Lack of Controllability of Solid fuels once ignited.

Bhavesha, Sushrutia, Sanskara,Manasia and Yuvaraj S.a*

a
Department of Aeronautical Engineering, Annasaheb Dange College of Engineering
and Technology, Ashta, Sangli, 416301, Maharashtra, India.

Abstract:

Solid fuels are used for our day-to-day application to modern propulsive systems. Because of
its reliable characteristics, a solid propellant consists of several chemical ingredients such as
oxidizer, fuel, binder, plasticizer, curing agent, stabilizer, and cross-linking agent. As it is
used in very precise machines like rockets and micro propulsive systems, they are expected to
be safe to use, storable, carriable, continuously burning, mouldable, and rocket-motor
development programs. Hence, here our research is focused to study the existing solid
propellants used in rockets and micro propulsive systems and making it more feasible to
explosive organic compounds with different radicals used in solid propellants so far or other
oxidizing fractions, are also incorporated into the molecular structure for ease of its operation
in all range of altitudes.

Keywords:

Introduction:

The increasing importance of solid propellants has led to their consideration in various
applications, including long-range missiles, multiple-step rockets, satellite missiles, sounding
rockets, and armature rocketry. These propellants offer advantages such as being simpler to
handle and store than liquid propellants, and being smaller due to their high propellant
density. They are ideal for military and space applications due to their simplicity and low
cost. Standard propellant rockets provide significant thrust for their size, and they are less
expensive to develop, test and manufacture than liquid propellant boosters. They are used in
various military and civilian applications, such as satellite launchers, observational rocket
engines, and civil engineering. However, solid propellants also face disadvantages, such as
the potential for unpredictable and uncontrollable combustion when ignited, as well as
challenges in management and modification once ignited. Additionally, solid propellants burn
at a set pace and are difficult to control, unlike liquid propellants, which can be throttled and
turned off.

This paper acknowledges these problems and offers a solution to address them. By addressing
these challenges, solid propellants can continue to be a valuable tool in various applications,
including military and civilian applications.
Solid propellants are categorized into homogeneous and heterogeneous types based on their
constituent ingredients and their chemical linkage. Homogeneous propellants have a
homogeneous physical structure, consisting of single-base, double-base, and triple-base
propellants. In contrast, heterogeneous propellants have a physically mixed structure,
consisting of crystalline particles acting as oxidizers and organic plastic fuels acting as
binders. Common oxidizers include AP, AN, ADN, RDX, and HMX. Binders are either inert
(HTPB, ballistic modifiers, cross-linking agents) or active (NG and NC, polyether polymer,
and azide polymers).

Homogeneous solid propellants include single-base propellants, which are gelatinized with
ethyl alcohol as the solvent, and double-base propellants, which are one of the oldest
propellants known for their nearly smokeless exhaust. These propellants have a plasticized
gel network and physicochemical properties such as energy density, mechanical properties,
and combustion characteristics and stability depend on the proportions of NC, nitrate ester,
stabilizers, plasticizers, and other catalysts. Cast-modified double-base propellants can be
improved by adding crystalline nitramines, aluminium, or azides (GAP), which can increase
the energy density of the propellant. Composite-modified double-base propellants start with a
nitrocellulose/nitro-glycerine double-base propellant as a binder and add solids like
ammonium perchlorate and powdered aluminium. The ammonium perchlorate makes up the
oxygen deficit introduced by nitrocellulose, improving the overall specific impulse. High-
performing propellants like NEPE-75 use HMX to increase specific impulses. The mixing of
composite and double-base propellant ingredients has blurred the functional definition of
double-base propellants, with the physical structure being somewhat heterogeneous and
physicochemical properties intermediate between composite and homogeneous propellants.
Triple-base propellants can be formed by adding NQ to a double-base propellant, which
contains a high amount of hydrogen atoms within its molecular structure, lowering the
average molecular weight of the propellant combustion products. If crystalline AP, HMX, or
RDX particles are used instead, the propellant is called CMDB.

