Steven J. Schneider

Glenn Research Center, Cleveland, Ohio

George A. Boyarko and Chih-Jen Sung

Case Western Reserve University, Cleveland, Ohio

Catalyzed Ignition of Bipropellants

in Microtubes

NASA/TM—2003-212206

February 2003

AIAA–2003–0674
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Steven J. Schneider

Glenn Research Center, Cleveland, Ohio

George A. Boyarko and Chih-Jen Sung

Case Western Reserve University, Cleveland, Ohio

Catalyzed Ignition of Bipropellants

in Microtubes

NASA/TM—2003-212206

February 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Aerospace Sciences Meeting and Exhibit

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Reno, Nevada, January 6–9, 2003

AIAA–2003–0674
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NASA/TMó2003-212206 1 

Catalyzed Ignition of Bipropellants in Microtubes 

Steven J. Schneider*

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

George A. BoyarkoÜ and Chih-Jen Sungá

Case Western Reserve University 

Cleveland, Ohio 44106 

Abstract 

This paper addresses the need to understand the physics and chemistry involved in propellant combustion 

processes in micro-scale combustors for propulsion systems on micro-spacecraft. These spacecraft are 

planned to have a mass less than 50 kilograms with attitude control estimated to be in the 

10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro Electrical 

Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching 

diameter of the propellants. Combustors of this size are expected to benefit significantly from surface 

catalysis processes. Miniature flame tube apparatus is chosen for this study because microtubes can be 

easily fabricated from known catalyst materials and their simplicity in geometry can be used in 

fundamental simulations to more carefully characterize the measured heat transfer and pressure losses for 

validation purposes. Experimentally, we investigate the role of catalytically active surfaces within 0.4 and 

0.8 mm internal diameter micro-tubes, with special emphases on ignition and extinction processes in fuel 

rich gaseous hydrogen and gaseous oxygen. Flame thickness and reaction zone thickness calculations 

predict that the diameters of our test apparatus are below the quenching diameter of the propellants in 

sub-atmospheric tests. Temperature and pressure rises in resistively heated platinum and palladium microtubes are used as an indication of exothermic reactions. Specific data on mass flow versus preheat 

temperature required to achieve ignition are presented. 

Introduction 

In recent years, there has been a tendency 

toward shrinking the size of spacecraft. New 

classes of spacecraft called micro-spacecraft have 

been defined by their mass, power, and size ranges 

[1]. Spacecraft in the range of 20 to 100 kg 

represent the class most likely to be utilized by 

most ìsmall satî users in the near future. There are 

also efforts to develop 10 to 20 kg class spacecraft 

for use in satellite constellations. More ambitious 

efforts will be to develop spacecraft less than 

10 kg, in which MEMS fabrication technology is 

required. These new micro-spacecraft will require 

new micro-propulsion technology. Although 

micro-propulsion includes electric propulsion 

approaches, the focus of this paper is microchemical propulsion which requires the 

development of micro-combustors. As combustors 

are scaled down, the surface to volume ratio 

increases. The heat release rate in the combustor 

scales with volume, while heat loss rate scales 

with surface area. Consequently, heat loss 

eventually dominates over heat release when the 

combustor size becomes smaller, thereby leading 

to flame quenching. The limitations imposed on 

chamber length and diameter has an immediate 

impact on the degree of miniaturization of a 

micro-combustor. Before micro-combustors can 

be realized, such a difficulty must be overcome. 

One viable combustion alternative is to take 

advantage of surface catalysis. 

ññññññññññññññññññññññññññññññññññ 

* Aerospace Engineer, On-Board Propulsion Branch, Steven.J.Schneider@grc.nasa.gov, Associate Fellow AIAA. Ü

 Graduate Student, Department of Mechanical and Aerospace Engineering, gab16@po.cwru.edu. Student Member AIAA. 

á

 Associate Professor, Department of Mechanical and Aerospace Engineering, cjs15@po.cwru.edu. Senior Member AIAA. 
NASA/TMó2003-212206 2 

Micro-chemical propulsion for small 

spacecraft can be used for primary thrust, orbit 

insertion, trajectory-control, and attitude control. 

Grouping micro-propulsion devices in arrays will 

allow their use for larger thrust applications. By 

using an array composed of hundreds or thousands 

of micro-thruster units, a particular configuration 

can be arranged to be best suited for a specific 

application. Moreover, different thruster sizes 

would provide for a range of thrust levels (from 

μNís to mNís) within the same array. Several 

thrusters could be fired simultaneously for thrust 

levels higher than the basic units, or in a rapid 

sequence in order to provide gradual but steady 

low-g acceleration. These arrays of micropropulsion systems would offer unprecedented 

flexibility and redundancy for satellite propulsion 

and reaction control for launch vehicles. 

A high-pressure bi-propellant micro-rocket 

engine is already being developed [2] using 

MEMS technology. High pressure turbopumps and 

valves are to be incorporated onto the rocket 

ìchip.î High pressure combustion of methane and 

O2 in a micro-combustor has been demonstrated 

without catalysis, but ignition was established with 

a spark. This combustor has rectangular 

dimensions of 1.5 mm by 8 mm (hydraulic 

diameter 3.9 mm) and a length of 4.5 mm and was 

operated at 1250 kPa with plans to operate it at 

12.7 MPa. These high operating pressures enable 

the combustion process in these devices, but these 

pressures are not practical for pressure fed satellite 

propulsion systems. Note that the use of these 

propellants requires an ignition system and that the 

use of a spark would impose a size limitation to 

this micro-propulsion device because the spark unit 

cannot be shrunk proportionately with the thruster. 

