American Institute of Aeronautics and Astronautics

1

Green Propellant Infusion Mission Plume Impingement

Analysis

John T. Yim1 and Brian D. Reed

2

NASA Glenn Research Center, Cleveland, OH, 44135

and

Christopher H. McLean3

Ball Aerospace & Technologies Corporation, Boulder, CO, 80301

The Green Propellant Infusion Mission (GPIM) is a Technology Demonstration Mission

(TDM) project, sponsored by NASA’s Space Technology Mission Directorate (STMD). The

goal of GPIM is to demonstrate the capability of a green propulsion system, specifically, one

using the monopropellant, AF-M315E. The GPIM propulsion system will be flown as a

payload on a Ball Aerospace BCP-100, a small, standardized spacecraft. The propulsion

system will have one 22 N thruster for primary divert manuevers and four 1 N thrusters for

attitude control. One of the risks identified for GPIM is potential contamination of sensitive

areas of the spacecraft from the effluents in the plumes of AF-M315E thrusters. NASA

Glenn Research Center (GRC) is conducting activities to mitigate the AF-M315E plume risk.

The plume risk mitigation activities include modeling the plume flow fields of the AF-M315E

thrusters, assessing the plume impingement on the BCP-100 spacecraft including the impact

on the power generating capabilities of the solar arrays, and conducting ground-based

plume measurements on an AF-M315E thruster to correlate the plume modeling with plume

data. This paper describes the preliminary results from the first activity, plume modeling

and plume impingement analysis. Plume flow fields of the 22 N and 1 N thrusters have been

modeled using both the method-of-characteristics based Reacting And Multi-Phase

(RAMP2) code and the Hypersonic Aerothermodynamics Particle (HAP) Direct Simulation

Monte Carlo (DSMC) code. The density and temperature plume flow fields are presented, as

well as species concentration at different locations in the plume. Chamber pressures from

400 psia to 100 psia are examined, simulating the GPIM propulsion system operating in

blowdown mode. Both equilibrium and frozen flow assumptions are also investigated. The

plume impingement on the BCP-100 spacecraft are evaluated, using both the PLume

IMPingement program (PLIMP) and also HAP. The heating rates on the spacecraft surfaces

are found to be fairly benign. Species flux on the surfaces are also examined, with hydrogen

gas found to be the dominant species in the backflow region. The simulations from both

models are planned to be compared to forthcoming plume measurement data, when AF-

M315E thruster testing is conducted.

I. Introduction

ONOPROPELLANT propulsion systems offer greater simplicity, reliability, and lower cost over higher

performance bipropellant systems. Hydrazine (N2H4) has been the standard monopropellant of choice for the

past several decades for attitude control and small delta-V maneuvers. However, hydrazine’s toxicity and

flammability hazards have led to research of “green” monopropellant formulations that will reduce environmental,

health, and safety impacts.1-2

In addition, certain green monopropellants are being pursued which have higher

storage densities and specific impulses as compared to hydrazine.3 The Green Propellant Infusion Mission (GPIM)

will demonstrate one such monopropellant, AF-M315E, through integration of a green monopropellant propulsion

1 Aerospace Engineer, Space Propulsion Branch, 21000 Brookpark Rd. MS 86-8, AIAA Member.

2 GPIM Co-Investigator, Space Propulsion Branch, 21000 Brookpark Rd. MS 86-8.

3 GPIM Principal Investigator, Mission Systems Engineering, 1600 Commerce Street, AIAA Senior Member.

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49th AIAA/ASME/SAE/ASEE Joint PropulsionConference

July 14 - 17, 2013, San Jose, CA

AIAA 2013-3850

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

American Institute of Aeronautics and Astronautics

2

system onto a BCP-100 spacecraft.4-6

The mission will demonstrate two classes of AF-M315E thrusters, one main

22 N thruster for primary divert maneuvers and four smaller 1 N thrusters for attitude control.

One concern with any new propellant is how the resultant plume may potentially interact with the rest of the

spacecraft. Thus, within the GPIM project, an effort is underway to characterize the AF-M315E plumes through

both modeling and test-based approaches at the NASA Glenn Research Center (GRC).6 Preliminary results of the

plume modeling effort are the focus of this paper, with subsequent work planned to experimentally characterize the

plume as well as further efforts to model the specific impact on the solar arrays.