Heterogeneous solid propellants are mixtures of crystalline oxidizer particles bound within a
polymeric fuel matrix. Common oxidizers like AP and AN produce high oxygen
concentration during thermal decomposition. Hydrocarbon-based polymers like HTPB,
CTPB, and PBAN are used, with high concentrations of oxidizers providing high specific 14
impulses. Aluminium particles are often added to increase specific impulses.

Literature review:

Solid propellants have been a subject of extensive research since rocketry's invention, with
ISRO focusing on optimizing their performance and adaptability to various environments.
This study utilized fundamental solid grain configurations like bates, finocyl, and star grain.
Some solid propellants are good at burn rate and configuration, but some lack fuel, making
them unsuitable for future missions. Liquid propellants, which provide time-to-time thrust,
cannot overcome solid propellants. From 1975-2022, ISRO's fuel had a good burn rate, but
11 satellites failed due to early ignition or lack of orbit transfer over time. The root causes of
these failures were early ignition or overconsumption of fuel over time.

This literature addresses the issue of lack of controllability in solid propellants, stating that
once ignited, these fuels are uncontrollable and burn out until the end. This is a significant
issue in rocket fuels.

The problem statement led to the development of solid rocket motor fuel properties, which
are both environmentally friendly and long-lasting. The fuel ignited in solid rocket fuel can
be helpful in space, but it also poses hazards. To effectively use the fuel, it is necessary to
change the grain configuration at different stages. This research utilized basic solid grain
configurations like bates, finocyl, and star grain configurations.

Selection of Solid propellant:

Composite propellants are made up of a mixture of oxidizer, fuel, and binder. They are
widely used in solid rocket motors. The composite propellant used to address this issue of
“Lack of Controllability of Solid Fuels once Ignited” is followed by the equation: Aluminium
(Al) + Ammonium Perchlorate (AP) + Polyurethane (PU) The specification about the above-
mentioned fuel oxidizer and binder is given below.

In our report, we’ve referred to different research papers with information about solid
propellants, burning rate, grain configuration, materials used so far, and calculations. With
the help of materials used in solid propellants so far, we’ve come up with our chemical
composition which has aluminium powder, polyurethane, ammonium perchlorate
(NH4ClO4), and ferric oxide (Fe2O3). Aluminium powder is a metallic fuel in rocket
propellants, enhancing energy output and combustion temperature through rapid oxidation
with ammonium perchlorate oxygen. Polyurethane binder is a preferred fuel and binding
agent in solid rocket propellants due to its mechanical properties, low burn rates, and ease of
processing. Ammonium perchlorate (AP) is a stable, high-oxygen powder used as an oxidizer
in solid rocket propellants for combustion. Iron (III) oxide, a reddish-brown pigment,
enhances combustion in propellants by acting as a burn rate modifier and enhancing
combustion processes.

Chemical Properties Table:

Sr. No Composite Propellant Density Molecular Weight

1. Aluminium 2.7 g/cm³ 26.982 g/mol
2. Ammonium Perchlorate 1.95 g/cm³ 117.49 g/mol
3. Polyurethane 0.033 g/cm3 88.109 g/mol

Chemical composition:
The Balanced chemical reaction between Aluminium, Ammonium Perchlorate, and
Polyurethane is given below:
8 Al + 3 NH4ClO4 + 2 PU → 4 Al2O3 + 3 HCl + 3 CO2 + 2 H2O

In this reaction, Aluminium reacts with Ammonium Perchlorate and Polyurethane to produce
Aluminium Oxide (Al2o3), Hydrogen Chloride (HCl), Carbon Dioxide (CO2), and Water
(H2O). The reaction's solid by-product aluminium oxide (Al2O3) provides structural
integrity.

