In the past years, Propulsion industries have refrained from seriously considering hybrids propellants as a viable alternative to liquid or solid because of three principal deficiencies for which hybrids have been noted. These include low combustion efficiencies, low regression rates and low volumetric loading. These three deficiencies may be attributed to the slow diffusion flames that are typically sustained along the interface between the solid fuel and gaseous oxidizer.
As a result, polymeric fuels used in hybrids are known to display week burning characteristics because they regress at about one order of magnitude more slowly than solids. To compensate for the sluggish mass flow rate of pyrolyzed fuel, large and expensive pressure cases are required. These cases are needed to hold the complex grain shapes comprising wide ports and wet surface areas that are sufficiently large to produce the desired thrust distribution. The gist: diffusion flames are no match to premixed flames unless a mechanism can be conceived to overcome their adverse characteristics.
Despite those long standing and well defined impediments, the three pronged problem affecting hybrids may not be too difficult to solve.
Cost effectiveness of hybrids may be achieved if only there is a three to four fold increase in their fuel regression rate. Such an increase would obviate the need for sophisticated grain designs by promoting major reductions in inert mass. (1)
With this in mind, an innovative vortex injection hybrid rocket engine (VIHRE) has been developed based on a cyclonic flow concept (2).
It is hoped to find widespread use for this engine in both commercial and military applications due to favourable features that it offers. One advantage of the VIHRE is its ability to produce a seven fold increase in regression rates by comparison to classical hybrids (3).
The improved performance granted by VIHRE can be attributed to its unique internal flow field that is dominated by swirling bidirectional motion .Fig1.
The corresponding coaxial, counter rotating vortex pair increases surface erosion while promoting mixing and turbulence.
Another feature of the VIHRE that sets it apart from a conventional hybrid conceptualization is the aft-end injection of the gaseous oxidizer just upstream of the nozzle (between the aft edge of the fuel grain and the inlet to the nozzle).By the alignment of the injector ports tangentially to the inner circumference, a strong vortex is produced that can travel helically along the fuel grain surface. Fuel particles swept in this manner are compelled to spiral around the chamber axis, thereby crossing the chamber length twice before exiting.
Naturally, combustion efficiency is considerably ameliorated due to the marked increase residence time and intense mixing between fuel and oxidizer. The injected oxidizer is prevented from short circuiting and draining out the nozzle by the swirl-induced radial pressure gradients that press the incoming stream against the side wall; the induced centrifugal forces cause this outer stream to cling to and climb the chamber's cylindrical wall.
In addition to the improved regression rate and combustion efficiency, the hybrid vortex engines utilize hollow, cylindrical grain cartridges that are simple to mass manufacture.
The corresponding web perforation reduces volumetric loading and precludes the need for large and therefore, expensive case housing. Another advantage is the increased effective length of the chamber owing to the convoluted path traced by fuel and oxidizer particles.
For a given thrust level the increased effective length permits an appreciable reduction in overall size and unused weight.
BACKGROUND
ROCKET PROPULSION
Propulsion is defined as the act of changing the motion of a body. Rocket Propulsion is a type of propulsion that produces thrust but ejecting stored matter, called the propellant (1). The energy source most commonly used in rocket propulsion is chemical combustion energy .The first high powered rockets developed would use this as their primary source of propulsion. Nowadays there are various other sources of propulsion which come in the form of solar radiation and nuclear reactions. Propulsion can be categorized into three major groups for ease of study; chemical propulsion, nuclear propulsion and solar propulsion.
Chemical Rocket Propulsion
It can be categorized by the type of propellants used i.e. Solid propellants, Liquid propellants, Gaseous Propellants and Hybrid Propellants.
Solid propellant rocket motor
This rocket motor consist of a solid propellant charge called the grain that on ignition burns smoothly at a predetermined rate on all exposed internal surfaces which then is forced out through the supersonic nozzle. It contains all chemical elements needed for burning. This is inherently a very powerful form of propulsion. The hot gas resulting from the burning fuel grain exit through the nozzle providing the required propulsive force known as thrust. Once ignited, the grain will burn in an orderly manner until all the propellant has been consumed.
The advantage of such a motor is that there are no feed systems, tanks or valves. This reduces the complexity of the motor.
The major disadvantage of this motor is that there is no control set in place, and it will stop only when the complete fuel grain is consumed.
