US3872664A - Swirl combustor with vortex burning and mixing - Google Patents

Swirl combustor with vortex burning and mixing Download PDF

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Publication number
US3872664A
US3872664A US406771A US40677173A US3872664A US 3872664 A US3872664 A US 3872664A US 406771 A US406771 A US 406771A US 40677173 A US40677173 A US 40677173A US 3872664 A US3872664 A US 3872664A
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United States
Prior art keywords
combustion chamber
set forth
swirling
fuel
main combustion
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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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US406771A
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English (en)
Inventor
Robert P Lohmann
Stanley J Markowski
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
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United Aircraft Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Aircraft Corp filed Critical United Aircraft Corp
Priority to US406771A priority Critical patent/US3872664A/en
Priority to CA208,868A priority patent/CA1004481A/en
Priority to AU73295/74A priority patent/AU482091B2/en
Priority to IL45684A priority patent/IL45684A/en
Priority to SE7412667A priority patent/SE400348B/xx
Priority to FR7434207A priority patent/FR2247674B1/fr
Priority to CH1369274A priority patent/CH585373A5/xx
Priority to NO743677A priority patent/NO145081C/no
Priority to JP49118007A priority patent/JPS5858563B2/ja
Priority to DE2449084A priority patent/DE2449084C2/de
Priority to IT28422/74A priority patent/IT1022870B/it
Priority to GB44643/74A priority patent/GB1482145A/en
Application granted granted Critical
Publication of US3872664A publication Critical patent/US3872664A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/346—Feeding into different combustion zones for staged combustion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
    • 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

