of the diverging section lb to the transition plane between the converging and diverging sections intersect along a vertex line 53, 4 which is offset from the central axis ZL of the body 1 on the opposite side of the axis relative to that indicated in FIGS. 2, 2a, 2b. The formation of the planes for cutting the cooling channels 2 from the end of the diverging section lb to its point of intersection with the converging section lc is similar to that described above. The radial planes R R are disposed in parallel relationship with the cutting planes V V however, the radial planes are located outwardly from the cutting planes relative to the location of the cooling channel, as distinct from the arrangement of the cutting planes V, and V which were located outwardly from the radial planes R,, R Initially, the tool reference plane W3 is congruent with the plane of symmetry KT of the cooling channel to be formed and then the body I is rotated about its central axis ZL until the tool reference plane is co-planar with the radial plane R By displacing the tool W in the direction of the arrow 1', the tool is placed in the cutting plane V in parallel relationship with the radial plane R and the cutting operation can be commenced. As shown in FIG. 5, the cutting operation commences at position c and progresses in the direction indicated by the arrow to position d corresponding generally to position b of FIG. 4 which is in the transition plane between the converging and diverging sections 10, 112. After the formation of the side wall plane V the opposite side wall of the channel located in the cutting plane V, is formed in a similar mirror image fashion as indicated by FIG. 3b.

In FIGS. 6 and 7 a combustion chamber-thrust nozzle body is positioned on a rotatable mandrel which is supported on a movable bed plate. A cutting device 13 with a tool W is movably positionable relative to the combustion chamberthrust nozzle body for cutting the cooling channels. The body is displaced in the axial direction relative to the cutting device by moving the bed plate in the direction of the arrow m. The variation in the cutting depth of the tool W is obtained by moving the cutting device 13 in the direction of the arrows 0-p and the tool itself is displaceable in the direction of the arrows h-i.

Since the divergence of the cooling channels in the direction of the diverging section lb is greater than in the direction of the cylindrical section la, it is possible to let the tool W extend from the cylindrical section into the converging section beyond the thrust nozzle neck or transition plane as indicated by position b in FIG. 4, as a result, the channel side walls corresponding to the cutting planes V V, are not damaged in the diverging section of the thrust nozzle. However, when the channel walls corresponding to the cutting planes V V, are being formed care must be taken that the tool W does not extend beyond the transition plane, that is, the nozzle neck indicated by the position d in FIG. 5 to avoid damage to the continuing side walls of the channel located on the opposite side of the transition plane at position d.

Accordingly, a very thin cutter tool W, that is, of the small width or a cutter tool having a very small diameter may be used in the formation of a cooling channel in the range of the nozzle neck and thereby avoiding any contact between the cutter and the side walls in the cutting planes V V beyond the position d.

It is possible to form the channel side walls in the converging section and the cylindrical section in the opposite direction to that indicated in FIG. 4, that is, from position b to position a. It is also possible to form the cooling channels in the diverging section ranging from the position d to the position 0.

What is claimed is:

l. A method of forming coolant channels in a rocket engine combustion chamber formed by an elongated monolithic tubular wall section symmetrically arranged about a longitudinally extending central axis and having a cylindrically shaped section at one end, a diverging section at the other end and a converging section tapering inwardly from and extending between the cylindrically shaped section to the small diameter end of the diverging section, in consideration of the determined dimensions of the coolant channels and the location of the cutting planes defining the lateral sides of the coolant channels in the cylindrically shaped, converging and diverging sections of the tubular wall section as based on the thermal conditions existing in the combustion chamber and with the channel widths in the different sections being based on a diameter proportionality of the sections, the steps of forming the channels comprising locating a cutting member outwardly from the tubular wall section and in a plane of symmetry bisecting the angle formed between the opposite cutting planes defining the longitudinally extending sides of the channel to be formed and extending through the longitudinally extending central axis of the tubular wall section, rotating the tubular wall section about its longitudinally extending central axis for displacing the cutting member into a radial plane of the tubular wall section adjacent to the cutting plane of the channel, displacing the cutting member laterally from the radial plane into the cutting plane and locating the side of the cutting member more remote from the plane of symmetry in the cutting plane while maintaining the cutting member in parallel relationship with its position in the adjacent radial plane, and operating the cutting member along the exterior surface of the tubular wall section while maintaining the cutting member in the cutting plane parallel to the radial plane between one end of the tubular wall section and the transition point between the converging and diverging sections, repeating the positioning of the cutting member into the opposite side wall cutting plane of the coolant channel and operating the cutting member for forming the other side wall of the cooling channel to the required depth, repositioning the cutting member intermediate the side wall cutting planes for removing any portion of the tubular wall section within the channel remaining between the cutting planes, and repeating the positioning and cutting steps for forming the coolant channels between the opposite end of the tubular wall section and the transition plane between the converging and diverging sectrons.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2397086 *Sep 27, 1943Mar 26, 1946Gear Grinding Mach CoMethod and apparatus for cutting keyways
US2633776 *Aug 14, 1948Apr 7, 1953Kellogg M W CoMethod of manufacturing turbine blades integral with turbine rotor
CH129351A * Title not available
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US4778314 *Jul 17, 1987Oct 18, 1988Klein, Schanzlin & Becker AktiengesellschaftMethod of machining grooves in shafts and like workpieces
US7334333 *Jan 26, 2004Feb 26, 2008United Technologies CorporationMethod for making a hollow fan blade with machined internal cavities
US7458780Aug 15, 2005Dec 2, 2008United Technologies CorporationHollow fan blade for gas turbine engine
US7993105Aug 9, 2011United Technologies CorporationHollow fan blade for gas turbine engine
US8448335 *Dec 19, 2006May 28, 2013Volvo Aero CorporationMethod of manufacturing a wall structure and a machining tool
US20050160599 *Jan 26, 2004Jul 28, 2005Palazzini Christopher M.Hollow fan blade for gas turbine engine
US20070036652 *Aug 15, 2005Feb 15, 2007United Technologies CorporationHollow fan blade for gas turbine engine
US20070128042 *Dec 6, 2005Jun 7, 2007United Technologies CorporationHollow fan blade for gas turbine engine
US20100058586 *Dec 19, 2006Mar 11, 2010Volvo Aero CorporationMethod of manufacturing a wall structure and a machining tool
EP2094420A1 *Dec 19, 2006Sep 2, 2009Volvo Aero CorporationA method of manufacturing a wall structure and a machining tool
Classifications
U.S. Classification409/132, 60/267
International ClassificationF02K9/64, F02K9/00, F02K9/97
Cooperative ClassificationF02K9/64, F02K9/972
European ClassificationF02K9/64, F02K9/97B