The document summarizes research on using paraffin wax as the solid fuel and N2O/N2O4 as the oxidizer in a hybrid rocket system. It discusses the advantages and disadvantages of hybrid systems compared to liquid and solid fuel systems. The document reviews literature on hybrid rocket regression rates and combustion characteristics. It outlines the objectives and methodology of experiments to study the regression rate, mass consumption rate, motor pressure-time history, thrust, and specific impulse of the paraffin wax/N2O hybrid system.

1. Study of Hybrid Combustion using
Paraffin Wax Fuel and N2O/N2O4 as Oxidizer
BY
Sandeep Patnala
ME/SER/1022/09
UNDER THE GUIDANCE OF
Dr. P.C. JOSHI
DEPARTMENT OF SPACE ENGINEERING AND ROCKETRY
BIRLA INSTITUTE OF TECHNOLOGY, MESRA

2. INTRODUCTION
 In Hybrid Propulsion Systems one component of the propellant is stored in
liquid/gaseous phase while the other is stored in solid phase
 Depending on the physical states of fuel and oxidizer the hybrid system is
divided in to Direct and Reverse hybrid propulsion system.
Direct Hybrid Reverse Hybrid
Oxidizer : Liquid Oxidizer : Solid
Fuel : Solid Fuel : Liquid
 The specific impulse for hybrid rockets can range from 275 - 350 sec.

3. Hybrid Rocket Configuration

4. Advantages of Hybrids
Compared to SOLIDS LIQUIDS
Simplicity -Chemically Simpler -Mechanically Simpler
Safety -Reduced Chemical Explosion -Reduced fire hazards
Hazards
-Thrust Termination and abort
possibility
Performance Related -Better Isp performance -Higher Fuel Density
-Throttling/Restart Capability -Easy inclusion of solid
performance additives
Other -Reduced environmental -Reduced mass of the liquid
impact
Cost -Reduced development costs are expected

5. Disadvantages of Hybrid Rockets:-
Hybrid rockets also exhibit some disadvantages when compared to
Liquid and Solid Rockets
 Oxidizer-to-Fuel ratio shift.
 Low regression rate.
 Uneven regression rate along the length of the grain.
The primary hazards associated with hybrids are:
 Blow back.
 Hard starts.

6. Objective
 To study the following combustion characteristics of paraffin wax and
nitrous oxide/nitrogen tetroxide hybrid system
1. Regression rate at different injection pressures
2. Mass consumption rate at different injection pressures
3. Pressure-Time history of hybrid rocket motor
4. Thrust and Specific impulse of above hybrid system

7. Methodology
 Literature review
 Design and fabrication of the rocket motor, mould with
mandrel and nozzle
 Theoretical Analysis
 Preparation of N2O4
 Experimental work
 Results

8. Literature Review
M.A.Karabeyoglu, D.Altman ---> Ref (1)
i)For several classes of liquefying fuels, the layer is hydro dynamically
unstable leading to substantial droplet entrainment from the melt
layer into the gas stream.
ii) The susceptibility of a given fuel to this shear driven instability
increases with decreasing viscosity and surface tension of the melt
layer.
iii)For practical oxidizer flux levels encountered in hybrid rocket
applications, droplet entrainment can dominate direct gasification.
iv)Several methods for increasing regression rates in hybrids were
tested.

9. Liquid Layer Combustion Theory
A) Liquid Layer Thickness Estimation :
The thickness of the energy liquid layer can be determined by the
energy transfer relations in the solid and liquid phases.
To estimate the liquid layer thickness following assumptions are made
1) Steady state regression rate of the slab
2) Physical properties of the material in both liquid and
solid phase are uniform.
3) The effect of convection in the liquid layer is ignored.
The possibility of the penetration of the thermal radiation into the fuel slab is
analyzed and several assumptions are made in the radiation treatment, which
are
1) Radiative flux is one dimensional.
2) Since the temperature levels in the slab are small, the
contribution of radiation emitted by internal material
to the radiative intensity is negligable.
3) The absorption coefficient of both liquid and solid
material behaves like a gray body.

10. Under these simplifying assumptions, the thermal analysis yields
the melt layer thickness as
-----> (1)
The thickness parameter ψ mainly depends on : thermo physical
properties and radiative characteristics of the fuel.
The characteristic thermal thickness in the liquid phase is defined as
-----> (2)
Ψ can be found as a solution of the following nonlinear equation
-----> (3)
Here the following definitions of the effective heating
parameters are introduced for convenience.
----> (3a,3b,3c)

11. Ta, Tm, Tv =Initial, Melting and Vaporization
temperatures of the fuel, respectively.
Qr and Qw =Radiative and Total heat fluxes to the fuel
surface.
Cl and Cs =Average specific heats of the liquid and solid.
Lm and Lv =latent heats for melting and vaporization.
r and rv = Total and vaporization components of the
regression rate
Another parameter that appears in the thickness expression is
the non-dimensional radiation parameter, Rl, which is defined as
the ratio of the thermal thickness to the radiative thickness in
the liquid phase.
------> (4)
where, ‘a’ is the average gray body absorption coefficient in the
liquid phase.