  • ABSTRACT This disclosure sets forth an engine wherein the combustion section includes a combustor with a main combustion burner having a hot fluid injected thereinto from a pilot burner. A multiplicity of swirling jets of a cooler oxidizer fluid are directed into the combustor for engaging with the hot pilot fluid. The swirling jets are directed through tubes built into the combustor wall. Fuel is directed into the main combustion burner into the area of mixed pilot hot flow and cooler swirling jet flow. This fuel can be injected to flow in mixed with the hot flow from the pilot burner or mixed with the cooler oxidizer flow through the tubes.
  • mixing of two dissimilar fluids may be substantially augmented by making the interface between said fluids unstable to centrifugal forces.
  • an unstable interface may be produced by having the two fluids flow in a concentric swirling configuration with the aerodynamic properties of the streams selected so as to have the relation p Vt outer p Vt inner be satisfied.
  • p is the density and Vt is the tangential velocity of the appropriate stream while the inner and outer refer to the radial position of the particular stream relative to the interface.
  • a combustor having a swirling flow therein has been formed employing a multiplicity of small swirling jets, rather than having a single interface ofa larger characteristic radius, which increases substantially the rapidity of the mixing and the burning process since the centrifugal force, the driving force for a rapid mixing, is inversely proportional to the radius of the interface between the fluids.
  • a combustor having a hot fluid injected therein with a multiplicity of cooler jets having engagement therewith provides a combustion device which makes it readily adaptable for replacing a combustor of a more conventional design.
  • An object of this invention is to provide a combustion device which will optimize burning conditions to reduce NOx, permit all residual reactions to go to completion greatly reducing CO and unburned hydrocarbons, and greatly reduce any trace of smoke.
  • Another object of this invention is to provide a combustion device having a pilot burner for directing hot gases into a main burner wherein cooler swirling columns of an oxidizer can be mixed therewith, with fuel being directed into the mixture.
  • Another object of this invention is to provide a combustion device wherein the fuel can be carried into the main combustion chamber by the hot gases entering from the pilot burner.
  • a further object of this invention is to direct fuel into a column of swirling air thereby providing a vaporized
  • premixed fuel-air mixture in one or more of the columns are premixed fuel-air mixture in one or more of the columns.
  • a further object of this invention is to provide a combustion device for a jet engine having two stages of inline burning wherein one stage can be used for idle operation of the engine.
  • FIG. 1 is a view of a gas turbine engine showing the location of the combustion chamber
  • FIG. 2 is an enlarged sectional view of the combustion section showing a combustion chamber therein;
  • FIG. 3 is a view taken along the line 33 of FIG. 2;
  • FIG. 4 is a modification of the rear portion of the combustion chamber of FIG. 2;
  • FIG. 5 is a modification of the front part of the combustion chamber of FIG. 2;
  • FIG. 6 is a modification of the combustion chamber
  • FIG. 7 is a view taken along the line 77 of FIG. 6;
  • FIG. 8 is another modification of the combustion chamber.
  • a turbojet engine 2 comprising a compressor section, combustion section, turbine section and exhaust section.
  • the engine 2 is of the conventional type described in greater particularity in US. Pat. No. 2,747,367.
  • a combustor 4 is shown mounted in a chamber 6 formed between an inner casing 8 and outer casing 10.
  • This chamber 6 is annular and connected at its forward end between the forward portions of the inner and outer casing to the exit of the compressor section.
  • the downstream end of the chamber 6 is connected to an annular exit duct 12 containing a plurality of turbine inlet vanes 14.
  • the inner casing 8 and outer casing 10 have annular flanges I6 and I8, respectively, which extend outwardly and inwardly. respectively to the annular exit 12 to enclose the rear portion of the annular chamber 6.
  • a combustor 4 comprises a pilot burner 20 and a main combustion burner 22.
  • the pilot burner 20 is shown as a conventional swirl stabilized burner employing an annular opening 24 around the end of a fuel nozzle 26 located in the forward end of the pilot burner, as in conventional combustor designs. Swril vanes 28 are located in the annular opening 24.
  • the swirling flow entering the pilot burner section 20 acts only to stabilize the recirculation region in the pilot combustion zone 30 and the tangential motion is essentially dissipated by the time the flow leaves the pilot burner 20 of the combustor 4.
  • the fuel nozzle 26 is of the conventional type and is connected by conduit 32 to a suitable external manifold 23 and fuel control means 25 not shown.
  • Ignition means 27 provides for lgnition of the mixture in the pilot burner section 20.
  • FIG. 3 shows the front view of one of the combustors 4 of the can type within the inner casing 8 and outer casing 10.
  • the main combustion burner 22 is shown as an extension of the pilot burner 20 and has its rearward end exhausting into the annular exit duct 12 across the turbine inlet vanes 14.
  • the forward end of the main combustion burner 22 is connected to the rearward end of the pilot burner 20.
  • the forward part of the main combustion burner 22 includes an outwardly extending funnel-shaped transition member 34.
  • the rearward part ofthe main combustion burner 22 includes a rearwardly extending transition member 36. Since this is the construction including a plurality of combustor cans the transition member 36 is formed of a plurality of circular projections 38 which blend into an annular rearward end 40 which is connected to the annular exit duct 12.
  • An intermediate substantially circular section 42 of the main combustion burner 22 connects the rearward part of the member 34 to the forward part of the member 36.
  • the members 34, 36 and section 42 are formed ofa louvered construction to provide cooling of the combustor walls. Openings 44 direct cooling air into the louvers. It is noted that this same type construction is used for the wall of the pilot burner 20. Openings 45 direct dilution air into the main combustion sections.
  • a plurality of swirl tubes 50 each containing swirl vanes 52 at the forward end thereof are arranged around the periphery of the main combustion burner 22 in the funnel-shaped transition member 34. These tubes produce a multiplicity of small swirling jets discharging into the main combustion burner 22 and into engagement with the hot pilot stream.
  • Fuel nozzle means 33 comprising a manifold 35 and nozzle 37 delivers fuel to the hot pilot stream entering the main combustion section 22.
