Julie Kleinhenz, Charles Sarmiento, and William Marshall

Glenn Research Center, Cleveland, Ohio

Experimental Investigation of Augmented Spark 

Ignition of a LO2/LCH4 Reaction Control 

Engine at Altitude Conditions

NASA/TM—2012-217611

June 2012
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Julie Kleinhenz, Charles Sarmiento, and William Marshall

Glenn Research Center, Cleveland, Ohio

Experimental Investigation of Augmented Spark 

Ignition of a LO2/LCH4 Reaction Control 

Engine at Altitude Conditions

NASA/TM—2012-217611

June 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

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Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217611 iii 

Contents 

1.0 Introduction ............................................................................................................................................ 1

Nomenclature ................................................................................................................................................ 2

2.0 Hardware ................................................................................................................................................ 3

2.1 Reaction Control Engine ............................................................................................................... 3

2.2 Test Facility .................................................................................................................................. 3

2.3 Exciter Units ................................................................................................................................. 4

2.3.1 Variable Energy Exciter.................................................................................................... 4

2.3.2 Champion.......................................................................................................................... 5

2.3.3 Unison ............................................................................................................................... 5

2.4 Spark Plug and Ignition Cable ...................................................................................................... 5

2.5 Diagnostics .................................................................................................................................... 5

3.0 Experiment ............................................................................................................................................. 6

3.1 Test Procedure/Operations ............................................................................................................ 6

3.2 Test Matrix .................................................................................................................................... 7

3.3 Data Reduction Method ................................................................................................................ 8

3.3.1 Energy Calculation............................................................................................................ 8

3.3.2 Identification of Ignition Spark....................................................................................... 10

4.0 Results .................................................................................................................................................. 11

4.1 Variable Energy Exciter .............................................................................................................. 11

4.1.1 Waveforms...................................................................................................................... 12

4.1.2 Energy Results ................................................................................................................ 12

4.2 Champion .................................................................................................................................... 15

4.3 Unison ......................................................................................................................................... 17

4.3.1 Waveforms...................................................................................................................... 18

4.3.2 Energy Results ................................................................................................................ 19

4.3.3 Ignition Spark.................................................................................................................. 20

5.0 Discussion ............................................................................................................................................ 21

5.1 Engine Environment—Effect on Spark ...................................................................................... 21

5.2 Ignition Types ............................................................................................................................. 24

5.2.1 Rumbling Ignition—I-Bit Test Series............................................................................. 26

6.0 Conclusions .......................................................................................................................................... 27

Appendix A.—Test Logs ............................................................................................................................ 31

References ................................................................................................................................................... 35

NASA/TM—2012-217611 1 

Experimental Investigation of Augmented Spark Ignition of a 

LO2/LCH4 Reaction Control Engine at Altitude Conditions 

Julie Kleinhenz, Charles Sarmiento, and William Marshall 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The use of nontoxic propellants in future exploration vehicles would enable safer, more costeffective mission scenarios. One promising “green” alternative to existing hypergols is liquid methane 

(LCH4) with liquid oxygen (LO2). To demonstrate performance and prove feasibility of this propellant 

combination, a 100-lbf LO2/LCH4 engine was developed and tested under the NASA Propulsion and 

Cryogenic Advanced Development project. A series of three test programs were performed at the NASA 

Glenn Research Center Altitude Combustion Stand in a low pressure environment: specific impulse and 

impulse bit performance tests (covered in previous publications) and ignition margin testing. 

High ignition energy is a perceived drawback of this propellant combination; one goal of the ignition 

margin test program was to explore ignition performance and reliability versus delivered spark energy. 

The other goal was to examine the sensitivity of ignition to spark timing and repetition rate. Three 

different exciter units were used with the engine’s augmented (torch) igniter. Captured waveforms 

indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks 

(in quiescent room air). This suggests that rising pressure and flow rate increase spark impedance and 

may at some point compromise an exciter’s ability to complete each spark. If encountered prior to 

ignition, the reduced spark energies of such quenched deliveries results in more erratic ignitions and 

adversely affects ignition probability. Although nonessential, sparks delivered after ignition were 

particularly susceptible to quenching. This occurred when the exciter output, or driving voltage, was 

inadequate to tolerate the accelerated pressure rise effects on spark impedance. However, ionization 

created by the flame did appear to lower the breakdown voltage, thus decreasing the high-voltage pulse 

requirements to initiate sparks after ignition. 

The timing of the sparks relative to the pressure/flow conditions also impacted the probability of 

ignition. Sparks occurring early in the flow could trigger ignition with energies as low as 1 to 6 mJ, 

though multiple, similarly timed sparks of 55 to 75 mJ were required for reliable ignition. Delayed spark 

application and reduced spark repetition rate both correlated with late and occasional failed ignitions. An 

optimum time interval for spark application and ignition therefore coincides with propellant introduction 

to the igniter and engine. Shifts of ignition timing were manifested by changes in the characteristics of the 

resulting ignition. 

1.0 Introduction 

Next generation propulsion systems are currently being developed to enable future exploration 

missions. Cryogenic bi-propellants are attractive for their high performance. Traditionally hydrogen has 

been favored as the fuel, but its thermal storage requirements (low boiling point) create additional vehicle 

dry mass, reducing potential payload capacity. Hypergolic propellants, such as hydrazine, are a common 

alternative but are highly toxic. The rigid handling and containment procedures are not cost efficient. 

Hydrocarbon fuels, methane in particular, are gaining appeal for their high performance potential (when 

paired with oxygen) and moderate thermal storage requirements. This combination is also nontoxic, 

enabling simpler handling and storage methods. Methane is also available from local resources on some 

exploration destinations, such as Mars. For example, it could be produced from the atmospheric Carbon 

Dioxide on Mars using Sabatier processes. 
NASA/TM—2012-217611 2 

Prior work with this propellant combination is limited, so a goal of the NASA Propulsion and 

Cryogenic Advanced Development (PCAD) project was to examine the feasibility and performance 

characteristics of these systems (Refs. 1 to 3, and 13). A 100-lbf Reaction Control Engine (RCE) was 

developed by Aerojet Corporation to study liquid oxygen/liquid methane (LO2/LCH4) combustion 

(Ref. 4) and was tested at altitude conditions using the Altitude Combustion Stand (ACS) at NASA Glenn 

Research Center. In particular, there was interest in demonstrating repeatable and reliable ignition of an 

engine for a wide range of propellant inlet temperatures. 

The first series of tests (Refs. 5 and 6) explored the specific impulse (Isp) performance with burn 

times up to 7 s. The engine met its Isp goal, and demonstrated that performance improved as propellant 

temperature increased or as mixture ratio decreased, which is consistent with previous studies. The next 

test series examined the impulse bit (I-bit) performance for pulsed mode operation (Ref. 7). Again, 

the engine met its goals, achieving I-bits <4 lbf-s with pulse durations under 80 ms. As many as 30 

consecutive pulses at a 5-percent duty cycle were repeatedly achieved. 

In the final test series, which will be covered in this document, ignition characteristics were examined 

during engine hot fires at altitude conditions. Since one of the perceived drawbacks of this propellant 

combination is a need for higher ignition energy (relative to oxygen/hydrogen), the goal was to explore 

ignition performance and reliability versus delivered spark energy. Aerojet adopted an augmented, or 

spark torch, igniter configuration in this engine to enable better control of the spark environment (versus a 

direct spark ignition within the main chamber), as demonstrated in their previous oxygen/hydrocarbon 

engine designs. Based on this previous engine work (Refs. 8 and 9), it was expected that 40- to 50-mJ 

sparks would be required to obtain reliable ignition of the oxygen/methane propellants, especially at low 

inlet temperatures where liquid/liquid spray injection can prevail. A Variable Energy Exciter unit was 

developed in-house to explore the ignition energy limits. Exciters used in the previous Isp and I-bit test 

series were also tested in order to help interpret the ignition behaviors observed in those tests. Here, the 

exciter units were instrumented to obtain high fidelity spark current and voltage waveform information. 

By coupling such data to the pressure and flow measurements in the engine, it was possible to identify the 

ignition spark in terms of energy and timing. 

Nomenclature 

ACS Altitude Combustion Stand 

BDV breakdown voltage 

DSO digital storage oscilloscope 

HV high voltage 

Isp specific impulse(s) 

I-bit impulse bit (lbf-s) 

MR Mixture Ratio (oxidizer/fuel) 

PCAD Propulsion and Cryogenic Advanced Development 

PCFS Propellant Conditioning Feed System 

PLC Programmable Logic Controller 

RBD reverse breakdown 

RCE Reaction Control Engine 

sps sparks per second 

Vdc Volts direct current 

VEE Variable Energy Exciter 

 
NASA/TM—2012-217611 3 

Figure 1: The 100-lbf RCE operating at altitude 

conditions in the ACS facility. 

 Figure 2: A schematic of 

the igniter cavity. 

2.0 Hardware 

2.1 Reaction Control Engine 

All ignition tests were performed using the Aerojet 100-lbf RCE (Ref. 4 and Figure 1). This radiativecooled 45:1 columbium nozzle was designed to operate at a nominal mixture ratio of 2.5 and a chamber 

pressure of 175 psia. The details of the configuration are described in detail in Reference 5. With the 

exception of exciter units, the tested engine configuration remained unchanged throughout all the test 

series. 

The engine is designed for use with an augmented, or spark torch, igniter with a capacitive discharge 

exciter (Figure 2). Ignition therefore consists of lighting the spark torch, which then ignites the main 

chamber. Unlike traditional augmented ignition systems, propellant flow to the igniter in this engine was 

integrated into the main feed system and not independently controlled. This approach minimized engine 

mass and manifold volumes, while reducing dribble volumes and increasing pulse performance. 

Therefore, ignition tests necessarily involved a main chamber burn. Approximately 2.5 percent of the 

total propellant flow to the engine was directed to the igniter cavity, where the nominal mixture ratio was 

approximately 1.82. 

2.2 Test Facility 

The ACS facility is located at the NASA Glenn Research Center. The recently restored facility 

became operational with this RCE test program. The facility can accommodate a 2000-lbf-class engine 

with chamber pressure up to 1000 psia and can accommodate liquid and gaseous oxygen, liquid and 

gaseous hydrogen, LCH4, and other hydrocarbon propellants. A water spray cart and a water-cooled 

diffuser/multistage ejector system are used to condense the products of combustion and draw vacuum 

conditions inside the test capsule to simulate altitudes up to 130,000 ft. 

The system is well instrumented, with over 100 channels dedicated to the research hardware alone. 

The data is logged at 1000 Hz using a National Instruments module. Data can be streamed to multiple 

displays simultaneously for quick turnaround. All system timing is controlled using a Programmable 

Logic Controller (PLC) which has a 10 ms resolution. 

To accommodate the propellant temperature control needed in this investigation, the standard feed 

system was replaced with two propellant conditioning feed systems (PCFSs). These were developed by 

Sierra Lobo, Inc., to provide precise control of propellant conditions (Refs. 10 and 11). The oxygen PCFS 

utilized liquid nitrogen as the conditioning fluid whereas the methane PCFS used liquid argon with a 

recirculation loop. 

 
NASA/TM—2012-217611 4 

Table 1: Specifications of the three exciters used in the Ignition Margin Program. 

 Variable Energy Exciter 

(VEE) 

Unison Champion 

Spark rate, sparks/sec Variable, 125 to 250 200 100 

Delivered spark energy,a mJ 30 to 50 55 to 65 12 

Spark discharge time, μs 100 to 400 200 50 

Time to first spark, ms 21 6.5 3 to 9 

a

As measured for quiescent air spark gap contents. 

2.3 Exciter Units 

As mentioned, three exciter units were used in the course of the ignition margin test program. All 

were variants of the capacitive discharge ignition type. The first unit, called the Variable Energy Exciter 

(VEE), was developed in-house to permit energy and spark rate variation. The other two units, made by 

Champion and Unison, had been used in the Isp performance and I-bit test programs, respectively. The 

three exciters provided a good mix of exciter types and capabilities. Note that the Champion and Unison 

units were chosen based on availability, and the results presented here should not be construed as an 

endorsement of these units. The same spark plug was used with all exciters. 

All units were instrumented with voltage and current probes, which were located on the ignition cable 

near the exciter. The same cable was used for all exciters. These probes were not vacuum compatible, 

therefore, the exciters were located outside the vacuum vessel. The high-voltage ignition cable penetrated 

the vessel within a pressurized sleeve to prevent corona discharge. The voltage and current data were 

captured by a digital storage oscilloscope (DSO) for high-speed waveform data acquisition. 

Historically, spark ignition systems have been operated using spark trains, which is a series of sparks 

triggered over a specified time. This is done to ruggedize the system against variations in propellant flow. 

