Analytical Study Of Paraffins C12H26 And It’sAuto-Cumbustion Process Using Decomposed Hydrogen

Peroxide

NAME OF RESEARCHER

ADDRESS:

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Abstract

This study recognizes the automatic burning sequence of paraffin with the use of

hydrogen peroxide. The rocket combustors uses the decomposed hydrogen peroxide as the

oxidizer, a liquid fuel is injected into the hot decomposition products comprising

oxygen and water vapor. The oxidizer is at a sufficiently high temperature to vaporize

and to autoburn the liquid fuel. Although the need for a separate ignition system is

eliminated with this configuration, two other issues arise: it is difficult to efficiently mix a

relatively small amount of liquid fuel into a large volumetric flow of oxidizer at the

performance-optimized mixture ratios of about eight; and the combustor design must

provide residence times sufficient for autocombustion. The latter issue typically results in

the use of high combustion chamber contraction ratios with their attendant higher weight

and surface area cooling requirements. In this study a transverse injector was used in a

dump combustor configuration, which incorporates a rearward- facing step, to

investigate the autoburning characteristics of JP-8 in decomposed hydrogen peroxide.

The goals of the investigation were to develop a greater understanding of the

autocombustion process and, if possible, develop auto-combustion model for a staged

combustor. The chamber contraction ratio was varied between three and five to

evaluate the effects of chamber gas Mach number, and the hydrogen peroxide

concentration was varied from 85 to 98% to evaluate the effects of oxidizer temperature.

Results showed that as hydrogen peroxide concentration and/or contraction ratio was

increased the fuel-rich equivalence ratio which defined the autocombustion boundary

increased as well. At a contraction ratio of 3.0, no autocombustion was achieved down to

an equivalence ratio of 1.37 using 85% hydrogen peroxide, but at 98% hydrogen

peroxide autocombustion occurred up to an equivalence ratio of 2.06. When the

contraction ratio was increased to 5.0 autocombustion was achieved at an equivalence ratio

of 1.38 using 85% hydrogen peroxide. More data is needed rega rding the effects of

pressure and decomposed gas Mach number to develop an accurate auto-combustion

model.

Introduction:

Hydrogen peroxide and kerosene rocket engines have a long history of use in

propulsion systems dating back prior to World War II.Although the performance of this

propellant combination is not as high as liquid oxygen/liquid hydrogen, LOx/LH2, or

nitrogen tetroxide/mono- methyl hydrazine, NTO/MMH, systems it is still a very

appealing option for a number of reasons. Hydrogen peroxide, H2O2, is a very versatile,

highly reactive, high density, storable, and non-toxic oxidizer. The versatility of

hydrogen peroxide is, in a way, a result of its reactivity. It can be decomposed and used

as a monopropellant for reaction control, to drive a turbine, or as a pressurant. Kerosene

based fuels such as Jet-A, JP-8, and RP-1 are very commonly used in the aviation and

rocket industries. These fuels are also storable and non-toxic. An important feature of

this propellant combination is its high density specific impulse, which is defined as the

total impulse delivered per unit volume of propellant. The density specific impulse of

H2O2/kerosene when compared to typical rocket propellant combinations is exceeded

only by the NTO/MMH system. Table below outlines performance parameters of

common rocket systems operating at similar conditions.

Spacecraft reaction control, RCS, and orbital maneuvering systems, OMS, have

typically used hydrazine and NTO/MMH rocket systems since the 1960’s. This was due to

the storability of the propellants, hydrazine’s high performance as a monopropellant, and

the fact that nitrogen tetroxide and mono- methyl hydrazine are hypergolic or ignite on

contact. These factors made hydrazine and NTO/MMH systems very simple and reliable.

However, all three propellants are toxic and corrosive while hydrazine is a carcinogen.

This creates significant safety hazards when trying to handle the propellants. As a result,

there is significant interest in developing rocket systems using non-toxic propellants to

replace hydrazine and NTO/MMH systems.There is also increasedinterest to develop

low cost, reusable satellite launch vehicles to replace the expendable vehicles currently used

in industry.Many of these expendable vehicles use toxic or cryogenic propellants, such as

liquid oxygen and hydrogen. Storable, non-toxic propellants are also preferred for these

launch vehicle applications for ease of handling on the ground.

Performance comparison of 90% H2O2/JP-8 system to common rocket propellant

combinations. Specific impulse calculated assuming a chamber pressure of 1000 psia

and equilibrium expansion to sea level pressure of 14.7 psia.