Currently, cold gas and small monopropellant 

N2H4 systems represent the state-of-the-art in micropropulsion. It is generally recognized that reducedscale versions of these conventional systems will 

not be practical for micro-satellites. In addition 

these propulsion systems will benefit greatly from 

the use of non-toxic propellants, which will require 

ignition. Systems fabricated using MEMS 

technology and catalyzed ignition/combustion are 

proposed. Once catalyzed ignition is established in a 

micro-tube, the reacting gases can be used for 

propulsion in the micro-propulsion device or they 

can be propagated into the combustion chamber for 

ignition in larger thrust class rockets. This benefits 

all rocket propulsion systems through the 

elimination of high voltage electrical discharges to 

achieve spark ignition. 

Results presented in this paper consist of an 

experimental evaluation of the minimum catalyst 

temperature for initiating/supporting combustion in 

sub-millimeter diameter tubes. The tubes are 

resistively heated and reactive premixed gases are 

passed through the tubes. Tube temperature and 

inlet pressure are monitored for an indication of 

exothermic reactions and composition changes in 

the gases. 

Flame Characteristics and Quenching Diameter 

Studies of premixed combustion in a tube have 

shown that if the tube diameter is progressively 

decreased, the flame eventually cannot propagate 

[3,4]. Thus, the so-called quenching diameter or 

distance is the critical dimension of the tube below 

which flame propagation is not possible. This is 

because diffusion of species and heat transfer from 

the flame front to the tube wall quenches the 

reaction. It is noted that wall quenching affects not 

only flammability limits, but also ignition 

phenomena [3,4]. 

Theoretical analysis has shown that the 

quenching diameter is on the same order as the 

flame thickness [3]. The flame thickness 

represents the overall non-equilibrium processes 

of reaction and diffusion in the flame, and hence 

provides an indication of the residence time within 

the flame. Since the flame thickness is inversely 

proportional to the burning rate, it decreases with 

increasing burning intensity. As such, nearstoichiometric flames would have a relatively 

smaller flame thickness as well as quenching 

distance due to stronger burning intensity. 

Spatially resolved profiles of temperature and 

heat release rate for a freely-propagating H2/O2

flame in the flame coordinate can be computed 

using a detailed mechanism of Mueller et al. [5], 

which is incorporated into the PREMIX Code [6]. 

PREMIX is a program for modeling steady, 

laminar, one-dimensional, premixed flames. 

Calculations were run for freely propagating H2/O2

premixed flames under adiabatic and non-adiabatic 

conditions. It should be noted that the fuel rich 

extinction limit for atmospheric H2/O2 flames is at 

an equivalence ratio of φ = 12, mixture ratio 

(oxygen/fuel mass ratio) of MR = 0.67, and a H2

mole fraction of 0.96 [7]. 
NASA/TMó2003-212206 3 

Flames and their characteristic dimensions can 

be characterized using various methods of analysis. 

In the first method, we analyze the H2/O2 flame by 

plotting its temperature profile in Figure 1. 

A flame thickness, δ, can be obtained by 

calculating the difference between the final and 

initial temperatures and dividing by the maximum 

value of the temperature derivative. As expected, 

δ increases with increasing equivalence ratio. This 

is also shown by plotting the temperature 

derivative in relation to the flame coordinate in 

Figure 2. A wider thermal structure is produced by 

mixtures of greater equivalence ratios. A plot of 

flame thickness as a function of equivalence ratio 

for atmospheric pressure combustion is given in 

Figure 3. Since the combustion process explored in 

this paper is at low pressure in an altitude test 

facility for rocket engines, the pressure effect on 

flame thickness is also calculated as a function of 

equivalence ratio. The flame thickness at a 

pressure of 34.5 kPa (approximately 5 psia) is also 

plotted in Figure 3. It is seen that flame thickness 

increases with decreasing pressure and the 

calculated flame thickness at the test pressure of 

34.5 kPa and an equivalence ratio of 10 (mixture 

ratio 0.8) is 4 mm, which is greater than the 

diameter of either the 0.4 mm Platinum or the 

0.8 mm Palladium micro-tubes used in this 

catalytic micro-combustion experiment. 

In the second method, we define a reaction 

zone as the full width of the half maximum of the 

heat release profile. Figure 4 shows the results for 

two different equivalence ratios with the reaction 

zone thickness increasing with increasing 

equivalence ratio. A plot of reaction zone thickness 

as a function of equivalence ratio for atmospheric 

pressure combustion is given in Figure 5. The 

pressure effect on reaction zone thickness is also 

calculated and shown in Figure 5 as a function of 

equivalence ratio for 34.5 kPa pressure. At the 

equivalence ratio of 10 (mixture ratio = 0.8), the 

reaction zone thickness is ~1 mm at atmospheric 

pressure and initial temperature of 300 K and 

increases to ~3 mm when the pressure is reduced to 

34.5 kPa. Experimental tests in the micro-tubes 

were run mostly between 1.2 and 0.4 mixture ratios 

(equivalence ratios of 6.7 to 20, respectively). At 

these mixture ratios and the low pressure of the 

tests the reaction zone thicknesses are greater than 

the diameter of either the 0.4 mm Platinum or the 

0.8 mm Palladium micro-tubes used in this 

catalytic micro-combustion experiment. 

Temperature vs. Flame Coordinate

0

500

1000

1500

2000

2500

3000

0 0.05 0.1 0.15 0.2 0.25 0.3

x-coordinate (cm)

Deg. K

φ = 5.5

φ = 2

φ = 10

Figure 1. Plot of the temperature versus the flame 

coordinate for atmospheric H2/O2 freely propagating 

flames of three equivalence ratios with an initial 

temperature of 300 K. 

dT/dx vs. Flame Coordinate

0

20000

40000

60000

80000

100000

120000

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4

cm

dT/dx (K/cm)

φ = 2

φ = 5.5

φ = 10.5

Figure 2. Plot of the temperature gradient versus the 

flame coordinate for atmospheric H2/O2 freely 

propagating flames of three equivalence ratios with an 

initial temperature of 300 K. 