The presence of a flow field, especially a turbulent one, can reduce ignition probability for a single spark 

(Ref. 12). For most of these tests, the exciter was active for 60 ms, where the number of sparks occurring 

in that time varied depending on the spark rate of the exciter (Table 1). 

2.3.1 Variable Energy Exciter 

The VEE was an unsealed, breadboard version ignition unit. It used capacitor energy storage and 

discharge to drive unipolar sparks at its output. Distinct from most such capacitive discharge designs, it 

employed inductive (transformer flyback) high-voltage (HV), or ionization pulse generation to initiate the 

sparks. The VEE’s open-circuit HV pulses (approx. 20 kV) were reduced to 9 to 10 kV at the spark plug 

due to the capacitive loading of the ignition cable. Once initiated, sparks were driven by the voltage and 

energy available from a 4.4-μF storage capacitor via a resistor-inductor path bypassing the output 

transformer. Available spark energy was adjusted by variation of the storage capacitor charge voltage 

(160 to 280 Vdc (volts direct current)) using a digipot control. After transmission losses, the 

corresponding stored energy (56 to 172 mJ) yielded a 30 to 50 mJ delivered spark energy range at the 

spark plug tip in benchtop tests. Deliberate selection of lower (<160 V) charge voltages and stored 

energies was avoided due to an inadequacy of such low capacitor voltages to sustain sparks until most of 

the stored energy could be discharged. The duration of individual sparks, or spark discharge time, 

increased as the spark energy setting was decreased. Spark rate was adjusted by variation of capacitor 

charging duty cycle and ionization pulse repetition rate using a second digipot. The VEE was powered by 

a preapplied 28 Vdc input and triggered to fire by a separate command signal that resulted in a first spark 

21 ms later. 

This exciter unit was developed specifically for this ignition test program. With variable spark rate 

and energy, ignition margins (limits) could be explored. This unit was not used in any prior engine 

testing, but had undergone checkout testing in a vacuum environment. 
NASA/TM—2012-217611 5 

2.3.2 Champion 

The Champion exciter (model number CH92111, which was a modified 305131 single channel 

design) was a hermetically sealed, vacuum-compatible, capacitive-discharge unit. Operable from a 26 to 

30 Vdc input, it generated bipolar sparks from a capacitor with 94 mJ of stored energy. Delivered spark 

energies were 12 mJ, as measured at the spark gap in room air conditions. The unit produced a spark rate 

of 100 sparks per second (sps) as it periodically generated ionization pulses rated for 20 kV, but reduced 

to 9 to 10 kV at the spark plug due to ignition cable loading. It was triggered to fire (commanded) by 

application of 28 Vdc input power that resulted in a somewhat jittery 3 to 9 ms time to first spark. 

This exciter unit was used during the Isp test program where several nonignition events occurred 

(Ref. 5). Potential causes for these nonignition events were sought using the additional instrumentation 

available in this ignition program. 

2.3.3 Unison 

The Unison exciter (part number 500700–1) was another hermetically sealed, vacuum-compatible, 

capacitive-discharge unit. It had a welded 1.37 m length of shielded ignition cable that was spliced to the 

unshielded facility cable. From a 23 to 36-Vdc input voltage, it delivered 55 to 65 mJ unipolar sparks 

from a capacitor with 160 mJ of stored energy at a rate of 200 sps. Sparks were initiated by the unit’s 

13-kV ionization pulses (5 to 6 kV at spark plug due to cable loading). The Unison was commanded by a 

switched application of 28 Vdc input power, yielding a 6.5 ms time to first spark. 

This exciter unit was used during the I-bit test program, where a few nonignition events were 

observed (Ref. 7). Like the Champion exciter, the Unison exciter was characterized here with additional 

instrumentation to explore possible causes for these events. 

2.4 Spark Plug and Ignition Cable 

A single modified Champion spark plug that featured a button-tipped electrode with a 0.64-mm 

igniter annular spark gap was used for all exciters. It was connected to the exciters using a 3.65-m, 

unshielded, silicone-insulated, and atmosphere-sleeved ignition cable. Cable shielding was eliminated to 

facilitate insertion of diagnostic probes and to avoid its substantial capacitive load on the exciter output. 

Despite this measure, the remaining capacitive load on the exciter outputs by the long cable still 

significantly attenuated open-circuit HV pulse magnitudes delivered to the igniter spark gap. Spark 

current return path was provided by a low-inductance engine and facility ground structure to which the 

exciters were externally bonded. 

2.5 Diagnostics 

A Tek P6015A, 1000:1 voltage probe was used to monitor spark discharge voltage. Its large 

attenuation ratio was needed to protect the oscilloscope amplifiers from the multi-kV breakdown, or 

ionization, pulses that initiate each spark. Because the probe was not vacuum compatible, it was necessary 

to locate it exterior to the test cell, hence at the exciter output rather than at the spark plug, during tests 

under altitude conditions. The voltage drop across this ignition cable length for the 30- to 80-V sparks 

were verified to be negligible (≤2 V) by room air gap spark calibrations using a pair of probes that 

monitored exciter and spark gap ends of cable. 

Spark current pulses were monitored with a Pearson model 110 current transformer threaded by a 

single pass of the ignition cable. It had a 0.1 V/A output and 20 MHz bandwidth. Spark pulses were 

verified to be well within the probe’s current-time maximum limit, which ensures that the coil does not 

saturate or distort the waveform. 

 
NASA/TM—2012-217611 6 

Waveform data were recorded using a Tek 3054B, 500-MHz bandwidth, four-channel DSO with a 

9-bit vertical resolution and a memory length of 10,000 samples per channel. Both exciter output 

(discharge voltage and current) and input signals were logged. DSO vertical and horizontal time scales 

were set to capture spark trains with adequate resolution to discern individual spark details. As an 

example, the 4 ms/division, or a 40-ms span used for most runs corresponded to a 250-kHz sample rate. 

For the Unison and VEE exciters, seven sparks were captured for each spark train. For the champion, only 

one or two sparks could be captured with adequate resolution. The oscilloscope was triggered by the 

spark command signal from the PLC so that the data could be synchronized with the engine data. 

However, the accuracy of this synchronization was limited to the nearest millisecond as a result of the 

lower 1 kHz engine data sample rate. 

3.0 Experiment 

3.1 Test Procedure/Operations 

Ideally, the engine would be fired in pulsed operation mode, where multiple ignition opportunities 

would be possible per test. However, the data from the digital oscilloscope could not be output to memory 

quickly enough to permit this. Therefore, each ignition test was a single engine cycle. Typically, 10 tests 

could be performed in 1 day, though as many as 23 were possible depending on the test matrix. Once 

ignited, the burn duration was 1 s at a simulated altitude of 75,000 to 120,000 ft (0.2 to 0.5 psia). 

Timing of exciter command signals was constrained by the facility PLC to 10 ms steps coinciding 

with, or offset from, the thruster valve commands that activated propellant flow. As with the previous test 

series (Refs. 5 and 7), the LO2 flow was initiated first, with LCH4 flow initiation 10 ms later. The exciter 

was commanded using four different timing scenarios: simultaneous with the LCH4 valve command, 

10 ms after LCH4 valve command, 20 ms after LCH4 valve command, and simultaneous with the LO2

valve command. These scenarios were chosen to help identify the igniting spark within the spark train and 

to explore different ignition behaviors. Note that in the previous test series, the exciter was always 

initiated prior to propellant flow (10 ms before the LO2 valve command). 

Because spark timing and ignition events occur on time scales of milliseconds, it is important to 

consider the system response. Figure 3 and Table 2 illustrate the timing relationships accounting for the 

response times of the exciter and flow valves. Feedback signals from the propellant valves are shown by 

the dashed lines. The rise in voltage (nonzero) indicates the valve has been commanded open, while the 

dip (approx. 10 ms later) indicates the valve response. When the voltage leveled off about 6 ms later, the 

valve was fully open. This coincided with a pressure rise in the propellant manifolds, shown by the thick 

solid lines, which confirmed that propellant was flowing. The vertical lines show the exciter command 

timings for the Unison and VEE. The symbols represent the average first spark times for the VEE (circle) 

and Unison (square), while the arrowed lines indicate the time to first spark for each timing scenario. As 

can be seen, the different exciter command timings yielded first sparks that occurred at various stages of 

propellant flow, including during LO2-only flow, and various mixed propellant conditions. 

Historically, spark ignition systems have been operated using spark trains, which is a series of sparks 

triggered over a specified time. This has presumably been done to ruggedize the system against variations 

in propellant flow and local mixture ratio. Typically, the exciter was commanded ‘on’ for 60 ms. The 

number of sparks occurring in that time varied depending on the spark rate of the exciter and the time to 

first spark. Spark command durations longer than 60 ms were not used in order to avoid potential late, or 

hard-start ignitions, particularly when employing delayed spark initiation timings. In addition, DSO 

memory limitations would have precluded capture of any longer spark trains with adequate resolution to 

discern individual spark details. 

 
NASA/TM—2012-217611 7 

Figure 3: The command timing of the propellants and exciter are shown. The dotted lines represent the valve 

feedback signals. It takes approx. 15 ms from the command until the valve is fully open. The solid lines show 

the manifold pressure response, where an increase indicates propellant flow. The vertical lines indicate the 

exciter activation and the arrows represent the time until the first spark occurs. 

Table 2: Command times are shown for propellant flow (top) and exciter (bottom). All numbers were taken from the 

data, as opposed to the programmed setpoint. Test to test variations are ±2 ms for any given value. The champion 

exciter experienced variability in response time and will be discussed later. 

Propellants Valve command, s Valve full open, s Manifold pressure rise, s

Oxygen 31.058 31.073 31.071 

Methane 31.068 31.084 31.082 

Spark commands Command signal, s VEE first spark, s Unison first spark, s

With oxygen command 31.058 31.080 31.064 

With methane command 31.068 31.092 31.074 

10 ms after methane command 31.078 NA 31.084 

20 ms after methane command 31.088 NA 31.094 

3.2 Test Matrix 

A total of 87 runs were performed during the ignition margin test program using the three exciter 

units. While the exciter conditions were varied according to each exciter’s capabilities, the engine 

conditions were kept constant. The mixture ratio, which is governed by the run tank pressures, was set 

according to the Isp test program. Therefore, if the engine was fired long enough to reach steady-state flow 

conditions, these setpoints would result in a mixture ratio (O/F) of 2.5. Likewise, the propellant timing 

scenarios were retained from the previous test programs; the oxidizer flow was initiated 10 ms prior to the 

fuel, and the fuel was shut down 20 ms after the oxidizer. The oxidizer lead was shown in early testing to 

reduce the risk of hard starts, and was maintained as a safety precaution. Likewise, the fuel lag on 
NASA/TM—2012-217611 8 

Figure 4: The propellant temperature conditions, as measured at the thruster valve, are shown for 

all tests. These are average temperatures over the ignition interval: starting at the initiation of 

methane flow (31.082 s) and ending 60 ms later. The target temperature ranges are shown, 

including the expanded target from the specific impulse test program (Ref. 7). 

shutdown reduced the risk of oxidation on the hot chamber walls. During tests with the Unison exciter, 

propellant temperature (measured at the fire valve) was varied among the previously established setpoint 

conditions; cold (170 R LCH4 and 163 °R LO2), nominal (204 °R), and warm (224 °R). Figure 4 shows 

the temperatures that were achieved for the tests. 

The test record for all tests is listed in Appendix A. The VEE was used in 28 hot fire tests over the 

course of 2 days, encompassing run numbers 293 to 329 (span includes cold flow and calibration runs). 

During the first day, only the energy setpoint was varied, while both energy and spark rate were varied the 

second day. The Champion exciter was tested for just 1 day, with 15 hot fire tests encompassing run 

numbers 330 to 348. Since only one to two sparks could be captured within the data resolution of the 

oscilloscope, the oscilloscope timing was the sole variable in these tests. The final 3 test days employed 

the Unison exciter, using a different propellant temperature each day. A total of 44 hot fire tests were 

performed with run numbers from 349 to 401. Spark command timing was the varied within each test 

day. 

3.3 Data Reduction Method 

3.3.1 Energy Calculation 

Spark energies were calculated using the discharge voltage and current data from the digital 

oscilloscope. The following method is applicable for the VEE and Unison exciters, but is not practical for 

the Champion exciter due to the difficulties of adequately sampling and processing its short-duration 

bipolar waveform. Although single spark capture could be resolved using a high DSO sample rate, the 

Champion’s erratic time to first spark in this application, the low duty cycle of its spark pulse trains, and 
NASA/TM—2012-217611 9 

the limited DSO memory (10 kS/waveform) precluded a workable approach for obtaining spark energies 

in ignition test runs (section 4.2). 

