Field of the invention
The present invention relates to a method of the development of segmented mandrel for the casting of composite propellant and formulation having energetic ingredients. The invention relates to achieve desired thrust time profile for large size case bonded motors having intricate grain geometry. This invention also relates to obtain desired shaped of groove in the cured propellant without machining or drilling. More particularly, this invention relates to develop a technique by which core support can easily be taken out after curing of propellant.
Background of the invention
Composite propellant is a heterogeneous mixture of an inorganic oxidizer and a metallic fuel held in a continuous matrix of an organic polymeric binder. The ballistic behavior of any propellant grain is dependent on the shape of the cavity in the part of propellant grain. The most commonly used port cavities are, cylindrical shape, star shape, finocyl type, etc. The shape of port cavity is obtained by inserting a metallic core followed by casting and curing of propellant. After curing, the metallic core is removed to get the desired shape. However, during removal of metallic core / de-coring, it produces frictional heat, tearing or pitting of the ignition surface of propellant, undesirable separation of propellant from the case walls, and undue aggravation of already existing bond imperfection between the propellant and the case. Therefore, this practice is not safe towards large size case bonded motors containing energetic ingredients. To overcome such problems, different methods have been claimed by many researchers in the form of patents or reports to suit the present requirements.
US3001363 describes a method to develop a solid propellant rocket motor comprising a pair of hemispheroidal casing rigidly secured together to form a substantially spherical casing, an aperture formed in one of the said hemispherical casing, an internally threaded cylindrical annular plate secured to hemispherical casing to achieve spherical shape.
US3196735 relates to rocket propellant grains and methods to have developed a foam cored rocket propellant grain in-situ in a rocket motor utilizing a mandrel of any configuration fabricated from an organic foam composition.
US3345693 describes a method claimed to have developed a channeled post position and fore and aft in rocket motor, a plurality of laminar components comprising flexible star points longitudinally positioned on post by means of fixing. However, the method claimed is star shaped and does not mention about the formation of groove.
US3952627 describes a method to developed a slot former assembly for use in fabricating the solid propellant in a motor case consist of an outer hollow core, an inner tubular core and a flexible slot former retained in space by means of a cord woven thereon in a basket weave provided with a cover of flexible material.
US3357189 describes a method to developed a precast, hollow core propellant charge for a solid propellant rocket motor, where core is supported by a collapsible supporting structure and removable from the core after cure through nozzle opening of the motor.
US4233020 & US4582275 describe a method to developed the mandrel of cantilever type mounting system consist of a rotatable longitudinally slit tubular steel shell providing a molding surface in which thermosetting resin and glass fiber reinforcing materials were used in production of reinforced plastic pipe.
US5262121 relates to a method of making an elastomeric mandrel for use in formation of parts from composite materials and using reusable flexible mandrel which can be made rigid according to the present method for use as a tool for formation of composite parts.
US5310159 describes a method to developed destructible cores which permit the casting of components having at least one internal cavity open to outside only by one or more restricted orifices opposing the extraction of a core by simple displacement.
US6325958 describes a method which claims to have developed to manufacture a mandrel or other shaped core items which may be used as mould form and which further has a utility as a carriers by virtue of the porosity of the mould form. In this process, hollow ceramic microsphere are mixed together with water soluble organic binder to form billet. The billet is dried to water for its solidification.
US6101948 describes a method for manufacturing solid rocket motor comprises placing a mandrel having at least one collapsible solid slot former positioned thereon substantially centrally along the axis of rocket case, casting uncured solid rocket propellant about the mandrel and collapsible solid slot former, curing the solid propellant, removing the mandrel and pressurizing the propellant and slot former, causing the slot former to collapse, and easily removed from the cured propellant.
In view of foregoing, the inventors of the present invention conducted an extensive patent search regarding the manufacture/use of segmented mandrel suitable for composite propellant and formulation having energetic ingredients with a definite size of grooves at the center of mandrel for large size case bonded motors. However, none of these references individually or separately suggest the present invention. Accordingly, the present invention takes the advantages to fabricate a segmented mandrel of 200mm dia. with annular dia. of 500mm having length 1.2m with the help of 14 Nos. of separable fins to provide desired groove in a propellant port fitted with bolt at the middle part of cylindrical portion of mandrel.
Object of the invention:
The main object of the present invention is to develop a segmented mandrel for composite propellant and formulation containing energetic ingredients, can be easily fitted in the motor body, cast, cured and de-cored safely.
