J.T. Yim and D.A. Herman

Glenn Research Center, Cleveland, Ohio

J.M. Burt

Air Force Research Laboratory, Glenn Research Center, Cleveland, Ohio

Modeling Analysis for NASA GRC Vacuum

Facility 5 Upgrade

NASA/TM—2013-216496

February 2013
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J.T. Yim and D.A. Herman

Glenn Research Center, Cleveland, Ohio

J.M. Burt

Air Force Research Laboratory, Glenn Research Center, Cleveland, Ohio

Modeling Analysis for NASA GRC Vacuum

Facility 5 Upgrade

NASA/TM—2013-216496

February 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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Level of Review: This material has been technically reviewed by technical management. 
Abstract

A model of the VF5 test facility at NASA Glenn Research Center was developed using the direct

simulation Monte Carlo Hypersonic Aerothermodynamics Particle (HAP) code. The model results

were compared to several cold flow and thruster hot fire cases. The main uncertainty in the model

is the determination of the effective sticking coefficient—which sets the pumping effectiveness of the

cryopanels and oil diffusion pumps including baffle transmission. An effective sticking coefficient of

0.25 was found to provide generally good agreement with the experimental chamber pressure data.

The model, which assumes a cold diffuse inflow, also fared satisfactorily in predicting the pressure

distribution during thruster operation. The model was used to assess other chamber configurations

to improve the local effective pumping speed near the thruster. A new configuration of the existing

cryopumps is found to show more than 2× improvement over the current baseline configuration.

1 Introduction

Vacuum Facility 5 (VF5) at NASA Glenn Research Center is one of the world’s premier electric

propulsion (EP) test facilities. The main chamber is 4.6 m (15 ft) in diameter and nearly 18 m

(60 ft) long. Six cryopumps and twenty oil diffusion pumps (ODP) can provide a base pressure

on the order of 1 × 10−7 torr and a theoretical maximum pumping speed over 3.5 million L/s on

air. However, due to the present baseline configuration of the chamber, the actual local pumping

speed can be far lower, particularly at the end of the chamber away from the cryopumps. The

conductance losses through the chamber are in part responsible for this [1]. As NASA moves toward

higher power EP systems, the need for facilities to provide adequate ground testing capabilities is

required. A sufficiently low background pressure is necessary for proper characterization of thruster

performance, plume properties, and lifetime [2–4]. Thus, an effort to improve the effective pumping

speed of VF5 is underway. As part of that effort, the capabilities of VF5 are modeled to provide

understanding of facility pumping and to assess new potential configurations. Both cold flow and

thruster hot fire test data are collected to provide comparisons for model validation. A description

of the model, the testing, and results and comparisons of both are discussed below.

2 Model and Test Description

The Hypersonic Aerothermodynamics Particle (HAP) code was used to perform 3D Direct Simulation Monte Carlo (DSMC) simulations of facility pressure distributions for VF5 [5]. The rarified

environment of an operating vacuum chamber makes DSMC an ideal approach for these simulations. The DSMC method is a particle-based gas flow simulation technique which is derived from

principles of kinetic theory, and is the most mature and widely used technique for rarefied flow

simulation. In a DSMC calculation, 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. 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.

For the VF5 models, the chamber geometry—namely the outer shell of the main chamber and

bell jar—is analytically defined within the code through spherical and cylindrical shells. Also

included are the graphite shielding panels placed approximately midway along the chamber length

to protect the cryopanels from direct plume impingement. VF5 contains six large cryopanels and

twenty 32” diameter ODPs. A diagram of the current baseline configuration is provided in Figure 1.

	

	 
Figure 1. The baseline configuration of VF5 with four markers indicating approximate gas source

locations. The pressure distribution coloring is based on a short rails source location and 60 mg/s

xenon gas flow.

The two vertical cryopanels along the chamber centerline are 2.8 m × 3.0 m × 0.5 m in size and

they pump along their left and right vertical surfaces. The four horizontal cryopanels are 0.4 m ×

4.6 m × 1.4 m in size and pump along their top and bottom horizontal surfaces. The ODPs are

modeled to have a square pumping surface rather than circular for simplicity; the surface area

is kept the same. No other details of the vacuum facility geometry are explicitly modeled. The

symmetrical nature of the vacuum chamber across the vertical plane along the chamber centerline

is utilized to reduce the simulated volume in half.