Figure 5 shows an example waveform for a single spark from the Unison exciter. (A VEE waveform 

was similar.) The discharge voltage (circles) and current (triangles) waveforms for a single spark are 

shown at the bottom. A spark discharge is indicated when the current level rises above its zero level. 

Energy was determined by calculating and then integrating the power level during this period, as shown 

in Equation (1) (where V is voltage, I is current, and t is time). For the Unison exciter, the “zero” level for 

the current was 3 amps (Figure 5 , upper right insert) and for the VEE it was 0.4 Amps. (Note that the 

Unison signal was unfiltered resulting in a higher zero offset.) 

   E  V  I dt spark (1) 

All data reduction was performed using IgorTM graphing software with its programming interface. 

Power was calculated only when the current exceeded a threshold value (zero offset), and the resulting 

array was scanned to determine spark times. A spark began when the power level exceeded zero for more 

than two time steps and ended when power dropped below zero for two time steps. These regions were 

then integrated to obtain energy for each spark. The squares at the top of Figure 5 show the power 

calculation for each time step and the shaded region indicates the integration time. The energy for this 

spark was 63 mJ. 

Figure 5: An example energy calculation is shown. The voltage and current of a single spark is shown at the 

bottom, with the power at the top. Energy is calculated by integrating the power curve during the time the 

current exceeds its noise floor, represented by the highlighted region. The insert graph illustrates the 

noise level for the Unison exciter.
NASA/TM—2012-217611 10 

3.3.2 Identification of Ignition Spark 

Since the individual spark energies in a spark train may vary, determining ignition energy required 

the identification of the ignition spark. For the Unison exciter, there were as many as seven consecutive 

sparks per test, while the VEE had up to nine sparks. Ignition was indicated by a rise in pressure in both 

the igniter cavity and the engine chamber. The spark waveform data was therefore synchronized to the 

engine pressure data to determine ignition. Figure 6 and Figure 7 show examples of this overlaid data. 

The sparks are indicated by spikes in the voltage and current signals (lower left axis), and the energy 

levels of each spark are indicated at the top. The dashed lines indicate the initiation of the propellant flow; 

at approximately 31.074 s for LO2 and approximately 31.084 s for LCH4. The squares and circles denote 

the chamber and igniter cavity pressures, respectively, while the open symbols show cold flow pressures 

for comparison. Ignition occurs when the igniter and chamber pressures deviate from the cold flow 

values. In Figure 6, the sparks began 20 ms before CH4 flow, and ignition occurred at approximately 

31.086 s, coinciding with the fifth spark. However, considering a finite response time for the pressure 

measurement, it is likely that the fourth spark triggered the ignition. This is evidenced by Figure 7, which 

is a VEE test using a low spark rate (the VEE is the only unit that permitted this variation). It is clear that 

the first spark triggered the ignition, but the pressure response lagged the spark by about 4 ms. Another 

indicator is the temperature in the igniter cavity, represented by the solid line, which begins to rise about 

8 to 10 ms after the ignition spark. When these delays are applied to Figure 6, it indicates that the fourth 

spark caused ignition. This is corroborated by other tests with the VEE at varying spark rates as well as a 

Unison test with a shortened spark command time. 

Figure 6: An example of a unison exciter run with pressure traces plotted against the waveforms. 

Ignition occurs when the chamber and igniter cavity pressures exceed their cold flow (nonignition) 

levels. This is correlated to the spark waveforms to determine which spark triggered ignition. 
NASA/TM—2012-217611 11 

Figure 7: An example data set taken from the VEE with a low spark rate. The pressures rise 

occurs between the first and second sparks, so the first spark must have triggered the 

ignition. The pressure response is delayed by approx. 4 ms relative to the ignition spark and 

igniter cavity temperature has an 8- to 10-ms delay. 

4.0 Results 

4.1 Variable Energy Exciter 

The VEE underwent a total of 28 runs over 2 test days. This unit was designed such that stored spark 

energy and spark rate could be adjusted using digital potentiometers (digipots). Originally, the digipot 

settings were calibrated using air spark gap tests performed at room conditions with no propellant flow, 

namely “dry spark” tests. These calibrations were performed on the engine test stand, thus the 

configuration was identical to hot fire tests. Although the spark rates shown in Figure 8 are accurate, the 

delivered spark energy exhibited a dependence on the engine (hot fire) environment. Therefore, the dry 

spark energy calibration was not transferable to the hot fire tests. In the following sections, test conditions 

are identified by the digipot setpoint. Energy was varied between “E_900” and “E_500” while spark rate 

was varied between “S_850” to “S_200.” The resulting spark energies and rates in test runs were then 

determined from the actual hot fire data. 

While the energy range shown in Figure 8 was smaller, and lower, than anticipated based on initial 

benchtop testing, it indicates that resulting spark energies should scale with the digipot setpoint for a fixed 

spark gap. This holds true as long as the setpoint is above a minimum capacitor charge voltage that is 

sufficient to sustain sparks without quenching. For the dry spark conditions of Figure 8, this threshold 

appears to be E_500, corresponding to 15 mJ delivered energy (from 180 V capacitor charge, 71 mJ 

stored). However, as the data presented below will suggest, under the variable hot fire conditions this 

minimum VEE setpoint changes substantially and renders the digipot settings ineffective for controlling 

delivered spark energy. This behavior is attributed to inadequate capacitor charge or driving voltage for 

the desired energy range. It could likely be mitigated by modifying the VEE to maintain higher storage 

capacitor voltages for the target energy range. 
NASA/TM—2012-217611 12 

Figure 8: Calibration of the VEE digipot settings for dry sparks. 

The energy calibration is shown on the left axis, but is not 

applicable to hot fire tests. The spark rate is shown at right 

(squares) and is universal. 

4.1.1 Waveforms 

Examples of waveform and pressure data from a VEE test are shown during a hot fire engine test 

Figure 9(a) and in stagnant air at room conditions, or dry spark, Figure 9(b). As with the previous graphs, 

voltage, current, and pressure data are displayed on the bottom and calculated spark energies at the top. 

Spikes in the voltage and current signals indicate a spark occurrence. Figure 10 shows expanded views of 

the individual sparks of Figure 9. Only the first seven sparks (out of nine) are shown due to size restrictions. 

Comparison of the hot fire and dry spark runs show significant irregularities in spark behavior during 

hot fire conditions relative to the repeatable spark energy and waveform behavior in the dry spark 

situation. For the nominal dry spark behavior in Figure 10(b), the voltage is steady during the discharge, 

with spikes (ringing) at the beginning and end. In Figure 9(b), the voltage is an “h” shape with the 

spike(s) representing the spark discharge and a gradual voltage decline (the hump of “h” shape) marking 

the capacitor recharge. These patterns are evident in all of the dry sparks, but only for the first five hot fire 

sparks (Figure 9(a)). The latter sparks of the hot fire run clearly exhibit an absence of capacitor recharging 

cycles and little to no spark discharge current. The lack of voltage recharge indicates that the capacitor 

never fully discharged. In the case of spark 6, there was no current draw or recharge cycle at the end, 

indicating a dropped or absent spark. However, spark 7 shows a brief current draw but no substantial 

voltage recharge. This indicates that the spark was quenched, or truncated, with only a partial discharge of 

the capacitor occurring. 

4.1.2 Energy Results 

 Figure 11 shows the energy results from two sparks in each VEE test: the ignition spark and the 

spark with the highest energy. The data are grouped by the energy setpoint (indicated at top) while the 

size of the symbol represents spark rate. Raising the energy setpoint appears to increase the range of 

observed spark energies, although the high energies themselves are inconsistent. Ignition spark energy is 

similarly inconsistent, making it difficult to establish whether the notion of minimum ignition energy 

applies, or could be resolved, within the available VEE energy range. Surprisingly, in some cases such as 

run 318, the ignition energy appeared to be <1 mJ. 

0

50

100

150

200

250

300

10

12

14

16

18

20

22

24

26

28

30

150 250 350 450 550 650 750 850 950

Spark rate,sparks/sec

Average Spark Energy, mJ

VEE Digipot set points

Energy

Spark Rate
NASA/TM—2012-217611 13 

Figure 9: Example waveforms from the VEE. (a) A hot fire test indicates inconsistent spark behavior and (b) a dry 

spark test (no propellant flow so no pressure data) shows repeatable, consistent spark performance. 

Figure 10: Closeup plots of individual spark waveforms for the VEE. (a) A hot fire run showing irregular behavior 

and (b) the dry spark data shown which is very repeatable. 

These energy fluctuations also are clear in Figure 12, which shows spark trains for a group of tests 

with the same spark rate, but different energy setpoints. The ignition sparks in each run are outlined. The 

inconsistency in spark behavior over the course of the spark train is evident. Overall, the spark energies 

appear to decrease later in the spark train. In several tests, this decrease seems to be triggered by the 

ignition spark. An example is shown in Figure 9 (left), where the ignition occurred on spark 5, and sparks 

6 to 9 were quenched. Also note that the igniting spark was not always the highest energy prior to 

ignition. 
NASA/TM—2012-217611 14 

Figure 11: Ignition spark energy is shown for each VEE run along with the highest spark 

energy. Data is divided according to the setpoint energy. The higher energy setpoint had a 

larger range of resulting energies. Ignition energy is very inconsistent. 

 

Figure 12: Spark energies are shown for a group of VEE tests at the same spark rate but 

varying setpoint energy. Ignition sparks are outlined. The energies over the spark train 

vary widely but appear to decrease late in the spark train and after ignition has occurred. 
NASA/TM—2012-217611 15 

Figure 13: The VEE spark occurrences for each run are shown as a function of test time. Vertical lines denote 

first indications of propellant flow. The ignition spark is circled, and dropped sparks are crossed out. 

Figure 13 illustrates the timing of the ignition spark relative to the propellant flow. The spark trains 

for all the tests are plotted relative to the elapsed test time. The ignition sparks are circled. The vertical 

lines represent the response of the manifold pressure transducers, indicating propellant flow initiation. 

Note that the propellant valve open command is sent approximately 14 ms prior to this (it takes approx. 

10 ms for the valve to respond and another approx. 4 ms until the manifold pressure rises). The dropped 

sparks, which are crossed out, often occurred immediately after ignition. 

4.2 Champion 

One day of testing was performed using the Champion exciter with a total of 15 tests. There were 

several logistical difficulties that limited the scope of this test matrix. First, the bipolar spark waveforms 

were difficult to resolve. Any noise in the signal made it difficult to distinguish the voltage shifts. 

Numerous dry spark tests revealed that the voltage probe needed to be within 6 in. of the spark plug 

igniter to obtain a clean signal. However, the voltage probe was not vacuum compatible and therefore 

could not be used in this position during hot fire testing. 

The other issue concerned the data resolution of the digital oscilloscope, which could store 10,000 

data points for a variable time-base span. With the VEE, this recording time was chosen as 40 ms 

(250 points/ms) so that seven of the maximum nine sparks could be resolved. However, the spark 

discharge time for the Champion exciter was much shorter (30 to 50 μs) and its spark rate lower (100 sps) 

than the VEE (200 μs, 250 sps). Therefore, using the same time base as the VEE would mean that four 

 
NASA/TM—2012-217611 16 

Figure 14: Examples of dry spark waveforms from the Champion exciter. The oscilloscope data rate was adjusted 

to resolve the waveform. (a) A fully resolved waveform is shown but only permitted 4 ms of data collection. (b) A 

20 ms window was sufficient to confirm an unquenched spark, but not for energy calculations. (c) The same data 

rate used with the VEE does not have sufficient resolution. 

spark waveforms could be obtained, but would only be represented by 12 data points each. Figure 14(c) 

illustrates a spark waveform captured with this sample rate. The bipolar voltage peaks cannot be 

distinguished. Figure 14(a) shows a similar spark waveform captured using a shorter time base (higher 

sample rate) resulting in multiple points for each polarization shift. Using this time base, only 4 ms of 

data was collected so only one spark in the train could be captured. However, due to a 7-ms variability in 

Champion exciter response to the spark command signal (i.e., in the time to first spark), it was not 

possible to repeatedly capture a single spark within this 4-ms window. 

Since significant modification would be required to correct the noise on the voltage probe and/or 

extend the memory of the oscilloscope, energy measurements for the Champion exciter were abandoned. 

The hot fire tests were ultimately performed with a 20-ms oscilloscope time base (Figure 14(b)) that 

generally captured two sparks at low resolution. This compromise sample rate was chosen solely for 

qualitative observation of spark fidelity (i.e., nominal vs. quenched or dropped sparks). DSO trigger delay 

was shifted in each test to capture different sparks in the train of five to six sparks, as shown in Figure 15. 