Two modeling approaches are used to analyze the plumes, particularly from density, species, and heat transfer

standpoints, and are further described below. The general composition and characteristics of the plumes are

described as well as the expected impacts on nearby spacecraft surfaces of the BCP-100.

II. Model Description

Two models were used to examine the plume properties of the GPIM thrusters operating on AF-M315E. The

first used the Reacting And Multi-Phase (RAMP2) / PLume IMPingment (PLIMP) program combination. The

second method used the Hypersonic Aerothermodynamics Particle (HAP) code. Both RAMP2 / PLIMP and HAP

were used to examine both the 22 N thruster with a 50:1 area ratio and the 1 N thruster with a 100:1 area ratio. For

the GPIM mission, the thrusters will operate on blowdown mode with expected chamber pressures at beginning of

life of approximately 400 psia (27.6 bar) and decreasing by end of life to approximately 100 psia (6.9 bar).

A. RAMP2 / PLIMP

RAMP2 is an axisymmetric method-of-characteristics based code that calculates the plume flowfield beginning

from the thruster nozzle throat to the exit plane within a user supplied nozzle profile.7 It also calculates the plume

flowfield expansion beyond the exit plane. It has the capabilities to account for nozzle boundary layers, multi-phase

flow, and flow chemistry among other options. A Chemistry Equilibrium with Applications (CEA) code equivalent

is used to generate the flow properties which are input into RAMP2.

The RAMP2 flowfield can be used as an input to the PLIMP code to subsequently calculate forces, moments,

pressures, heating rates, and species contamination on surfaces specified by the user.8 PLIMP has the capability to

apply the appropriate flow regime from turbulent viscous flow to free molecular, to determine the appropriate

surface impingement properties. RAMP2 and PLIMP have been used for a number of plume impingement studies

including Shuttle Reaction Control System (RCS) and solid rocket boosters, Orion Service Module (SM) RCS

engines, and Space Launch System (SLS) solid rocket boosters among others.9-11

B. HAP

HAP is an implementation of the Direct Simulation Monte Carlo (DSMC) method.12

The DSMC method is a

particle-based gas flow simulation technique which is derived from principles of kinetic theory. A large population

of simulated particles is tracked through the computational domain, with a combination of deterministic particle

movement routines and probabilistic binary collision routines, in order to simulate underlying physics in the

governing Boltzmann equation. Other DSMC implementations have been successfully used for plume flow and

plume impingement studies.13-14

For example, the DSMC Analysis Code (DAC), has been applied to model plumes

for several different programs including plume impingement on the International Space Station, the Orion SM RCS

engines, and the Altair lunar lander ascent engine among others.15-17

These efforts have shown that DSMC is a

viable method to simulate and analyze plume flows and resulting plume impingement.

The HAP code is a general Cartesian implementation of DSMC, with features including dynamic grid

adaptation; shared memory parallelization; and options to import externally defined surface geometries or to

automatically generate triangulated surfaces based on analytical geometry definitions. In particular, a useful ability

of the code is the use of a collision limiter scheme with a gradient-based local time step adaptation to selectively

reduce numerical errors from multiple sources (e.g. incomplete equilibration, operator splitting, finite collision

separation) within regions which are especially prone to inaccuracy.18

This allows for efficient calculation of the low

Knudsen number regions of the flow without being overly diffusive. With this capability, an inflow boundary

applied at the thruster exit plane was able to be implemented. The CEA program is used to generate the species

composition, plume density, temperature, and flow velocity at the appropriate area ratios to be used for the inflow

boundary conditions. HAP then calculates the resultant expansion flow of the thruster plume. A geometry model of

the BCP-100 spacecraft was also imported into the program and the particle flux and heat transfer were calculated at

the surfaces. For the 22 N and 1 N free plume flow studies and the 22 N case for spacecraft interaction, symmetry

planes were utilized in the lateral directions to reduce the computational requirements. Due to the inherent

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American Institute of Aeronautics and Astronautics

3

asymmetry of the 1 N thruster position and vector on the BCP-100, no symmetry advantages were applied for that

particular model. Typical runs averaged on the order of a million simulated particles at the steady state condition.