The gas source is modeled as an annulus with a uniformly distributed inflow number flux to

simulate the flow exiting a Hall thruster discharge channel. The inflow number flux is calculated to

achieve the desired mass flow rate. The rest of the thruster and support structure geometry are not

modeled in these simulations. The inflow gas is modeled as a diffuse inflow at room temperature,

which is appropriate for cold gas flow conditions. Thruster hot fire is not directly modeled as Hall

thruster propellant particle exit velocities are high enough to require a prohibitively small time

step to resolve their motion and physics. Results presented below, however, will show that the cold

diffuse flow inflow conditions still approximate hot fire conditions to a fair degree. Other research

work has shown only minor differences noted for adding electrostatic fields and/or charge exchange

physics to DSMC flow simulations [6].

The model of VF5 was correlated to experimental pressure data to assess confidence in the

simulation results. Several different available data cases were used to check the validity of the

model. They are listed in Table 1. Four xenon gas source locations were used for the various

tests and their approximate positions are shown in Figure 1. The “short rails” location placed the

source, typically a thruster, approximately at the end cap / main chamber interface axial location.

The “extended rails” location is approximately 1.5 m further into the main chamber. The “mid

chamber reverse” placed the source just in front of the vertical shielding panels facing upstream

back towards the bell jar. The pressure color contour distribution in Figure 1 is based on 60 mg/s

xenon gas flow from the short rails location. Unless explicitly stated otherwise, all gas flow rates,

pressures, and pumping speeds reported in this document are for xenon gas.

The measured chamber pressure data are obtained using Granville-Phillips ion gauges and have

been corrected for xenon gas, which is used for all tests considered here. The ion gauge locations

are largely in the same locations for all tests and reported in Table 2. For the azimuthal locations,

	

	 	
Table 1. Test data sources

# Date Source Source location Active pumps

1 May 2011 Cold flow Bell jar & mid chamber Cryos only

2 Oct 2011 Cold flow Short rails ODPs only

3 Oct 2011 Cold flow Short rails Cryos + ODPs

4 Nov 2011 Cold flow Short rails Cryos + ODPs

5 Nov 2011 Hot fire Short rails Cryos + ODPs

6 Jan 2012 Hot fire Mid chamber reverse Cryos + ODPs

7 Apr 2012 Hot fire Extended rails Cryos only

Table 2. Ion gauge locations

Gauge Uncertainty Radial Azimuthal Axial

G-P 347 Stabil-Ion Module ±25% Along main chamber wall ∼8:00 ∼1 ft from end cap

G-P 390 Micro-Ion Module ±16% Along main chamber wall ∼8:00 ∼1 ft from end cap

G-P 390 Micro-Ion Module ±16% Along main chamber wall ∼4:00 ∼26 ft from end cap

G-P 390 Micro-Ion Module ±16% Along main chamber wall ∼4:00 ∼46 ft from end cap

G-P 370 Stabil-Ion Gauge ±5% ∼0.75 m from centerline ∼4:00 Along thruster exit plane

noon is along the vertical and the clockwise direction is based on looking into the chamber from

the bell jar end. The 95% confidence intervals of the chamber pressure reading uncertainties for

each ion gauge are also provided in Table 2. They are based on the accuracy and repeatability

values reported by the manufacturer. Error bars for the chamber pressure measurements in the

plots within Section 3 are set to these 95% confidence level values.

The primary unknown variable in these simulations is the sticking coefficient—the fraction of

incident particles that are effectively pumped through cryocondensation, cryosorption, or cryotrapping. The sticking coefficients of gases on a bare cryosurface have been found to lie generally within

a range of 0.6 – 1.0 depending on the gas and surface temperature [7]. However, the VF5 cryopumps

have liquid nitrogen-cooled chevron baffles in front of the cryosurfaces. The transmission probabilities through these baffles—which were provided to be 0.30 to 0.35—are accounted for through a

modified effective sticking coefficient. A series of computational runs was performed to assess the

sensitivity of the modeling results to the applied effective sticking coefficient. The model was set

up to correspond to tests #3 and 4 in Table 1. The effective sticking coefficient was varied from

0.15 to 0.40 in increments of 0.05. The pressure was assessed at four locations corresponding to the

bottom four gauges listed in Table 2. The averaged measured pressures with their corresponding

uncertainty are plotted as shaded bands in Figure 2. The model results are plotted as symbols.