The solid circle symbols represent sparks that were captured by the oscilloscope (±2 ms), while the open 

symbols are the predicted times for the remaining sparks based on a 100 sps rate. The ignition time, as 

indicated by a rise in the chamber pressure, is indicated by the black squares. No quenched sparks or 

dropouts were observed in the oscilloscope data, and the waveform behavior appeared repeatable. 

Nevertheless, 2 of the 15 tests (runs 337 and 347) failed to ignite. The energy of the spark shown in 

Figure 14(a) is approximately 16 mJ. The results suggest this spark energy is insufficient for repeatable 

ignition in an igniter/thruster environment, at least for a bipolar type of spark with a 100-sps rate. 

 
NASA/TM—2012-217611 17 

Figure 15: The spark timing of the Champion exciter is shown. The solid symbols represent those 

recorded by the oscilloscope (±2 ms), while the open symbols are inferred assuming a 100 sparks 

per second rate. The black squares show the time of the chamber cavity rise, indicating that ignition 

was likely caused by the preceding spark. 

It is also evident in Figure 15 that the spark timing was inconsistent (especially time to first spark). 

Likewise, ignitions tended to occur later relative to fuel introduction than with the VEE and Unison 

(section 4.3.3), where ignitions approximately coincided with the methane manifold pressure rise. Taking 

into account the approximately 4-ms LCH4 manifold response time (section 3.3.2), Figure 15 indicates 

that ignition usually was triggered by the first to third spark following methane entry (at approx. 

31.078 s). This ignition delay is likely attributable to the Champion’s lower spark rate combined with a 

less than 100 percent ignition probability for each spark. 

4.3 Unison 

The Unison exciter underwent a total of 44 tests in 3 test days. The characteristics of this exciter’s 

sparks proved to be the easiest to examine relative to the previous units. Therefore, this was the only unit 

to be tested at all three propellant temperature conditions. Because the unit was unipolar, like the VEE, 

energy calculations were straightforward. However, hot fire spark behavior was more consistent than with 

the VEE. The duration and duty cycle of the spark discharges were long enough such that the DSO could 

resolve the waveforms for all sparks in the train. 

Since the energy level was fixed for the Unison exciter, it was not useful for exploring energy 

margins. Therefore, the primary test goal was to examine ignition performance. The timing of the first 

spark with respect to propellant flow was varied in order to investigate its impact on ignition behavior. 
NASA/TM—2012-217611 18 

4.3.1 Waveforms 

Example waveforms are shown in Figure 16 and Figure 17. Figure 16 shows spark occurrence with 

respect to propellant flow, while Figure 17 shows closeups of each individual spark. As with previous 

plots, spikes in the voltage and current indicate a spark occurrence. In both figures, a hot fire test at 

nominal propellant temperature is shown in image (a) while a dry spark test is shown in image (b). The 

dry spark test shows very repeatable results, both in terms of energy (50 to 55 mJ) and discharge time. 

Indeed, individual dry spark waveforms in Figure 17(b) are nearly identical, but in the hot fire test, the 

waveforms are inconsistent. Spark 6, for example, almost appears to be quenched based on its short 

discharge time. However, its energy is quite high at 92 mJ whereas most of the others are around 60 to 

75 mJ. This suggests that a full spark did occur, but with a higher than usual spark impedance or 

discharge voltage (approx. 64 V). As a result, energy transfer from the storage capacitor to spark was 

more efficient, yielding the higher spark energy and reduced discharge time. 

Figure 16: Example waveforms from the Unison. (a) A hot fire test indicates inconsistent spark behavior and (b) a 

dry spark test (no propellant flow so no pressure data) shows repeatable, consistent spark performance. 

Figure 17: Closeup plots of individual spark waveforms for the Unison. (a) A hot fire run showing irregular behavior 

and (b) the dry spark data shown which is very repeatable. 
NASA/TM—2012-217611 19 

However, many of the highest spark energies were not valid, for example, the highest indicated 

energy in this test was the first spark, at 143 mJ (Figure 17(a)). The gap breakdown voltage (BDV) for the 

exciter HV/ionization pulse tends to be significantly higher for the first spark of train when there is no 

residual ionization from previous sparks or from a flame kernel. This voltage spike frequently overdrives 

the HV probe DSO amplifier, retarding its recovery after the overshoot induced by this abrupt gap voltage 

collapse. This artifact is apparent by the slowly relaxing, inverted (positive) polarity of the spark voltage 

for the initial approximately 1/3 of the discharge (Figure 17(a), spark 1). Given the unipolar spark current 

output of the Unison exciter, an inverted (positive) polarity for spark voltage cannot be valid. Since this 

period is when most of the spark energy is deposited, the initial inaccurate voltage reading can have a 

large impact on indicated spark energy. This effect was not always present due to the typical statistical 

fluctuations in first spark BDV. Likewise, this effect was also found to arise occasionally for other sparks 

in the train, presumably when their BDVs spiked either randomly or as a result of flow conditions. This 

first spark artifact was observed for the Unison and the VEE. For this reason, all spark waveforms were 

checked and their energies flagged as suspect when their initial polarity indicated a HV probe amplifier 

overdrive condition. 

4.3.2 Energy Results 

The spark energies for all the Unison exciter tests are shown in Figure 18, grouped by propellant 

temperature conditions. The dry spark energies are shown at each condition (thick black line) for 

reference. Some sparks energies were questionable due to the HV overdrive artifact discussed in a 

previous section. These are indicated by an “X” in the figure, while the sparks that triggered ignition are 

circled. In many cases it was an overdrive spark that caused ignition. In fact, the frequency of this 

coincidence suggests that higher BDVs and/or associated gap conditions may increase the ignition 

probability of a given spark. The energies in the cold and nominal cases were consistently higher than the 

dry spark baseline of 55 mJ, while the warm condition was similar to the dry spark. The data scatter is 

widest in the warm case, which may be due to the larger sample size (23 tests). Conversely, the energies 

were quite consistent at cold propellant conditions. There is a slight energy increase as the train proceeds, 

which is 

Figure 18: The spark energies are plotted for all the Unison exciter tests. Results are grouped by propellant 

temperature conditions with 9 cold runs, 12 nominal, and 23 warm. The thick line represents the dry spark 

energies, which are consistent at approx. 55 mJ. The igniting spark for each run is circled. Energies that are 

suspect due to a high-voltage overdrive issue are crossed out. 

 
NASA/TM—2012-217611 20 

particularly noticeable in the nominal temperature case. This may be due to the higher pressure/flow 

conditions late in the spark train, which drives up the spark impedance. This results in a higher energy 

transfer efficiency, and shorter spark duration. 

4.3.3 Ignition Spark 

The ignition sparks are highlighted in Figure 18. Many ignitions were triggered by sparks that 

experienced the HV overdrive artifact, which makes it difficult to draw conclusions about the ignition 

energy, therefore, instead of looking at energy, Figure 19 illustrates the dependency on spark timing. This 

plot is analogous to Figure 13 for the VEE where ignition sparks are circled and vertical lines indicate 

propellant flow initiation (the valve was commanded approx. 14 ms prior to flow initiation). The majority 

of igniting sparks occurred within ±3 ms of the methane manifold pressure rise, which indicates methane 

flow introduction. The pressure transducer response can lag the valve open command by as much as 3 to 

4 ms (section 3.1) even though it appears that some sparks occurred in an oxygen-only flow regime, this 

cannot be definitively proven within the uncertainties of the system response. All ignitions, but one, were 

triggered off the spark that immediately preceded or followed the methane manifold pressure rise. Even 

when the spark command was delayed until 20 ms after the methane valve (runs 371 to 373), it was the 

first spark, thus the first one after methane flow introduction, that triggered ignition. The exception was 

run 357, which will be discussed later as a unique ignition event termed a “rumbling ignition.” It is 

interesting to note that the first tests of the day were ignited by the spark immediately prior to methane 

flow, while the latter tests were ignited by the spark immediately following methane flow. At the warm 

propellant condition, most ignitions were triggered immediately after the methane flow, despite earlier 

opportunities. Such behavior possibly hints at an ignition pattern relating to propellant temperature, with 

warmer propellants tending to correlate with slightly later ignition. The trend seems to hold in the 

Figure 19: The Unison spark occurrences for each run are shown as a function of test time. Propellant 

flow initiation is indicated by vertical lines. Three propellant temperature conditions are represented, 

where each condition represents 1 test day. The ignition spark is circled. Dropped sparks are 

crossed. Generally, the spark at, or immediately following, the methane flow triggers ignition. 
NASA/TM—2012-217611 21 

Table 3: Unison tests with spark dropouts are summarized.

Run Command timing Time of 

first spark, s 

Duration, 

ms 

Total 

sparks 

Dropout 

spark no. 

386 10 ms after CH4 31.0843 60 7 1 

389 20 ms after CH4 31.0952 60 7 All 

392 20 ms after CH4 31.0652 60 7 1 to 3 

400 20 ms after CH4 31.0942 20 4 All 

nominal case, but the change in spark command timing after the third run of the cold propellant case 

excludes it as a reference to confirm this pattern of ignition behavior. This could be related to hardware 

temperature, which affects the injected propellant temperatures, and will be discussed in section 5.2. 

Only four tests failed to trigger ignition. All occurred at warm inlet conditions and all were 

attributable to deliberately unorthodox (early and late) spark command timings. Runs 397 and 398 both 

used a shortened spark train (only four sparks) applied prior to methane flow. Even though the sparks 

were successfully discharged, the propellant/flow conditions were insufficient for ignition. In runs 389 

and 400, no sparks were discharged. More detail is shown in Table 3, which indicates the spark command 

timing relative to the methane valve open command, the total duration of the spark train, and which 

sparks in the train were dropped. These tests were all performed at warm propellant conditions. The 

common factor is that these sparks were commanded to begin late in the test, well after propellant flow 

introduction. The igniter cold flow pressures already exceeded 20 psia before the first spark attempt. At 

this pressure, the gap’s estimated Paschen (DC) breakdown threshold is approximately 3.5 kV (for 

pressure*gap of 66 torr-cm) (Refs. 14 and 15), therefore, the cable-degraded Unison HV pulse capability 

of approximately 6 kV was marginal. This considers that microsecond HV pulse breakdowns, as used 

here, often substantially exceed Paschen thresholds due to the statistical time-lag effect1

 in absence of 

ambient ionization. Faster rise and shorter duration HV pulses can exaggerate this (fluctuating) overshoot 

of the Paschen limit. 

5.0 Discussion 

The following discussion primarily encompasses the VEE and Unison exciters. Discussion of the 

Champion exciter will be limited because of the previously described issues in obtaining high-resolution 

waveforms. Since its sparks were also bipolar, its waveforms could not be directly compared to the two 

unipolar units. 

5.1 Engine Environment—Effect on Spark 

It was observed in sections 4.1.1 and 4.3.1 that the spark behavior in an engine environment was 

significantly different than a dry spark (room air and no propellant flow). Both the VEE and Unison 

demonstrated repeatable spark waveforms in terms of discharge time, profile, and energy during dry spark 

tests. During hot fire engine tests, however, the waveforms and resulting energies were inconsistent 

within a single spark train. This suggests that the propellant flow field has a significant impact on spark 

behavior. 

 

 

1

Even with sufficient voltage, breakdown may be delayed due to the absence of avalanche initiating free electrons in 

the electrode gap. Without an artificial source (e.g., photoelectron emission by UV illumination), the rate at which 

free electrons randomly appear in this region is typically less than or comparable to the time scale of a 10 to 50 μs 

duration HV pulse. The HV pulses involved here are insufficient for electrode field emission of electrons. Thus, 

breakdown can be delayed until a free electron is available in the gap. This delay is known as statistical time lag. For 

fast-rising HV pulses, it often results in substantial “overshoot” of DC Paschen levels before breakdown is obtained. 
NASA/TM—2012-217611 22 

Figure 20: Average energy integration times indicate spark 

discharge time for the Unison and Variable Energy Exciter (VEE). 

Glassman (Ref. 12) noted that minimum ignition energy is higher in flowing mixtures due to a 

reduction in deposited energy density (energy per volume or per spark column length) as the spark is 

stretched downstream. Likewise, previous experiments (Refs. 16 and 17) of ignition in turbulent premixed 

methane and air mixtures noted that the strain rate fluctuations at the spark location prevented 

stabilization of the flame kernel. For turbulent nonpremixed flows, local mixture ratio fluctuations also 

affect ignitability near the spark. The randomness of these fluctuations causes ignition success to be 

probabilistic, therefore, a higher energy is typically needed to ensure ignition. 