OxidizerFuel

90% H2O2

JP-8NTOMMH

LOxLH2

LOxRP-1

OptimumO/F Ratio

7.8 2.2 3.5 2.6

Characteristic Velocity,C* (ft/s)

5300 5710 7940 5890

Chamber Temperature,Tc (°F)

4600 5650 4450 6160

Specific Impulse,Isp (sec)

267 288 386 300

Density Specific Impulse,Density Isp (sec)

344 346 101 308

Hydrogen peroxide and kerosene rocket systems have the potential to replace

their toxic and cryogenic predecessors. Table above shows that both NTO/MMH and

H2O2/kerosene systems have comparable density specific impulse, density Isp, and are

both higher than cryogenic systems. This means that per unit volume of propellant a

H2O2/kerosene systems offer similar if not superior performance compared to

conventional propellant combinations. On a per unit mass basis hydrogen peroxide systems

are not quite as good performers.As a monopropellant hydrogen peroxide has a lower

specific impulse, Isp, than hydrazine and a bipropellant H2O2/kerosene system also has

lower Isp than NTO/MMH, LOx/RP-1, and LOx/LH2. However, analyses have

shown that hydrogen peroxide/kerosene systems may be the most cost effective for future

launch vehicles regardless of mass-based performance.

There are some technical issues associated with these bipropellant systems that must

be resolved to make it a viable replacement for NTO/MMH and current launch vehicle

propellants. Since NTO and MMH are hypergolic it makes the system very simple in

design, it is desired that a H2O2/kerosene system have similar simplicity as

well.Hydrogen peroxide and kerosene are not hypergolic by themselves, and there is

research being done make these propellants ignite on contact.Alternatively, an

H2O2/kerosene engine can operate in a staged configuration. In this configuration the

hydrogen peroxide is decomposed in a catalyst bed and the kerosene fuel is injected into

the hot decomposed gases. If conditions are correct the oxidizer/fuel mixture can

autoignite eliminating the need for a complex ignition system. However, autocombustion

is dependent on a number of different factors such as fuel injector design, hydrogen

peroxide concentration, decomposed gas velocity, chamber pressure, and mixture

ratio. A better understanding of the autocombustion process in these staged

H2O2/kerosene rocket engines is required. The goals of the research described in this

thesis include; outlining a design method for a staged engine injector, generating

experimental data on autocombustion under varying engine operating conditions,

and creating a model to aid in the prediction of autocombustion. Results of this research

may make the staged-bipropellant H2O2/kerosene rocket a lighter, more reliable, and

higher performing engine in the future.

Hydrogen peroxide

Hydrogen peroxide is an inherently unstable chemical compound that

exothermically reacts, or decomposes, into hot oxygen gas and water vapor. Hydrogen

peroxide is miscible in water and is commercially manufactured as an aqueous solution in

a variety of concentrations. Concentrations are usually designated as percent H2O2 by

weight of solution. Propellant-grade H2O2, or HTP, is greater than 70% in concentration

and most modern engines tend to use 85, 90, or 98%. The decomposition rate of

propellant-grade H2O2 is less than 0.1% per year over normal atmospheric temperature

and pressure ranges.Decomposition is significantly accelerated as the temperature of the

H2O2 and/or its environment is increased and/or when the liquid is in contact with certain

materials or contaminants. These factors can potentially cause a chain reaction of

decomposition since the heat released during a reaction can provide the energy necessary

to decompose the surrounding H2O2 and so on. This is a very dangerous situation in most

cases, however, when controlled it can be advantageous quality.Table b e l o w outlines

the variation in physical and decomposition properties of hydrogen peroxide with

concentration.

Table : Properties of liquid and decomposed hydrogen peroxide with concentration.

Concentration 70 % H2O2 80 % H2O2 90 % H2O2 98 % H2O2Liquid Properties (@ STP)

Molecular Weight 26.86 28.89 31.29 33.42Specific Gravity 1.283 1.333 1.387 1.432Boiling Point (F) 257 -- 287 299

Vapor Pressure (psia) 0.137 -- 0.065 0.045Heat Capacity (Btu/lbm- R) 0.738 -- 0.663 0.633

Decomposed Gas PropertiesTemperature (F) 504 952 1393 1746

Molecular Weight 21.04 21.56 22.11 22.56Specific Heat Ratio 1.315 1.287 1.265 1.251Mass Fraction O2 0.341 0.376 0.423 0.461

Mass Fraction H2O 0.659 0.624 0.577 0.539

Hydrogen peroxide has a low vapor pressure, as Table above shows, on the order

of one-tenth of a psi. This is significantly lower than the vapor pressure of other

common oxidizers such as liquid oxygen, 735 psia at -193 °F, and nitrogen tetroxide,