For the pressure near the chamber walls, an effective sticking coefficient of 0.20 to 0.25 was found

to provide the best guess fit. The near thruster pressure data required a sticking coefficient closer

to 0.40. As a general compromise, with some emphasis on the chamber wall locations where most

of the test data is and will be measured, an effective sticking coefficient of 0.25 was chosen. This is

comparable to other published cryopumping simulation work which has found values around 0.3 –

0.4 for the effective coefficient to provide a good fit to their data [8]. An effective sticking coefficient

of 0.25 is used for all results presented below.

Some of the results are presented in terms of pumping speed instead of pressure. The pumping

speed is easily calculated from

P S = (  ̇m mg/s)(22.4 L-atm/mol)(760 torr/atm)

(1000 mg/g)(M g/mol)(P − Pbase torr) (1)

where  ̇m is the mass flow rate, M is the molecular weight of the gas, and Pbase is the base pressure

measured when there is no gas inflow. The model assumes a perfectly leak tight vessel with no

	

	 
Figure 2. The effect of the sticking coefficient on modeled chamber pressures at four locations.

The shaded bars correspond to the 95% confidence interval of the measured test data at those four

locations. The symbols plot the model results.

presence of non-condensable species such as helium, therefore Pbase would theoretically be zero for

the model. Since the pressure in the chamber will vary with location, the pumping speed too will

vary throughout the chamber.

3 Results and Discussion

The VF5 model is run for several cases corresponding to different thruster and chamber configurations. For all cases, the location and orientation of the cryopumps and ODPs are the same. A

set of tests were run that operated: 1) cryos only, 2) ODPs only, and 3) all pumps in the chamber.

The effect of the pumping scheme is evaluated with the model. Results are presented in Figure 3.

The model shows a fair approximation to the pressure data collected along the chamber wall. The

model results for the ODP-only case does not capture the drop in pressure at the far end of the

channel, however, the overall magnitude of the pressure tracks with the recorded data, particularly

near the thruster, which is the location of most interest. Different xenon gas flow rates were also

examined. Testing included adjusting the inflow rate of xenon cold gas inflow from 20 to 107 mg/s

with all of the pumps operating. Models were run at 20, 60, and 100 mg/s and the results are

compared in Figure 4. Again, the magnitude of pressure change is well captured by the model.

The model was also compared to pressure data obtained during thruster hot fire testing. The

model assumes a diffuse cold gas inflow, as simulating the high velocity plume particles requires

prohibitively small time steps. Test #5 in Table 1 corresponds to NASA-457Mv2 Hall thruster performance testing at flow rates ranging from 18.5 to 92.4 mg/s [9]. Pressure data were collected from

a screen-shielded Stabil-ion gauge located 0.75 m from the thruster centerline along the thruster

exit plane. Similar data were collected from cold flow test #4 with the same thruster, chamber,

and ion gauge configuration. The results from the two tests and the corresponding cold flow model

results are shown in Figure 5.

Two other thruster hot fire cases are examined. Both operate the thruster at different locations

than the standard location on the short rails. The thruster is placed on an extended set of rails

further into the main chamber for one test, while the other places the thruster near mid-chamber

	

	 
Figure 3. Comparison of model and test chamber pressure results for different pumping operations

in VF5.

Figure 4. Comparison of model and test chamber pressure results for different flow rates in VF5.

Figure 5. Comparison of chamber pressure results 0.75 m from thruster centerline for cold flow

test, hot fire test, and cold flow model.

	

	 
Figure 6. Comparison of chamber pressure along chamber walls for the extended rails thruster

location hot fire test.

Figure 7. Comparison of chamber pressure along chamber walls for the mid chamber reverse thruster

location hot fire test.

and points the thruster back towards the bell jar end. The model was run for both of those cases.

The comparison for the extended rails configuration is shown in Figure 6 and the mid chamber

reversed configuration is shown in Figure 7.

The test cases described above provide a confidence in the ability of the model to predict the

chamber pressure for the baseline configuration of VF5. The model was utilized to assess the

pumping speed of alternate VF5 configurations. The other configurations included moving the

existing panels as well as adding new ones. The initial set of configurations examined is presented

in Figure 8. It is seen that the presence of cryopanels along the length of the chamber significantly

reduces the pressure in the upstream regions of the chamber near the thruster, as compared to the

baseline configuration. This is due to reducing conductance losses before the gas reaches a pumping

surface [1].