Aluminium Oxide chemical structure

The burning surface of a rocket propellant grain moves back or further away from a previous
position in a direction perpendicular to this burning surface. The regression rate that is
typically measured in inches per second or mm per second is termed the burning rate or burn
rate. For the different propellants used and launched so far, the burning rate is different. The
difference is because of the composition and the formulation as for every propellant the
properties are different. Taking into consideration the burn rate of the solid propellant and
how it changes according to the conditions is the basic as well as important part of the study
of solid rocket motors.

On a small scale, one can take a composition of material after its fabrication we can arrange
the propellant grain in a specific way when while building it so that the thrust will be
different throughout the flight. This will help the flight change its thrust value. Hence the
specific impulse will change by the changing thrust value. With the help of different grain
configurations varying thrust can be obtained. Each grain configuration has its properties
regarding its physical behavior. The uniqueness of the different grain configurations will help
our project to acquire controllability of solid fuels once ignited and varying thrust will be
obtained. The project is completed by us regarding theoretical research and calculation. The
calculations are done using a computer program analytical software called CEA (Chemical
Equilibrium with Applications).

CEA and Openmotor:

A] CEA:

The software is launched by NASA for the computer program chemical calculations. The
computer program computes the characteristics and chemical equilibrium compositions of
complicated mixes. Applications include shock-tube characteristics for the incident and
reflected shocks, Chapman-Jouguet detonations, assigned thermodynamic states, and
theoretical rocket performance. A computer program analytical software called CEA
(Chemical Equilibrium with Applications). Gas mixture equilibrium compositions,
thermodynamic characteristics, and transport characteristics may all be calculated using CEA.
It takes into account a variety of chemical species and enables users to input reaction
processes and other parameters to precisely simulate certain systems.

B] OpenMotor:

For the purpose of simulating rocket motor internal ballistics, which entails forecasting the
behaviour of the rocket motor based on propellant characteristics, grain geometry, and nozzle
parameters, open motor software is utilized. For those who are experimenting with rocket
motors, OpenMotor is a free, open-source internal ballistics simulator, and OpenRocket is a
potent design and simulation tool for model rockets that may also be used for internal
ballistics modelling.
Analytical study and some performance parameters of fuel and oxidizer on CEA Software:

The software calculates chemical equilibrium product concentrations from any set of reactants
and determines thermodynamic and transport properties for the resulting mixture and also
shows the performance parameters of selected Fuel and Oxidizer.

Step 1: Selection of pressure value.

In the above section of pressure select the pressure from the sea level i.e., 1 ATM.

Step 2: Selection of Fuel.

In the above selection of fuel, section select the fuel using the periodic table the selected fuel
is Al.

Step 3: Selection of Oxidize.

In the selecting section of Oxidizer select the oxidizer using the periodic table NH4CLO4.

Step 4: Set the Oxidiser/Fuel (o/f) Ratio and Exit Conditions.

In the above setting of the O/F ratio set the ratio to 5.4

In the above setting of the Pc/Pe Pressure Ratios set the ratio to 39.47 and set the values of
Subsonic Area Ratio and Supersonic Area Ratio to 10 and 20.

Step 5: Final Analysis of the performance parameters of the given data of Fuel and Oxidiser.
Calculation:

Given:
Γ(Gamma)= 1.1
Pe=101325 Pascal
Po= 4 MPa
ε= 20

Formulae:
At=Ae/ε-----------------(1)

— (2)

= At x Po/C*------------------(3)
Fmax= Ve---------------------(4)
Isp= Isp/go--------------------(5)

Solution:
From Eqn. 1.
Ae= π * (0.02/2) ^2
Ae= 3.14*10^-4 m2
Therefore,
At= Ae/ε = (3.14*10^-4) / 20
At= 1.57*10^-5 m2
Now,
= (1.57*10^-5) * (4*10^6) / 1502.9 = 4.18*10^-2 kg/sec