  • a fuel control 29 delivers fuel to manifold 35 by conduit 31.
  • the pilot burner flow and swirling jet flow from tubes 50 have the flow split therethrough accomplished by the relative size of the pilot burner inlet passage 24 in relation to that of the swirl tubes 50.
  • This split would be dictated by the amount of energy required to initiate the combustion process in the swirling jet flow and would generally involve having the jets flow about 70 to 80 percent of the total air that will be vitiated while the remainder passes through the pilot burner.
  • the angular momentum of the individual jets dissipates by the time the bulk flow enters the turbine so there is no net rotation of the flow at this point either, and the swirling nature is restricted only to the immediate locality where it is employed to accelerate the combustion process.
  • approximately 20 percent of the fuel enters nozzle 26 while 80 percent enters through nozzle means 33.
  • approximately percent enters annular opening 24 approximately 30 percent enters tubes 50, approximately 30 percent enters dilution holes 45, and approximately 30 percent enters cooling holes 44.
  • FIG. 4 An alternative construction is shown in FIG. 4 which would involve introducing the secondary fuel into the main combustion air prior to its passing through the swirlers. In this way the flow entering the main combustion burner is in the form of a swirling premixed fuel-air mixture. The instability arising at the outer boundary of these jets permits the surrounding pilot gases to act as a source of igniter for the ensuing main combustion process.
  • FIG. 4 the changes from FIG. 2 are represented by the relationship between the swirl tubes 50B and the fuel nozzle means 338.
  • the swirl tubes 50B while being connected to the main combustion burner 22 in the same manner, have their forward ends curved to a point radially outwardly from, and just rearwardly of the manifold 35B of the fuel nozzle means 33.
  • the fuel nozzles 37B extend into the curved forward ends of swirl tubes 50B and are connected at their forward ends to the manifold 358.
  • the fuel nozzles 37B need not be located in each swirl tube 508. The amount of fuel desired can be placed through, for example, every other swirl tube 508.
  • FIG. 5 An alternative construction is shown in FIG. 5 where the forward part of the pilot burner is formed having the pilot fuel injected by fuel nozzles 26C into the air entering the pilot burner and where it then passes through a perforated plate flameholder 41C.
  • the flameholder 41C acted to regulate the quantity of air entering the pilot combustion zone and also stabilized the flame in the pilot section.
  • the function of the pilot burner is to generate a previtiated hot gas stream and to one skilled in the art there are other means of accomplishing this purpose within the content of this disclosure.
  • FIG. 6 shows a modification of the combustion section wherein a combustor 4A is shown mounted in a chamber 6A formed between an inner casing not shown and an outer casing 10A.
  • This chamber 6A is annular and connected at its forward end to the exit of the compressor section.
  • the downstream end of the chamber 6A is connected to an annular exit duct containing a plurality of turbine inlet vanes as shown in FIG. 2.
  • the combustor 4A comprises a pilot burner 20A and a main combustion burner 22A.
  • the pilot burner 20A is formed having an annular pilot combustion zone 30A formed between outer and inner wall members 60 and 62.
  • the outer wall member 60 is connected to and spaced from the outer casing 10A by a plurality of struts 64.
  • Inner wall 62 is connected to and spaced from a centerbody 68 by a plurality of struts 70, forming an annular passageway 72.
  • Outer wall member 60 extends forwardly in an annular passageway 66, approximately at the center thereof. formed by the wall of the centerbody 68 and outer casing 10A.
  • the wall member 60 is bent rearwardly at its forward end 61 forming a streamline flow splitter for inlet air from the compressor section.
  • the centerbody 68 has a forward opening 69 to permit the entry of inlet air from the compressor section.
  • Inner wall member 62 is spaced approximately halfway between the outer wall member 60 and the wall of the centerbody 68.
  • the forward end-of the wall member 62 is bent outwardly and rearwardly at 63 forming a smaller annular inlet passageway 65 which permits air in the annular passageway formed between the centerbody 68 and outer wall member 60 to enter the pilot combustion zone 30A.
  • the rearward end of this bent back position is tapered inwardly at 67 forming an entry passage to the flameholder 74 which is placed between outer wall member 60 and inner wall member 62 to hold the flame at that point in the pilot burner.
  • the pilot fuel is directed from a control 76 to a manifold 78.
  • the fuel is carried from manifold 78 by a plurality of conduits 32A to fuel nozzles 26A, said nozzles 26A being located in the annular passage 65.
  • An igniter 80 provides ignition at 82, just rearwardly of the flameholder 74.
  • Secondary fuel is delivered by control 83 to a manifold 35A.
  • This fuel is carried from manifold 35A by a plurality of conduits to a plurality of fuel nozzles 37A where it is directed into the pilot burner section so that it can be carried with the hot gases from the combustion zone, thereby forming a hot fuel rich mixture at the exit 84 located between the outer wall member 60 and inner wall member 62.
  • the centerbody 68 extends downstream of the ends of the outer and inner wall members 60 and 62 and the portion extending rearwardly thereof comprises louvers 86 having cooling openings 88 and a rearward opening 90 for the exit of air therefrom.
  • a center hub 92 is positioned in the opening 90 while swirl vanes 94 extend therearound. It is noted that a solid plate could be used in lieu of the hub 92 and swirl vanes 94.
  • An intermediate wall member 96 is positioned between outer wall member 60 and outer casing A, said wall 96 being spaced from outer wall 60 approximately the same distance as the wall of centerbody 68 is spaced from inner wall member 62 forming an annular passageway 97.
  • This construction also forms an annular passageway 98 between the wall member 96 and outer casing 10A and this passageway permits cooling and dilution air to pass around the main combustion burner 22A.
  • Wall member 96 extends downstream, as referred to above, to an annular exit duct, such as shown in FIG. 2, with the wall formed as louver sections with cooling holes and dilution air holes.
  • Swirl tubes 50A containing swirl vanes 52A are located at the rear portion of the annular passageways 97 and 72, a short distance from the rear end of the outer and inner wall members 60 and 62.
  • Each swirl tube 50A has its swirl vanes 52A fixed to the interior thereof and they extend inwardly to a smaller center tube 53A.
  • Each tube 50A therefore emits a swirl column of air around a straight center column of air. This construction helps maintain the swirling column of air in its columnar form for a longer period of time.