Here, the flow effects are evident in the spark waveforms themselves, namely the spark discharge 

times. Figure 20 shows the average energy integration time for each spark over all VEE and Unison tests. 

In both cases, the discharge time decreased as the spark train proceeded. Since the spark train was applied 

coincident with the introduction and ramp up of propellant flow, this reduction in individual spark 

duration correlates with an increasing flow rate and pressure. This trend is more severe in the VEE, where 

ignition also seems to adversely impact spark behavior. 

As observed in section 4.1.2, quenched and dropped sparks frequently occurred immediately after 

ignition when using the VEE. These patterns of spark behavior may be primarily attributed to the effect of 

increasing pressure on spark discharge impedance and associated power dissipation. (While increased 

flow rate has a similar effect, flame kernel development after ignition would not be expected to accelerate 

flow through the spark gap.) Higher pressure raises spark voltage and power (I*V) for each discharge. 

Ultimately, there is a limit to an exciter’s capability to drive higher impedance sparks. This is governed 

by the combination of its storage capacitor voltage and internal series inductance, hereafter referred to as 

“capacitor driving voltage.” When exciter capacitor voltage is sufficient (as with Unison) to sustain 

higher voltage sparks, the corresponding higher spark power dissipation yields faster discharge of 

capacitor stored energy and shorter duration sparks. However, if exciter capacitor voltage is already 

marginal (as with VEE), then the higher spark voltages prevent full discharge of capacitor energy before 

sparks get interrupted, or quenched. This tendency is exacerbated when the VEE’s spark energy 

(capacitor voltage) is further reduced via the digipot setpoint. 

Another contributing factor, particularly for the VEE post-ignition apparent dropped sparks, could be 

flame-induced ionization. The level of flame-induced ionization is small compared to the ionization in the 

spark channels, therefore, it is unlikely to significantly affect discharge impedance. However, it is 

sufficient to significantly reduce pulse BDVs of the gap that initiate sparks. Under such circumstances, a 

weakness of the inductive flyback method (used by the VEE to produce HV pulses) is that it is 

susceptible to reverse polarity breakdown (RBD) at its output over the (pre-HV pulse) transformer 
NASA/TM—2012-217611 23 

“charging” interval. During this period, ramping of current through the transformer primary generates a 

moderate (approx. 800 V for VEE) reverse polarity voltage on the transformer secondary (VEE output). 

Normally, this remains open-circuit while its polarity flips upon deliberate, abrupt interruption of the 

transformer primary current. The induced collapse of transformer magnetic field then produces a desired 

HV pulse on the VEE output. However, if the VEE output is shorted (e.g., by a premature gap 

breakdown) during the transformer “charging” interval, this has been found to prevent the subsequent 

intended (correct polarity) HV pulse generation, gap breakdown, and discharge of the storage capacitor 

energy into a spark. (A possible example of this is evident in Figure 9 (left). The typical “h” shape of the 

voltage that would indicate a capacitor recharge is not present in later sparks. This indicates that the 

capacitor never discharged.) For these igniter gap dimensions, these moderate voltage RBDs have only 

been identified to occur for attempted sparking in post-ignition situations with the VEE. It has not been 

observed with flow prior to ignition, when gap cold flow pressures typically preclude triggering of any 

breakdowns by sub-kV transient voltages. (Note: RBV was observed in preflow conditions, when the 

low/vacuum pressures similarly reduced the Paschen limit.) Although a RBD still produces a shortduration (approx. 2 μs), weak (<1 mJ), reverse polarity spark powered by collapse of the exciter output 

transformer’s magnetic field (and not by a capacitor discharge), it manifests as an apparent dropped spark 

for the data acquisition sample rates (1 sample/4 μs) used here. 

Therefore, while reduced spark durations and quenched sparks are readily explained by pressureinduced elevation of spark impedance, it is also possible that post-ignition ionization conditions 

facilitated some RBD (at <800 V) dropped sparks for the VEE tests. Although the latter supposition might 

be argued to conflict with Paschen breakdown limits for the gap and pressures involved, the level of postignition ionization was likely sufficient to suppress or render invalid such limits. The presence of the 

flame-induced ionization could be sufficient to counter the pressure increase induced either by flow ramp 

up or by the ignition flame kernel itself. This is supported by the Unison data, for which no dropped 

sparks occurred after ignition, despite its lower delivered HV pulses (approx. 6 kV). 

The few Unison spark dropouts all occurred prior to ignition, typically only in late spark application 

conditions (Figure 15, runs 386, 389, 392, and 400). By these times, pressure in the igniter cavity had 

risen to a level that rendered the cable-attenuated Unison HV pulses (approx. 6 kV) marginal relative to 

the effective Paschen breakdown threshold (Refs. 14 and 15) for the gap. A similar explanation may apply 

to VEE dropped sparks that occurred late in instances of nonignition (Figure 15, runs 317, 319, 320, and 

323 to 325). However, unlike the VEE, the pressure rise caused by ignition did not cause dropped sparks 

for the Unison. This suggests that the post-ignition ionization may have been sufficient to make the gap 

mildly conductive, facilitating breakdown at the higher pressures by lowering BDVs substantially below 

the Paschen curve levels. The Unison had a higher capacitor driving voltage than the VEE, so provided its 

lower HV pulse could achieve breakdown and thereby initiate each spark, it was able to sustain these 

sparks to completion despite their elevated discharge impedances caused by the ignition pressure rise. 

In terms of minimum ignition energy, the lowest observed ignition was approximately 1 mJ with the 

VEE (run 318). This suggests a well-timed spark, even at low energy, can trigger ignition. However, the 

mix of ignitions and nonignitions for similarly timed 1 to 6 mJ quenched VEE sparks indicates that 

ignition for these low-energy sparks is less than reliable. Other studies involving turbulent and/or 

nonpremixed flows have noted a similar probabilistic nature with regard to minimum ignition energy 

(Ref. 16). The limitations of the current test program did not permit statistical significance to quantify 

such probabilities. While the Unison exciter did not permit energy variation, it was able to reliably ignite 

the flow. The majority of sparks occurred at or above the dry spark baseline of 55 mJ. The timing of the 

sparks was critical; the only failed ignitions occurred when sparks were limited to an interval before 

methane flow (runs 397 and 398) or delayed until after igniter pressure exceeded a level corresponding to 

Paschen breakdown constraints (runs 389 and 400). 

The limited Champion exciter appears to support these observations. None of the captured sparks 

from this unit appeared to be quenched, yet 2 of the 15 runs failed to ignite. The dry spark energy was 

lower than the other units at approximately 16 mJ, with the hot fire spark unlikely to be significantly 
NASA/TM—2012-217611 24 

higher. This, combined with the low spark rate, would decrease the probability of ignition. It should also 

be noted that the ignitions that were achieved frequently occurred late in the train. 

5.2 Ignition Types 

The timing of spark train relative to propellant flow was shown to influence spark performance, but 

also could impact the nature of the ignition itself. Four distinct ignition patterns were identified based on 

the pressure rise in the chamber and igniter cavity upon ignition. These are represented in Figure 21, 

which shows the igniter cavity pressure and chamber pressure rise. The open symbols indicate the cold 

flow pressures, so deviation from this indicates an ignition event. The first ignition type, termed a 

“gradual” ignition, is represented in Figure 21(a). The deviation from cold flow is marked by a gradual 

slope increase of the pressures in both the chamber and igniter cavity. This contrasts with an “abrupt” 

ignition (Figure 21(b)) where there is a sudden, sharp slope change in both pressure traces. A “late” 

ignition (Figure 21(c)) has the same basic characteristics as an abrupt ignition, but occurs at least 10 ms 

after fuel entry, when the cold flow pressure is higher. The pressure spike is also more severe than for the 

abrupt ignition. The final scenario is a “rumbling” ignition (Figure 21(d)), which was an unusual case. In 

all other scenarios, the chamber and igniter cavity pressure departures are nearly simultaneous. In the 

rumbling ignition, igniter cavity pressure begins to deviate from cold flow while main chamber pressure 

continues to track its cold flow profile. Rumbling ignition was only observed in one test of this series. 

However, it was also observed in the previous I-bit test series, where the behavior was more pronounced. 

Rumbling ignition will be discussed in more detail in section 5.2.1. 

Figure 21: Ignition types are represented by the igniter cavity pressure and 

chamber pressure traces. Deviation from the cold flow indicates ignition. 

(a) Gradual ignition, (b) Abrupt ignition, (c) Late ignition, and (d) Rumbling 

ignition. 
NASA/TM—2012-217611 25 

Figure 22: The ignition types are shown for each test of the (a) Unison and (b) VEE. Data are grouped according to 

test day. The first tests of each day have a gradual ignition while late tests have an abrupt ignition.

In Figure 22 the different ignition types are correlated to spark timing (symbols) and time of day 

(X-axis) for both the Unison (a) and VEE (b) exciters. For the Unison, there were 3 test days, each at a 

different propellant temperature. Both of the VEE test days were at nominal temperature. Generally, the 

first test of every day yielded a gradual ignition. Relating back to Figure 13 and Figure 19, these were 

also the runs that ignited on the spark just prior to methane flow. At the nominal and cold conditions this 

persisted for the first 3 to 4 runs, typically the point when a change in spark train parameters was 

implemented. Ignition then usually shifted to the spark after methane flow, which resulted in an abrupt 

ignition. When spark initiation was delayed beyond approximately 8 ms after methane entry (e.g., spark 

command with LCH4 valve, unless prevented by timing as with VEE day 2, or 20 ms after CH4 valve as 

in Unison warm and cold tests), late-type ignition resulted. Other changes made to spark train settings as 

each test day progressed, including reductions in capacitor driving voltage (reduced VEE spark energy), 

spark rate, and spark train duration were generally associated with more erratic ignition timing and some 

nonignitions. For example, VEE late ignitions occurred intermittently despite early spark initiation when 

lower spark energies (capacitor driving voltages) were selected (as in runs 300, 303, and 306). 

The ignition behavior shifts observed appears to be primarily a consequence of test sequence. While 

propellant temperature may have contributed a secondary influence, it had no clear effect on ignition 

characteristics for the ignition margin testing. Yet, for the pulse strings (sequential short engine burns) of 

I-bit testing (Ref. 7), propellant temperature did exhibit a distinct effect on ignition behavior, especially in 

regard to the incidence of rumbling ignition (section 5.2.1). The differing influence of propellant 

temperature between the two test series can be attributed to the hardware temperature operating 

conditions. 

In the single-shot ignition margin tests, a warm nitrogen purge through engine manifolds and igniter 

was maintained between each engine firing. This kept the hardware at relatively warm temperatures 

(approx. 480 to 500 °R) prior to each run, effectively resetting hardware conditions. Given the short 

duration of each run and the hardware temperatures relative to the much colder fluid temperatures, it is 

likely the propellant injected into igniter was ultimately gaseous regardless of initial propellant 

temperature or state upstream. Additionally, only minor MR variations would result. Valve propellant 

temperatures therefore had little effect on ignition. 

In contrast, pulse strings in the I-bit test series were commanded without intervening purge cycles. 

Igniter hardware was able to chill-in as each pulse string progressed. This allowed the hardware to 

achieve temperatures more comparable to the propellants, permitting the possibility of transition, or 

mixed-phase propellant injection into the igniter. Consequent changes of ignition behavior that occurred 

during the progression of each pulse string could thereby have been triggered by the developing spray 

injection and/or associated MR fluctuations. 
NASA/TM—2012-217611 26 

Figure 23: These three images show the igniter pressures from 18 consecutive pulses. The first six pulses in the left 

image have fluctuations indicative of a rumbling ignition. In the middle image the behavior becomes more 

consistent, culminating in very repeatable abrupt ignitions by the last six pulses at the right. 

These ignition behavior transitions are evident in Figure 23, which shows ignitions for 20 sequential 

pulses at nominal propellant conditions. The igniter cavity pressure traces for each pulse are plotted 

against a relative time scale in milliseconds. Cold flow pressure is also shown as the solid line. 

Figure 23(a) shows the first six pulses. The first pulse was a stable, gradual ignition but the next three 

pulses showed fluctuations in the igniter cavity, indicative of a rumbling ignition. The behavior then 

began to stabilize as abrupt ignitions for pulses 7 to 12, shown in Figure 23(b). The last six pulses 

(Figure 23(c)) show very repeatable abrupt ignitions. The transition from gradual ignition early in the test 

to abrupt ignition late for a pulse train in the I-bit test series exhibits the same trend observed for ignition 

tests over the course of a test day (Figure 22). 