Flame Width vs. Equivalence Ratio

0.00000

0.05000

0.10000

0.15000

0.20000

0.25000

0.30000

0.35000

0.40000

0.45000

0.50000

0 2 4 6 8 10 12

Equivalence Ratio, φ

cm

atmospheric

34.5 kPa

Figure 3. Plot of flame thickness versus equivalence 

ratio for atmospheric pressure and 34.5 kPa. 
NASA/TMó2003-212206 4 

Heat Release Profile

0

2E+10

4E+10

6E+10

8E+10

1E+11

1.2E+11

0 0.1 0.2 0.3 0.4 0.5

x-coordinate, cm

Energy Release, ergs/(cm 2*sec)

φ=9.5

φ=5.5

Figure 4. Heat release profile versus the flame 

coordinate for H2/O2 flames of two different 

equivalence ratios at atmospheric pressure. 

Reaction Zone Thickness vs. Equivalence Ratio

0.00000

0.10000

0.20000

0.30000

0.40000

0.50000

0.60000

0 2 4 6 8 10 12 14

Equivalence Ratio, φ

cm

atmospheric

34.5 kPa

Figure 5. Plot of reaction zone thickness versus 

equivalence ratio for atmospheric pressure and 

34.5 kPa. Reaction zone thickness increases with 

increasing equivalence ratio and increases with 

decreasing pressure. 

In order to determine the flammability limits 

and capture the extinction turning point, the 

PREMIX code was modified to include radiative 

heat loss and the capability of solving the 

equations at the singular points. When continuing 

the calculation from the upper (stable) branch to 

the lower (unstable) branch by passing through the 

turning point, numerical difficulty, especially with 

detailed chemistry, is encountered because the 

Jacobian matrix used in the Newtonís method 

becomes singular at the turning point. To solve this 

problem, the flame controlling continuation 

method [8] is employed in this study. Radiative 

heat loss from H2O is considered since it is the 

primary radiative species in hydrogen flames. The 

corresponding Planck mean absorption coefficient 

is a function of temperature and is obtained from 

Ref. [9]. The analysis was done to test the rich 

flammability limits of the H2/O2 flames at low 

pressure. Figure 6 shows the variation in the mass 

burning rate with equivalence ratio at pressure of 

34.5 kPa. It is seen that the burning rate decreases 

with increasing equivalence ratio. The turning 

point, beyond which steady flame propagation is 

not possible, defines the rich flammability limit. 

Calculations at 34.5 kPa show that the rich limit of 

the H2/O2 flame occurs at an equivalence ratio of 

12.78, which corresponds to a mixture ratio of 

0.626 and a hydrogen mole fraction of 0.962. We 

further note that for near rich limit hydrogen 

flames, in which the Lewis number is greater than 

unity, pulsating instability has been observed to 

take place in mixtures with fuel concentrations 

slightly below that of the rich limit [7,10,11] and 

the flame is found to oscillatorily extinguish well 

ahead of the static extinction limit [10,11]. 

Therefore, the flammability range of a H2/O2

mixture is expected to be narrowed when the 

intrinsic oscillatory nature of flame propagation is 

taken into account. 

Equiv. Ratio vs. Burning Rate (.34 atm)

0

0.00025

0.0005

0.00075

0.001

11 11.5 12 12.5 13

Equivalence Ratio (φ)

Burning Rate g/(cm^2*sec

Figure 6. Plot of mass burning rate versus equivalence 

ratio for H2/O2 flames at 34.5 kPa (5 psia). The turning 

point of the response curve defines the rich flammability 

limit. 

Catalytic Combustion 

The principal advantage of using catalysts for 

primary combustion reactions is their ability to 

sustain reactions that are well beyond the 

flammability limit for homogeneous gaseous 

combustion. During the last three decades, 

catalysts have been investigated for both 

augmenting primary heat release in gas turbines 

and post-combustion treatment of pollutant 
NASA/TMó2003-212206 5 

emission in aircraft, automobile, and power 

generation applications. For generating power, the 

focus is on conditions at which catalytic reactions 

are used to promote the primary combustion 

reactions. 

By fundamental definition, a ìcatalystî is a 

substance that alters the reaction rate, while there 

is no change in its properties or concentration 

during the catalytic process. In a broader term, the 

so-called catalyst includes substances that generate 

catalytic species. A heterogeneous catalyst 

typically adsorbs the reactants and subsequently 

promotes critical reactions at certain adsorption 

sites. Surface reactions occur and then the 

products are desorbed and along with heat are 

transported into the flow by diffusion, thereby 

increasing the overall reaction rate. 

Heterogeneous catalyst types can be classified 

into three types: metals, non-metals (metal oxides 

and metal sulfides), and salts/acids. The main role 

of metal catalysts is hydrogenation and 

dehydrogenation reactions, which generally occur 

at relatively low or moderate temperatures. Noble 

metal catalysts or metal oxide catalysts are likely 

choices for the complete oxidation of an organic 

compound. Metals are chosen for this study 

because they can be easily fabricated into microtubes from known catalyst materials and they can 

be resistively heated. For oxidation reactions, only 

the noble metals, such as platinum and palladium, 

retain their integrity for long periods of time. Most 

metals in high-temperature, oxygen-enriched 

environments are prone to undergo bulk oxidation 

or form thick oxide layers on the surfaces. 

The reacting flow processes within the 

catalytic bed are complicated. Take a premixed 

fuel/air mixture flowing through a long channel 

with catalytic surfaces as an example [12]. Near 

the inlet section, the fuel conversion is controlled 

by the heterogeneous reactions at the surface of 

the catalyst. In this region, the bulk gas 

temperatures are typically too low to support 

homogeneous combustion. After a certain distance 

downstream of the inlet, the heterogeneous 

reaction rates increase with increasing 

temperatures of both near-wall gas and substrate. 

