Xiao-Yen J. Wang

Glenn Research Center, Cleveland, Ohio

Forrest E. Lumpkin III

Johnson Space Center, Houston, Texas

Frank Gati, James R. Yuko, and Brian J. Motil

Glenn Research Center, Cleveland, Ohio

Orion Service Module Reaction Control

System Plume Impingement Analysis 

Using PLIMP/RAMP2

NASA/TM—2009-215601

March 2009

AIAA–2009–0834
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Xiao-Yen J. Wang

Glenn Research Center, Cleveland, Ohio

Forrest E. Lumpkin III

Johnson Space Center, Houston, Texas

Frank Gati, James R. Yuko, and Brian J. Motil

Glenn Research Center, Cleveland, Ohio

Orion Service Module Reaction Control

System Plume Impingement Analysis 

Using PLIMP/RAMP2

NASA/TM—2009-215601

March 2009

AIAA–2009–0834

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

47th AIAA Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Orlando, Florida, January 5–8, 2009
Acknowledgments

Xiao-Yen Wang thanks Sheldon Smith for his valuable help in using the RAMP2 and PLIMP codes, Julien du Castel and 

Ries Smith at Lockheed Martin for providing the PIDYN and RPM solutions, Ananda Himansu for his insightful discussions, 

Timothy Roach for his help with the ProE model, and Shane Malone for reviewing the paper, and Colleen McGraw for her 

great support to the Orion SM Passive Thermal Control System (PTCS) under the Orion project.

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Level of Review: This material has been technically reviewed by technical management. 

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2009-215601 1

Orion Service Module Reaction Control System 

Plume Impingement Analysis Using PLIMP/RAMP2 

Xiao-Yen J. Wang 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Forrest E. Lumpkin III 

National Aeronautics and Space Administration 

Johnson Space Center 

Houston, Texas 77058 

Frank Gati, James R. Yuko, and Brian J. Motil 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The Orion Crew Exploration Vehicle Service Module Reaction Control System engine plume 

impingement was computed using the plume impingement program (PLIMP). PLIMP uses the plume 

solution from RAMP2, which is the refined version of the reacting and multiphase program (RAMP) 

code. The heating rate and pressure (force and moment) on surfaces or components of the Service Module 

were computed. The RAMP2 solution of the flow field inside the engine and the plume was compared 

with those computed using GASP, a computational fluid dynamics code, showing reasonable agreement. 

The computed heating rate and pressure using PLIMP were compared with the Reaction Control System 

plume model (RPM) solution and the plume impingement dynamics (PIDYN) solution. RPM uses the 

GASP-based plume solution, whereas PIDYN uses the SCARF plume solution. Three sets of the heating 

rate and pressure solutions agree well. Further thermal analysis on the avionic ring of the Service Module 

was performed using MSC Patran/Pthermal. The obtained temperature results showed that thermal 

protection is necessary because of significant heating from the plume. 

I. Introduction 

The Orion project is under the Constellation program for the new space exploration vision initiated by 

President Bush in 2004. The Constellation program is responsible for providing the elements that will 

transport humans and cargo to both the International Space Station (ISS) and the Moon. These elements 

are the Crew Exploration Vehicle (CEV, or Orion), the Crew Launch Vehicle (Ares I), the Lunar Surface 

Access Module (Altair), and the Cargo Launch Vehicle (Ares V). Orion, with a crew of up to six 

astronauts, will launch on Ares I and then use its main engine to insert itself into a safe orbit to either 

dock with the ISS or with Altair. For the ISS mission, Orion will be responsible for separation, entry, 

descent, and landing. For the lunar missions, Orion also will have to maintain itself in low lunar orbit and 

perform the trans-Earth injection maneuver to return from the Moon. Orion consists of the Launch Abort 

System (LAS), Crew Module (CM), Service Module (SM), and Spacecraft Adapter (SA). The CM is a 

capsule design that provides the primary structure for crew support, incorporates the bulk of the avionics 

systems, and provides the capability for entry and parachute landing. The LAS will safely extract the CM 

from the launch configuration in the event of an early launch abort. The SM, the structure that interfaces 

with Ares I, will perform in-space flight propulsion operations and power generation. 
NASA/TM—2009-215601 2

Here, the SM was studied for its Reaction Control System (RCS) engine-plume-impingement effects. 

In a space environment, the exhaust plume of a rocket engine may expand so as to impinge upon the 

spacecraft structure or its components. Plume impingement will result in surface heating, pressure, and 

perhaps contamination. The SM RCS will operate in nearly all mission phases; the longest steady-state 

firing is 120 s. The SM radiator panel and solar arrays are the major concern and require a safe 

operational environment. Other surfaces or components might need thermal shields or protection from the 

plume impingement. Thrust loss is also a major concern for the guidance, navigation, and control of the 

vehicle. 