5.2.1 Rumbling Ignition—I-Bit Test Series 

Perhaps the most interesting ignition case is that of the “rumbling” ignition. The data suggest that 

sparking initiated a reaction in the igniter torch cavity, possibly an unstable flame kernel that did not 

immediately induce ignition of the main chamber. While this ignition type only occurred once in this 

ignition margin test series, a revisit of the previous I-bit test series (Ref. 7) provides more evidence. 

During the I-bit test series, the engine was pulsed (short fired) up to 30 consecutive times. While no 

spark waveform data were captured for those tests, the igniter cavity pressure and chamber pressures were 

recorded and can be examined. Figure 24 shows a closeup of several rumbling ignitions during one of 

these pulse trains. Igniter cavity and chamber pressures are shown near the time of ignition, along with 

the cold flow baselines, for five successive pulses. A cold flow departure and fluctuation in the igniter 

cavity pressure is evident in all five pulses, but in the top three plots (pulses 8, 9, and 10) there is a 10- to 

15-ms delay before a corresponding rise in chamber pressure. For these cases, chamber pressure remains 

at the cold flow pressure, indicating the activity is initially isolated to the igniter cavity. Since these 

ignitions typically display mild oscillations in igniter pressure until main chamber ignition, they are 

referred to as rumbling ignitions. 

The majority of the rumbling ignitions observed during the I-bit test series occurred within the first 

10 pulses. It was most commonly encountered for cold propellant conditions where 15 percent of the 

131 pulses performed were classified as rumbling ignitions. Only two of those occurred at the end of the 

pulse train (after pulse 20). A total of 218 pulses were performed at nominal temperature, with 11 percent 

demonstrating this ignition type. At warm conditions, only 1 of the 170 total pulses exhibited a rumbling 

ignition, suggesting that propellant temperature and its relation to injection hardware temperature are key 

factors affecting its incidence. 
NASA/TM—2012-217611 27 

Figure 24: Ignition pressure traces illustrate rumbling ignitions (top three plots). Fluctuations in 

the igniter pressure indicate a reaction, but the chamber pressure does not respond. The data 

are taken from a pulsed fire engine test from the previous test program (run 255), where this 

type of ignition was more commonly observed. 

6.0 Conclusions 

The ignition margin test series was performed at NASA Glenn Research Center’s Altitude 

Combustion Stand using the 100-lbf LO2/LCH4 RCE. With the goal of exploring minimum energy 

requirements for this propellant combination, a total of three exciters were tested. All were instrumented 

with a high-speed digital oscilloscope to resolve current and voltage waveforms of the spark discharges. 

These data were synchronized with the engine pressure and temperature data to enable identification of 

the ignition spark. 

For all exciter units, spark waveforms in the hot fire, engine environment differed from those of dry 

spark tests (room condition, no-flow environment). The dry sparks demonstrated repeatable waveforms 

and consistent spark energies; whereas the hot fire tests had changing waveforms and inconsistent 

energies within a single spark train. As igniter pressure and flow rate increased during progression of each 

spark train, spark discharge voltages tended to increase and result in shorter spark durations. This 

indicates that rising pressures and flow rates substantially increase spark impedance, which, at some 

point, can affect an exciter’s ability to complete sparks without quench. That is, higher pressures and 

cross flows in the spark gap region can raise spark voltage to levels that exceed the limit that a particular 
NASA/TM—2012-217611 28 

exciter can tolerate. When an exciter is able to sustain sparks of increased voltage, the energy transferred 

to the sparks is marginally improved over that for baseline spark voltage levels. However, if spark 

voltages reach exciter limits, premature extinctions and/or reductions of spark energy (i.e., quenched 

sparks) typically result. If these features occur prior to ignition, this has an adverse impact on ignition 

promptness and reliability. 

After ignition, the accelerated pressure rise caused by the flame kernel had the most obvious effect on 

spark discharge characteristics. The Variable Energy Exciter (VEE), which had a higher (9 to 10 kV) 

high-voltage (HV) pulse, was able to achieve breakdown and initiate sparks at elevated pressures, but its 

low capacitor driving voltage was often unable to sustain, or complete each spark. Sparks occurring after 

ignition therefore tended to quench. This was not the case for the Unison which, with its higher capacitor 

driving voltage, always delivered full sparks after ignition. It might be surprising that the Unison’s lower 

HV pulse (6 kV) was sufficient to obtain breakdown in the higher post-ignition pressure environment. 

However, this may be attributed to ignition-induced flame ionization, which facilitated reduced 

breakdown voltages (BDVs), effectively invalidating Paschen-law pressure effects. By this same 

mechanism, flame ionization may have caused some post-ignition dropped sparks for the VEE, due to its 

circuit design. During preparation for HV pulse generation, moderate voltage on the VEE’s output 

transformer secondary may have been sufficient to yield breakdown in the mildly ionized gap 

environment. As a result, a very weak, inductive, reverse polarity spark occasionally replaced and 

prevented the intended capacitor discharge spark, resulting in an apparent dropped spark. 

Spark energies were obtained for two of the three units. With the third unit, the Champion exciter, 

spark energy capture proved to be incompatible with the data system capabilities. While the VEE was 

designed to allow precise control of spark energy, the engine environment caused a great deal of 

variability in waveform behavior and delivered energies. The unit’s energy adjustment only served to 

limit the overall range of spark energies, but the sparks themselves were not repeatable, primarily due to 

erratic spark quenches. The individual sparks that triggered ignition were identified, but also exhibited a 

large scatter in energies. In some cases ignition was triggered by quenched sparks of energies as low as 

1 to 6 mJ, although ignition by such sparks (in same time interval) was not reliable. By comparison, 

similarly timed, higher energy Unison sparks (55 to 75 mJ) regularly yielded ignition. 

Yet, a spark of this higher energy did not always trigger ignition, so no distinct energy threshold can 

be presented for reliable ignition by a spark in a given time interval. Despite this, relatively reliable 

ignition was experienced with an appropriately timed train of Unison sparks. Evidently, a sequence of 

several sparks over an appropriate time interval is required such that their cumulative probability of 

ignition approaches 100 percent. Ignition probability thus depends not just on spark energy, but also on 

spark timing relative to propellant flow introduction. This was further supported by the Unison exciter 

tests in which the timing of the spark train was adjusted. The spark nearest to the initiation of propellant 

flow (either immediately before or after) tended to trigger ignition. 

This ignition timing also impacted the nature of the ignition, as indicated by the pressure behavior in 

the igniter cavity and chamber. Four ignition scenarios were identified; the earlier ignitions caused a 

gradual pressure rise, while the later ignitions caused a more severe pressure rise. This was supported by 

data from the previous test series, which examined engine performance in pulsed mode operation. The 

pulses early in the train showed a gradual ignition, while the later pulses had abrupt ignitions. In the 

unique case of a “rumbling” ignition, which was primarily observed during pulse testing, the chamber 

pressure was unaffected by pressure fluctuations in the igniter cavity. In most cases, pressure response in 

the igniter cavity and chamber were simultaneous. The rumbling ignition suggests that an unstable flame 

kernel can develop in the igniter cavity. However, all rumbling ignitions did ultimately trigger main 

chamber ignition, albeit delayed. 

From the standpoint of future LO2/LCH4 engine design, these RCE tests have demonstrated that 

reliable ignition of the propellants is possible. While spark energy was found to exhibit a significant 

influence on this reliability, no distinct minimum spark energy for reliable ignition was found. Rather, 

ignition displayed a stochastic dependence on spark energy. Note that it is delivered spark energy that is 

relevant in this regard. For the dynamic environment of the engine torch igniter, delivered spark energy 
NASA/TM—2012-217611 29 

did not always correspond to the rated output spark energy of the exciter. It is therefore important to 

verify that exciter design characteristics, such as capacitor driving voltage, are sufficient to guarantee 

unquenched spark delivery under actual igniter conditions. While higher spark energies tend to increase 

ignition probability, spark timing and repetition rate also play a critical role. Lower energy sparks that are 

well timed with respect to propellant entry often achieve ignition. Observations indicate that sparks early 

in the flow ramp-up (in this case –4 to +10 ms relative to methane flow initiation) have the highest 

probability of success. Yet, a single well-timed higher energy spark alone does not guarantee ignition. 

Several sparks over the optimum time interval appear to be necessary to ensure an adequate cumulative 

probability of ignition. Thus, a high spark rate (here, 200 to 300 sparks per second) yielding more sparks 

in this optimum interval helps facilitate ignition reliability. Later sparks, timed after the optimum interval, 

can still trigger ignition, but do so with reduced probability. 

NASA/TM—2012-217611 31 

Appendix A.—Test Logs 

The following tables list all tests that were performed as part of this study. The information is 

abbreviated with respect to JANNAF standards to facilitate easier viewing of the parameters referenced in 

this document. Raw data is available upon request. 

 Run: The number of the test run as referenced to the first facility tests in the 100-lb RCE test 

program. 

 Test description: This column indicates whether or not an ignition occurred within each run. 

“Hotfire” refers to a successful ignition. 

 Target temperature: This column indicates the propellant temperature setpoint condition: cold 

(170 °R LCH4, 163 °R LO2), nominal (204 °R), and warm (224 °R). 

 Exciter timing: When the exciter was activated, relative to the propellant valve commands. 

“Simultaneous with LOX” or “with Lox cmd” indicates the exciter was commanded on at the 

same time the LOX valve was commanded open. The actual response times of the valves are 

described in section 3.1. 

 First captured spark: This indicates the time in which the first recorded spark occurred. For the 

VEE and Unison, this is the first spark of the train. But since only one or two sparks could be 

captured for the Champion, this indicates when the first captured spark occurred. This is the 

elapsed time relative to the PLC start. For all tests, both propellants were flowing at 31.082 s. 

 Spark duration: The exciter remained active for the amount of time specified here. 

 Number of sparks: The total number of sparks in the train. 

 (VEE ONLY) Energy Digipot setting: A digital potentiometer was used to vary spark energy. 

The setpoint value on the digipot dial is indicated here. (The resulting energies varied, so a 

calibration between setpoint and energy could not be represented.) 

 (VEE ONLY) Spark rate Digipot setting: A digital potentiometer was used to vary spark rate. 

The setpoint value on the digipot dial is indicated here as well as the resulting spark rate (in 

parentheses). 

 (Champion ONLY) Scope delay: This is the trigger delay used on the digital oscilloscope so 

the different sparks in the train could be captured. A higher delay would allow the capture of 

a spark late in the train. This number does not correlate directly to other DAQ times used in 

this paper, and is shown for completeness. 

 (Champion ONLY) Scope resolution: This is another oscilloscope setting, displayed here for 

completeness. This setting governs the data rate. 

 (Champion ONLY) Captured sparks: Only one to two sparks could be captured for the 

champion. This indicates which sparks in the train were captured. 

 Run tank pressures: The pressure in the propellant holding tanks. This is a pressure-driven 

flow system, so these pressures govern the mass flow rate of the propellants. Engine flow 

rates were highly transient during the ignition time. 

 Run tank temperature: The temperature of the propellant in the holding tanks. The desired conditions (cold, nominal, and warm) were judged based on the temperature at the thruster valve. 

Thus, the run tank temperatures were chosen to account for line losses. Note the temperatures 

are given in Fahrenheit, since this was the convention of the propellant conditioning systems. 

 Valve temperature: The temperature at the propellant thruster valves. 

 Ignition type: These designate the ignition characteristics as described in section 5.2. 

 Notes: Test specific notes. 

 Likely ignition spark: This indicates which spark in the train is believed to have triggered ignition. 