As the reaction rate increases to an extent, the fuel 

conversion is no longer controlled by the surface 

reaction rate. Instead, the fuel consumption is now 

controlled by the rate at which new gas phase 

reactants are transported to the surface, and the 

catalytic reactions at the wall will be limited solely 

by the rate of mass transfer. Further downstream 

and beyond this diffusion-limited region, thermal 

runaway of homogeneous or gas phase reactions 

occur as heat continues to diffuse into the bulk gas 

flow. These homogeneous reactions are typically 

initiated at bulk gas temperatures in the range of 

1200 to 1300 K. 

Based on the above description, it is 

imperative to identify how the catalyzed processes 

scale with the combustor size [13]. In the surface 

kinetics controlled region, the process is primarily 

dependent on the catalytic Damkˆhler number, 

which is the ratio of system residence time to the 

characteristic heterogeneous reaction time. For a 

given residence time, the catalytic Damkˆhler 

number is inversely proportional to the combustor 

size due to the dependence of the heterogeneous 

reaction rate on surface area. The transition from 

surface reaction- to diffusion-limited processes 

and the progression of the latter are expected to 

scale with the mass transport number that is the 

ratio of the characteristic residence time of the gas 

in the reactor to the mass/heat transport time. 

Since both the heat and mass transfer times scale 

inversely with the hydraulic diameter, both 

transport processes are favored when the 

combustor dimension is reduced, as is the overall 

progression to the homogeneous reaction region. 

Therefore, although the increase of the surface 

to volume ratio of the combustor is problematic 

for gas phase combustion, it favors catalytic 

combustion. Despite the catalytic reaction being 

generally slower than the gas phase reaction, due 

to the relative increase of surface area and the 

lower temperature of the catalytic reaction, microcombustors utilizing catalytic combustion may be 

easier to implement than those utilizing gas phase 

reactions. In fact, it has been demonstrated [14] 

using a heat-recirculating engine that with a 

catalyst located in the combustion volume 

hydrogen/air mixtures can exhibit thermal 

runaway at room temperature and butane (or 

propane) can self-start at less than 500 K. 

Many unique technology issues are related to 

MEMS-based micro-propulsion systems. They 

include overall issues such as materials selection, 

fabrication, sealing, assembly, pressurization, 

testing, and characterization, as well as the 

combustion related issues. In view of the above 

discussion the combustor remains one of the most 
NASA/TMó2003-212206 6 

challenging components to miniaturize. Therefore, 

the fundamental understanding of various microcombustion phenomena constitutes the long range 

goal of the present study. 

The preceding discussion of the physical 

processes and scaling of catalytic combustion 

systems, identifies increases in both the 

heterogeneous Damkˆhler number and the nondimensional heat and mass transfer rates as 

combustor size is reduced, resulting in the 

favorable effects for heterogeneous catalysis due 

to increased surface-to-volume ratio. These effects 

can offset the wall quenching effects. 

Apparatus and Test Procedures 

Miniature flame tube apparatus is chosen for 

study because microtubes can be easily fabricated 

from known catalyst materials and their simplicity 

in geometry can be used in fundamental 

simulations to more carefully characterize the 

measured heat transfer and pressure losses for 

validation purposes. Experimentally, we 

investigate the role of catalytically active surfaces 

within microtubes, with special emphases on 

ignition and extinction. The experimental 

apparatus involves the microtube, along with a 

mass flow control system, electrical power supply 

to heat a section of tube, and a data acquisition 

system. A schematic of the instrumented 

microtubes is shown in Figure 7. Two capillary 

tubes were available for testing. One was a 

0.4 mm I.D. platinum tube with a 0.15 mm wall 

thickness and an overall length of 100 mm of 

which 70 mm was resistively heated. The other 

tube was 0.8 mm I.D. and was made of palladium 

with a wall thickness of 0.1 mm and an overall 

length of 190 mm of which 160 mm was 

resistively heated. Three thermocouples were spot 

welded to the outside of the tubes at the 20, 50, 

and 80 percent positions of the heated section as 

shown in the schematic. Figure 8 is a picture of 

the 0.8 mm I.D. experimental microtube 

apparatus. The tubes were tested in a small rocket 

test facility at the NASA Glenn Research Center. 

The facility maintains an ambient pressure of 

1.38 kPa (10 torr) during the tests [15]. 

These initial tests were conducted with 

gaseous oxygen and gaseous hydrogen propellants 

at low mixture ratios (high equivalence ratios) to 

test the rich flammability limits of the propellants, 

to avoid melting the catalysts, and to avoid 

Figure 7. Schematic of microtube test articles. 

Figure 8. Microtube test apparatus (0.8 mm I.D.). 

oxidation loss of catalyst material. These 

propellants are non-toxic and can be produced onorbit by water electrolysis [16]. They also have 

detailed gas-phase chemistry, transport properties, 

and surface kinetics available for numerical 

simulation.

Results and Discussion 

A Fanno flow calculation was conducted as a 

first step to assess the effects of friction on the 

flow velocity in the constant area tubes. A more 

rigorous Rayleigh-Fanno flow calculation needs to 

be done in the future to assess the additional 

effects of resistive heat transfer to the gases and 

heat release due to reaction in the tube. In order to 

bound the effects of the change in gas properties 

on the Fanno flow calculation, flow with both the 

reactant and equilibrium product properties is 

evaluated. For example, at a fuel rich mixture ratio 

of 0.8 the following equilibrium reaction can be 

written for the propellants 

H2 + 0.0504 O2 → 0.1008 H2O + 0.8992 H2

X - Marks Thermocouple Locations 

X X X 

20% 

50% 

80% 

Heated Length 

Overall Length 

100% 

Flow 

In 

TC1 

TC2 

TC3 
NASA/TMó2003-212206 7 

If this reaction goes to equilibrium, the gas 

temperature rises from 298 to 1100 K [17], the 

average molecular weight changes from 3.455 to 

3.629, the viscosity rises from 0.094 to 

0.268 millipoise, and the ratio of specific heats 

changes from 1.405 to 1.353. This information is 

used in the Fanno flow calculation for the 0.8 mm 

I.D. tube with a measured mass flow of 

0.00123 g/sec. The ambient reactants enter the tube 

at 48.0 m/sec velocity. The flow is laminar at a 

Reynolds Number of 208 yielding a friction factor 

of 0.307. This friction factor chokes this flow at 

the exit of the 190 mm length of tube. The 

estimated average residence time of the unreacted 

gases in the tube would be 0.403 msec. In the limit, 

if the equilibrium products of the reaction are 

introduced at the entrance of the tube, the inlet 

velocity could rise up to a maximum of 55.3 m/sec. 