Specifically, the object of the present invention is to have a desired groove in the port of propellant with the help of segmented mandrel.
Another object of the present invention is to achieve desired groove geometry without machining and drilling.
Further object of the present invention is to have a desired thrust time profile for large size case bonded motors having intricate grain geometry.
Yet another object of the present invention is to develop a technique by which core supports can easily be taken out.
Still another object of the present invention is to de-core the mandrel without compromising with safety.
Yet another object of the present invention is to establish a technique for segmented mandrel which can be readily adapted.
Statement of the invention:
The present invention relates to a novel method/technique to fabricate segmented mandrel suitable for composite propellant and the compositions/formulations containing energetic ingredients. The method for fabrication of segmented mandrel suitable for composite propellant comprising the steps of (1) dividing segmented mandrel in cylindrical portions and fins portions having intricate grain geometry comprising:
• a cylindrical portion of mandrel made up of mild steel
• aluminum fins of trapezium shape with curved surfaces
• aluminium fins of rectangular shape
• bolts for holding fins on the surface of cylindrical portion of mandrel
(2) casting the segmented mandrel; (3) curing of said propellant; and (4) decoring the mandrel. The present invention describes a method to achieve desired shaped of groove in the middle portion of the mandrel and can provide specific thrust time curve for large size case bonded rocket motors.
Brief description of the Drawings:
Figure 1 illustrates complete setup of segmented mandrel. Figure 2 (a) illustrates fin design for segmented mandrel.
Figure 2 (b) illustrates rear portion of fin assembled with cylindrical section.
Figure 3 illustrates single fin of segmented mandrel.
Detailed Description of the invention:
The present invention describes a method for fabrication of segmented mandrel suitable for composite propellant comprising the steps of dividing segmented mandrel in cylindrical portion (101) and fins portion(102) having intricate grain geometry comprising:
• a cylindrical portion of mandrel made up of mild steel;
• aluminum fins of trapezium shape (103) with curved surfaces;
• aluminium fins of rectangular shape(104);
• bolts fixed in holes (104) for holding fins on the surface of cylindrical portion of mandrel.
Further, said fins are bolted on the outer surface of cylindrical portion of mandrel from the center. The said assembled mandrel is suitable for composite propellant and compositions containing energetic ingredients. The assembled segmented mandrel suitable for propellant composition with or without energetic ingredients is cast using gravity/ vacuum casting technique. The said propellant is then cured in water jacketed oven. Then the de-coring of the cured propellant composition along with assembled segmented mandrel is carried out by loosening the bolt inside the cylindrical portion of mandrel. The cylindrical portion and the cured propellant composition along with assembled segmented mandrel are de-cored with the help of crane. First two rectangular shaped fins are taken out first from central recess from the cured propellant composition containing fins. Thereafter, the remaining trapezium shaped fins are taken out one by one. The de-coring process does not produce frictional heat. In this process, said fins provided groove of annular shape whereas said annular groove cannot be obtained using conventional star shaped/finocyl shaped mandrel. Said annular groove is obtained without machining or drilling and also provides required thrust time profile. The desired burn surface area and mass flow rate can be achieved for large size case bonded rocket motor.
The present technique is very much suitable for HD 1.1 propellant where machining is totally prohibited.
One embodiment of present invention relates to a novel technique used to fabricate a segmented mandrel suitable for composite propellant and formulation containing energetic ingredients with desired shape of groove at the middle portion of mandrel. The fabrication method comprises following steps:
Initially, segmented mandrel is divided into two parts, i.e., cylindrical portion (101) and fin portion (102) Fig 1. The said cylindrical portion is 200±2mm in diameter and 3mm in thickness and made of mild steel. The said fin portion is trapezium shape with curved surface having thickness of 3mm and diameter of fins part is 500±2 mm. The cylindrical portion is made up of mild steel while fins (Fig 3) are made up of aluminium. The cylindrical portion of mandrel is holed at twenty eight places spherically to hold the fins. The holes are made in such a way that each fin is supported by two bolts having diameter of 8 mm both at upper side and lower side (Fig 1 (b) and Fig 2)
The fins (103) are made up of aluminum sheet having thickness of 3mm with hollow trapezium shape with inner and outer curved surfaces. Each fin (103) is provided with tapped holed (104) at two places to bolt it properly at the outer surface of cylindrical portion(lOl) of mandrel (Fig 2). The total number of fins used for the fabrication of segmented mandrel is 14 of which two fins are smaller in size having almost rectangular shape (105). The fully assembled segmented mandrel (Fig 3 a, b and 4) is kept inside the motor for further processing.