12. An explicit solution for the algebric nonlinear equation, eqn(3),
for the general case could not be obtained.
The two limiting cases are :
1) Rl >>1, the absorption of the radiation in the liquid layer is very large. The
thickness can be solved explicitly as
----> (5)
All the thermo physical properties of the fuel material are lumped in
the logarithmic term and the Qr/Qc ratio does not affect the thickness. This
case is important for propellants that are loaded with strongly absorbing
material such as carbon black.
2) Rl <<1, the absorption of the radiation in the liquid phase is small. The
thickness of the thermal layer in the liquid is much smaller than the radiative
thickness in the liquid and as a consequence all the radiative flux is absorbed
in the solid. Unlike the other extreme, in this case the film thickness depends
on the ratio of radiative heat flux to the convective heat flux and it can be
expressed as
• ----> (6)

13. B) Liquid Entrainment Relationships:
The entrainment component of the regression rate depends on the
parameters expressed below
According to Gater and L’Ecuyer entrainment mass transfer rate,
-------> (7)
Where rent – entrainment portion of the regression rate
Xe -- entrainment parameter
mL -- liquid flow rate per unit area with in the melt layer
In the light of the experimental findings and the results of the linear
theory, a general empirical expression for the entrainment rate of
liquid droplets in terms of the relevant properties of the hybrid motor
is,

14. Mass Transfer Mechanism

15. L.M.C.Santos , L.A.R.Almeida ---> ref (2)
I. This paper deals with the combustion of ultra-high molecular
weight polyethylene, and paraffin as the solid fuels burning with
gaseous oxygen as well as N2O as the oxidizers.
II. HTPB has been the preferred solid fuel for hybrid propulsion as the
majority of experimental work shows.
III. High density polyethylene in the pressure range of 20-40 bar
resulted in regression rates no higher than 0.3mm/sec, while
paraffin and oxygen with oxidizer mass flux ranging from 2 to
10g/cm2 s resulted in 1.0 mm/sec.

16. Tsong-Sheng Lee and Hsin-Luen Tsai ----> ref(3)
I. A series of Paraffin-HTPB based nitrous oxide hybrid rocket fuels
have been studies experimentally in a laboratory scale motor.
II. About 3 to 4 times the regression rate is increased by burning
paraffin wax as compared to regression rates of HTPB swirling O2
hybrid system as oxidizer mass flux at 90kg/ m2 sec.
III. Isp of 50p fuel can reach up to 220sec wit oxidizer mass flux of
110kg/ m2 sec. Lower regression rate of HTPB fuel resulted in a
lower Isp value and presenting fuel-lean combustion as oxidizer
mass flux went beyond 80kg/ m2 sec due to part of the oxygen
ejecting through nozzle directly without reacting with fuel vapors.

17. Alexandra Lazarav, Alon Gray ----> ref (4)
I. The paper presents an experimental investigation of a laboratory
hybrid motor employing paraffin fuel and gaseous oxidizer.
II. Several types of paraffin has been tested and it is said that paraffin
fuels have poor mechanical characteristics. This was over come by
adding polymer to paraffin .
III. Plain Paraffin :-
Some fuel was blown through the nozzle, unburned, thus reducing
the performance, while some fuel was accumulated in the motor.

18. Results of Static firing experiments with plain paraffin as fuel
Paraffin and Polymer mixture fuel :
Several ideas were proposed to improve paraffin based fuel characteristics,
performance and regression rate. Among those ideas adding nano sized aluminum particles and
carbon black to the plain paraffin improved mechanical properties. Thus improved density
specific impulse and enhanced linear regression rate.
Results of static firing experiments with paraffin wax and polymer mixture as fuel

19. Work Done
 Literature survey
 Design and modification of Combustion Chamber
 Design and fabrication of Mandrel, Case, Cover, Base Plate
 Design of Nozzle
 Evaluation of the performance parameters
 A small quantity of N2O4 has been Prepared

20. Work to be done
 Preparation of N2O4
 Preparation of fuel Grain
 Experimental work
 Results and discussion

21. Procedure for the preparation of N2O4
 Step 1 :-
NaNO3
1. NaNO3 is placed in an open pan in an oven at approx 315°C for
20minutes or at a lesser temperature for a longer period.
2. Care has to be taken to ensure that the temp remains below the
decomposition temperature of NaNO3 ( 380°C).
MnSO4
1. MnSO4 is roasted in an open pan in an oven for an hour at 315°C.
 Step 2 :-
1. The powders are placed in a separate bag and are pound into
pieces.
2. 155gm of MnSO4 and 170gm of NaNO3 are mixed thoroughly and
they are rolled to ensure proper mixing.