  • the swirl tubes 50A are located in pairs around the circumference of annular passageway 97 and annular passageway 72.
  • the pairs of swirl tubes 50A are spaced apart to allow chutes 102 to be placed downstream thereof and not interfere in a detrimental way with flow from the tubes.
  • the chutes are provided to divert part of the hot pilot gases into the void regions between the pairs of swirlers 50A.
  • the pairs of swirl tubes 50A in both annular passageways 72 and 97 have their vanes 52A directed so that fluid is swirled as it passes therethrough in opposite directions, that is air will be swirled in a clockwise direction through one swirl tube 50A while it will be swirled in a counterclockwise direction in the adjacent swirl tube 50A of the pair.
  • Blockage means are provided to prevent flow from passing around the swirl tubes 50A. Blocking means are shown at 104, 106 and 108. In annular passageway 72 blockage means are provided for the same purpose. They are shown at 110 and 112.
  • FIG. 8 shows a modification of the construction of FIG. 6 wherein a combustor 4D is shown mounted in chamber 6D formed between an inner casing not shown and an outer casing 10D.
  • This chamber 6D is annular and connected at its forward end to the exit of the compressor section.
  • the downstream end of the chamber 6D is connected to annular exit duct containing a plurality of turbine inlet vanes as shown in FIG. 2.
  • a combustor 4D comprises a pilot burner 20D and a main combustor burner 22D.
  • the pilot burner 20D is formed having an annular pilot combustion zone 30D formed between outer and inner wall members 60D and 62D.
  • the outer wall member 60D is connected to and spaced from the outer casing 10D by a plurality of struts 64D.
  • Inner wall 62D is connected to and spaced from a short forward centerbody 68D by a plurality of struts 70D.
  • Struts 71D connect the short centerbody to wall member 60D.
  • the pilot burner is formed in the same manner as that in FIG.
  • annular flange member 124 extends outwardly and rearwardly from the rear end of member 120 and an annular flange member 126 extends inwardly and rearwardly from the end of 122.
  • Swirlers 50D are mounted around each of the flanges 124 and 126 with the swirl tubes being directed inwardly at an angle towards each other.
  • the swirl tubes 50D are located around the flanges in much the same manner as in FIG. 7.
  • the main combustion burner 22D extends rearwardly from the outer edge of the flange 124 and a short centerbody extends rearwardly from the inward end of the flange 128.
  • This centerbody 128 is formed as the rear section ofthe centerbody 68 in FIG. 6.
  • FIGS. 6 and 8 were employed in a single annular combustor, the centerbody would not be truncated as shown but extended rearwardly to act as the inner wall of the annular inlet passage to the turbine while the outer wall of the main combustion chamber would become the outer wall of the annular inlet passage to the turbine.
  • a method of carrying out a combustion process comprising the steps of:
  • step 3 A method as set forth in claim 1 wherein the fuel is introduced in step 4 along with the hot gases of step 1.
  • step 4 A method as set forth in claim 1 wherein the fuel is introduced in step 4 along with the oxidizer of step 2.
  • a method as set forth in claim 1 including the step of:
  • step 3 the swirling columns of an oxidizer are directed from both sides of said annular combustion chamber into said hot gases.
  • step 5 the hot gases of step 1 have sufficient temperature to provide spontaneous ignition of the combustible mixture of step 4.
  • step 2 an individual column of an oxidizer if formed by swirling an oxidizer around a straight jet of oxidizer.
  • step 4 is introduced into the swirling oxidizer and straight jet of oxidizer.
  • step 1 the confined volume is a combustion chamber.
  • steps 1 and 3 the oxidizer in the swirling columns and the hot gases provide all of the oxygen for combustion in the confined volume.
  • a combustor including a main combustion chamber, a pilot combustion chamber connected at the upstream end thereof for directing a hot gas flow thereinto, means for forming and directing a plurality of swirling jets of air each swirling individually about its own axis into the upstream end of said main combustion chamber and hot gas flow so as to establish intimate contact with the hot gas flow from said pilot combustion chamber, means for directing a fuel into said main combustion chamber for mixing with the hot gas flow and intermixing swirling jets of air.
  • a combination as set forth in claim 12 wherein said means for directing a fuel into said main combustion chamber includes means for directing a fuel into said pilot combustion chamber so as to enter said hot gas flow and be carried from the pilot combustion chamber into the main combustion chamber.
  • said means for directing a fuel into said main combustion chamber includes means for directing fuel into a plurality of each of the swirling jets ofair prior to entering said main combustion chamber.
  • said means for forming and directing a plurality of swirling jets of air into said main combustion chamber comprises individual tubes fixed to said main combustion chamber adjacent its upstream end. said tubes hav' ing swirling means located therein to impart rotary motion to air passing therethrough.
  • pilot combustion chamber has an annular discharge which is concentric with the annular main com bustion chamber.
  • said main combustion chamber is formed having a diverging transition member at its forward end, said diverging transition member having a small forward opening and a larger rearward opening which is connected to the outer wall of the combustion chamber, said pilot combustion chamber being connected to the small forward opening of said diverging transition member, fixed to the diverging transition member a plurality of tubes of said main combustion chamber, said tubes having swirling means located therein.
  • pilot combustion chamber forms its hot gas flow at a temperature which will provide spontaneous ignition when said hot gas, swirling jets of air and fuel mix within the combustion chamber.
  • a combination as set forth in claim 15 wherein said means for forming and directing a plurality of swirling jets of air into said main combustion chamber comprises smaller tubes positioned within said individual tubes for forming a straight jet of air at the center of said swirling jet.
  • a combination as set forth in claim 24 wherein said means for directing a fuel into said main combustion chamber includes means for directing fuel into air upstream of said swirling means.
  • a combination as set forth in claim 12 wherein said means for forming and directing a plurality of swirling jets of air into said main combustion chamber has means to direct said swirling jets at an angle to the hot gas flow from said pilot combustion chamber as it enters into the main combustion chamber.
US406771A 1973-10-15 1973-10-15 Swirl combustor with vortex burning and mixing Expired - Lifetime US3872664A (en)

