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Publication numberUS3798902 A
Publication typeGrant
Publication dateMar 26, 1974
Filing dateJun 13, 1973
Priority dateAug 21, 1968
Publication numberUS 3798902 A, US 3798902A, US-A-3798902, US3798902 A, US3798902A
InventorsButter K
Original AssigneeMesserschmitt Boelkow Blohm
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Arrangement of cooling channels for rocket engine combustion chambers
US 3798902 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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United States Patent [191 Butter ARRANGEMENT OF COOLING CHANNELS FOR ROCKET ENGINE COMBUSTION CHAMBERS [75] lnventor: Karl Butter, Munchen, Germany [73] Assignee: Messerschmitt-Bolkow-Blohm Gesellschatt mit beschrankter Haftung, Munich, Germany [22] Filed: June 13, 1973 [21] Appl. No.1 369,458

Related US. Application Data [60] Continuation-in-part of Ser. No. 225,143, Feb. 10, 1972, abandoned, which is a division of Ser. No. 852,121, Aug. 21, 1969, Pat. No. 3,678,802.

3,190,070 6/1965 Neu 60/3966 1111' 3,798,902 [451 Mar. 26, 1974 Jones et a1. 60/39.66 Butler 60/267 Stanger 5 7] 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 andradial 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 PATENTEIIIIAR26 1925 3.798302 sum 2 BF 4 FIGZc This is a continuation-in-part of application Ser. No.

225,143, filed Feb. 10, l972 now abandoned, which was a division of application Ser. No. 852,121 filed Aug. 21, 1969, and U.S. Pat. No. 3,678,802.

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 chamber-thrust 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 temperture 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 cooling 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 dimensioning of 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 nozzles still provides considerable difficulties apart from the fact that the manufacturing methods used up to the present time are expensive.

In the patent to Stockel et al., U.S. Pat. No. 3,595,025 issued July 27, l97l, the side walls of the cooling channels are defined by planes passing through the center line of the monolithic tubular wall section, accordingly, it is not possible to vary the widths of the coolant channel to accommodate varying operating conditions.

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 nozzle of a rocket engine in which the cooling channels are easily and economically formed. Further, the dimensions of the root or base of the cooling channels and also the dimensions of the channels on the outer surface of the tubular wall section can be varied to obtain the optimum operating conditions both for heat transfer and for the closures of the coolant channels on the radially outer surface of the wall section.

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 planeswhich 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 the 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. With respect to cooling channels defined by planes extending through the center line of the combustion chamber-thrust nozzle, by displacing the center either closer to or further from the wall in which the channels are cut it is possible to maintain the crossrequired at the base of the channels, the base surfaces can be increased without appreciably altered the crosssectional areas of the channels. In cooling channels, while the differences in dimensions may be small, by displaying the center through which the side walls of the channels pass compared to the center of the tubular wall section, it is possible to achieve optimum operating conditions.

By employing the present invention the number of coolant channels can be reduced while maintaining the same total channel cross-sectional area, and, of course, the reverse is also true.

In view of the extremely high heat transfer rates involved where thecombustion chamber temperature is at about 3,000C and the pressure is very high, the variation afforded by the present invention assures optimum operating conditions.

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. I

BRIEF DESCRIPTION OF THE DRAWING In the drawing:

FIG. I 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 are transverse 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 chambenthrust nozzle;

FIG. 20 is a transverse section, similar to FIGS. 2, 2a and 2b indicating the variation in cooling channel dimensions obtained in accordance with the present invention;

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

FIG. 3c is a transverse section, similar to FIGS. 3, 3a and 3b indicating the variation in cooling channel dimensions obtained in accordance with the present invention; and

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

DETAILED DESCRIPTION OF THE INVENTION In FIG. 1 a combustion chamber thrust nozzle body I is shown containing a longitudinally extending cooling channel 2, though not shown the body 1 contains a. plurality of such cooling channels and the completed combustion chamber-thrust nozzle member is provided with an exterior covering about the body 1 which forms the outer wall of the cooling channels 2. The body 1 is formed of a monolithic tubular wall section of variable diameter along its length. One end of the body 1 is formed by a cylindrically shaped section 1a, the opposite end is formed by a diverging section lb and a converging section lc 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 10 and the diverging section 1b at which the minimum diameter of the combustion chamber exists. I

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 though mechanical stresses must also be considered. 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. However, to provide the maximum heat transfer surface it is important to increase the width of the channels adjacent the inner surface of the tubular wall section.

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 10, lb and the width y for the cylindrical section la. 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 means duct width measured at half channel height; Bisecting 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 I and disposed in parallel relationship with the adjacent cutting planes.

In the step of forming the channels, 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 h 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 1 for the depth of the channel and, as indicated in FIG. 4, the tool progresses from the position a in the cylindrical section 10 of. the combustion chamber to the position b located at the transition plane between the converging section 1c 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 i 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 necessry 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 as additional cutting operation may require one or more individual longitudinally extending cutting steps.

As indicated previously and as shown in FIG. 3, 4 and 5, the mean channel width z at the end of the diverging section lb remote from the converging section 1c 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 transition plane between the converging and diverging sections intersect along a vertex line 83.4 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 1b to its point of intersection with the converging section 1c 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 W8 is congruent with the plane of symmetry KT of the cooling channel to" be formed and then the body 1 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 i, 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 (1 corresponding generally to position b of FIG. 4 which is in the transition plane between the converging and diverging sections 10, lb. After the formation of the side wall plane V,,, the oppositie side wall of the channel located in the cutting plane V is formed in a similar mirror image fashion as indicated by FIG. 312.

