US3260045A - Propelling charge for solid-fuel rockets - Google Patents

Propelling charge for solid-fuel rockets Download PDF

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Publication number
US3260045A
US3260045A US3260045DA US3260045A US 3260045 A US3260045 A US 3260045A US 3260045D A US3260045D A US 3260045DA US 3260045 A US3260045 A US 3260045A
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Prior art keywords
solid
notch
propelling charge
crack
propelling
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32—Constructional parts; Details not otherwise provided for
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/10—Shape or structure of solid propellant charges
    • F02K9/18—Shape or structure of solid propellant charges of the internal-burning type having a star or like shaped internal cavity
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/26—Burning control

Definitions

  • This invention relates to a propelling charge for solidfuel rockets. It is a requirement in particular for the propelling charges of solid-fuel rockets that the rocket should exhibit a constant impact diagram or target pattern within a certain temperature range, so as to enable the projectile to be used within the required temperature range with uniformly good hitting results, independently of any temperature control system. This temperature requirement generally extends over a fairly large zone of the negative and positive temperature range.
  • This property of the powder must be taken into consideration when the firing table is drawn up.
  • This temperature dependence of the burning rate of the powder is particularly disadvantageous in rockets, since if the constriction ratio (ratio of the surface of the powder to the nozzle cross-section) is planned to be constant for the entire temperature range, there is obtained at low negative temperatures a greater burning time of the rocket which, moreover, results in a poorer specific impulse owing to the lower combustion chamber pressure developed. The firing range is thereby decreased and the dispersion of the impact diagram becomes greater.
  • This eliect is of even greater importance, for example, in the case of a double-base powder (Ngl-Nc powder) which, in certain circumstances, owing to various additions, for example of lead salts for correcting the burning rate, exhibits a zone of unsteadiness in the burning rate as a function of the pressure, it being possible for the burning rate above and below this zone to be very different. For this reason, the operating range of the rocket should be located either above or below the zone of unsteadiness.
  • the rocket is so designed that, by reason of the temperature requirement, the equilibrium pressure of the combustion chamber for a certain temperature range is above and, for the remainder, in or below, the range of unsteadiness, the drawbacks described make themselves noticeable to a particularly serious extent.
  • the rocket may become totally unusable in the temperature range in which the unsteadiness lies because of too marked a dispersion.
  • the drawbacks described can be reduced in rockets having loosely inserted internally burning propelling charges in that changes due to temperature in the physical properties of the propellant (dimensions, solidity etc.) are used to compensate changes in the chemical behavior (burning) which are likewise due to temperature.
  • the invention can be carried into effect very advantageously in that the changes in the dimensions of the propellant body are used through the coefiicient of expansion to change the construction ratio. This can be effected, for example, by changes in the dimensions of the propelling charge body varying the cross section of the nozzle directly or by way of suitable intermediate elements, whereby a change in the constriction ratio is likewise obtained.
  • changes in the dimensions of the propelling charge body which are due to temperature can also be used to build up tensile and/or bending stresses in the wall of the body which result in deliberate breaking up of the latter and thereby enlarge the effective, i.e. burning, surface, whereby the constriction ratio is likewise aflected.
  • This can be brought about in a very simple and suitable manner by providing the propelling charge body at its inner and/or outer periphery with one or more longitudinal grooves which, when the pressure increases inside the propelling charge body, promote, as predetermined breaking point or points, the formation of cracks.
  • FIGURE 1 is across-section of a rocket
  • FIGURE 2 is a partial longitudinal section thereof
  • FIGURE 3 is a cross-section of the rocket showing the propelling charge after formation of a crack
  • FIGURE 4 is a perspective view of a sealing strip.
  • the reference 1 designates the cylindrical wall casing of the combustion chamber in which is arranged the propelling charge 2 formed as an internal burner.
  • annular gap 3 Between the wall of the combustion chamber and the outer periphery of the propelling charge body 2 there is an annular gap 3, the width of which is subject to changes due to temperature, since the wall of the combustion chamber and the propelling charge body have different coeflicients of expansion.
  • the width of the annular gap 3 is so calculated that, at normal temperature at approximately 10 C. or 15 C. as examples, the propelling charge body can easily be pushed into the combustion chamber. In the upper temperature range, the wall of the propelling charge body is then applied tightly against the wall of the combustion chamber, while at falling temperatures a widening annular gap is formed.
  • the hollow propelling charge body has internally a star-shaped crosssection in order to obtain a large surface (burning area) and is provided, for example, with two opposite notches (predetermined breaking points) 5, S.
  • a further notch 6 into which there is provided an elastic sealing strip 7 with a V-shaped profile.
  • the sealing strip is pressed against the Wall of the combustion chamber by the pressure obtaining therein, the two sides of the seal spreading open and maintaining their bearing action against the propelling charge body. In this way, the interior of the propelling charge body continues to be kept under pressure and any entry of the combustion chamber pressure into the annular gap is prevented. Furthermore, it is made possible in this way to obtain an area increasing in size by stages for corresponding temperature ranges.
  • the eflect can be so controlled that in the positive temperature range above the normal temperature mentioned no crack is formed through the body bearing against the wall of the combustion chamber, one crack, for example, is formed as from a certain lower temperature, below the normal temperature mentioned and another crack at even lower temperatures through the further embrittlement and more intensive bending open of the propelling charge body.
  • FIGURES 3 and 4 better illustrate particularly the elastic sealing strip 7 which has a V-shaped profile.
  • FIG. 3 thus shows in cross-section a view similar to FIG. 1 which is a cross-section of a rocket and which illustrates the propelling charge 2 immediately after the formation of a crack in the direction from the notch 5 down into the further notch 6.
  • the sealing strip 7 is V-shaped as clearly shown in FIG. 4 which is a perspective view of a portion of the strip and tis sealing strip has been forced in FIG. 3 against the cylindrical wall casing 1 with the side portions of the strip forced against the sides of notch 6.
  • a rocket propelling means for solid-fuel motors 1.
  • a casing comprising a casing, a propellant charge of the internal burning solid type in the casing and spaced by a gap therefrom, a longitudinally extending notch provided on the outer periphery of the propellant charge, said notch being of suflicient radial depth so that a crack is formed due to combustion pressures between the inner and outer surfaces of the solid propellant charge adjacent the notch, and sealing means in the notch to seal off the outer surface of the solid propellant charge from the burn area, said crack providing increased burn area.
  • a rocket propelling means for solid-fuel rocket motors according to claim 1, in which the sealing means comprises an expansible elastic strip.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Testing Of Engines (AREA)
  • Gasket Seals (AREA)
US3260045D 1962-01-30 Propelling charge for solid-fuel rockets Expired - Lifetime US3260045A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DER31998A DE1161453B (de) 1962-01-30 1962-01-30 Feststoffraketen-Treibsatz