Further to this, the propellant formulation can be prepared by well known methods, i.e., the binder, plasticizer, bonding agents, antioxidants and ballistic modifiers can be blended in a vertical planetary mixer, after which inorganic fuel and inorganic oxidizer can be added in increments and mixing can be continued until homogeneity is achieved. The curing agents, cross linking agents or other additives are added at this junction and thoroughly mixer with /without vacuum and cast under vacuum to avoid any voids in the propellant grain.
Conventionally, hydroxyl terminated poly butadiene (HTPB) is used as a binder, procured from M/s Anabond Ltd, Chennai having OH value 42 mg KOH/g and viscosity at 30°c is 4000-6500 Centi Poise and number average molecular weight 2600, in the range of 10-12 % by weight. The mixing temperature is maintained at 45±2 °c to have satisfactorily low viscosity during casting.
Further, the bonding agents in this system consist of a mixture of 1, 4-butanediol and 1, 1, 1-tri methylol propane in a ratio of 2:1 and antioxidant preferred is 2 phenyl naphthyl amine (NONOX-D) are in the range of 0.1-0.15 % by weight in present invention.
Metallic powders are generally used to enhance the ballistic performance of propellant formulations. The most commonly used metallic fuel is Al, used in composite propellant formulation up to 18 % level having avg. particle size 15±3 urn with assay min 99 %.
The energetic ingredients are generally added into composite propellant formulation to enhance its energy output. The increase in energy output is due to positive heat of formation of ingredients as well as strained rings. The commonly used energetic ingredients in propellant and explosive formulation are 1,3,5- tri nitro 1,3,5- tri aza cyclo hexane (RDX), 1,3,5,7-tetra nitro 1,3,5,7-tetra aza cyclo octane (HMX) and 2,4,6,8,10,12-hexa nitro 2,4,6,8,10,12-hexa aza iso wurtzitane (HNIW or CL-20). Out of these three, RDX and HMX are preferred as it is readily available and cost effective. Although CL-20 has better performance than HMX in view of velocity of detonation, heat of formation and density but cost does not permit to use it on large scale. Therefore, for the present invention only RDX and HMX has been incorporated in the range of 10-12 % by weight.
Oxidizers, which are applicable in the solid rocket propellant formulation of this invention, are that oxygen containing solid which can readily give up oxygen and are basically inorganic perchlorates and nitrates salts. However, ammonium perchlorate is preferred oxidizer in solid propellant formulation for this invention. In the
preparation of solid propellant formulation, oxidizer is ground to the required particle size i.e 45-50 urn for better packing of the ingredients in the composition. The unground coarse ammonium perchlorate having avg. particle size 300±10 urn is used as such in the coarse / fine ratio of 3:1 in present invention. The quantity of oxidizer used in the present invention is about 65-70 % by weight in bimodal form and when energetic ingredients are incorporated, the content of oxidizer become 55-60 % by weight.
Generally hydroxyl group containing binders are cured with a curative having isocyanate groups like MDI, TDI, IPDI, HMDI etc. However, for the present invention TDI (Toluene di isocyanate) is used exclusively.
The cast compositions is cured at 50 ±2°C for 7 days in water jacketed oven. The quantity of curative used in the present invention is in the range of 0.7-0.75 % by weight.
The other process aids such as dioctyl adipate (DOA) as plasticizer is used in the range of 3-5 % by weight in the present invention to achieve desired viscosity during casting.
The following examples further illustrate the best mode contemplated for the practice of the invention.
Example-I
To a planetary mixer (Capacity 1000 I) 50.45 kg of pre-polymer resin, i.e., HTPB, 15 kg DOA as a plasticizer along with 0.5 kg of Nonox-D and 0.60 kg of bonding agents except curative were charged and the whole system was mixed well for half an hour followed by mixing under vacuum for another half an hour to drive out entrapped air. After this, 90 Kg Al powder particle size 15±3 urn was added. After complete addition of aluminum powder, it was again mixed for another 20 minutes. After this, 228.5 kg of ammonium perchlorate (avg particle size 300 urn) and 110 kg of ammonium perchlorate ((avg particle size 50 urn) were added and mixed in such a way that homogeneous mix was obtained. The overall mixing temperature was
maintained at 45±2°C. After addition of complete solid ingredients , the temperature of the mix was maintained at 40 ± 2°C. At this stage, 3.5 kg of toluene di isocyanate as a curative was added and further mixed for another 40 minutes.