Liquid Propellant rocket engines
These engines use liquid propellants that are fed from tanks to the thrust chamber. The tanks used are relatively very heavy when carrying the liquid fuel. There are 2 systems used in these type of rocket engines
Bipropellant: In this arrangement the liquid oxidizer (e.g. liquid oxygen) and the liquid fuel (e.g. kerosene) are stored separately and mixed in the thrust chamber.
Monopropellant: A single liquid containing both oxidizing and fuel compliment (e.g. Hydrazine) are burnt in the presence of a catalyst in the thrust chamber.
The liquids in the tanks are kept under high pressure and are released into the rocket thrust chamber by remote controlled valves. This concept increases the control applied upon the system. It can be prone to mechanical failures due its complexity and the weight of the rocket is increased drastically due to the additional control elements.
The thrust development of a liquid propellant engine is low for a gas pressure feed system as shown in figure ().By the addition of feed pumps the thrust can be increased, this also means a significant increase in weight.
Gas Propellant rocket engines
This type of engine also works on similar concepts used in the liquid propellant engines with just one major difference being that the propellant is converted to gaseous form before entering the thrust chamber.
Warming up the gases prior to entering the thrust chamber proves useful in increasing the thrust due to its efficient burning.
HYBRID PROPELLANT ROCKET ENGINES
Hybrid motors differ from solid and liquid rockets, by combining solid fuel with a liquid oxidiser. Hence; it's said to be a "hybrid" of a solid and a liquid rocket. The solid fuel grain is contained within the combustion chamber, and the oxidizer is fed under pressure from an oxidizer tank. The oxidiser tank can be part of the same structure, or a separate component (2).
In a typical hybrid rocket motor, the fuel is a solid and the oxidizer is a liquid. In operation the hybrid is distinctly different from solid or liquid rocket motors. It burns a macroscopic turbulent diffusion flame where the oxidizer to fuel ratio varies down the length of the chamber, the final composition determines the motor performance. The hybrid uses lesser plumbing than as seen in a liquid rocket thus providing operational flexibility as well as avoiding the explosiveness of the solid rocket motor.
By providing a valve between the oxidizer tank and the solid fuel grain, the pressurized liquid oxidizer supply can be throttled and thus controlling the burn of the solid fuel.
The only downfall in the conventional hybrid configuration is that a low volumetric loading is observed and the oxidizer residence time is short, which in turn provides an inadequate burn rate or regression rate of the solid fuel grain.
OTHER FORMS OF PROPULSION
NUCLEAR PROPULSION
Three different types of nuclear energy sources have been investigated for delivering heat to a working fluid, usually liquid hydrogen, which subsequently can be expanded in a nozzle and thus accelerated to high ejection velocities (6000 to 10,000 m/sec). However, none can be considered fully developed today and none have flown. They are the fission reactor, the radioactive isotope decay source, and the fusion reactor. All three types are basically extensions of liquid propellant rocket engines. The heating of the gas is accomplished by energy derived from transformations within the nuclei of atoms. In chemical rockets the energy is obtained from within the propellants, but in nuclear rockets the power source is usually separate from the propellant.
ELECTRIC ROCKET PROPULSION
In all electric propulsion the source of the electric power (nuclear, solar radiation receivers, or batteries) is physically separate from the mechanism that produces the thrust. This type of propulsion has been handicapped by heavy and inefficient power sources. The thrust usually is low, typically 0.005 to 1 N. In order to allow a significant increase in the vehicle velocity, it is necessary to apply the low thrust and thus a small acceleration for a long time (weeks or months).There are 3 basic types of electric propulsion, electro thermal rocket propulsion, which heats gas electrically .The other 2 types, the electrostatic or ion propulsion engine and the electromagnetic or magneto plasma engine. Both have different working principles to accomplish propulsion and thermodynamic expansion of gas in nozzle. Both will work only in a vacuum.
ADVANTAGES & DISADVANTAGES OF HYBRIDS
Safety: Since the fuel is inert thus can be manufactured, transported and handled safely.
Simplified throttling and shutdown: Since fuel flow rate automatically adjusts to the oxidizer flow rate, this can be controlled with the help of a valve. By shutting down the valve, thrust termination is achieved.
Grain Robustness: If crack exist in the grain, it's not catastrophic because it burns only where the oxidizer is present.
Propellant Variety and Low Cost: The type of propellant used can vary, dense energetic materials can be added to the fuel to increase performance or a cheap easy to manufacture material can be used.
Temperature Sensitivity: Since temperature effect on burn rate is small as in liquids the ambient launch temperature variations have little effect on operating chamber pressure.