The rise in viscosity, however, dominates the 

Reynolds Number dropping it to 73 and the friction 

factor increases to 0.876. With reactions, therefore, 

the Fanno flow remains choked in the 190 mm 

length of tube and the inlet pressure is driven up by 

the rise in the friction factor. 

A Fanno flow calculation for the 0.4 mm I.D. 

tube with a measured mass flow of 0.000412 g/sec 

gives an inlet velocity for the unreacted gases of 

38.4 m/sec. The flow is laminar at a Reynolds 

Number of 140 and the friction factor is 0.459. 

Again, with no reactions in the gas, this friction 

factor chokes the flow at the exit of the 100 mm 

length of tube. The estimated average residence 

time of the unreacted gases in the tube is 

0.214 msec. In the limit, if the equilibrium 

products of the reaction are introduced at the 

entrance of the tube, the inlet gas temperature rises 

and the inlet velocity could rise up to a maximum 

of 43.2 m/sec. The rise in viscosity dominates the 

Reynolds Number dropping it to 49 and the friction 

factor increases to 1.31. With reactions, the flow 

remains choked in the 100 mm length of 0.4 mm 

I.D. tube and the inlet pressure is again driven up. 

Actual experimental data with combustion are 

expected to lie somewhere between these two 

extremes. This analysis, however, is helpful in 

evaluating the pressure data. 

Experimental results consist of a determination 

of the minimum catalyst preheat temperature for 

initiating/supporting combustion in the two submillimeter diameter tubes. Data are obtained with 

the tubes resistively heated by the power supply in 

300

400

500

600

700

800

900

1000

0 50 100

Time ~ seconds

Temperature ~ K

TC1

TC2

TC3

Figure 9. Temperature data for 0.8 mm I.D. palladium 

tube, MR = 0.8, 0.00123 g/sec. 

0

10

20

30

40

50

60

70

0 50 100

Time ~ seconds

Pressure ~ kPa

P in

Figure 10. Pressure data for 0.8 mm I.D. palladium tube, 

MR = 0.8, 0.00123 g/sec. 

voltage control mode and with reactive premixed 

gases passed through the tubes. Tube temperature 

and inlet pressure are monitored for an indication 

of exothermic reactions and composition changes 

in the gases. 

Temperature data for the 0.8 mm I.D. tube 

preheated to a peak temperature of 477 K and 

showing an exothermic reaction is given in 

Figure 9. The total mass flow is 0.00123 g/sec of 

gaseous oxygen and gaseous hydrogen at an 

oxidizer to fuel mixture ratio of 0.8 (equivalence 

ratio 9.92). The thermocouples are located as 

shown in Figure 7. The resistivity of palladium at 

these temperatures is 17.2×10-8 Ω-m and the 

voltage drop across the tube is 1.21 Volts, giving 

an estimated power input of 15 Watts for the 

deferred Rayleigh-Fanno flow calculation. At 

10 seconds into the test TC1 has dropped to 338 K, 

TC2 has dropped to 412 K as the gas picks up heat, 

but TC3 has risen to 511 K as the gases are heated 

by catalyzed reactions between TC2 and TC3. At 

10 seconds into the test TC3 starts a rapid rise to 

750 K. Heat from this reaction is conducted 

through the tube upstream toward TC2 which starts 

a rapid temperature rise to 815 K at 16 seconds 
NASA/TMó2003-212206 8 

into the test. Further conduction to the TC1 

position is indicated by its rapid temperature rise at 

40 seconds into the test. Steady state is reached at 

TC3 at approximately 60 seconds into the test. The 

test was terminated at 100 seconds, however, 

before steady state was reached at TC1. 

The pressure in a plenum upstream of the tube 

is given in Figure 10. The pressure at the start of 

the test is 2.7 kPa in the altitude test facility. The 

pressure rises rapidly to 34 kPa at 3 seconds into 

the test and stays there until about 10 seconds into 

the test when catalyzed reactions start to drive the 

pressure up as discussed in the Fanno flow 

discussion above. The pressure then rises to 59 kPa 

at 71 seconds into the test and stays there until the 

test is terminated at 100 seconds. This pressure rise 

is due primarily to a rise in viscosity and friction 

factor as the gases are heated by the exothermic 

reaction. The cold flow pressure predicted by the 

Fanno Line at the 0.00123 g/sec mass flow is 

34.9 kPa which agrees with the pressure measured 

at 3 seconds into the test. If equilibrium 

combustion were achieved, the Fanno Line would 

predict a pressure rise to 112 kPa, instead of the 

measured 59 kPa. 

300

320

340

360

380

400

420

440

0 50 100

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 11. Non-ignition temperature data for 0.8 mm 

I.D. palladium tube, MR = 0.8, 0.00123 g/sec. 

0

5

10

15

20

25

30

35

40

0 50 100

Time ~ sec

Pressure ~ kPa

P in

Figure 12. Non-ignition pressure data for 0.8 mm I.D. 

palladium tube, MR = 0.8, 0.00123 g/sec. 