 Spark energy: The calculated energy for each spark in the train. 
NASA/TM—2012-217611 32 

 Variable Energy Exciter Run Test description Target temperature condition Exciter timing First captured spark, s Spark duration, ms Number of sparks Energy - Digipot setting Spark rate digitpot and (spark rate) Ignitio n type LO2 LCH4 LO2 LCH4 LO2 LCH4 NOTES Likely Ignition Spark 1 2 3 4 5 6 7 8 9 8/12/2010

293 Abort thrust cal294 Thrust cal295 Hotfire Nominal simulataneou with LOX 31.0802 60 9 900 850 (250spark/s) grad 325 265 -263 -264 193.3 195.0 Start at a high energy, and adjust timing 1 18.00 12.23 3.21 0.29 5.91 28.52 7.08 53.56 17.55296 Hotfire Nominal simulataneou with LOX 31.0802 60 9 900 850 (250spark/s) grad 325 265 -263 -264 201.5 194.9 1 21.18 15.45 0.00 0.00 7.68 27.08 10.95 8.38 4.93297 Hotfire Nominal simulataneou with LOX 31.0801 60 9 900 850 (250spark/s) grad 325 265 -262 -267 203.1 219.4 Methane temp was high 1 74.91 11.71 0.55 12.70 7.93 7.08 6.91 11.80 0.48298 Hotfire Nominal simulataneou with LOX 31.0802 60 9 800 850 (250spark/s) grad 325 265 -262 -268 198.6 199.5 1 0.88 0.00 0.00 0.00 0.00 12.08 8.98 0.00 3.54299 Hotfire Nominal simulataneou with LOX 31.0799 60 9 800 850 (250spark/s) abrupt 325 265 -262 -268 209.7 202.6 2 28.31 15.46 0.00 0.00 5.19 5.04 0.00 7.13 0.00300 Hotfire Nominal simulataneou with LOX 31.0802 60 9 800 850 (250spark/s) late 325 265 -262 -268 196.0 201.8 7 49.22 21.74 33.30 63.80 63.22 75.14 11.32 1.48 6.40301 Hotfire Nominal simulataneou with LOX 31.0801 60 9 700 850 (250spark/s) abrupt 325 265 -263 -266 196.8 200.4 3 23.73 12.03 41.53 10.33 14.86 5.46 4.76 7.02 3.55302 Hotfire Nominal simulataneou with LOX 31.0811 60 9 700 850 (250spark/s) abrupt 325 265 -263 -266 203.5 201.0 3 28.43 29.59 22.41 0.00 8.29 9.21 7.61 4.04 6.04303 Hotfire Nominal simulataneou with LOX 31.0802 60 9 700 850 (250spark/s) late 325 265 -263 -265 195.2 203.6 5 38.19 41.88 27.61 31.27 31.48 0.00 5.99 9.15 4.42304 Hotfire Nominal simulataneou with LOX 31.0802 60 9 500 850 (250spark/s) abrupt 325 265 -261 -264 202.4 201.3 3 13.21 14.96 20.77 0.00 4.84 5.90 3.97 3.15 4.44305 Nolight Nominal simulataneou with LOX 31.0801 60 9 500 850 (250spark/s) non 325 265 -261 -264 199.6 203.1 0 9.07 21.12 24.33 5.40 26.53 29.59 7.81 1.09 0.05306 Hotfire Nominal simulataneou with LOX 31.0801 60 9 500 850 (250spark/s) late 325 265 -261 -265 200.9 200.7 4 16.76 26.06 22.11 19.89 0.00 0.00 3.39 5.58 3.30307 Thrust cal

8/26/2010

308 Aborted Thrust cal309 Thrust cal310 Cold flow Nominal simulataneou with LOX 325 265 -265 -265 Pcign purge is high... 14psi, should be 5psi311 Cold flow Nominal simulataneou with LOX 325 265 -265 -265 Corrected Pcign purge312 Hotfire Nominal simulataneou with LOX 31.0802 60 8 900 850 (250spark/s) grad 325 265 -264 -268 211.4 198.8 1 13.72 0.00 0.00 32.90 28.90 4.98 7.63 26.40313 Hotfire Nominal simulataneou with LOX 31.0813 60 4 900 200 (125spark/s) grad 325 265 -264 -268 199.4 201.0 Ignited on first spark 1 58.94 0.00 0.00 4.32 314 Hotfire Nominal simulataneou with LOX 31.0812 60 4 900 200 (125spark/s) grad 325 265 -263 -269 203.0 204.0 Repeat previous 1 12.32 0.49 18.66 8.48 315 Hotfire Nominal simulataneou with CH4 31.0903 60 8 900 850 (250spark/s) late 325 265 -264 -266 206.5 199.1

Spark command with methane; still ignite 1st spark 1 61.29 0.00 4.91 27.40 5.62 30.55 6.42 0.00

316 Hotfire Nominal simulataneou with CH4 31.0902 60 7 900 850 (250spark/s) late 325 265 -264 -260 208.6 199.8 repeat 1 47.42 14.32 0.13 4.16 13.95 8.58 0.00 0.00317 Nolight Nominal simulataneou with CH4 31.0903 60 7 500 850 (250spark/s) non 325 265 -263 -264 211.5 199.7 Energy setpoint reduced 0 6.39 3.34 0.00 0.00 0.00 0.00 0.00 318 Hotfire Nominal simulataneou with CH4 31.0904 60 8 500 850 (250spark/s) late 325 265 -264 -267 216.6 202.0 repeat 1 0.00 4.17 0.00 3.59 0.00 3.32 3.02 0.00319 Nolight Nominal simulataneou with CH4 31.0903 60 4 500 200 (125spark/s) non 325 265 -264 -266 212.6 204.9

Reduce spark rate to clarify which ignites. NO sparks! 0 0.00 0.00 0.00 0.00 

320 Nolight Nominal simulataneou with CH4 31.0903 60 4 500 200 (125spark/s) non 325 265 -263 -264 206.6 203.4 try again 0 5.21 0.00 0.00 0.00 321 Hotfire Nominal simulataneou with LOX 31.0802 60 4 900 200 (125spark/s) grad 325 265 -263 -266 206.8 203.5

Go back to look at previous (sucessful) conditions.Try one here then reduce spark energy 1 16.63 0.00 4.74 5.42 

322 Hotfire Nominal simulataneou with LOX 31.0802 60 4 500 200 (125spark/s) grad 325 265 -263 -265 204.0 198.9 Energy setpoint reduced 1 5.40 0.00 0.00 0.00 323 Nolight Nominal simulataneou with LOX 31.0812 60 4 500 200 (125spark/s) non 325 265 -263 -263 204.7 200.9 0 4.10 0.00 0.00 0.00 324 Nolight Nominal simulataneou with LOX 31.0802 60 4 700 200 (125spark/s) non 325 265 -263 -264 206.8 200.4

Tried to get late ignition again, like in previous test day 0 7.39 0.00 0.00 0.00 

325 Nolight Nominal simulataneou with LOX 31.0802 60 4 700 200 (125spark/s) non 325 265 -262 -264 207.2 200.6 repeat 0 6.54 1.27 0.00 0.00 326 Hotfire Nominal simulataneou with LOX 31.0812 60 8 700 850 (250spark/s) abrupt 325 265 -259 -264 205.5 202.5 repeat condition from earlier day 2 6.71 3.08 0.00 0.00 0.00 4.07 2.20 2.54327 Hotfire Nominal simulataneou with LOX 31.0802 60 5 700 425 (143spark/s) abrupt 325 265 -262 -262 205.2 202.8

try mid range spark rate too see if rate affects ignition 2 5.36 14.75 4.39 3.33 0.00 

328 Cold flow Nominal simulataneou with LOX 325 265 -260 -262329 Thrust cal

Valve 

Temperature, °R

Run Tank Pressure, 

psig

Run Tank 

Temperature, °F Spark energy, mJ
NASA/TM—2012-217611 33 

 Champion Exciter

Run #

Test 

description

Target 

temperatu

re 

condition Exciter timing

First captured 

spark, s

Spark 

duration, ms

Scope delay, 

ms

Scope 

resolution, 

ms/division

Captured 

sparks, # in 

sequence LO2 LCH4 LO2 LCH4 LO2 LCH4 NOTES

9/16/2010

330 Dry spark331 thrust cal332 Facility abort Nominal simulataneous with CH4 60 12 2 325 265 -260 -263 Set pressure error on the ejector train333 Hotfire Nominal simulataneous with CH4 31.0823 60 12 2 2 325 265 -260 -263 206.4 200.8

Only 1 spark captured scope, looks like the next was igniting spark. Will shift scope timing 

334 Hotfire Nominal simulataneous with CH4 31.0774 60 14 2 1, 2 325 265 -261 -262 203.3 199.1 2 sparks captured, include ignition spark. Both look good -no drop out335 Hotfire Nominal simulataneous with CH4 31.0843 60 14 2 2 325 265 -261 -261 204.5 199.8 Repeat: 1 captured spark, not ignition spark. 336 Hotfire Nominal simulataneous with CH4 31.0783 60 16 2 2,3 325 265 -260 -262 200.2 200.3

Shifted scope timing again to catch likely ignition sparks. Late ignition. 2 captured sparks, neither ignition

337 No light Nominal simulataneous with CH4 31.0844 60 16 2 2 325 265 -261 -261 197.6 202.2 No ignition. Tign is decreasing so not even late ignition. Gas in both manifolds.

338 Hotfire Nominal simulataneous with CH4 31.0779 60 16 2 2,3 325 265 -261 -260 203.1 203.1

Increased input voltage to 29.5V (was 28V) to see if this increases spark rate & try to increase number sparks we see. Spark rate did not change signficantly. Late ignitions seem more regular now, so shift scope delay on next test

339 Hotfire Nominal simulataneous with CH4 31.0906 60 22 2 3 325 265 -260 -260 212.2 203.2 Increased scope delay. 1 spark captured after ignition.340 Hotfire Nominal simulataneous with CH4 31.0835 60 22 2 2,3 325 265 -260 -261 201.7 201.7

Turned off powersupply between tests to discharge. See if it effects 1st spark timeing. Did not effect 1st spark. 

341 Hotfire Nominal simulataneous with CH4 31.0882 60 22 2 3,4 325 265 -260 -261 216.9 201.8 repeat: 2 captured sparks, 2nd ignition spark. Both sparks look good.342 Hotfire Nominal simulataneous with CH4 31.0891 60 30 2 3,4 325 265 -259 -260 203.1 204.1 Increased scope delay. 2 captured sparks, 2nd ignition, both clean sparks. 343 Hotfire Nominal simulataneous with CH4 31.0711 60 30 2 1,2 325 265 -258 -260 201.0 204.2

Input voltage( trigger signal) spiked early. Scope time window wrong (too early). Don't know why, repeat to see

344 Hotfire Nominal simulataneous with CH4 31.0964 60 30 2 3,4 325 265 -259 -260 197.7 202.8 2 sparks both after ignition, both good.345 Hotfire Nominal simulataneous with CH4 31.1091 60 40 2 5 325 265 -257 -257 206.9 202.9 Shifted scope window to catch last sparks. 1 spark captured346 Hotfire Nominal simulataneous with CH4 31.0998 60 40 2 4,5 325 265 -259 -256 206.9 203.0 2 captured sparks, ignition spark, both good347 No light Nominal simulataneous with CH4 31.1044 60 40 2 4,5 325 265 -258 -256 202.9 202.6 2 captured sparks look fine348 thrust cal

Valve 

Temperature, °R

Run Tank Pressure, 

psig

Run Tank 

Temperature, °F
NASA/TM—2012-217611 34 Unison "brick" Exciter Run Test description Target temperatu re condition Exciter timing First captured spark, s Spark duration, ms Number of sparks Ignitio n type LO2 LCH4 LO2 LCH4 LO2 LCH4 NOTES Likely Ignition Spark 1 2 3 4 5 6 7 9/23/2010 349 Thrust cal 350 Abort Nominal simulatanous with Ch4 60 non 325 265 -263 -265 There were no sparks-power supply was off 351 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 grad 325 265 -262 -262 205.5 197.0 2 64.61 60.70 76.64 69.01 73.95 71.38 74.48 352 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 grad 325 265 -263 -261 206.6 201.3 2 143.73 63.07 74.52 72.80 76.17 92.74 68.89 353 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 grad 325 265 -259 -264 209.1 199.3 2 64.34 61.25 75.51 67.81 72.31 80.62 82.32 354 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 grad 325 265 -262 -264 202.8 200.0 2 60.69 63.09 74.47 70.97 71.33 86.70 82.70 355 Hotfire Nominal simulatanous with Ch4 31.0753 60 7 abrupt 325 265 -262 -264 208.6 199.5 3 59.25 68.87 47.42 71.58 71.87 68.88 68.89 356 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 abrupt 325 265 -262 -262 205.6 199.5 3 187.70 64.50 77.77 83.73 79.45 97.92 78.82 357 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 rumble 325 265 -260 -264 209.1 198.5 5 47.47 67.82 70.89 73.48 141.24 92.69 87.18 358 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 abrupt 325 265 -261 -264 212.1 202.5 3 96.84 66.29 136.08 76.31 76.77 87.47 80.29 359 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 abrupt 325 265 -261 -262 196.9 201.7 3 123.29 64.23 48.85 73.19 74.71 78.28 79.04 360 Hotfire Nominal simulatanous with Ch4 31.0743 60 7 abrupt 325 265 -262 -263 209.1 201.1 3 55.22 64.63 123.11 72.90 67.01 75.42 75.28 361 steady state run 3 362 Thrust cal 9/30/2010 363 Thrust cal 364 Cold flow Cold 325 285 -296 -281 365 Hotfire Cold simulatanous with Ch4 31.0743 60 7 grad 325 285 -296 -281 174.2 180.1 Looks like possible"rumbling ignition". Definate quench on spark #3. 2 55.81 54.11 66.12 65.31 64.12 63.26 71.53 366 Hotfire Cold simulatanous with Ch4 31.0742 60 7 grad 325 285 -269 -283 173.6 179.6 2 49.38 59.08 68.20 65.77 78.45 59.25 74.40 367 Hotfire Cold simulatanous with Ch4 31.0753 60 7 grad 325 285 -296 -282 176.3 179.0 2 167.16 56.00 69.89 62.16 65.90 64.58 76.06 368 Hotfire Cold 10ms after Ch4 31.0843 60 7 abrupt 325 285 -296 -278 177.1 178.2 1 146.95 63.10 61.36 66.93 67.01 71.14 83.08 369 Hotfire Cold 10ms after Ch4 31.0843 60 7 abrupt 325 285 -296 -282 175.8 177.8 1 71.26 65.94 60.24 63.38 81.56 69.57 69.56 370 Hotfire Cold 10ms after Ch4 31.0842 60 7 abrupt 325 285 -296 -283 177.5 177.7 1 76.81 63.88 61.60 65.32 67.89 64.06 63.19 371 Hotfire Cold 20ms after Ch4 31.0952 60 7 late 325 285 -295 -281 175.8 177.5 1 172.54 65.57 79.22 66.37 69.82 74.75 65.70