Classical hybrids are plagued by certain disadvantages (3)
Low Regression rate: Due to a small fuel web, combustion chamber size is limited and multiple ports are needed to provide adequate burning to meet thrust requirements.
Low Bulk density: Due to low regression rate large grain surface area is needed to provide extra thrust which is generally done with multiple ports which leads to low volumetric fuel loading or bulk density.
Combustion efficiency: The large diffusion flame results in lower degree of mixing hence low impulse efficiency.
Oxidizer to Fuel ratio: It shifts with burning time which lowers the theoretical performance.
HISTORY
AIM & OBJECTIVES
METHODOLOGY
In order to successfully design a vortex hybrid, one has to overcome various hurdles that impede the development of this type of rocket motor. The objective here was to design a flight capable vortex hybrid rocket engine with a Specific impulse range of 1280Ns-2560Ns that also exhibits a high solid fuel regression rate while producing the thrust necessary for the given impulse range.
To achieve the necessary thrust and solid fuel regression rate, a bidirectional vortex would have to be created within a cylindrical chamber made of polypropylene fuel grain. In order to drive the vortex without allowing it to decay, a pressurised oxidizer distribution system is designed such that the vortex is formed at the downstream end near the nozzle inlet and moves its way up towards the head end, unlike a classical hybrid configuration that starts at the head end and moves downstream.
The entire oxidizer distribution system is sealed with o-rings, to avoid any leakages in various parts of the system and at the entrance to the cylindrical chamber. The oxidizer distribution system would be designed to withstand highly pressurized liquid nitrous flowing at a very low temperature.
Before the oxidizer enters the combustion chamber it travels through ports which augment an increase in the velocity of the liquid oxidizer as it exits into the combustion chamber. These ports have been designed such that the flow into the chamber is tangential to the chamber walls.
A hybrid rocket motor for the given impulse range comes under the classification of a K Class Rocket motor which limits the outer diameter of the motor to 54 mm. Hence the entire oxidizer distribution system and fuel had to be within this set limit.
On finalization of the design, the prototype was manufactured using commercially available and low cost materials such as Commercial grade Al Alloy, Brass and Stainless Steel.
The prototype was assembled and made air tight .Three types of tests are carried out ,firstly the a pressure test was carried out to make sure the oxidizer distribution system is completely sealed to withstand highly pressurized liquid nitrous. This is done under no chamber pressure conditions.
After which the time taken by the tank to empty out completely is measured, this will help in ascertaining the oxidizer mass flow rate by use of Bernoulli's equations. The fill stem functionality is also simultaneously tested for proper sealing and uninterrupted flow of liquid nitrous and the oxygen required for initial combustion.
Finally the load cell that is attached to the top of the tank is tested manually by applying force to get an accurate thrust reading from the picoscope.
For the Firing, the test motor is setup vertically on a vertical test rig this is done due to the nature of the fill stem mechanism being used to transport the high pressured liquid nitrous to the tank. The igniter is connected to the fill stem too such that it can ignite the oxygen that enters the chamber once the solenoid valve is switched on.
Once the experimental setup is complete the oxygen would be used to light up the nitrous as it enters the chamber resulting in a bidirectional vortex being formed. The flame would burn of the thread holding the fill stem in place and it would drop out just before the nitrous catches fire.
By analysing the rate of uniform depletion of the walls of the fuel grain with respect to the time, the solid fuel regression rate can be found.
Problems
Based on previous literature, a model of the bidirectional vortex that occurs in the combustion chamber would be constructed and studied. In order to successfully generate a bidirectional vortex in the combustion chamber a few parameters would have to be found.
Even though the bidirectional vortex hybrids have good scalability the Chamber aspect ratio would be required to define the intensity of the vortex generated .A sensitivity analysis would be done in order to get a good chamber ratio based on tests carried out in literature () for regression rates variation on different chamber sizes. Once the chamber size is decided the inner bore and length of the Fuel grain can be cut to match.
The diameter of each the port will influence the mass flow rate of the oxidizer entering the chamber. Thus theoretical mass flow rate is calculated and the port dimensions can be altered accordingly. The velocity of the swirling flow called the swirl velocity is defined by the pressure difference between the chamber and the oxidizer tank pressure and the port diameter.
LITERATURE REVIEW
MATHEMATICAL MODEL OF BIDIRECTIONAL VORTEX
The mathematical model for the bidirectional vortex engine uses the nomenclature and coordinate axes used by Majdalani ( ).The engine is modelled as a cylindrical chamber of length (Lo) and radius (a) with a closed head end called the headwall and partially open downstream end that is connected to a nozzle with a throat radius (b). At the base, the fraction of radius permits outflow given as .In order to simplify the mathematical model the field of interest is limited from the headwall to a base plane just at the nozzle entrance such that it remains incompressible. Since flow would accelerate while expanding through the nozzle.