The flow field inside the rocket engine and plume expansion into the ambient were computed using 

RAMP2, in which the flow is in two-dimensional or axisymmetric geometry with frozen or equilibrium 

chemistry. Boundary layer correction is also included in RAMP2 to account for viscous effects inside the 

nozzle. Details about the code are provided in Reference 1. 

The PLIMP code, (Ref. 2) developed at the NASA Marshall Spaceflight Center, was used to compute 

the heating rate and pressure on the surfaces that are subject to plume impingement. Thus, the force and 

moment on the components are also available. In PLIMP, surfaces are represented using simple geometry 

surfaces, such as a cylinder, flat plate, or circular plate. The flow angle toward the surface and the flow 

regime, such as continuum, transitional, or free molecular, are considered in calculating the heating rate 

and pressure. Multiphase flows, such as solid particles and gases, can also be modeled. PLIMP can be 

used for general plume impingement analysis at both low and high altitudes. 

The RPM code was used to predict the space shuttle orbiter Primary Reaction Control System plume 

impingement (forces and heating) on the ISS and other spacecraft near the shuttle orbiter. It uses a source 

flow assumption for the plume flow field where adjustable parameters in the formulation are set to 

achieve a best fit to a computational fluid dynamics (CFD) solution generated by the GASP code. The 

source flow assumption is generally accepted to be valid when the impingement distance is larger than 10 

nozzle exit diameters. The source flow formulation in RPM divides the plume into inviscid and viscous 

regions in order to account for the boundary layer’s affect on the plume. It uses the extended version of 

Simon’s model in Reference 3. Plume impingement forces and moments are determined by an 

engineering model that bridges between Newtonian and free-molecular analytic expressions for the 

pressure and shear coefficients. Plume impingement heating is also computed with an engineering model 

that bridges between continuum and free-molecular expressions for aerodynamic heating. This code was 

developed at the NASA Johnson Space Center and is described in Reference 4. 

The PIDYN code*

 uses a vacuum plume Newtonian approach for forces and moments, and basic heattransfer calculation excerpts from MINIVER, the miniature version of the JA70 General Aerodynamic 

Heating Computer Code (Ref. 5). The SCARF code provides the plume flow field to PIDYN. SCARF 

also is based on Simon’s model. In SCARF, the plume flow field is modeled much as in RPM by 

employing the source flow assumption and an adjustment for boundary layer expansion at an off-axis 

region. The primary difference between RPM and PIDYN is that the various adjustable parameters in 

PIDYN are based on characteristics of the engine (the most significant being the assumed ratio of specific 

heats), whereas in RPM the parameters are adjusted to match a CFD solution. This code was developed at 

Lockheed Martin and had heritage use at Lockheed Martin Space Systems for many spacecraft plumeimpingement applications. 

In summary, the RPM and PIDYN codes use a similar approach based on using a point source 

assumption to compute the plume field. However, the flow inside the engine nozzle changes dramatically 

from the nozzle throat to the exit, the plume solution will be sensitive to the specified ratio of the specific 

heat (γ) of the hot gas. The correct γ should be used for an accurate plume field. Both RPM and PIDYN 

codes can read in the mesh files generated by ProE for the geometries, which saves users from having to 

create the geometry from scratch. PLIMP will use a more accurate plume solution computed by RAMP2, 

but users will need to use an analytical approach and only structured mesh to define the geometry as a 

 

*

Castel, J.D., Private communication, Lockheed Martin, Denver, CO, June 2008. 
NASA/TM—2009-215601 3

part of input to PLIMP. There is no interface between PLIMP and computer-aided design (CAD) 

software. 

In the following sections, first the flow field inside the engine nozzle and its plume are presented; 

then the plume impingement results are presented for both forward-facing and after-facing thrusters, 

along with thermal analysis on the avionic ring of the SM. Finally, conclusions are drawn. 

II. Solution of the Flow Inside the Engine Nozzle and Its Plume 

RAMP2 uses the chemical equilibrium compositions and applications (CEA) code to compute the 

thermodynamic and flow properties inside an engine nozzle. In Table 1, the CEA results of the pressure 

(P), temperature (T), density (ρ), molecular weight, specific heat (Cp), ratio of the specific heat (γ), sonic 

velocity, Mach number (Ma), and the mole fractions are listed for different locations inside the thruster. 