Temperature data for the 0.8 mm I.D. tube 

preheated to a peak temperature of 422 K and 

showing no reaction is given in Figure 11. The 

total mass flow is 0.00123 g/sec of gaseous oxygen 

and gaseous hydrogen at an oxidizer to fuel 

mixture ratio of 0.8 (equivalence ratio 9.92). Some 

reactions between TC2 and TC3 cause the 

temperature of TC3 to rise to a peak of 430 K at 

6 seconds into the test, otherwise all of the 

temperatures fall to a steady state profile of TC1 at 

321 K, TC2 at 351 K and TC3 at 377 K. The 

pressure in a plenum upstream of the tube is given 

in Figure 12 for this case. The pressure rises 

rapidly and reaches a pressure of 33 kPa at 

15 seconds into the test and is in agreement with 

the cold flow Fanno Line calculation. 

Temperature data for the 0.8 mm I.D. tube 

preheated to a peak temperature of 589 K at a 

mixture ratio of 0.6 (equivalence ratio 13.2) and a 

mass flow of 0.00160 g/sec is given for a reacting 

case in Figure 13. The corresponding pressure data 

is given in Figure 14. This equivalence ratio is 

above the extinction limit predicted by theory. 

If the preheat temperature is reduced to 

approximately 530 K, the temperature data shown 

300

400

500

600

700

800

900

1000

0 50 100

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 13. Temperature data for 0.8 mm I.D. palladium 

tube, MR = 0.6, 0.00160 g/sec. 

0

10

20

30

40

50

60

70

80

0 50 100

Time ~ sec

Pressure ~ kPa

P in

Figure 14. Pressure data for 0.8 mm I.D. palladium tube, 

MR = 0.6, 0.00160 g/sec. 
NASA/TMó2003-212206 9 

300

350

400

450

500

550

600

0 50 100

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 15. Non-ignition temperature data for 0.8 mm 

I.D. palladium tube, MR = 0.6, 0.00160 g/sec. 

0

10

20

30

40

50

0 50 100 150

Time ~ sec

Pressure ~ kPa

P in

Figure 16. Non-ignition pressure data for 0.8 mm I.D. 

palladium tube, MR = 0.6, 0.00160 g/sec. 

in Figure 15 results indicating that the reactions 

have not occurred. The corresponding pressure 

data is shown in Figure 16 and is in agreement 

with the cold flow Fanno Line calculation at 

0.00160 g/sec indicating that there are no 

reactions. 

Temperature data for the 0.4 mm I.D. platinum 

tube preheated to a peak temperature of 640 K 

which shows an exothermic reaction is given in 

Figure 17. The total mass flow is 0.000412 g/sec of 

gaseous oxygen and gaseous hydrogen at an 

oxidizer to fuel mixture ratio of 0.8 (equivalence 

ratio 9.92). The resistivity of platinum at these 

temperatures is 23.4×10-8 Ω-m and the voltage 

drop across the tube is 1.41 Volts, giving an 

estimated power input of 31 Watts for the deferred 

Rayleigh-Fanno flow calculation. At 30 seconds 

into the test, the thermocouples have dropped to 

their minimum values of TC1 at 400 K, TC2 at 480 

K and TC3 at 550 K. The remaining 30 seconds of 

the test show an exothermic reaction taking place 

giving a rise in temperature at TC3 to 650 K at the 

end of the test. This test was terminated before 

steady state was reached. The pressure in a plenum 

upstream of the tube is given in Figure 18. The 

300

400

500

600

700

0 20 40 60

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 17. Temperature data for 0.4 mm I.D. platinum 

tube, MR = 0.8, 0.000412 g/sec. 

0

10

20

30

40

50

60

0 20 40 60

Time ~ sec

Pressure ~ kPa

P in

Figure 18. Pressure data for 0.4 mm I.D. platinum tube, 

MR = 0.8, 0.000412 g/sec. 

pressure rises to 43 kPa at 30 seconds into the test 

and then rises to 48 kPa at the end of the test due 

primarily to a rise in viscosity as the gases are 

heated by the exothermic reaction. 

Temperature data for the 0.4 mm I.D. tube 

preheated to a peak temperature of 585 K which 

shows only a very mild reaction is given in 

Figure 19. The total mass flow is 0.000412 g/sec of 

gaseous oxygen and gaseous hydrogen at an 

oxidizer to fuel mixture ratio of 0.8 (equivalence 

ratio 9.92). Convection between TC2 and TC3 

cause the temperature at TC3 to rise to a peak of 

584 K at 10 seconds into the test, otherwise all of 

the temperatures fall to a steady state profile of 

TC1 at 380 K, TC2 at 450 K and TC3 at 501 K at 

25 seconds into the test and remain these levels 

except for a slight rise in TC3 until the test is 

terminated at 69 seconds. The pressure in a plenum 

upstream of the tube is given in Figure 20 for this 

case. The pressure at the start of the test is 2.5 kPa 

and rises rapidly to 43 kPa at 30 seconds into the 

test as it did in the reacting case shown in Figure 

17. The pressure then rises to a plateau of 47 kPa at 

65 seconds into the test indicating a mild viscosity 

change and remains there until the test is 

terminated at 69 seconds. 
NASA/TMó2003-212206 10

300

350

400

450

500

550

600

650

0 20 40 60

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 19. Temperature data for 0.4 mm I.D. platinum 

tube, MR = 0.8, 0.000412 g/sec. 

0

10

20

30

40

50

0 20 40 60

Time ~ sec

Pressure ~ kPa

P in

Figure 20. Pressure data for 0.4 mm I.D. platinum tube, 

MR = 0.8, 0.000412 g/sec. 

300

350

400

450

500

550

600

650

700

0 50 100

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 21. Temperature data for 0.4 mm I.D. platinum 

tube, MR = 0.6, 0.000537 g/sec. 

