NON-DEGASSING SOLID PROPULSION SYSTEM FOR EXTENDED SPACE MISSIONS
Field of the invention
This invention relates to a non-degassing solid propulsion system for extended space missions. In particular, the present invention relates to solid propulsion systems, which are more suitable for solid rocket motors used in the deorbiting of a spacecraft for landing and spin stabilization of spacecraft.
Background and Prior art
Non-degassing composite solid propellant, free from particulate exhaust, is an essential requirement for extended space missions and other interplanetary missions. A propellant containing an ingredient that outgases in space environment results in a weight loss and porosity in the propellant, thereby affecting its performance. This can also lead to the contamination of sophisticated electronic components on board the spacecraft. Similarly in the case of interplanetary and maimed missions, where the view ports, cameras and other sensitive analytical instruments are exposed to the exhaust fumes of the motors, the alumina particles created on firing the motors loaded with propellant containing aluminium can cause obscuration of vision, damage to the cameras as well as other instruments. Hence, a need exists in developing suitable subsystems of the solid propulsion system including a metal free propellant, low outgoing insulation and inhibition systems under vacuum levels for use in critical space missions.
Extensive studies in the recent past have been made in the process of developing varied types of subsystems of a propulsion system for outer space missions. For instance, US 7,461,503, US 7,371,784, US 7,070,705, US 6,787,586, US 6,691,505, US 5,830,384 and US 5,498,649 describe rocket motor insulation systems to be used in space missions. However, the insulation system used in all these studies is Ethylene-propylene-diene monomer (EPDM) based insulation

material, which is not suitable for outer or extended space missions because of the associated degassing problems.
US 6,265,330 discloses an insulation material composed of a glass fabric, a resin binder and a fire retardant. The resin binder preferably is composed of a phenolic resin and a Buna-N rubber, and the fire retardant is preferably alumina trihydrate. This material is also not space qualified as it contains phenolic resin as one of its ingredients, which imparts volatility.
US 5,399,599 describe a composition based on polyamide-ether thermoplastic elastomer, which also has degassing problems.
Several other studies directed towards the designing of different motor and propellant systems however indicate the presence of varied amounts of aluminium in their propellant composition. Since these systems produce alumina particles in the exhaust, they remain inappropriate for use in extended and interplanetary space missions.
It is evident form the above mentioned prior art that although there are reports on non-degassing solid propellants, an entire solid propulsion system comprising a non-degassing solid propellants. free from aluminium and plasticizer, a liner, insulation and inhibition systems suitable for outer space missions is not known in the art. The present invention described herein addresses such a propulsion system, which can overcome above said problems.
The present invention is advantageous over the prior art in many ways. Firstly, the invention uses hydroxyl terminated polybutadiene (HTPB) as a propellant, which is free from aluminium and plasticizer. Secondly, the liner system developed for bonding the propellant with refractory oxide based insulation in de-orbit solid motors has a bond strength of >7 ksc compared to the required value of 5 ksc. Thirdly, a new insulation system based on fluoroelastomers of the present invention is space qualified

and more suitable for tiny sized, spin stabilized solid motors, where the conventionally used refractory oxide insulation is practically not feasible. Finally, a carefully designed inhibition system has resolved the combustion instability problem that is otherwise common with the non-aluminized propellants.
Objective of the invention:
The principal object of the present invention is to provide a non-degassing solid propulsion system for extended and interplanetary space missions comprising a space qualified composite solid propellant, an adhesive system constituting a liner, a rocket motor chamber insulation system and an inhibition system to achieve non-degassing properties.
Another object of this invention is to provide a method of processing a solid propellant, a liner, an insulation system and an inhibition system, all of said non-degassing solid propulsion system.
A further object of this invention is to provide a solid rocket motor having said non-degassing solid propulsion system, comprising a non-aluminized and non-plasticized propellant, a liner, an insulation system and an inhibitor system.
Detailed Description of the Invention
Accordingly, the present invention provides a non-degassing solid propulsion system for extended and interplanetary space missions comprising a space qualified composite solid propellant, an adhesive system constituting a liner, a rocket motor chamber insulation system and an inhibition system to achieve non-degassing properties.