22.  Step 3 :-
1. The ingredients( reactants) are pound in a high temperature
resistant vessel.
2. The vessel is heated to produce vapors which are cooled using a
condenser and an ice bath to get liquid N2O4.

23. Dinitrogen Tetroxide Preparation Apparatus

24. Test Rocket Motor

25. Base Plate

26. Mandrel

27. Case

28. Cover

29. Convergent –Divergent Nozzle

30. References
 Karabeyogul, M.A., Cantwell, B.J., and Altman,D., “ Development and
Testing of Paraffin-Based Hybrid Rocket Fuels”, AIAA/SAE/ASME/ASEE,
37th Joint Propulsion Conference and Exhibit, July 8-11, 2001, Salt Lake
City, Utah, AIAA-2001-4503.
 Alexandra, L., Alon Gany “Experimental Investigation of Paraffin-Fueled
Hybrid Combustion”, Third European Combustion Meeting 2007.
 L.M.C.Santos., L.A.R.Almeida, A.M.Fraga and C.A.G.Veras
“Experimental Investigation of a Paraffin based Hybrid Rocket” .
 Karabeyoglu, M.A., Altman, D., Cantwell, B.J., Patent Pending, “High
Regression Rate Hybrid Rocket Fuels”, US Patent,1999.
 McCormick, Anthony et al.,“Design, Optimization, and Launch of a 3”
Diameter N2O/ Aluminized Paraffin Rocket”, 41st AIAA/ ASME/ SAE/
ASEE Joint Propulsion Conference & Exhibit; Tucson, AZ; USA; 10-13
July 2005. pp. 1-14-2005.
 Tsong-Sheng Lee., Hsin-Luen Tsai.,”Fuel Regression Rate in a Paraffin-
HTPB Nitrous Oxide Hybrid Rocket”, 7th Asian-Pasific Conference on
Combustion; Nation Taiwan University, Taipei, Taiwan;24-27 May2009.

31.  Martin J. Chiaverini; George C. Harting; Yen-Cherng Lu; Kenneth
K. Kuo.,” Combustion of Solid Fuel Slabs with Gaseous Oxygen in
a Hybrid Motor Analog” NASA-CR-201116.
 Chatterjee,A.k., and Joshi,P.C., “Length and Port Size Effect on
Combustion of PVC Plastisol Hybrid Fuel,” Propellant and
Explosive Journal, Vol.15, 1980, pp.163-167.
 Chatterjee,A.K., “Some Combustion Studies of PVC-O2 Hybrid
Systems”, Thesis of Masters of Science in Space Engineering and
Rocketry, Birla Institute of Technology, Mesra, Ranchi,
India,1972.
 Matthas Grosse., Bad Reichenhall., “ HERA experimental rocket
project”.
 Marxman,G.A., Wooldrige,C.E., and Muzzy,R.J., “Fundamentals
of Hybrid Boundary Layer Combustion”, Progress in Aeronautics
and Astronautics, Vol.15,AIAA,New York., 1966,pp.269-289.

32.  Martin J.Chiaverini., Kenneth K.Kuo., Aeretz., and George
C.Harting., “ Regression-Rate and Heat-Transfer Corellations
for Hybrid Rocket Combustion”, Journal Of Propulsion And
Power, Vol.17.No1, January-February2001.
 Phimon George., Krishnan.S., Varkey.P.M., Ravindran.M., and
Lalitha Ramachandran., “ Fuel Regression Rate in Hydroxyl-
Terminated-Polybutadiene/Gaseous-Oygen Hybrid Rocket
Motors”, Journal Of Propulsion And Power, Vol.17.No.1,
January-February2001.
 Rajeshwar Dayal Swami., Alon Gany., “ Analysis and Testing of
Similarity and Scale Effects in Hybrid Rocket Motor”, Science
Direct, Acta Astronautics 52(2003),pp.619-628.
 Potapkin.A.V., and Lee.T.S., “ Experimental Study of Thrust
Performance of a Hybrid Rocket Motor with Various Methods
of Oxidizer Injection”, Combustion, Explosion, and Shock
Wave. Vol.40, No.4,pp.386-392,2004.
 Michael Harris., Bruce Helming., Jacob Teufert “PANTHR
Paraffin and Nitrous-Oxide test Hybrid Rocket”,MAE
4291,2006.