III. Results and Discussion

Both sets of codes were run for the 22 N thruster (50:1 area ratio) and a comparison of the free plume flowfield

is performed. Two different chamber pressures (400 and 100 psia), and both equilibrium and frozen flow

assumptions are examined with each code. Sample contour plots of the plume density and temperature in vacuum

for the 22 N thruster at 400 psia chamber pressure and equilibrium flow are shown in Figure 1. Going more in detail

of the plume flowfield variations, the overall number density of the plume are shown for both axially along the

thruster centerline and azimuthally at a 10 cm radius from the center of the thruster exit plane in Figure 2. The

RAMP2 density results are higher than the HAP results, particularly near the thruster and in the backflow region.

However the results between the two codes are generally within the same order of magnitude. As would be

expected, the number density is strongly correlated to the chamber pressure, with about a four to one ratio

corresponding to the 400 and 100 psia cases. The differences between the equilibrium and frozen flow cases are

minor, though HAP predicts a higher density in the backflow region of the plume for equilibrium flow as compared

to frozen.

Figure 1. Plume density (left) and temperature (right) for the 22 N thruster with a chamber pressure of

400 psia and equilibrium flow as calculated by RAMP2 (top) and HAP (bottom). Grid lines use a 10 cm

spacing.

Figure 2. Plume number density along the thruster centerline (left) and at a 10 cm radius from the center of

the thruster exit plane (right).

Part of the reason for the differences in the backflow region of the plume is due to differences in the chemical

species makeup of the plumes between frozen and equilibrium flow. Figure 3 plots the major species of the

AF-M315E plume at a 10 cm radius from the center of the thruster exit plane as calculated by HAP. Trace species,

including ammonia, nitrogen oxide, atomic hydrogen, and the hydroxyl radical, are not shown for clarity. One of the

main differences seen between the chemical compositions of frozen and equilibrium flow is the significant presence

of carbon monoxide in frozen flow, while for an equilibrium assumption, it comprises less than 0.15% of the plume

by number density. Methane becomes the fifth main constituent instead, though it is below 0.5% itself. It can be

seen at large angles from the thrust axis that hydrogen is the dominant species by far, and that the difference in the

1E+20

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American Institute of Aeronautics and Astronautics

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hydrogen density between frozen and equilibrium flow accounts for the difference in the overall density in that

region. These trends are further highlighted in Figure 4, which shows area plots of the relative molar fractions of the

major species of the plume for both equilibrium and frozen flow. Species with a mole fraction below 0.01 (1%) are

not shown for clarity. The dominance of hydrogen is again seen in the wide angles from the thrust centerline. At

smaller angles, within 45° of the thrust axis, water is seen to comprise the majority of the plume with notable

fractions of carbon dioxide and nitrogen. Condensation of water on sensitive surfaces is a potential concern of the

use of AF-M315E, and thus its propagation as it exits the thruster is an area of interest. This analysis indicates water

comprises less than 1% of the plume by mole fraction past 75°, and less than 0.1% beyond approximately 90° from

the thrust axis. Methane comprises less than 0.1% of the plume by mole fraction past approximately 60°, and the

other trace species are well under 0.1% of the plume throughout the entire domain. The dominance of hydrogen—

primarily of H2, and to a much lesser extent, atomic hydrogen—in the backflow region is likely due to their low

atomic mass compared to the other species. Indeed, the relative rate of drop off in number density of the species at

larger angles appears to be correlated to their molecular weight. Gas molecular collisional events in the plume are

more likely to give lower weight atoms and molecules a greater scattering angle compared to the other species, and

thus subsequently disproportionately populate the backflow region.

Figure 3. Plume species number density at a 10 cm radius from the center of the 22 N thruster exit plane for a

400 psia chamber pressure case.

Figure 4. Plume species molar fractions assuming equilibrium flow (left) and frozen flow (right) at a 40 cm

radius from the center of the 22 N thruster exit plane.

1E+15

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In addition to the chemical makeup of the plume, an area of potential concern is the heating due to plume

impingement. As a first order approximation, the total kinetic energy of the plume is calculated, and the cumulative

amount of energy within certain half-angle cones around the thrust axis is summed. Figure 5 presents the percentage

of the plume energy within a given cone half-angle as calculated by HAP, and it should be noted that the RAMP2

results are very similar. A rough rule of thumb that is sometimes quoted for hydrazine thrusters assumes

approximately 92% of the plume thermal energy is contained with a 30° cone. This same rule of thumb is seen to

hold for AF-M315E as well in Figure 5, as examined across a number of cases. Differences in chamber pressure or

plume chemistry do not appear to have a significant effect on these trends.

Figure 5. Cumulative percent of total plume kinetic energy for the 22 N thruster.