There is also provided a method of processing a solid propellant, a liner, an insulation system and an inhibition system, all of said non-degassing solid propulsion system.
Conventional solid rocket motors used in launch vehicles are not suitable in spacecrafts meant for outer space missions, as they use aluminized and plasticized propellants. Outer space missions to lunar surface or Mars require special purpose rocket motors for de-orbiting or de-boosting or spin stabilization of the spacecraft and should use non-degassing solid propulsion system comprising non-aluminised and non-plasticized propellant and suitable liner, insulator and inhibition systems.
The key requirement of the propellant is its low particle content during combustion and its ability to survive in high vacuum of the order of 10"^ to 10"''' torr for long duration and non-degassing properties. The composite solid propellant formulation can be fine-tuned to exhibit negligible weight loss and constant mechanical properties such as tensile strength, elongation and modulus under these high vacuum conditions.
The composite solid propellant composition of the present invention comprises of hydroxyl terminated polybutadiene (HTPB), ammonium perchlorate and copper chromite (0.1 to 0,5% by weight) as ballistic modifier and at least one aliphatic/alicyclic isocyanate as curing agent. HTPB is produced by free radical polymerization of butadiene using free radical initiators. Propellant composition has 70-90% by weight of solid loading containing coarse and fine ammonium perchlorate in the weight ratio of 9:1 to 1:1. Ammonium perchlorate powder is used as an oxidizer, which has a bimodal distribution. The propellant contains slow reacting curing agent, which is an aliphatic di isocyanate like isophorone diisocyanate (IPDI). Essentially, it does not contain conventional additives like plasticizers/processing aids as they lead to outgassing in outer space. The absence of plasticizer leads to processing difficulties, which are overcome by stoichiomeric tailoring of the composition and the choice of curative. The slow reacting curing agent enables better

processability in the absence of plasticizer and bonding agents. Any known diol-triol mix comprising of diols such as ethylene glycol or 1,4 butane diol and triols such as 1,1,1-trimethylol propane or glycerol in ratios ranging from 3:1 to 1:3 may be added to the composition to control the chain extension and cross-linking of HTPB.
The liner composition is carefully designed to ensure good bonding of propellant with the refractory oxide based insulation in the deorbit motor chamber. It comprises HTPB as binder, triols such as 1,1,1 tri methylol propane or glycerol in the concentration range of 0.5 to 5% by weight as cross linker and furnace black of grade N 660 as the reinforcing filler at a loading of 5 to 18% and Vanadium pentoxide.
The insulation system composition comprises of a fluoroelastomer binder, filler consisting of silica and zinc oxides, and a curating agent consisting of calcium oxide.
The inhibition system comprises a composition similar to a liner material.
The present invention also includes a process for preparing a composite solid propellant composition which comprises the steps of mixing hydroxyl terminated polybutadiene, ammonium perchlorate and adding copper chromite and at least one aliphatic diisocyanate to form a slurry and subsequently casting the said slurry. By suitable tuning of the NCO to OH ratio, a low modulus propellant for case bonding applications as well as high modulus propellant for machining freestanding grains can be formulated.
Propellant mixing operations can be carried out in conventional horizontal sigma kneader or vertical change can mixer. The sliury may then be subjected to de-aeration and cast into plastic cartons/motors at a residual pressure of 3 to 15 torr. The cast propellants are then cured at temperatures ranging from 40-70" C for duration of 5-10 days. The cured propellants are then evaluated for mechanical properties and bum rates.

One embodiment of the present invention involves the method of processing a Uner composition. Accordingly, the solid ingredients of the liner are treated to have an average particle size, less than 1 micron. These ingredients are later mixed in a stainless steel reactor equipped with planetary rotors at two different speeds under vacuum for 4 to 6 hours. The required quantity of the mixture is cured with slow curing diisocynates such as isophorone di isocyanate (IPDI), so that cure compatibility is achieved with that of propellant. The IPDI concentration can be such that the reactant stoichiometry (NCO/OH ratio) is in the range of 0.9 to 2.0. When IPDI is used as the curing agent, a polyurethane cure catalyst such as dibutyl tin dilaurate (DBTDL) or triphenyl bismuth is used, 0.02 to 0.8% by weight being the catalyst concentration. The final composition with the curing agent is diluted to a consistency suitable for application by using dichloromethane in the concentration, 28 to 37 by weight percentage. The diliJted liner composition is applied over the insulation surface in the deorbit motor chamber.
In a preferred embodiment, a premix is prepared wherein the triol cross linker is melt and dissolved in the binder hydroxyl terminated polybutadiene, at a temperature of 70-80°C and dispersed using a stainless steel reactor equipped with planetary rotors at two different speeds to which the other solid ingredients like furnace black, vanadium pentoxide and the catalyst dibutyl tin dilaurate-triphenyl bismuth are added; the said premix is degassed to remove the moisture content, cooled to ambient temperature, packed and preserved in an air tight container, prior to application in motor.
The lined motor chamber is sent to the propellant casting station for casting. Spin motors use elastomeric insulation material in the form of molded cups. Liner preparation steps for spin motor are similar to those explained above for deorbit motor.
Another embodiment of the present invention involves a method of processing an insulation system suitable for a non-degassing solid propulsion system. The insulation system is composed of an elastomeric insulation sleeve moulded from a