110 psia at 160 °F. It is advantageous to use a propellant with a low vapor pressure

in rocket

system for several reasons. In turbo-pump systems the propellant can be fed to the

pumps at a low pressure without risking cavitation. In addition, only a low absolute

pressure is required in the propellant tank to prevent the liquid from vaporizing. As a

result, use of a propellant with a low vapor pressure leads to low tank and system

pressures which reduce tank and system mass. Another attractive feature of hydrogen

peroxide is its high heat capacity, 0.66 Btu/lbm- R for 90% H2O2 as shown in Table

above This is comparable to the heat capacity of water, 1.0 Btu/lbm- R, which is

considered a very good coolant and is used for a number of applications. The high heat

capacity of hydrogen peroxide suggests that it would be an excellent coolant for a rocket

system.

Hydrogen peroxide also possesses the properties of a storable propellant. It is a

stable liquid over a reasonable range of temperature and pressure, and it is sufficiently

non-reactive with tank material, when properly passivated, for significant lengths of time,

although the concentration will gradually decrease. It is considered to be a non-toxic

propellant as well. Toxic propellants are poisonous to humans through inhalation or

contact with the body tissue. However, hydrogen peroxide solutions and vapors are

irritating to body tissue. Solutions can cause skin burns and vapors can inflame the

respiratory tract, however, exposure is only lethal in extremely high doses especially

through ingestion.

The most important feature of hydrogen peroxide as propellant is its reactivity.

The hot gases produced when H2O2 is decomposed contain a significant amount of

energy, see Table above .This makes hydrogen peroxide an excellent monopropellant.

Monopropellant thrusters are typically used for low thrust applications such as reaction

control systems (RCS). Hydrogen peroxide of 85 or 90% conc entration has been used in

RCS systems in the past, such as the Mercury space capsule, and new systems using

H2O2 are currently in development. These gases can also be expanded through a row of

turbine blades imparting its energy to generate turbine rotation. This is important

since many rocket systems use turbo-pumps driven by turbines to feed propellants to the

combustion chamber. Decomposed hydrogen peroxide could potentially be used as a

tank pressurant as well. The reactivity of H2O2 also makes it a versatile propellant that

can be used for a number of different propulsion systems. Using hydrogen peroxide as an

oxidizer in a bipropellant main engine as well as a monopropellant for reaction control

and turbine power eliminates the need for separate systems for each of these applications.

This greatly simplifies the overall propulsion system design.

Purpose of the study:

o To investigate the auto-combustion characteristics of paraffin-based JP-8 fuel in

decomposed hydrogen peroxide.

o To perform testing using staged-bipropellant rocket engine in a dump combustor

configuration.

o A fuel jet trajectory analysis was performed during injector design to model jet

breakup and fuel distribution in the oxidizer port and to prevent jet impingement.

o Downstream of the injection point a rearward- facing step was used to provide flame

stabilization at the entrance to the combustion chamber.

o Testing was structured to study the affects of gas temperature, gas velocity, and

equivalence ratio on autocombustion.

o To perform the test at three stsges.

1)Strong Autocombustion

2)Week Autocombustion

3)No Autocombustion

o To analyze the test cases by collecting the data from Mestre and Ducourneau and

Walder.

o To check that the variations in both the pressure and equivalence ratio combined to

influence the autocombustion temperature at a constant contraction ratio not just one

or the other.

o However, the test conditions were such that the variations could not be isolated from

one another.

o To investigated the effect of gas velocity on autocombustion.

o Changes in contraction ratio will also affected the trajectory of the fuel jet in the

oxidizer port or not.

o As the contraction ratio was increased the shear layer residence time created by

the rearward- facing step was increased as well or not .

o Mainly to proves that the stable autocombustion is possible at contraction ratios as

low as 3.0 and equivalence ratios less than 1.4 using 90% H2O2.

METHODOLOGY:

The test conducted in various phases to determine the actual performance using hydrogen

peroxide in monopropellant and bipropellant system.

1. Test Facility verview

2. Cavitating Venturi Flow Control

3. Data Acquisition and Control

4. Instrumentatio

5. Test Article and

etup.

6. Hydrogen Peroxide Dilution.

7. Pressure

Budget

8. Firing Sequence

Test Procedure

Four people are required to conduct a rocket test at APCL, and each of these people

has a specific set of responsibilities during the test. The test conductor is responsible for all

test operations, reads the test procedures, and maintains the list of test conditions. The test

operator loads propellants, operates manual valves and regulators, and performs other

functions related to propellant or pressurant as dictated by the test conductor. The data

system operator runs the LabVIEW program, monitors and records test data, and

maintains operability of all instrumentation and controls for each test. The site safety

director maintains functionality of safety equipment, keeps site clear of unauthorized

personnel, and ensures that test personnel follow safety procedures at all times.