Publications (1)

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US3260045A true US3260045A (en) 1966-07-12

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ID=7403717

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Application Number Title Priority Date Filing Date
US3260045D Expired - Lifetime US3260045A (en) 1962-01-30 Propelling charge for solid-fuel rockets

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US (1) US3260045A (US06534493-20030318-C00166.png)
BE (1) BE627648A (US06534493-20030318-C00166.png)
CH (1) CH410527A (US06534493-20030318-C00166.png)
DE (1) DE1161453B (US06534493-20030318-C00166.png)
GB (1) GB1019043A (US06534493-20030318-C00166.png)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2235283A1 (en) * 1973-06-29 1975-01-24 Poudres & Explosifs Ste Nale Propellant blocks for rockers - with central combustion channel having unequal branches

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2500149B1 (fr) * 1981-02-17 1985-12-06 Poudres & Explosifs Ste Nale Chargement propulsif biregime a canal en trompette comportant une section en etoile
US4781117A (en) * 1987-07-20 1988-11-01 The United States Of America As Represented By The Secretary Of The Navy Fragmentable warhead of modular construction

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2957307A (en) * 1956-11-06 1960-10-25 Amcel Propulsion Inc Powder propellant rocket motors
US2957309A (en) * 1957-07-22 1960-10-25 Phillips Petroleum Co Rocket motor
GB844245A (en) * 1958-11-25 1960-08-10 Ici Ltd Improvements in or relating to rocket motors

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2235283A1 (en) * 1973-06-29 1975-01-24 Poudres & Explosifs Ste Nale Propellant blocks for rockers - with central combustion channel having unequal branches

Also Published As

Publication number Publication date
GB1019043A (en) 1966-02-02
CH410527A (de) 1966-03-31
BE627648A (US06534493-20030318-C00166.png)
DE1161453B (de) 1964-01-16

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