Brian Reed

Glenn Research Center, Cleveland, Ohio

Decomposing Solid Micropropulsion Nozzle

Performance Issues

NASA/TM—2003-212225

May 2003

AIAA–2003–0672

https://ntrs.nasa.gov/search.jsp?R=20030060681 2018-05-15T18:31:38+00:00Z
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Brian Reed

Glenn Research Center, Cleveland, Ohio

Decomposing Solid Micropropulsion Nozzle

Performance Issues

NASA/TM—2003-212225

May 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno, Nevada, January 6–9, 2003

AIAA–2003–0672
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7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

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Available electronically at http://gltrs.grc.nasa.gov
Decomposing Solid Micropropulsion Nozzle Performance Issues

Brian Reed

National Aeronautics and Space Administration 

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Micropropulsion technology is essential to the success

of miniaturized spacecraft and can provide ultra-precise

propulsion for small spacecraft. NASA Glenn 

Research Center has envisoned a micropropulsion 

concept that utilizes decomposing solid propellants 

for a valveless, leak-free propulsion system. Among 

the technical challenges of this decomposing solid 

micropropulsion concept is optimization of 

miniature, rectangular nozzles. A number of flat 

micronozzles were tested with ambient-temperature 

nitrogen and helium gas in a vacuum facility. The 

thrusters were etched out of silicon and had throat 

widths on the order of 350 microns and throat depths 

on the order of 250 microns. While these were halfsections of thrusters (two would be bonded together 

before firing), testing provided the performance trend 

for nozzles of this scale and geometry. Area ratios 

from 1 to 25 were tested, with thrust measured using 

an inverted pendulum thrust stand for nitrogen flows 

and a torsional thrust stand for helium. In the 

nitrogen testing, peak nozzle performance was 

achieved around area ratio of 5. In the helium 

testing, nozzle performance peaked for the smallest

nozzle tested, area ratio 1.5. For both gases, there 

was a secondary performance peak above area ratio 

15. At low chamber pressures (< 1.6 atm), nitrogen 

provided higher nozzle performance than helium.

The performance curve for helium was steeper, 

however, and it appeared that helium would provide 

better performance than nitrogen at higher chamber 

pressures.

Introduction

In recent years, there has been interest (primarily in 

government and academia sectors) toward shrinking 

the size of spacecraft, so that the costs and risk of 

missions can be reduced.1,2 Spacecraft in the range of 

20- to 100-kg represent the class most likely to be 

utilized by most “small satellite” users in the near 

future. There are efforts to develop 10- to 20-kg 

class spacecraft for use in satellite constellations. 

There are more ambitious efforts to develop 

spacecraft less than 10-kg, employing microelectricalmechanical systems (MEMS) fabrication 

technology. Obviously, as spacecraft size decreases 

the difficulty in miniaturizing spacecraft systems 

increases (size is given here in terms of mass, but the 

implications for the volume envelope are even more 

severe). 

Propulsion represents the most challenging spacecraft 

system to miniaturize. There is a point where 

reduced-scale versions of conventional propulsion 

systems will no longer be practical. This restriction 

is not only a result of practical volume, mass, and 

power limitations, but also of the physics at reduced 

scale, where forces ignored at conventional scale 

become dominant.3

 At this point, a fundamentally 

different approach to propulsion must be taken. To 

address this need various efforts are underway to 

develop unique miropropulsion technologies for 

miniature spacecraft applications.4,5,6

Micropropulsion systems not only have to be reduced 

scale in mass and volume, but also use little or no 

power, be relatively simple yet flexible enough to use 

in multiple functions, and be cost effective for what 

is likely to be a relatively low-cost spacecraft.

Micropropulsion applications would not be necessary 

restricted to MEMS spacecraft as micropropulsion 

can be used to conduct precision maneuvers for 

larger spacecraft. Functions for larger spacecraft 

include spin-up and spin-down, precision positioning 

and pointing, constellation maintenance, and 

deorbiting. Grouping micropropulsion systems in 

arrays will allow their use for larger thrust 

applications. These “macroscale” functions of 

micropropulsion may ultimately be the first 

application of this technology. 

This report concerns a valveless, flexible 

micropropulsion system that uses decomposing solids 

as propellants.7

 The concept would use an array 

composed of hundreds (or thousands, as necessary) 

of MEMS thruster units, arranged in the 

NASA/TM—2003-212225 1
configuration best suited for a particular application. 

Different thruster sizes (throat areas, nozzle area 

ratios) would provide for a range of thrust levels 

(from μN’s to mN’s) within the same array. Several 

thrusters could be fired simultaneously for thrust 

levels higher than the basic units, or in a rapid 

sequence in order to provide gradual but steady low-g 

acceleration. Decomposing solid propellant pellets 

(slow-burning propellants referred to as gas 

generators) would provide long-term, leak-free 

storage. Initiation of the propellants would be 

accomplished with a diode laser-based, fiber-optic 

network. Silicon is transparent at wavelengths above 

1.3 microns, allowing laser penetration and initiation 

without structure-compromising feedthroughs. 

Power requirements for initiation would be less than 

100 milliwatts.