Conclusion:

A total of 24 tests were conducted to investigate the autocombustion characteristics

of kerosene-based JP-8 fuel in decomposed hydrogen peroxide.Testing was performed

using staged-bipropellant rocket engine in a dump combustor configuration.The engine

used a catalyst bed to decompose the hydrogen peroxide and a transverse injector to

inject the JP-8 into the decomposed gas stream. A fuel jet trajectory analysis was

performed during injector design to model jet breakup and fuel distribution in the

oxidizer port and to prevent jet impingement. Downstream of the injection point a

rearward- facing step was used to provide flame stabilization at the entrance to the

combustion chamber.This design is commonly referred to as a dump combustor

configuration.Testing was structured to study the affects of gas temperature, gas

velocity, and equivalence ratio on autocombustion.

Each test was classified into one of three groups based on visual observations and

measured chamber pressure data.Tests classified as strong autocombustion produced a

stable, red-orange flame at the nozzle exit and bipropellant C* efficiencies of greater than

90%. The chamber pressure in these tests rose sharply within one-tenth of a second

following the initiation of fuel flow. The second classification, weak autocombustion,

was typified by highly unstable flames that varied in color from red-orange to green. The

C* efficiencies for these tests ranged from 65 to 90%. The delay between fuel initiation

and chamber pressure rise was on the same order as the strong autocombustion case, but the

rise was not as sharp. The third test classification was no autocombustion. During these

tests the fuel was vaporized in the chamber but did not autoignite producing either a thick,

white vapor cloud or a clear plume at the nozzle exit. The biprop C* efficiencies for these

tests were less than 65% in most cases and the chamber pressure rise was minimal

resulting from the vaporization of the fuel.

It was determined that severe chamber pressure instabilities present during weak

autocombustion tests caused the unstable flame structure. The instability caused the

pressure in the chamber to oscillate, and in some cases the average amplitude of the

oscillation was almost 30% of the chamber pressure. It is believed that the fuel and

decomposed gas mixture was initially autoignited at a high pressure point in the oscillation

producing a bright, red-orange flame. As the chamber pressure fell to a low point in the

oscillation it most likely altered the path of the combustion reaction causing a change in the

color and possibly the temperature of the flame.In some cases the pressure may have

fell far enough to quench the flame completely. This was seen during some tests when a

flame was visible at one instant and then a vapor cloud the next instant. As the pressure

continued to oscillate after the initial point of autocombustion so too did the flame.

The gas temperature, gas velocity, and equivalence ratio were controlled during

testing by varying the H2O2 concentration, chamber contraction ratio, and oxidizer mass

flow rate respectively. Each test series was set up such that the concentration and

contraction ratio remained constant while the equivalence ratio was varied to determine the

boundary between strong autocombustion and no autocombustion for fuel rich

conditions. Once the boundary was determined at a particular concentration it was

increased for the next test series and the process was repeated again. Results showed that

as the concentration, or decomposition temperature, was increased the equivalence ratio at

the boundary between strong and no autocombustion increased as well. At a contraction

ratio of 3.0 and a concentration of 85% H2O2 no autocombustion was achieved

down to an equivalence ratio of 1.37 while at a concentration of 98% strong

autocombustion was achieved up to an equivalence ratio of 2.06. This trend agrees with

past data from Mestre and Ducourneau for kerosene in air as well as Walder for kerosene

in decomposed hydrogen peroxide. Both show that higher temperatures are required for

autocombustion as equivalence ratio increases, or as the mixture becomes more fuel rich,

for a specified mixture residence time. Other studies done with kerosene fuel in air at

oxidizer rich equivalence ratios suggest that equivale nce ratio plays a negligible role in

autocombustion.

However, due to the fact that the oxidizer flow rate was varied to change

equivalence ratio the monoprop chamber pressure was altered as well.There is

nearlyuniversal agreement from past autocombustion studies that the autocombustion

temperature decreases with increasing pressure at a fixed residence time or contraction ratio

in this case. Data from Mestre and Ducourneau suggests that a pressure increase from

100 to 115 psia can alter the autocombustion temperature of kerosene fuel in air by

approximately 90°F at an equivalence ratio of 2.0. Data from Walder suggests a

temperature difference of about 20°F for kerosene fuel in decomposed hydrogen peroxide

at the same pressures and a stoichiometric equivalence ratio. It is believed that the

variations in both the pressure and equivalence ratio combined to influence the

autocombustion temperature at a constant contraction ratio not just one or the other.