372 Hotfire Cold 20ms after Ch4 31.0942 60 7 late 325 285 -295 -280 177.4 177.4 1 45.63 76.71 61.28 65.53 60.80 66.59 60.26373 Hotfire Cold 20ms after Ch4 31.0942 60 7 late 325 285 -294 -280 178.5 177.7 1 116.27 65.04 63.58 74.76 60.89 63.79 59.96374 Cold flow Cold -290 -284375 Thrust cal

10/21/2010

376 Thrust cal377 Cold flow -240 -239378 Hotfire warm simulatanous with Ch4 31.0743 60 7 grad 370 265 -240 -238 225.5 220.9 Zero shift in current 2 42.48 53.56 46.19 55.82 58.60 62.78 55.78379 Hotfire warm simulatanous with Ch4 31.0753 60 7 abrupt 370 265 -239 -240 226.2 228.6 Zero shift in current 3 43.40 47.75 63.65 51.13 56.43 55.20 56.82380 Hotfire warm simulatanous with Ch4 31.0743 60 7 abrupt 370 265 -240 -238 226.4 225.7 3 44.32 52.99 69.15 49.74 54.51 61.28 60.46381 Hotfire warm simulatanous with Ch4 31.0743 60 7 abrupt 370 265 -239 -238 228.6 218.1 3 50.29 42.25 31.94 50.18 49.58 54.71 53.02382 Hotfire warm simulatanous with Ch4 31.0753 60 7 abrupt 370 265 -240 -239 226.7 216.3 3 58.34 48.50 62.67 60.65 49.20 59.49 53.80383 Hotfire warm simulatanous with Ch4 31.0742 60 7 abrupt 370 265 -239 -238 225.1 226.3 3 33.68 43.74 101.05 53.37 55.54 56.20 61.01384 Hotfire warm simulatanous with Ch4 31.0742 60 7 abrupt 370 265 -239 -239 219.0 226.3 3 48.49 43.71 131.03 54.47 60.40 63.68 55.10385 Hotfire warm simulatanous with Ch4 31.0733 60 7 abrupt 370 265 -239 -238 225.8 221.7 3 149.20 47.04 117.57 57.72 55.54 64.19 56.81386 Hotfire warm 10ms after Ch4 31.0843 60 7 abrupt 370 265 -239 -239 219.5 218.9 2 0.00 72.52 47.72 56.37 55.12 56.18 53.32387 Hotfire warm 10ms after Ch4 31.0843 60 7 grad? 370 265 -239 -237 227.9 221.8 1 77.91 60.73 57.90 52.55 60.29 49.20 53.95388 Hotfire warm 10ms after Ch4 31.0843 60 7 abrupt 370 265 -239 -238 225.2 222.5 2 127.06 50.26 52.61 55.71 46.08 55.37 61.88389 Nolight warm 20ms after Ch4 31.0952 60 7 non 370 265 -239 -238 222.8 223.5 No current draw on any spark -all dropout 0 0.00 0.00 0.00 0.00 0.00 0.00 0.00

390 Hotfire warm 20ms after Ch4 31.0943 60 7 late 370 265 -237 -237 226.7 223.2 1 50.70 51.93 57.51 47.92 57.14 57.71 55.95391 Hotfire warm 20ms after Ch4 31.0943 60 7 late 370 265 -239 -237 215.3 222.5 1 66.02 68.93 55.43 58.19 62.41 52.52 67.51392 Hotfire warm 20ms after Ch4 31.0652 60 7 late 370 265 -238 -238 226.6 224.9

First sparks dropped out... late ignition. Looks like later ignition means less reliable sparks 4 0.00 0.00 0.00 67.54 50.70 48.76 45.59

393 Hotfire warm simulatanous with LOX 31.0653 60 7 grad 370 265 -238 -237 219.5 217.2

The sparks before propellant flow don’t tell much. Had hoped to see timing shift with respect to manifold press rise, but not 4 151.64 45.55 42.87 53.78 44.03 54.47 47.63

394 Hotfire warm simulatanous with LOX 31.0653 60 7 abrupt 370 265 -238 -236 234.2 219.4 5 39.25 48.85 49.75 27.46 124.55 47.90 47.88395 Hotfire warm simulatanous with LOX 31.0643 40 7 abrupt 370 265 -238 -236 224.9 225.4 (limited train duration). Didn't work because the 5 58.96 47.06 47.72 50.91 150.23 49.53 53.16396 Hotfire warm simulatanous with LOX 31.0943 20 4 grad 370 265 -238 -234 221.4 218.9 4 136.66 47.32 46.31 51.22397 Nolight warm simulatanous with LOX 31.0943 20 4 non 370 265 -237 -241 229.4 218.6

Voltage on exciter increased from 28 to 32v to try to decrease delay to 1st spark 0 41.55 36.56 47.00 47.96 0.00 0.00 0.00

398 Nolight warm simulatanous with LOX 31.0643 20 4 non 370 265 -237 -240 228.5 218.2 0 151.99 44.53 35.68 44.74 0.00 0.00 0.00399 Hotfire warm 20ms after Ch4 31.0653 20 4 late 370 265 -237 -237 220.8 234.5

Return previous conditions that had drop outs to try to recreate. Keep reduced spark duration to minimize hard start risk 1 27.42 53.91 52.56 55.23 0.00 0.00 0.00

400 Nolight warm 20ms after Ch4 31.0942 20 4 non 370 265 -237 -235 239.2 223.7 No current draw on any spark -all dropout 0 0.00 0.00 0.00 0.00 0.00 0.00 0.00401 Thrust cal

Spark energy, mJ

Run Tank Pressure, 

psig

Run Tank 

Temperature, °F

Valve 

Temperature, °R
NASA/TM—2012-217611 35 

References 

1. Melcher, J.C.; and Allred, J.K.: Liquid Oxygen/Liquid Methane Testing of the RS–18 at NASA 

White Sands Test Facility. AIAA 2008–4843, 2008. 

2. Melcher, J.C.; and Allred, J.K.: Liquid Oxygen/Liquid Methane Test Results of the RS–18 Lunar 

Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility. AIAA 2009–

4949, 2009. 

3. Stone, R.; Tiliakos, N.; Balepin, V.; Tsai, C.-Y.; and Engers, R.: Altitude Testing of LOX-Methane 

Rocket Engines at ATK GASL. AIAA 2008–3701, 2008. 

4. Robinson, P.J.; Veith, E.M.; Hurlbert, E.A.; Jimenez, R.; and Smith, T.D.: 100-lbf LO2/LCH4-

Reaction Control Engine Technology Development for Future Space Vehicles. 59th International 

Astronautical Federation, Glasgow, Scotland, United Kingdom, Sept. 29–Oct. 3, 2008. 

5. Marshall, W.M.; and Kleinhenz, J.E.: Performance Analysis of Duration Tests of a 100 lbf 

LO2/LCH4 Reaction Control Engine at Altitude Conditions at NASA Glenn Research Center. 

NASA/TM—2011-217131, 2011. 

6. Marshall, W.M.; and Kleinhenz, J.E.: Hot-Fire Testing of 100 lbf LOX/LCH4 Reaction Control 

Engine at Altitude Conditions. JANNAF 57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee 

Meeting, Colorado Springs, CO, May 3–7, 2010. 

7. Marshall, W.M.; and Kleinhenz, J.E.: Analysis of 100 lbf LOX/LCH4 Reaction Control Engine in 

Pulsed, Altitude Operation at NASA Glenn Research Center. NASA/TM—2012-217613. 

8. Lawver, B.R.; and Rousar, D.C.: Ignition Characterization of the GOX/Ethanol Propellant 

Combinations. AIAA–84–1467, 1984. 

9. Lawver, B.R.; Rousar, D.C.; and Wong, K.Y.: Ignition Characterization of LOX/Hydrocarbon 

Propellants. NASA CR–171884, 1985. 

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Methane/Liquid Oxygen Propellant Conditioning Feed System (PCFS) Test Rigs. JANNAF 3rd 

Spacecraft Propulsion System Joint Subcommittee Meeting, Orlando, FL, Dec. 8–12, 2008. 

11. Grasl, S.J.; Skaff, A.F.; Nguyen, C.; Schubert, J.; and Arrington, L.: Liquid Methane/Liquid Oxygen 

Propellant Conditioning and Feed System (PCFS) Test Rigs—Preliminary Test Results. JANNAF 

57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee Meeting, Colorado Springs, CO, May 3–7, 

2010. 

12. Glassman, I.: Combustion. 3rd Ed., Academic Press, 1996, p. 347. 

13. Stiegemeier, B.R.; and Marshall, W.M.: Sea-Level Testing of a 100 lbf LOX/Methane Reaction 

Control Engine. JANNAF 57th JPM/7th MSS/5th LPS/4th SPS Joint Subcommittee Meeting, 

Colorado Springs, CO, May 3–7, 2010. 

14. Bury, K.M.; and Kerslake, T.W.: The Effect of Reactor Control System Thruster Plume 

Impingement on Orion Service Module Solar Array Power Production. NASA/TM—2008-215423, 

2008. 

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Chemical Physics, vol. 25–5, 1956, pp. 1053–1058. 

16. Mastorakos, E.: Ignition of Turbulent Non-Premixed Flames. Progress in Energy and Combustion 

Science, vol. 35, 2009, pp. 57–97. 

17. Huang, C.C.; Shy, S.S.; Liu, C.C.; and Yan, Y.Y.: A Transition on Minimum Ignition Energy for 

Lean Turbulent Methane Combustion in Flamelet and Distributed Regimes. Proceedings of 

Combustion Institute, 2007, vol. 31, pp. 1401–1409. 
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Experimental Investigation of Augmented Spark Ignition of a LO2/LCH4 Reaction Control 

Engine at Altitude Conditions 

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Kleinhenz, Julie; Sarmiento, Charles; Marshall, William 

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14. ABSTRACT

The use of nontoxic propellants in future exploration vehicles would enable safer, more cost-effective mission scenarios. One promising 

“green” alternative to existing hypergols is liquid methane (LCH4) with liquid oxygen (LO2). A 100 lbf LO2/LCH4 engine was developed 

under the NASA Propulsion and Cryogenic Advanced Development project and tested at the NASA Glenn Research Center Altitude 

Combustion Stand in a low pressure environment. High ignition energy is a perceived drawback of this propellant combination; so this 

ignition margin test program examined ignition performance versus delivered spark energy. Sensitivity of ignition to spark timing and 

repetition rate was also explored. Three different exciter units were used with the engine’s augmented (torch) igniter. Captured waveforms 

indicated spark behavior in hot fire conditions was inconsistent compared to the well-behaved dry sparks. This suggests that rising pressure 

and flow rate increase spark impedance and may at some point compromise an exciter’s ability to complete each spark. The reduced spark 

energies of such quenched deliveries resulted in more erratic ignitions, decreasing ignition probability. The timing of the sparks relative to 

the pressure/flow conditions also impacted the probability of ignition. Sparks occurring early in the flow could trigger ignition with energies 

as low as 1 to 6 mJ, though multiple, similarly timed sparks of 55 to 75 mJ were required for reliable ignition. Delayed spark application and 

reduced spark repetition rate both correlated with late and occasional failed ignitions. An optimum time interval for spark application and 

ignition therefore coincides with propellant introduction to the igniter.

15. SUBJECT TERMS

Spark ignition; Chemical propulsion; Cryogenic rocket propellants; Combustion 

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