Similar to the analyses performed on the 22 N thruster, the 1 N thruster (with a 100:1 area ratio) was also

examined. The simulated plume flowfield density and temperature of the 1 N thruster in a vacuum environment is

shown in Figure 6. The results shown correspond to a 400 psia chamber pressure and equilibrium flow. As with the

22 N thruster simulation results, the two codes are fairly comparable downstream, but with differences in the very

near field and at large angles from the thrust axis.

Figure 6. Plume density (left) and temperature (right) for the 1 N thruster with a 400 psia chamber pressure

and equilibrium flow as calculated by RAMP2 (top) and HAP (bottom). Grid lines are at a 10 cm spacing.

The chemical makeup of the 1 N thruster plume is very similar to the 22 N plume. Figure 7 shows a semi-log

plot of the species number density at a 10 cm radius from the center of the thruster exit plane. Water again

comprises over half of the plume within an approximately 45° cone, with notable presence of carbon dioxide,

nitrogen, and to a lesser extent, carbon monoxide in the frozen flow assumption. Hydrogen again dominates at large

angles. The amount of methane is also relatively larger in the equilibrium case, where it now comprises 1 – 2% of

the plume in the downstream region within approximately 60° of the thrust axis.

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Figure 7. Plume species number density at a 10 cm radius from the center of the 1 N thruster exit plane for a

400 psia chamber pressure case.

The thermal energy of the 1 N thruster plume follows closely to the 22 N results presented above in Figure 5.

Again, the approximate rule of thumb of 92% of the plume thermal energy being contained within a cone with a

half-angle of 30° from the thrust axis appears to hold for the AF-M315E 1 N thruster as well. This is observed

across different chamber pressures, frozen vs. equilibrium flow, and the two codes as also mentioned for the 22 N

thruster above.

In addition to the free plume flow studies, the interaction with the BCP-100 spacecraft as configured for GPIM is

examined. Two cases are examined, the main 22 N thruster firing and a single 1 N thruster firing. The 22 N thruster

is aligned normal to the backplane of the spacecraft. The heating rate on the spacecraft is calculated by PLIMP using

the 22 N RAMP2-calculated flowfield as an input with the results shown in Figure 8. The expected worst case

assumptions, namely the maximum chamber pressure of 400 psia and equilibrium chemistry, are used to generate

the results. The maximum heating rate on the spacecraft bus is calculated as 27 W/m2. The solar arrays see a

maximum heating rate of approximately 1 W/m2. The 400 psia chamber pressure, equilibrium flow, 1 N RAMP2

results are also input into PLIMP, and are aligned to match the 50° canting angle away from the backplane of the

spacecraft. The results for a single 1 N thruster are also shown in Figure 8. Here the maximum heating rate on the

bus is now 127 W/m2, and on the solar arrays, 2 W/m

2. The increase in heating rate on the backplane surface is due

to the closer thrust vector alignment of the 1 N thrusters to the spacecraft compared to the 22 N. The interaction with

the GPIM spacecraft is also examined using HAP. The 22 N and 1 N thruster cases are shown in Figure 9. The

calculated heating rates are somewhat lower than that calculated by RAMP2 and PLIMP. For the 22 N case, the

maximum heating rate on the bus is 10 W/m2

and approximately on the order of 0.1 W/m2 on the solar arrays. For

one 1 N thruster firing, the heating rate on the bus is lower, at 8 W/m2 with negligible heating on the solar arrays.

The reason behind the heating rate discrepancies are still unclear. Future testing will help to provide additional data

on the backflow region and to determine which value is a better suited for this scenario.

Figure 8. Heating rates calculated by RAMP2/PLIMP on the BCP-100 spacecraft surfaces from the central 22

N thruster firing (left) and the lower left 1 N thruster firing (right).

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Figure 9. Heating rates calculated by HAP on the BCP-100 spacecraft surfaces from the central 22 N thruster

firing (left) and the lower left 1 N thruster firing (right).

The fluxes of various plume chemical species on the spacecraft surfaces are also examined. The number fluxes

on the GPIM spacecraft surfaces of the major species from the 22 N thruster with a chamber pressure of 400 psia are

calculated. The worst-case, or maximum flux, for hydrogen gas and water vapor arise from the equilibrium

assumption and those cases are shown in Figure 10. As was seen earlier, hydrogen gas is by far the dominant species

in the backflow region impinging on the spacecraft surfaces, likely due to its notable presence in the main plume

flow and low molecular weight as speculated above. For the frozen flow case, atomic hydrogen is also notably

present, and is also shown in Figure 10. The 1 N thruster case is also examined and results presented in Figure 11.