Along the base plane, an incompressible fluid is assumed to enter the cylindrical chamber, tangentially hitting the inner surface of the chamber at a defined volumetric rate
. The tangential velocity is considered large enough to prevent flow from short circuiting, a condition by which the flow would drift towards, and out of the nozzle ().The resulting bidirectional vortex is formed is considered from cases of vortex tubes and cyclonic separators.
The sidewall injection velocity is used to capture solid fuel regression rate. The sidewall injection is considerably smaller than due to the rate of pyrolysis along the chamber wall. The flow sweeps up the propellant surface until it reaches the head end where the outer vortex switches axial polarity, reverses inwardly and continues down through the nozzle .This inner spiral also known as "mantle" is a fluid layer that separates the outer vortex from the inner vortex. The mantle is meant to rotate about the chamber axis. It demonstrates no axial translation but rather continues to spiral towards the nozzle. This Bi-directional cyclone is formed by the strong angular momentum of the incoming tangential flow. (4)
In order to get an ideal flow through the nozzle the mantle dimensions are matched to the nozzle throat.
Assumptions
A bulk gas motion, nonreactive model was used by Majdalani () with the following assumptions made to the flow.
The flow was assumed to be steady, inviscid, incompressible, rotational and axisymmetric and based on these assumptions three components of velocity were obtained the axial and radial components ().
The wall-tangential boundary layer obtained earlier by Majdalani and Chiaverini (6) .The axial and radial boundary layers are now found to be of equal size and consistent with the tangential layer. Having obtained the three components of the velocity, essential flow characteristics, such as pressure, vorticity, swirling intensity, and wall shear stresses, could be sought.
Boundary Condition Assumptions
There were 2 sets of boundary conditions used
First set were with regards to the axisymmetry and headwall impermeability and the second pertained to the inlet conservation and bulk mass conservation.
Fully tangential inflow
Zero axial flow at headwall
Zero radial flow at the centre line
Prescribed radial inflow at sidewall
Inflow matches the outflow at the base.
Boundary Conditions
Normalized Formulae
Vortex Formation
In order to prove that the inflow tangential velocity also called the swirl velocity supports vortex formation the θ-momentum equation is solved.
Where , hence:
Which confirms a vortex motion that is a characteristic of swirling inviscid flow stated by Majdalani ().
Mass Conservation
The mass balance was globally applied by Majdalani () to account for the radial inflow along the side wall. Hence the input mass flow should be equal to the flow leaving the nozzle.
Where
The parameter is used to compensate for the constantly increasing chamber diameter and is a function of the solid fuel regression rate.
By the varying of the parameters and used in Majdalani's normalized equations the mantle size can be varied and calculated.
Axial Equation
Radial Equation
Flow Characteristics
The wall injection velocity can be estimated using the solid regression rate .Since there is a presence of vortex formation in the combustion chamber the propellant is consumed at a constant rate along the circumference. The solid regression rate is discussed in later chapters. Thus when mass conservation was applied along the burning fuel grain surface
Where subscripts and stand for the solid and gaseous phases. An estimate is made with regards to the gas density at the surface of the fuel grain using the ideal gas equation of state.
The solid fuel density for Polypropylene (PP) can be calculated accordingly. The regression rate can be obtained from the upcoming sections. The sidewall velocity is meant to vary from 0.3m/s to 2.3m/s for a typical hybrid vortex.()
Mantle Matching
SOLID REGRESSION RATE
Regression Rate Behaviour in Classical Hybrids
Throughout their development Classical Hybrids have been plagued with low regression rates and inaccurate measurement thus making it harder to predict the operational performance of a hybrid rocket engine . Classical hybrids are constructed with a head end injector. In order to increase the regression rate various types of injectors such as swirl, x flow injectors.() have also been utilized. Depending on the type of oxidizer and fuel the basic expression for regression rate can me modified.