The composite propellant slurry, thus, obtained was cast using gravity/ vacuum casting technique in a rocket motor equipped with segmented mandrel containing 14 fins at the center of mandrel in an annular space. . The cast grain was cured in a water jacketed oven at 50°C. for a period of 7 days. The segmented mandrel was de-cored by following steps
• Removal of bolts to loosen the cylindrical portion of mandrel
• Removal of cylindrical portion of mandrel with the help of crane
• Removal of rectangular shape of fins (2 nos) to farther loosen the remaining fin
• Removal of trapezium shaped fins one by one
• Cleaning of annular groove
Example: II
The method of processing of propellant composition is analogous to that example 1 except the incorporation of energetic ingredients such as RDX/HMX in the weight percentage of 10-12 % by replacing coarse ammonium perchlorate. The grain cured as per normal curing period. Moreover, incorporation of RDX/HMX in the composition beyond 10 % enhance the sensitivity of composition, hence it was not tried further.
Although the invention has been described with reference to specific embodiments, this description is not meant to be construed in a limiting sense. Various modifications of the disclosed embodiments, as well as alternate embodiments of the invention, will become apparent to persons skilled in the art upon reference to the description of the invention. It is therefore contemplated that such modifications can be made without departing from the spirit or scope of the present invention as defined.

We claim:
1. A method for fabrication of segmented mandrel suitable for composite propellant
comprising the steps of:
a. dividing segmented mandrel in cylindrical portions and fins portions having
intricate grain geometry comprising:
• a cylindrical portion of mandrel made up of mild steel
• aluminum fins of trapezium shape with curved surfaces
• aluminium fins of rectangular shape
• bolts for holding fins on the surface of cylindrical portion of mandrel
b. casting the segmented mandrel;
c. curing of said propellant; and
d. decoring.
2. The method for fabrication of segmented mandrel as claimed in claim 1 wherein said cylindrical portion is 200±2mm in diameter and 3mm in thickness and made of mild steel.
3. The method for fabrication of segmented mandrel as claimed in claim 1 wherein said fin portion is trapezium shape with curved surface.
4. The method for fabrication of segmented mandrel as claimed in claim 1 wherein the thickness of fins are 3mm of aluminium sheet.
5. The method for fabrication of segmented mandrel as claimed in claim 1 wherein said fins are bolted on the outer surface of cylindrical portion of mandrel from the center.
6. The method for fabrication of segmented mandrel as claimed in claim 1 wherein the diameter of fins part is 500±2 mm.
7. The method for fabrication of segmented mandrel as claimed in claim 1 wherein said assembled mandrel is suitable for composite propellant and compositions containing energetic ingredients.
8. The assembled segmented mandrel as claimed in claim 7 wherein said propellant composition with or without energetic ingredients is cast using gravity/ vacuum casting technique.
9. The cast composition using assembled segmented mandrel as claimed in claim 8 wherein said propellant is cured in water jacketed oven.
10. The cured propellant composition along with assembled segmented mandrel as claimed in claim 9 wherein said de-coring is carried out by loosening the bolt inside the cylindrical portion of mandrel.
11. The cured propellant composition along with assembled segmented mandrel as claimed in claim 10 wherein said cylindrical portion is de-cored with the help of crane.
12. The cured propellant composition containing fins as claimed in claim 11 wherein said first two rectangular shaped fins are taken out first from central recess.
13. The cured propellant composition containing fins as claimed in claim 12 wherein the remaining trapezium shaped fins are taken out one by one.
14. The fully de-cored propellant grain as claimed in claim 14 wherein the de-coring process does not produce frictional heat.
15. The fully de-cored propellant grain as claimed in 14 wherein said fins provided groove of annular shape.
16. The fully de-cored grain as claimed in claim 14 wherein said annular groove cannot be obtained using conventional star shaped/finocyl shaped mandrel.
17. The fully de-cored grain as claimed in claim 14 wherein said annular groove is obtained without machining or drilling.
18. The fully de-cored grain as claimed in claim 14 wherein said annular shape provides required thrust time profile.
19. The annular groove is obtained without machining or drilling as claimed in claim 17 wherein the desired burn surface area and mass flow rate can be achieved for large size case bonded rocket motor.
20. The annular groove obtained without machining or drilling as claimed in 19 wherein present technique is very much suitable for HD 1.1 propellant where machining is totally prohibited.