Again the 400 psia chamber pressure is assumed for worst case fluxes. For this case, a significant increase in water

vapor is seen, though it is still at least two orders of magnitude lower than hydrogen gas, which still dominates the

impingement on the spacecraft surfaces.

Figure 10. Plume species number flux on the GPIM spacecraft surfaces for the 22 N thruster with 400 psia

chamber pressure and equilibrium flow for H2 (left) and H2O (center), and frozen flow for H (right).

Figure 11. Plume species number flux on the GPIM spacecraft surfaces for the 1 N thruster with 400 psia

chamber pressure and equilibrium flow for H2 (left) and H2O (center), and frozen flow for H (right).

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IV. Conclusion

A characterization of the AF-M315E green monopropellant plume was examined for the GPIM mission. Two

sets of modeling tools, RAMP2 / PLIMP and HAP, were used to simulate the thruster plume in vacuum and the

resultant impingement properties on the BCP-100 spacecraft bus. Two thruster classes were examined, the 50:1 area

ratio 22 N main thruster and the 100:1 area ratio 1 N attitude control thrusters. The density profiles in the axial and

azimuthal directions were calculated as well as the chemical species composition of the plume. Water vapor is noted

to comprise more than half of the plume by molar fraction within approximately 45° of the thrust axis, with a

notable presence of both carbon dioxide and nitrogen gas. At large angles from the thrust axis, hydrogen gas was

found to be the dominant species. The azimuthal distribution of chemical species was found to correlate to the

molecular weight of the species, with the lighter molecules having a greater probability of being scattered at larger

angles. These same trends were observed for the species impinging on the GPIM spacecraft bus and solar array

surfaces.

The thermal aspects of the plume and plume impingement were also examined. The AF-M315E thrusters were

found to generally obey a rule of thumb also seen for hydrazine where 92% of the thermal energy of the plume is

contained within a cone with a half-angle of approximately 30° from the thrust vector. The heating rate of the GPIM

spacecraft due to the AF-M315E plume impingement was also calculated. The modeled heating rates were benign; a

maximum heating rate of 127 W/m2 was observed with the RAMP2 / PLIMP results.

The goal of this work is to help the adoption of AF-M315E as a viable replacement for hydrazine due to its

superior handling characteristics and higher density specific impulse. From the preliminary results presented here,

no obvious issues were immediately identified from a thermal or contamination standpoint, though further

assessment is required including a full analysis of the consequences, if any, of the plume deposition on the solar

arrays and other sensitive surfaces. NASA GRC also plans to pursue experimental characterization of the AF-

M315E plume within the next year, which will also help to verify the results of these models.

Acknowledgments

The first author would like to thank Jonathan Burt for the HAP code he originally developed for the Air Force,

and for insight into its application for plume impingement studies. The authors also acknowledge excellent support

by Adam Sexton at Ball Aerospace with regards to CAD model assistance. The programmatic support of David

Frate at NASA GRC is also gratefully noted.

References 1Jankovsky, R. S., “HAN-Based Monopropellant Assessment for Spacecraft,” 32nd AIAA/ASME/SAE/ASEE Joint Propulsion

Conference, Lake Buena Vista, FL, AIAA-96-2863, July 1996. 2Zube, D. M., Wucherer, E. J., and Reed, B. D., “Evaluation of HAN-Based Propellant Blends,” 39th AIAA/ASME/SAE/ASEE

Joint Propulsion Conference and Exhibit, Huntsville, AL, AIAA-2003-4643, July 2003. 3Hawkings, T. W., Brand, A. J., McKay, M. B., and Tinnirello, M., “Reduced Toxicity, High Performance Monopropellant at

the U.S. Air Force Research Laboratory,” 4th International Association for the Advancement of Space Safety Conference,

Huntsville, AL, May 2010. 4McLean, C. H., Spores, R., and Sheehy, J. A., “Green Propellant Infusion Mission, Program Construct, and Mission

Objectives,” 60th JANNAF Propulsion Meeting, Colorado Springs, CO, May 2013. 5McLean, C. H., Hale, M. J., Deininger, W. D., Spores, R. A., Frate, D. T., Johnson, W. L., and Sheehy, J. A., “Green

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