However, the test conditions were such that the variations could not be isolated from one

another.

The effect of gas velocity on autocombustion was investigated by varying the

contraction ratio of the engine. Increasing the contraction ratio decreases the Mach

number of the gases in the chamber as well as the decomposed gas in the oxidizer port.

As previously discussed, at contraction ratio of 3.0 no autocombustion was achieved at all

the tested conditions using 85% H2O2. At this contraction ratio the Mach number in the

chamber is about 0.20 and 0.45 in the oxidizer port.When the contraction ratio was

increased to 5.0, however, strong autocombustion was achieved at an equivalence ratio

of 1.36 and weak autocombustion was achieved up to an equivalence ratio of 2.31. At

this contraction ratio the Mach number in the chamber is about 0.12 and 0.20 in the

oxidizer port. Therefore, as the contraction ratio is increased, or the gas velocity decreased,

the temperature required to achieve autocombustion at a particular equivalence ratio

decreases. This result also agrees with past data from Mestre and Ducourneau with regards

to residence time of a kerosene/air mixture. Walder made a similar conclusion and chose

to correlate the temperature decrease with the characteristic length of a rocket combustion

chamber instead of residence time. Intuitively this result makes sense because as the

available reaction time of the mixture increases the probability of autocombustion should

increase as well at a certain temperature. Both studies suggest that the affect of residence

time, characteristic length, or gas velocity on the autocombustion temperature is only

significant up to a point after which the effects are negligible. Since the fuel flow rate

was kept constant during contraction ratio variations as well the chamber pressure

increased with increasing contraction ratio. Based on the previous discussion on pressure,

it is believed that pressure affects also contributed to the decrease in

autocombustion temperature at a larger contraction ratio.

Changes in contraction ratio also affected the trajectory of the fuel jet in the

oxidizer port. Fuel trajectory analysis was performed using momentum ratios calculated

from measured test data. All of the calculated trajectories at a contraction ratio of 3.0

penetrated halfway to the centerline of the duct or less, which includes both strong

autocombustion tests and no autocombustion tests. This may suggest that the fuel

trajectory and atomization did not play a critical role in autocombustion.However, the

variations in momentum ratio were accompanied by changes in the equivalence ratio and

flow rate of hydrogen peroxide.Therefore, the affect of varying jet trajectory at a

constant equivalence ratio and monoprop chamber pressure was not determined.

As the contraction ratio was increased the shear layer residence time created by the

rearward- facing step was increased as well. Past studies have shown that the shear layer

time can be correlated with ignition delay to predict the stability of a flame. In this study,

the initial intention was to deve lop a correlation or model for autocombustion relating the

shear layer residence time to an ignition delay parameter, which was similar in form to

that of past studies. The ignition delay parameter included the effects of temperature and

equivalence ratio, while velocity effects were included in the residence time

parameter. However, the correlation was not attempted due to inconclusive data

separating the effects of pressure, equivalence ratio, and contraction ratio. Most likely a

term would need to be added to the ignition delay portion of the correlation to include

pressure effects. In addition, large uncertainties were present in some of the calculated

data, including shear layer residence time. The large uncertainty originated from the

chamber pressure measurements, which were made using 3000 psia range transducers

with an accuracy of ±7.5 psia. During testing the monoprop chamber pressure was on the

order of 100 psia making the measurement uncertainty about 7.5%. The uncertainty

increased from this point through the rest of the calculations.

Despite this, it is believed that the rearward-facing step did provide enough

residence time to improve the autocombustion limits based on past data. Many flight-

rated staged-bipropellant engines that used hydrogen peroxide and kerosene have

had contraction ratios of seven of higher and ran at stoichiometric equivalence ratios.

A study by Walder on the autocombustion of kerosene in hydrogen peroxide used

contraction ratios of six and higher at stoichiometric conditions. In addition, data from

Wu et al using a similar engine design running 85% H2O2 at an equivalence ratio

somewhere between 0.8 and 1.4 did not achieve autocombustion at contraction ratio of

approximately 5.0 even at a monoprop chamber pressure of 340 psia. Data from this

study proved that autocombustion was possible at an equivalence ratio of 1.4 using 85%

H2O2 at a monoprop chamber pressure of approximately 100 psia. The data from this

study also proves that stable autocombustion is possible at contraction ratios as low as

3.0 and equivalence ratios less than 1.4 using 90% H2O2.

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