fluoroelastomer composition that is formulated to be free from any liquid ingredient. The formulation is initially obtained by treating the fluoroelastomer with silica, zinc oxide and calcium oxide in a two roll mixing mill with a friction ratio of 1.1 to 1.4. The formulation is further moulded into insulation cups using a hydraulic press at a pressure of 75-100 ksc and a temperature of 150-180°C for 40 to 90 minutes and finally post curing the molded cups for two to four hours at 180-220°C,
In a preferred embodiment, the composition for the elastomeric insulation is based on a copolymer of tetra fluoroethylene and hexafluoropropylene with a molecular weight in the range 120000 - 180000. It is reinforced with precipitated silica with an average particle size of less than 4 microns. The silica filler is loaded in the range 7.0 to 29.4% by weight. The composition is formulated with calcium oxide with a weight percentage of 2.1 to 8.2. The composition also contains zinc oxide powder with an average particle size less than 4 microns at a concentration range of 2 to 5 weight percentage. The sleeves are then compression moulded in a mould designed for 1.88% cure shrinkage. This is followed by a post curing operation at 180 to 220°C for 2 hours.
Yet another embodiment of the present invention relates to a method of processing an inhibition system. The process conditions and catalyst concentration are suitably modified to obtain the required inhibition material. When used for inhibition in any of the solid rocket motors, the premix is mixed with tolylene diisocyanate (TDI) curative in place of IPDI. In particular, the deorbit motors are inhibited from both the ends. The nozzle end is inhibited after the propellant is cast, as is the normal course with the current propellant motors, while the head end inhibition is done in a unique way. The head end inhibition is a pre-cast operation. The head end inhibition is a tubular solid with an internal diameter flushing with the port and a wall thickness of 4 to 6 mm. This inhibition sleeve is cast from the liner/inhibition premix detailed earlier and is matured for 72 hours prior to this positioning inside a lined motor chamber. The inhibition sleeve exterior also is

smeared with the inner material before the motor chamber goes for preservation and casting.
A further embodiment of the present invention relates to solid rocket motors having a non-degassing solid propulsion system wherein said propulsion system comprises a non-aluminised and non-plasticized propellant, a liner, an insulation and an inhibition system processed in a manner to achieve the non-degassing properties. A solid rocket motor can be of two different classes. One class of motors weighing about 1 kg is used for deorbiting of a spacecraft for landing. The second type of solid rocket motor is tiny in size and is used for spin stabilization of spacecraft. In this tiny motor, ceramic insulation laying is practically not feasible and hence a non-degassing, elastomeric insulation and the corresponding liner as well as inhibition systems are preferred.
In particular, in the case of spin motors, tiny sized tubular propellant grains are machined and are bonded into elastomeric insulation sleeves. These are then charged into the metallic chambers. The spin grains are subjected to severe most vacuum conditions compared to the de orbit motors and hence these sub systems have to be highly non-degassing.
These motors could withstand deep space environment of -160°C to +130°C under vacuum of 10'^ to 10'''* torr for more than 20 days, which was proven in Chandrayaan-1 mission.
Examples
Example 1: Properties of the propellant
The products of the present propellant and aluminized propellant combustion plume were analyzed and it was found that the present propellant formulation is

particle free, in spite of the fact that small quantities of copper chromite was added to the propellant for enhancing the burn rates.
Conventional composite said propellant with HTPB as binder and Ammonium Perchlorate (AP) as oxidizer containing plasticizers have end of mix viscosity in the range of 4500-6500 Ps. The present propellant without any plasticizer has comparable viscosities in the range 5000-6500 Ps at 50°C.
Curing takes place when the composite propellant is kept for 5 to 10 days in a hot air oven at the cure temperature in the range 40 to 70°C. Mechanical properties such as tensile strength, elongation, modulus and hardness may be adjusted by varying ratio isocyanate to hydroxyl groups (NCO/OH) of the composition in the range of 0.55 to 1.00.
The ballistic characteristics determined with respect to the cured strand burn rate (CSBR) and motor performance also remain unchanged with vacuum aging.
The outgassing properties such as total mass loss (TML) and collected volatile condensable matter (CVCM) and Water Vapor Regain (WVR) are evaluated for the propellant as per ASTM E-595 procedure under a vacuum of 10'^ to 10'* torr in the ground facility for all the components of solid propulsion system. The value for total mass loss is in the range 0.12 to 0.13 %, CVCM values in the range 0.01 to 0.02% and WVR 0.09%. The conventional solid propellants containing plasticizers have TML values of 1.394%, CVCM of 0.01% and WVR 0.021%.
The mechanical properties of the components of solid propulsion system are evaluated after a vacuum aging of 30 days in the ground facility. Table 1 shows the results of the endurance tests passed by the propellant.