Priority Applications (12)

Application Number Priority Date Filing Date Title
US406771A US3872664A (en) 1973-10-15 1973-10-15 Swirl combustor with vortex burning and mixing
CA208,868A CA1004481A (en) 1973-10-15 1974-09-10 Swirl combustor with vortex burning and mixing
AU73295/74A AU482091B2 (en) 1973-10-15 1974-09-13 Aswirl combustor with vortex burning and mixing
IL45684A IL45684A (en) 1973-10-15 1974-09-19 Swirl combustor with vortex burning and mixing
SE7412667A SE400348B (sv) 1973-10-15 1974-10-09 Brennare
CH1369274A CH585373A5 (de) 1973-10-15 1974-10-11
FR7434207A FR2247674B1 (de) 1973-10-15 1974-10-11
NO743677A NO145081C (no) 1973-10-15 1974-10-14 Fremgangsmaate til utfoerelse av en forbrenning, og brenner for utfoerelse av fremgangsmaaten
JP49118007A JPS5858563B2 (ja) 1973-10-15 1974-10-14 カリユウネンシヨウホウホウ オヨビ ネンシヨウキ
DE2449084A DE2449084C2 (de) 1973-10-15 1974-10-15 Brennkammer
IT28422/74A IT1022870B (it) 1973-10-15 1974-10-15 Metodo per la combustione e miscelazione del vortice di flussi vorticosi e relativa camera di combustione
GB44643/74A GB1482145A (en) 1973-10-15 1974-10-15 Combustor and method of carrying out a combustion process