0

10

20

30

40

50

60

0 50 100

Time ~ sec

Pressure ~ kPa

P in

Figure 22. Pressure data for 0.4 mm I.D. platinum tube, 

MR = 0.6, 0.000537 g/sec. 

300

350

400

450

500

550

600

650

0 50 100

Time ~ sec

Temperature ~ K

TC1

TC2

TC3

Figure 23. Temperature data for 0.4 mm I.D. platinum 

tube, MR = 0.6, 0.000537 g/sec. 

0

10

20

30

40

50

60

0 50 100

Time ~ sec

Pressure ~ kPa

P in

Figure 24. Pressure data for 0.4 mm I.D. platinum tube, 

MR = 0.6, 0.000537 g/sec. 

Temperature data for the 0.4 mm I.D. tube at a 

mixture ratio of 0.6 (equivalence ratio 13.2) and a 

mass flow of 0.000537 g/sec is given for a reacting 

case in Figure 21. The peak preheat temperature of 

640 K was used. The corresponding pressure data 

is given in Figure 22. This equivalence ratio is 

above the extinction limit predicted by theory. If 

the peak preheat temperature is reduced to 585 K, 

the temperature data shown in Figure 23 

indicate that less reactions have occurred. The 

corresponding pressure data is shown in Figure 24. 

These data and data at other mass flows and 

fuel rich mixture ratios are shown on mass flow 

versus preheat temperature plots in Figure 25 for 

the 0.8 mm I.D. palladium tube and Figure 26 for 

the 0.4 mm platinum tube. Using temperature and 

pressure rise as an indication of reactions the solid 

symbols indicate those test conditions where 

reactions were observed. As mass flow decreases 

the preheat temperature required to initiate 

reactions decreases. For the 0.8 mm palladium tube 

the lowest preheat temperature for reactions was 

340 K at a mixture ratio of 0.8 and a mass flow 

of 0.000128 g/sec. Above a mass flow of 

0.0022 g/sec reactions did not occur at any preheat 

temperature in the 0.8 mm I.D. palladium tube, 

190 mm long with 160 mm resistively heated. 
NASA/TMó2003-212206 11

300

350

400

450

500

550

600

650

700

750

800

0 0.001 0.002 0.003

Total Mass Flow ~ g/sec

Tube Temperature ~ K

MR 0.4

MR 0.6

MR 0.8

MR 1.0

MR 0.4

MR 0.6

MR 0.8

MR 1.0

Figure 25. Ignition data for 0.8 mm I.D. palladium tube 

ñ open symbols are non-ignition and solid symbols are 

ignition as indicated by tube temperature rise. 

300

350

400

450

500

550

600

650

700

750

0 0.0005 0.001 0.0015 0.002

Total Mass Flow ~ g/sec

Tube Temperature ~ K

MR 0.6

MR 0.8

MR 1.0

MR 1.2

MR 0.6

MR 0.8

MR 1.0 

MR 1.2

Figure 26. Ignition data for 0.4 mm I.D. platinum tube ñ 

open symbols are non-ignition and solid symbols are 

ignition as indicated by tube temperature rise. 

For the 0.4 mm I.D. platinum tube, the lowest 

preheat temperature for reactions was 480 K at 

a mixture ratio of 1.0 and a mass flow of 

0.000187 g/sec. Above a mass flow of 

0.00054 g/sec reactions did not occur at any 

preheat temperature in the 0.4 mm I.D. platinum 

tube, 100 mm long with 70 mm resistively heated. 

Summary 

New micro-spacecraft will require new micropropulsion technology, which will require the 

development of micro-combustors. The heat 

release rate in the combustor scales with volume, 

while heat loss rate scales with surface area. 

Consequently, heat loss eventually dominates over 

heat release when the combustor size becomes 

smaller, thereby leading to flame quenching. The 

limitations imposed on chamber length and 

diameter has an immediate impact on the degree 

of miniaturization of a micro-combustor. One 

viable combustion alternative is to take advantage 

of surface catalysis. The favorable effects on 

surface catalysis associated with the high surfaceto-volume ratio as the combustor size is reduced 

can offset the wall quenching effects. Flame 

thickness and reaction zone thickness calculations 

predict that the diameters of our test apparatus are 

below the quenching diameter of the gaseous 

oxygen and gaseous hydrogen propellants in subatmospheric tests. 

Miniature flame tube apparatus is chosen for 

study because micro-tubes can be easily fabricated 

from known catalyst materials and their simplicity 

in geometry can be used in fundamental 

simulations to more carefully characterize the 

measured heat transfer and pressure losses for 

validation purposes. A Fanno flow calculation is 

used to predict the cold flow pressure rise in the 

tubes, where the experimental mass flow is shown 

to choke at the exit of the tubes due to friction. 

Data on the temperature and pressure rise as a 

function of time in resistively heated platinum and 

palladium micro-tubes is used as an indication of 

exothermic reactions in a series of tests varying 

mass flow, preheat temperature, and fuel rich 

mixture ratio. As mass flow decreases the preheat 

temperature required to initiate reactions is shown 

to decrease. 

For the 0.8 mm palladium tube, 190 mm long 

with 160 mm resistively heated, the lowest preheat 

temperature for reactions was 340 K at a mixture 

ratio of 0.8 and a mass flow of 0.000128 g/sec. 

Above a mass flow of 0.0022 g/sec reactions did 

not occur at any preheat temperature in the 

0.8 mm I.D. palladium tube. For the 0.4 mm I.D. 

platinum tube, 100 mm long with 70 mm 

resistively heated, the lowest preheat temperature 

for reactions was 480 K at a mixture ratio of 

1.0 and a mass flow of 0.000187 g/sec. Above a 

mass flow of 0.00054 g/sec reactions did not 

occur at any preheat temperature in the 0.4 mm 

I.D. platinum tube. 