Example 2: Properties of liner, inhibitions and insulation systems
The mechanical properties and the outgassing properties such as total mass loss (TML) and collected volatile condensable matter (CVCM) were evaluated for liner, inhibition and insulation systems.
Table 2 gives the results of endurance tests passed by the liner and inhibition system for extended space missions.

WE CLAIM:
1. A non-degassing solid propulsion system for extended space missions
comprising:
- a space qualified composite solid propellant,
- an adhesive system constituting a liner,
- a rocket motor chamber insulation system and
- an inhibition system to achieve non-degassing properties.

2. The solid propulsion system as claimed in claim 1, wherein said solid propellant is free from aluminium and plasticizer.
3. The soHd propulsion system as claimed in claim 1, wherein said solid propellant comprises a hydroxyl terminated polybutadiene binder, ammonium perchlorate as oxidizer, 0,1 to 0.5% by weight of copper chromite, and atleast one aliphatic/alicyclic isocyanate as curing agent.
4. The solid propulsion system as claimed in claim 3, wherein said ammonium perchlorate is a mixture of coarse and fine particles in the weight ratio of 9:1 to 1:1.
5. The solid propulsion system as claimed in claim 3, wherein a diol-triol mixture selected from ethylene glycol or 1, 4 butane diol and triols like 1,1,1 trimethylol propane or glycerol in the ratio of 3:1 to 1:3 is added as chain extension or cross linking agent to hydroxyl terminated polybutadiene.
6. The solid propulsion system as claimed in claim 1, wherein said liner and said inhibition system comprises a hydroxyl terminated polybutadiene binder, a triol cross linker (0.5 to 5%), a reinforcing filler consisting of 5 to 18% of furnace black and Vanadium pentoxide.

7. The solid propulsion system as claimed in claim 1, wherein said insulation system comprises a fluroelastomer binder, fillers consisting of silica and zinc oxide, and a curative consisting of calcium oxide.
8. The solid propulsion system as claimed in any of the preceding claims, wherein said hydroxyl terminated polybutadiene is produced by free radical polymerization of butadiene using free radical initiators.
9. The solid propulsion system as claimed in any of the preceding claims, wherein said solid propellant has viscosities in the range of 5000-6500 Ps at 50°C.
10. The solid propulsion system as claimed in claim 3, wherein said propellant is processed in a manner comprising the steps of:

- mixing of said hydroxyl terminated polybutadiene, ammonium perchlorate, copper chromite, and aliphatic diisocyanate in a mixer;
- casting said mixture into motor hardware or plastic cartons at a residual pressure of 3 to 15 torr;
- curing of said motor hardware or plastic cartons in a hot air oven at 40° - 70°C for 5-10 days;
- evaluating said cured propellants for mechanical properties and bum rates.
11. The solid propulsion system as claimed in claim 4, wherein said liner and
inhibition system is processed in a marmer comprising the steps of
- dissolving a meh of said triol linker in said hydroxyl terminated polybutadiene binder at a temperature of 70° - 80°C to form a blend;
- dispersing said blend in a stainless steel reactor;
- adding furnace black, vanadium pentoxide and dibutyl tindilaurate bismuth to the dispersed premix;
- degassing the above premix to remove the moisture content;
- cooling to ambient temperature;
- packing and preserving the cooled mixture in an air tight container.

12. The solid propulsion system as claimed in claim 7, wherein said insulation
system is processed in a manner comprising the steps of:
- Compounding the said fluroelastomer with silica, zinc oxide and calcium oxide
in a two roll mixing mill with a friction ratio of 1.1 to 1.4 to obtain a
formulation;
- moulding said formulation into insulation cups using hydraulic press at a pressure of 75-100 ksc and a temperature of 150-180°C for 40 to 90 minutes.
- Post curing the above molded cups for two to four hours at 180 - 220°C.
13. The solid propulsion system as claimed in claim 1, wherein said propulsion
system can withstand deep space environment of -160^*0 to +130°C under vacuum of
lO'^tolO-'Sorr.
14. A solid rocket motor having a non-degassing solid propulsion system wherein
said propulsion system comprises a non-aluminised and non-plasticized propellant, a
liner, an insulation and an inhibition system.
15. The solid rocket motor as claimed in claim 14, wherein said motor is designed
for deorbiting of a spacecraft for landing.
16. The solid rocket motor as claimed in claim 14, wherein said motor is designed