Given certain input parameters and CEA thermodynamic properties, RAMP2 can be used to compute the 

flow field inside the thruster and the plume. The contours of Ma, log10 ρ, T, log10 P, velocity (u), and flow 

angle are plotted in Figure 1. At the throat, the gas could get as hot as 5000 °R, and at the nozzle exit, the 

temperature will reach approximately 900 °R. The plume will expand significantly because of the vacuum 

ambient. 

TABLE 1.—SERVICE MODULE REACTION CONTROL SYSTEM ENGINE PERFORMANCE 

(CEA RESULTS) 

 Injector Combustion 

chamber end 

Throat Nozzle exit 

Pressure, P, psi 108 105.4 60.9 0.058 

Temperature, T, R 5147 5408 5055 1371 

Density, ρ, lbm/ft3 3.74×10–2 3.66×10–2 2.28×10–2 8.18×10–5

Molecular weight, 1/n 20.154 20.157 20.338 20.649 

Specific heat, Cp, Btu/lbm-F 1.058 1.056 0.889 0.477 

Ratio of the Specific heat, γ 1.157 1.157 1.1671 1.2528 

Sonic velocity, ft/s 3931.1 3927.5 3797 2033 

Mach number 0 0.146 1 5.225 

Mole fraction 

 CO 0.13447 0.13443 0.13334 0.03003 

 CO2 0.03378 0.03384 0.03644 0.14234 

 H2 0.16904 0.16904 0.17049 0.27943 

 H2O 0.31516 0.31533 0.32534 0.2377 

 N2 0.30201 0.30207 0.30528 0.31049 

Figure 2 shows a comparison of the dynamic pressure (1/2 ρu

2

) between the RAMP2 and GASP 

solutions, and Figure 3 shows a quantitative comparison along the axial direction at the circumferential 

angle (θ) = 0° and the circumferential direction at radius (R) = 15 ft. It can be seen that the RAMP2 plume 

solution agrees well with the GASP solution in the core of the plume; away from the core, the RAMP2 

solution is generally higher than the GASP solution. 
NASA/TM—2009-215601 4

Figure 1.—RAMP2 solutions of the flow inside the engine nozzle and its plume. 
NASA/TM—2009-215601 5

III. PLIMP Plume Impingement Results 

Once the plume solution is available as an input to PLIMP, PLIMP can be run with additional input 

parameters, such as the definition of the impinging surfaces, the RCS engine location and orientation, and 

the surface wall temperature. Figure 4 shows plots of the contours of the heat flux and pressure on the SM 

for the aft-facing thruster with the assumption of a cold wall (Twall = 70 °F). It shows that the dead panel 

underneath the RCS pod, the radiator panel, and the housing for auxiliary pod are subjected to plume 

impingement, where the flow is nearly transitional or free molecular. The maximum heating rate occurs 

on the housing of the auxiliary pod and reaches 0.35 Btu/ft2

-s. The heat flux on the radiator panel and the 

dead panel is not significant enough to cause any thermal concerns. 

Figure 2.—Comparison of RAMP2 and GASP solutions of the flow inside the engine and plume. 

Figure 3.—Comparison of RAMP2 and GASP solutions along the axial and circumferential directions. 
NASA/TM—2009-215601 6

Figure 4.—PLIMP results for the heating rate and pressure of the aft thruster. 

Figure 5.—PLIMP results for the heating rate and pressure of the forward thruster. 

Figure 5 shows plots of the corresponding PLIMP results for the forward-facing thruster, showing 

that the worst spot is the avionic ring right next to the thruster, where the maximum heating rate could 

reach 7.4 Btu/ft2-s. The flow around there is in the continuum flow regime and is turbulent. Since the heat 

flux is so significant, further thermal analysis was performed to compute the temperature to check 

whether any thermal protection or thermal shield is necessary for the avionic ring. The temperature results 

are presented in the following section. 

Table 2 shows comparisons of the maximum heating rate and pressure among the PLIMP, RPM, and 

PIDYN solutions. The RPM and PIDYN solutions were explained in private communications.*,†

 The 

PLIMP solution agrees well with the RPM solution for both forward- and after-facing thrusters. The 

heating rate and normal force from PLIMP are expected to be higher than those from RPM since the 

RAMP2 solution is more conservative. The PIDYN heating rate and normal force on the avionic ring are 

lower than the PLIMP and RPM solutions for the forward-facing thruster, whereas the PIDYN solution 

agrees better with the PLIMP and RPM solutions for the after-facing thruster. 

 

†

Smith, R., private communication, Lockheed Martin, Houston, Texas, June 2008. 

 
NASA/TM—2009-215601 7

TABLE 2.—COMPARISON OF PLIMP, RPM, AND PIDYN RESULTS 

[Twall = 70 °F.]