Ve= sqrt (2*1.11*8.314*3000/ ((1.11-1) *29.8))

Ve= 130m/sec
Thus,
Thrust= 4.18*10^-2*130
Thrust= 5432 N

Specific impulse, Isp = 2662.2/9.81

= 271.38 sec

Selection of Grain Configuration:

The formed mass of a solid-propellant rocket motor is known as the grain configuration. Grain
configurations are included in the casing as cast, moulded, or extruded bodies. Depending on
the size and form of the propellant grain, the burn time, volume of gas, and burn rate are
determined. The two types of grain storage are freestanding grains and grains placed into cases.
In the free-standing approach, the grains are produced separately, but in the bonded method,
the propellant is cast right into the casing. Grain tensions, strain, and loading need to be
analysed to ensure structural integrity. For the diverse functions of the solid rocket motor, there
are numerous sorts of grain configurations. The kinds of grain configuration utilized to address
this issue of "Lack of Controllability of Solid Fuels Once Ignited". The various grain
geometries listed below are listed after the:

1. Bates Grain Configuration:

The Bates grain configuration is a type of solid-fuel rocket motor grain geometry that
consists of one or more cylindrical grain segments with the outer surface forbidden but
the segment ends and cylindrical core allowed to burn. The Bates grain is dimensioned
to provide a flat-topped thrust curve (neutral burn), reduce the expense of propellant
characterization, and simplify data processing. This means that the burning area does
not fluctuate dramatically throughout the burn. Only the two outer ends and the centre
bore of the Bates grain burn.
2. Finocyl Grain Configuration:

The Finocyl grain configuration is a type of solid-fuel rocket motor grain geometry that
consists of a hollow cylindrical grain with a fin or 25 fins attached to the inner surface
that divides the grain into several segments that burn simultaneously to produce a high
thrust level. The Finocyl grain configuration is complex, requiring a number of
variables to define the geometry, which makes the geometrical construction and
optimization process more difficult.

3. Star Grain Configuration:

The star grain configuration provides a larger initial burn area without reducing
volumetric efficiency. Star-grain patterns improve volumetric effectiveness. Isp and
volumetric efficiency both boost total impulse. By reducing chamber pressure and
limiting flow separation in the nozzle, the star-shaped pressure vessel minimizes
pressure vessel collapse. Payload potential is increased by lowering rocket acceleration
and chamber pressure. A neutral-regressive burn can be provided using star grain. For
 best effectiveness, the neutral-regressive burn profile closely matches the ambient
pressure curve. All of these advantages enhance the effectiveness and flying
characteristics of a sounding rocket. The framework evaluates the design parameters in
this study using an original and specialized general design methodology. The results
show that the proposed framework for solid rocket engine design and development is
viable and effective.

Simulation result of grain configuration on OpenMotor software:

PROCEDURE FOR SIMULATION

1. Set the parameters for propellant which we deduced from the CEA software.

2. Set the grain geometries.

3. Apply the nozzle parameters and run the simulation. The result will display a graph of time
vs (thrust, chamber pressure, and Kn) as desired.

The above graph shows the relation between time v/s chamber pressure and thrust
as the time increases
Compared parameter:

Conclusion:

Following a comparison of the aforementioned parameters and the obtained results, it can be
said that after the solid propellant has been mixed and added grain configuration, the particular
parameters that were crucial to calculate and apply are in the desired and optimum range, which
is justified by the problem statement "Lack of Controllability of Solid Fuels Once Ignited"
above. The shift in grain configuration causes a change in thrust value and specific impulse.
As a result, the above-mentioned problem statement may be theoretically justified by
computation and the use of alternative grain arrangements. The assertion is supported by
OpenMotor software and computer programme analysis software CEA. The theoretical thrust
for various grain designs is derived by providing varied inputs and parameters. However, the
veracity of this study is purely speculative. The theoretical outcome is achieved by study,
material selection, grain configuration selection, computation, and data input.