Since the divergency of the cooling channels in the direction of the diverging section lb is greater than in the direction of the cylindrical section liz, 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 (I 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 a 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 c.

In FIG. 2c the cutting planes V, and V for the side walls of the channels intersect at a line Sl,2 offset outwardly from the inner wall of the tubular wall section relative to its central axis ZL. The planes R1 and R2 intersect the mid-height of the opposite side walls of the channels and intersect at the central axis ZL. It can be seen that the formation of the coolant channels by the cutting planes V, and V affords a variation in the width of the channels at their base or'root P adjacent the inner surface of the tubular wall section and at the radially outer surface 0 of the wall section through the I cross-sectional area of the channels formed by the cutting planes V, and V is approximately the same as that defined by the planes R1 and R2. The reference letter P indicates the increase in width along both sides of the base P of the channel relative to the outline formed by the planes R1 and R2 providing increased heat transfer surface, while reference letter 0 indicates a reduction in width at the outer surface 0 along both sides of the channel affording a greater area of support for the closure for the radially outer surfaces of the channels.

In FIG. 3c the line 83,4 is closer to the inner surface of the tubular wall section than the central axis ZL thereby reversing the width relationships at the base P and outer surface 0 of the cooling channels as compared to FIG. 2c. The cutting planes V and V intersecting at the line 53,4 define the side walls of the channel. The planes R3 and R4 intersecting at the line ZL also intersect the side walls of the channels at midheight. The cross-sectional area of the channels defined by the cutting planes V and V is approximately the same as defined by the planes R3 and R4.-The difference as comparedto FIG. 20 is that the base of the cooling channel isreduced by 2p and the width of the outer surface 0 is increased by 20. As a result, the cooling channels formed in accordance with the present invention afford accommodation of the thermal stresses experienced in the monolithic tubular wall section,

and, in addition, the adjustment of the surfaces at thev base and outer surface of the channels to meet the various operating conditions involved.

What is claimed is:

l. A wall member for a rocket engine combustion chamber comprising a longitudinally extending monolithic tubular wall section having a cylindrically-shaped section at one end, a diverging section at the other end, and a converging section extending between said cylindrically shaped section and said diverging section and tapering inwardly toward the axis of said tubular wall section from said cylindrically shaped section to the smaller diameter end of said diverging section, a multiplicity of separate coolant channels extending longitudinally along the outer surface of said wall section and being continuous for the axial length of said tubular wall member, said channels having a variable cross sectional area along their length in accordance with the 7 temperature conditions experienced within the rocket engine combustion chamber, the width of said flow channels being substantially uniform for the extent of said cylindrically-shaped section then diminishing to a minimum in the transition plane from said converging section to said diverging section and increasing from the transition plane to the end of said diverging section and width of the channels varying in direct relationship to the diameter of the tubular wall section in its different sections, the longitudinally extending side walls of said channels being disposed in planes parallel with and offset from radial planes extending through the longitufrom the longitudinal axis of said tubular wall section. t l

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US3190070 *Apr 5, 1950Jun 22, 1965Thiokol Chemical CorpReaction motor construction
US3267664 *Mar 19, 1963Aug 23, 1966North American Aviation IncMethod of and device for cooling
US3585800 *Jul 27, 1967Jun 22, 1971Aerojet General CoTranspiration-cooled devices
US3595025 *Jul 9, 1969Jul 27, 1971Messerschmitt Boelkow BlohmRocket engine combustion chamber
US3630449 *May 11, 1970Dec 28, 1971Us Air ForceNozzle for rocket engine
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4078604 *Apr 17, 1975Mar 14, 1978Messerschmitt-Bolkow-Blohm GmbhCooling channel surface arrangement for a heat exchanger wall construction
US4108241 *Mar 19, 1975Aug 22, 1978The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationHeat exchanger and method of making
US4245469 *Apr 23, 1979Jan 20, 1981The United States Of America As Represented By The Administrator Of The National Aeronautics And Space AdministrationHeat exchanger and method of making
US4890454 *Mar 30, 1989Jan 2, 1990Messerschmitt-Boelkow-Blohm GmbhWall surface structure having an improved radiant heat discharge capability
US4936953 *May 22, 1989Jun 26, 1990John AbbottCold trap vapor control device
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
US5832719 *Dec 18, 1995Nov 10, 1998United Technologies CorporationRocket thrust chamber
US8448335 *Dec 19, 2006May 28, 2013Volvo Aero CorporationMethod of manufacturing a wall structure and a machining tool
US8567061 *Sep 8, 2009Oct 29, 2013Eads Space Transportation GmbhCombustion chamber comprising a cooling unit and method for producing said combustion chamber
US20100058586 *Dec 19, 2006Mar 11, 2010Volvo Aero CorporationMethod of manufacturing a wall structure and a machining tool
US20100229389 *Sep 8, 2009Sep 16, 2010Eads Space And Transportation GmbhCombustion chamber comprising a cooling unit and method for producing said combustion chamber
Classifications
U.S. Classification60/260, 165/169, 60/267, 165/146
International ClassificationF02K9/97, F02K9/00, F02K9/64
Cooperative ClassificationF02K9/64, F02K9/972
European ClassificationF02K9/64, F02K9/97B