Ultimately, a configuration relocating all of the existing cryopanels was chosen based on pumping

effectiveness, cost, backsputter rate, and other facility considerations. The new proposed chamber configuration moves the long horizontal panels upstream closer to the bell jar as shown in

Figure 8(d). The two vertical panels will also be moved parallel to each other at the far end of

	

	 
(a) Baseline

(b) Add 2 new panels (c) Add 10 new panels

(d) Move horizontal panels (e) Move all existing panels, add 2 new panels

Figure 8. Initial set of chamber configurations examined.

Figure 9. The new proposed configuration of VF5 which rearranges all existing cryopanels. The

pressure distribution coloring is based on a short rails source location and 60 mg/s xenon gas flow.

	

	 
(a) Axial distribution of pressure and local pumping speed along chamber wall

(b) Radial distribution of pressure and local pumping speed along thruster exit plane

Figure 10. Comparisons of pressure and local pumping speed distributions between baseline and

new proposed chamber configurations for 60 mg/s xenon flow at the short rails location.

the chamber as configured in Figure 8(e). A model of the new configuration was developed and

is shown in Figure 9. Both Figure 9 and Figure 1, of the baseline configuration, show a pressure

distribution based on 60 mg/s of xenon flow from a thruster at the “short rails” location. A closer

look at the pressure distribution is shown in Figure 10. The axial distribution of the pressure along

the chamber wall again shows a marked reduction in the upstream pressure towards the thruster

location while the downstream pressure near the far end of the chamber is similar as before. The

radial pressure distribution along the thruster exit plane vertical axis also shows improvement over

the current baseline. The pumping speed is calculated for both configurations based on the pressure

0.75 m from the thruster centerline along the thruster exit plane and shown in Figure 11. More than

two times improvement, from 320 kL/s to 700 kL/s, is seen in the modeling results for a 60 mg/s

flow rate. The pumping speeds calculated from pressure measurements at the same location during

thruster hot fire tests are also shown in the figure.

	

	 
Figure 11. Comparison of calculated xenon pumping speeds 0.75 m from thruster centerline in exit

plane.

4 Conclusion

A model of VF5 was developed and run using the HAP DSMC code. The baseline cryopanel

configuration of VF5 was examined across several different source and pumping cases to assess

the validity of the model as compared to measured test data. VF5 was operated with different

pumps activated, different gas inflow rates, and different thruster locations. The model captured

the pressure changes along the length of the chamber and compared well with test results. The

main source of uncertainty in the modeling results was the estimation of the effective sticking

coefficient used to model the effects of the transmission probability through the baffles in front of

the pumping surfaces as well as the sticking probability of the gas particles to be pumped away after

they reach the pumping surface. A coefficient of 0.25 was found to achieve an acceptable comparison

to the measured test data for the cases examined. One other possible significant contributor to

model uncertainty is the approximate geometry of the chamber and its contents. A more detailed

geometrical description of the chamber, as perhaps represented in a high fidelity CAD model, would

help further reduce uncertainty in the model results.

In addition to the present baseline configuration of VF5, several new variants were examined to

assess potential improvements in pumping speed throughout the chamber. These options included

moving the current existing cryopanels as well as including additional new panels. Ultimately, a

new configuration that rearranges the existing panels was chosen based on a variety of factors. The

new configuration is expected to provide more than 2× the pumping speed near the thruster than

the present baseline setup.

References

1. Dankanich, J. W., Swiatek, M. W., and Yim, J. T., “A Step Towards Electric Propulsion Testing Standards: Pressure Measurements and Effective Pumping Speeds,” 48th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 2012, AIAA-2012-3737.

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on Stationary Plasma Thruster Testing,” 23rd International Electric Propulsion Conference,

1993, IEPC 93-93.

	

	 
3. Byers, D. and Dankanich, J. W., “A Review of Facility Effects on Hall Effect Thrusters,” 31st

International Electric Propulsion Conference, 2009, IEPC-2009-076.

4. Diamant, K. D., Spektor, R., Beiting, E. J., Young, J. A., and Curtiss, T. J., “The Effects

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Propulsion Conference and Exhibit, 2012, AIAA 2012-3735.

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Vacuum Chamber,” 29th International Electric Propulsion Conference, 2005, IEPC-2005-138.

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9. Soulas, G. C., Haag, T. W., Herman, D. A., Huang, W., Kamhawi, H., and Shastry, R., “Performance Test Results of the NASA-457M v2 Hall Thruster,” 48th AIAA/ASME/SAE/ASEE

Joint Propulsion Conference and Exhibit, 2012, AIAA 2012-3940.