Where G is the mass flux
Regression Rate Behaviour in Vortex Hybrids
With desire to increase fuel regression rates for obtaining volumetrically efficient fuel grain designs the Orbital Technologies Corporation experimented on scaled vortex hybrid designs with the swirl oxidizer injector located on the aft end of the fuel grain. The regression rate behaviour of HTPB was initially investigated by Knuth et al ().Using gaseous oxygen as the oxidizer, several other fuel blends were tested as well. The effects of the swirl injector design and fuel port geometry on the regression rate were also examined. Basic empirical and semi empirical formulas were developed for the regression rate and heat transfer correlations, along with additional numerical simulations of the bidirectional vortex flow field. In the tests HTPB solid fuel regression rate increased by 6 fold when compared with classical hybrid solid fuel regression rates ().For this particular type of flow an empirical regression rate correlation was developed to describe regression rate as a function of mass flux,
With and
When tested at average mass fluxes of around 100 kg/m2s, the regression rates were found to be about 2.3 mm/s which when compared to a classical hybrid for the same mass flux and axial flow come to around 0.4 to 0.6 mm/s.
The empirical power of 0.54 on the mass flux shows that the bidirectional vortex will burn at about the constant rate where the fuel surface area increase should balance the solid fuel regression rate over time such that the O/F ratio of the mixture remains constant over the course of the burn.
In the case of the vortex hybrid the low dependency on is shown to have other influencing factors affecting the regression rate. According to Knuth et al () this showed that the regression rate may be dependent on the injection velocity and the fuel port geometry. Thus higher injection velocities, which induce a stronger swirl in the combustion port, would have higher regression rates. Knuth et al() developed a semi empirical regression rate and heat transfer correlation to show that highly swirling flows in the near wall region was responsible for a large increase in the convective heat transfer rates.
Where a dimensionless regression rate, and is a function of the Reynolds number based on average port diameter. takes into account the enhanced heat transfer due to tangential injection. takes into account the effects of the wall blowing, combustion and presence of the inner vortex. For the case of a bidirectional vortex hybrid the regression rates are approximately independent of , the flame zone occurring along the surface of the chamber is said to alter the velocity profile of swirling flow even in presence of blowing.
Another aspect is the L/D ratio which is the reciprocal of the regression rate; this dictates the vortex strength dissipation along the fuel port due to the viscosity of the flow field. Engines with relatively large L/D ratio form a more uniform coaxial vortex but the vortex strength had to be larger in order for it to reach the head wall.
Differences between Classical hybrids and Vortex hybrids
Nozzle analysis
Introduction
The generic function of a rocket nozzle is an efficient conversion of thermal energy in the propellant to kinetic energy; in order to obtain high thrust velocity in the desired direction. It is an integral part of any rocket motor design. During its operating phase the surface of the graphite nozzle is exposed to very high heat fluxes. These fluxes can cause an increase in the reactivity of the nozzle material which is this case is made of graphite. Due to the high temperature ranges (2500 to 3600 K) that the nozzle inlet and nozzle throat may be subjected to, a graphite nozzle was chosen.
Graphite nozzles
There are various materials available for nozzles. Nozzles made from Molybdenum are excellent for rocket applications with little or no throat erosion but are very expensive and hard to manufacture.
On the other hand nozzles made from graphite are easily available and much easier to manufacture and cheaper in comparison. It's the common choice for small scale test motors.
The only drawback with a graphite nozzle is that it has limited uses due to the phenomena of nozzle erosion that takes place at the nozzle inlet and throat after every use. The recession rate of graphite nozzle throat by thermo chemical erosion is also dependent on the diffusion rate of oxidizing species near the surface region
Erosion in graphite nozzles
There are 2 types of erosion that occur in the graphite nozzle:
Thermo chemical erosion that happens along the inlet and the throat of the nozzle .This type of erosion occurs when the increase of chamber pressure results in a substantial increase of the heat-transfer rate from the high-temperature products i.e. the ejecting propellant to the nozzle material, especially at the nozzle throat. Furthermore, once the flame temperature of the propellants exceeds 3600 K, it enhances thermal loading and chemical reactions at nozzle surfaces.
In addition to thermo chemical erosion, the graphite material can also be consumed by a mechanical erosion mechanism. This is due to the high velocity of propellant being ejected out of the Convergent-divergent nozzle.
Any throat erosion can cause severe performance reduction. Therefore, it is vitally important to study the fundamental interactions between the propellant combustion products and nozzle materials under these severe operating conditions.
When the graphite reacts at such high temperature the electrons tend to shift and rearrange the bonds. This form of carbon is called carbine and its one of the causes for the soot found inside a burnt hybrid fuel grain.
Hence the prediction of the thermo chemical erosion rate of the graphite nozzle would be of great use while designing a rocket motor which houses a graphite nozzle. This is one of the topics under current investigations and it is considered to be beyond the scope of this project.
The recent