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US406771A US3872664A (en) 1973-10-15 1973-10-15 Swirl combustor with vortex burning and mixing

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US3872664A true US3872664A (en) 1975-03-25

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US406771A Expired - Lifetime US3872664A (en) 1973-10-15 1973-10-15 Swirl combustor with vortex burning and mixing

Country Status (11)

Country Link
US (1) US3872664A (de)
JP (1) JPS5858563B2 (de)
CA (1) CA1004481A (de)
CH (1) CH585373A5 (de)
DE (1) DE2449084C2 (de)
FR (1) FR2247674B1 (de)
GB (1) GB1482145A (de)
IL (1) IL45684A (de)
IT (1) IT1022870B (de)
NO (1) NO145081C (de)
SE (1) SE400348B (de)

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US5284019A (en) * 1990-06-12 1994-02-08 The United States Of America As Represented By The Secretary Of The Air Force Double dome, single anular combustor with daisy mixer
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AU7329574A (en) 1976-03-18
IT1022870B (it) 1978-04-20
FR2247674A1 (de) 1975-05-09
IL45684A (en) 1976-11-30
NO145081C (no) 1982-01-06
SE400348B (sv) 1978-03-20
SE7412667L (de) 1975-04-16
CH585373A5 (de) 1977-02-28
NO145081B (no) 1981-09-28
CA1004481A (en) 1977-02-01
DE2449084A1 (de) 1975-04-17
DE2449084C2 (de) 1984-09-06
NO743677L (de) 1975-06-02
JPS5858563B2 (ja) 1983-12-26
IL45684A0 (en) 1974-12-31
JPS5065708A (de) 1975-06-03
GB1482145A (en) 1977-08-03
FR2247674B1 (de) 1980-05-23

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