References 

1. Mueller, J., ìThruster Options for Microspacecraft: 

A Review and Evaluation of Existing Hardware and 

Emerging Technologies,î AIAA 97-3058, 1997. 

2. London, A. P., Epstein, A. H., and Kerrebrock, J. L., 

ìHigh-Pressure Bipropellant Microrocket Engine,î 

Journal of Propulsion and Power, Volume 17, 

No. 4, pp. 780-787, 2001. 
NASA/TMó2003-212206 12

3. Williams, F. A., Combustion Theory, 2nd edition, 

Addison-Wesley Publishing Company, Menlo 

Park, California, 1985. 

4. Glassman, I., Combustion, 3rd Edition, Academic 

Press, San Diego, CA, 1997. 

5. Mueller, M. A., Yetter, R. A., and Dryer, F. L., 

ìFlow Reactor Studies and Kinetic Modeling of 

the H2/O2/NOx and CO/H2O/O2/NOx Reactions,î 

International Journal of Chemical Kinetics 31, 

pp. 705-724 (1999). 

6. Kee, R. J., Grcar, J. F., Smooke, M. D., and Miller, 

J. A., A Fortran Program for Modeling Steady 

Laminar One-Dimensional Premixed Flames, 

Sandia Report SAND85-8240, 1985. 

7. L. He, P. Clavin, ìPremixed Hydrogen-Oxygen 

Flames. Part II: Quasi-isobaric Ignition Near the 

Flammability Limits,î Combustion and Flame 93, 

pp. 408-420, 1993. 

8. Nishioka, M., Law, C. K., and Takeno, T., ìA 

Flame-Controlling Continuation Method for 

Generating S-Curve Responses with Detailed 

Chemistry,î Combustion and Flame 104, pp. 328-

342, 1996. 

9. Soufiani, A., and Taine, J., ìHigh Temperature Gas 

Radiative Property Parameters of Statistical 

Narrowband Model for H2O, CO2, and CO and 

Correlated k Model for H2O and CO2,î 

International Journal of Heat Mass Transfer 40, 

pp. 987-991, 1997. 

10. Christiansen, E. W., Sung, C. J., and Law, C. K., 

ìPulsating Instability in the Fundamental 

Flammability Limit of Rich Hydrogen/Air 

Flames,î Proceedings of the Combustion Institute

27, pp. 555-562, 1998. 

11. Christiansen, E. W., Law, C. K., and Sung, C. J., 

ìSteady and Pulsating Propagation and Extinction 

of Rich Hydrogen/Air Flames at Elevated 

Pressures,î Combustion and Flame 124, pp. 35-49, 

2001. 

12. Groppo, G., Tronconi, E., and Forzatti, P., 

ìModeling of Catalytic Combustors for Gas 

Turbine Applications,î Catalysis Today 17, 

pp. 237-250 (1993). 

13. Waitz, I. A., Gauba, G., and Tzeng, Y.-S., 

ìCombustors for Micro-Gas Turbine Engines,î 

Journal of Fluids Engineering 120, pp. 109-117 

(1998). 

14. Maruta, K., Takeda, K., Sitzki, L., Borer, K., 

Ronney, P. D., Wussow, S., and Deutschmann, O., 

ìCatalytic Combustion in Microchannel for 

MEMS Power Generation,î Third Asia-Pacific 

Conference on Combustion, Seoul, Korea, 

June 24-27, 2001. 

15. Arrington, L. A. and Schneider, S. J., ìLow Thrust 

Rocket Test Facilityî, AIAA 90-2503, July 1990. 

16. deGroot, W. A. et., al., ìElectrolysis Propulsion for 

Spacecraft Applications,î AIAA 97-2948, 1997. 

17. McBride, B. J. and Gordon, S., ìComputer 

Program for Calculation of Complex Chemical 

Equilibrium Compositions and Applications,î 

NASA RP-1311, June 1996. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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Available electronically at http://gltrs.grc.nasa.gov

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AIAA–2003–0674

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18

Catalyzed Ignition of Bipropellants in Microtubes

Steven J. Schneider, George A. Boyarko, and Chih-Jen Sung

Chemical propulsion; Rocket engines; Low thrust propulsion; Spacecraft propulsion

Unclassified -Unlimited

Subject Category: 20 Distribution: Nonstandard

Prepared for the 41st Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno, Nevada, January 6–9, 2003. Steven J. Schneider, NASA Glenn Research Center; George A. Boyarko

and Chih-Jen Sung, Case Western Reserve University, Cleveland, Ohio 44106. Responsible person, Steven J. Schneider,

organization code 5430, 216–433–7484.

This paper addresses the need to understand the physics and chemistry involved in propellant combustion processes in micro-scale

combustors for propulsion systems on micro-spacecraft. These spacecraft are planned to have a mass less than 50 kilograms with

attitude control estimated to be in the 10 milli-Newton thrust class. These combustors are anticipated to be manufactured using Micro

Electrical Mechanical Systems (MEMS) technology and are expected to have diameters approaching the quenching diameter of the

propellants. Combustors of this size are expected to benefit significantly from surface catalysis processes. Miniature flame tube

apparatus is chosen for this study because microtubes can be easily fabricated from known catalyst materials and their simplicity in

geometry can be used in fundamental simulations to more carefully characterize the measured heat transfer and pressure losses for

validation purposes. Experimentally, we investigate the role of catalytically active surfaces within 0.4 and 0.8 mm internal diameter

micro-tubes, with special emphases on ignition and extinction processes in fuel rich gaseous hydrogen and gaseous oxygen. Flame

thickness and reaction zone thickness calculations predict that the diameters of our test apparatus are below the quenching diameter of

the propellants in sub-atmospheric tests. Temperature and pressure rises in resistively heated platinum and palladium micro-tubes are

used as an indication of exothermic reactions. Specific data on mass flow versus preheat temperature required to achieve ignition are

presented.