 PLIMP RPM PIDYN 

Forward thruster (continuum flow) 

 Maximum heating rate on avionic ring, Btu/ft2-s 7.4 6.43 6.47 

 Maximum pressure on avionic ring, psf 11.3 13.1 5.99 

Normal force on the avionic ring, lbf 3.7 3.6 2.81 

Aft thruster(free molecule flow) 

 Maximum heating rate, Btu/ft2-s 0.35 0.178 0.287 

 Maximum pressure, psf 0.039 0.04 0.05 

IV. Thermal Analysis on the Avionic Ring of the Service Module 

The finite-element model of the avionic ring on SM was built in MSC Patran/Pthermal using a 3500-

finite-element mesh. On the avionic ring, the outer face sheet (OFS) and inner face sheet (IFS) are 0.035-

in.-thick composite IM–7; in between is a 0.25-in.-thick aluminum honeycomb (H/C) core. In the thermal 

analysis, the plume impingement heating rate from PLIMP with Twall = 70 °F was imposed on the OFS; 

the contact resistance between the OFS and H/C core, and between IFS and H/C core, was 2 Btu/h-ft2

-F; 

the OFS radiated to deep space at T = –455 °F with an emissivity of 0.88. The IFS was adiabatic 

(insulated). The initial wall temperature was –455 °F. 

The computed temperature contour at t = 120 s is plotted in Figure 6, and the time history at the OFS, 

IFS, and H/C is plotted in Figure 7. It can be seen that the temperature at OFS reaches 1260 °F within 

40 s. Since the thermal resistance between OFS and H/C is high, not much heat goes through the H/C by 

conduction and most of the heat radiates to the ambient. Thermal protection or a thermal shield will be 

necessary for the area on the avionic ring that is subject to plume impingement since the temperature is 

far beyond the limit of the material. 

Figure 6.—Temperature contour of the avionic ring at t = 120 s. 
NASA/TM—2009-215601 8

Figure 7.—Temperature-time history of the outer face sheet (OFS), 

inner face sheet (IFS), and honeycomb (H/C). 

V. Conclusions 

Plume impingement of the Orion Service Module Reaction Control System was analyzed using the 

plume impingement program (PLIMP)/reacting and multiphase program 2 (RAMP2). The RAMP2 

solution of the plume was compared with the GASP solution and showed good agreement. Furthermore, 

the PLIMP results, including heating rate and pressure (force) on the impinging surfaces were compared 

with the corresponding results from RPM and PIDYN. A reasonable agreement was achieved. Further 

thermal analysis on the geometry of most concern (avionic ring) on the Service Module showed that 

thermal protection or thermal shields are necessary. 

References 

1. Smith, S.D., “High Altitude Chemically Reacting Gas Particle Mixtures,” LMSC–HREC TR 

D867400, Aug. 1984. 

2. Wojciechowski, C.J., and Penny, M.M., “Development of High Altitude Plume Impingement 

Analysis for Calculating Heating Rates, Forces and Moments,” LMSC–HREC D162867 I–III, 

Contract NAS8–25511, Mar. 1971. 

3. Simons, G.A., “Effect of Nozzle Boundary Layers on Rocket Exhaust Plumes,” AIAA Journal, 

Vol. 10, No. 11, 1972, pp. 1534, 1535. 

4. Lumpkin, F.E., LeBeau, G.J., and Kanipe, D.B., “Model for Predicting Orbiter PRCS Plume 

Impingement Loads and Heating: RPM Ver. 3.1.” 

5. Wurster, K.E., “MINIVER—a Versatile Aerothermal Analysis/TPS Design Tool,” Users Guide, 

Oct. 2000. 
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14. ABSTRACT

The Orion Crew Exploration Vehicle Service Module Reaction Control System engine plume impingement was computed using the plume

impingement program (PLIMP). PLIMP uses the plume solution from RAMP2, which is the refined version of the reacting and multiphase 

program (RAMP) code. The heating rate and pressure (force and moment) on surfaces or components of the Service Module were 

computed. The RAMP2 solution of the flow field inside the engine and the plume was compared with those computed using GASP, a 

computational fluid dynamics code, showing reasonable agreement. The computed heating rate and pressure using PLIMP were compared 

with the Reaction Control System plume model (RPM) solution and the plume impingement dynamics (PIDYN) solution. RPM uses the 

GASP-based plume solution, whereas PIDYN uses the SCARF plume solution. Three sets of the heating rate and pressure solutions agree 

well. Further thermal analysis on the avionic ring of the Service Module showed that thermal protection is necessary because of significant 

heating from the plume. 

15. SUBJECT TERMS

Thermal analysis; Plume impingement; Spacecraft 

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