For this decomposing solid micropropulsion concept, 

there are five major technology challenges that need 

to be addressed, summarized in Table I. Reliable, 

reproducible fabrication of miniature thrusters having 

widths on the order of 300 microns and aspect ratios 

of at least 2:1 is a critical aspect of many 

micropropulsion efforts. Since, in this concept, 

thrusters are likely to be composed of two halfsections to allow the propellant grain to be inserted in 

the chamber, bonding of half-sections without 

initiating the propellant is a critical procedure in 

thruster fabrication. Another technical challenge is 

selection of the decomposing solid propellant. The 

propellant needs to provide high performance (Isp > 

200 sec) in a low-heat flux decomposition process 

with benign decomposition products. The initiation

of propellant using a diode laser represents a 

technical challenge that is key to this 

micropropulsion concept. Initiation of propellant 

through the thruster walls provides for a valveless, 

leak-free system, but care must be taken not to create 

a burdensome system of fiber optics and power 

requirements need to be low. Finally, the individual 

components (microthrusters with propellant, 

initiation system) must be packaged into an array that 

can be configured for multiple and different 

spacecraft propulsion functions. 

The issue of optimization of rectangular nozzles, 

subjected to low Reynolds number, viscousdominated flows is the subject of this report. There 

have been experimental and computational efforts 

investigating the performance of miniature nozzles.8,9

Established approaches to optimization of nozzles 

breakdown at lower Reynolds numbers, where 

viscous effects begin to become more prevalent. 10,11

At the miniature scale, the predominance of viscous 

effects is amplified.3

 Furthermore, the nature of the 

MEMS manufacturing process results in rectangular 

rather than circular nozzles. Depending on the aspect 

ratio of the nozzle, the top and bottom wall 

boundaries are likely to heavily influence the 

flowfield behavior. These physical differences of 

nozzles at the miniature scale indicate a different 

optimization of performance.

Challenge Issue

Micro-Thruster 

Fabrication

Large depths (>200 

μm) needed with tight 

tolerances on throat 

dimensions and lowtemperature bonding

Micro-Nozzle 

Optimization

Viscous-dominated 

flow in rectangular 

nozzle

Propellant Selection Need highperformance, low-heat 

flux, benign propellant 

that is compatible with 

laser initiation

Diode Laser Initiation Need to establish 

parameters for lowpower, reliable 

initiation

Array System 

Integration

Need efficient 

packaging of 

microthrusters, 

switching of fiber 

optics, and sequencing 

of thruster firing

Table I: Technical Challenges for Decomposing 

Solid Micropropulsion Concept

This report summarizes the performance testing of 

miniature, rectangular nozzles on ambienttemperature nitrogen and helium gases. The nitrogen 

test data has been previously reported,12 although 

there was an error in reporting the flowrate data that 

will be corrected in this report. The helium test data

is being reported for the first time here. Modeling of 

these micronozzle contours has been conducted using 

ambient-temperature nitrogen13 and simulated 

combustion products of a decomposing solid 

propellant.14 The modeling and experimental are 

complimentary efforts aimed at determining the 

optimal area ratio and configuration of micronozzles 

over a range of gas flow conditions (molecular 

weight, temperature).

NASA/TM—2003-212225 2
Test Apparatus and Procedure

Thruster/Holder Assembly

All testing was conducted using miniature thruster 

half-sections fabricated from 500-μm thick silicon 

wafers. The basic thruster half-section contour is as 

shown in Figure 1. This contour is a laboratory 

design and doesn’t represent an engineering model 

configuration. The designed chamber diameter was 

6.35 mm. The chamber-to-throat contraction angle 

was 49.5 ̊. The design throat widths was 350-μm,

while the throat depth and length was 300 μm. The 

nozzle expansion angle was 27.7 ̊ and the nominal 

area ratios were 1, 5, 10, 15, 18.25, 20, and 25. 

When the thruster half-sections were actually 

fabricated, there was considerable variability in the 

dimensions.

Figure 1: Micronozzle thruster contour

The miniature thruster half-sections were installed in 

custom made thruster holders (Figure 2), which were 

then mounted to the thrust stand. The assembly had a 

small line for the gas inlet, with ports for pressure 

and temperature measurements. A method was 

developed to seal the thruster half section within the 

stainless steal thruster holder. With two hard 

surfaces, the smallest flaws on either piece could 

cause a fracture in the brittle silicon wafer. A thinfilm gel was used to seal the thruster to its holder, 

with moderate clamping action applied to obtain a 

good seal without cracking the brittle thruster. Under 

pressure, the gel will deform to the imprint of the 

thruster, and slightly decrease the flowpath cross 

section. The soft interface is also intended to 

distribute the loading on the wafer more evenly. A 

matching backing is applied on the back surface of 

the half thruster. However, at the parts where there is 

virtually no loading, i.e. the etched surface of the 

thruster, bowing will occur, stressing the silicon 

wafer. On a couple of occasions, the thruster cracked 

precisely down the center of the inner chamber, 

clearly due to this loading stress. For each 

thruster/holder assembly, leak tests were performed 

by pressurizing the feed system and thruster assembly 

to a pressure of 0.7 MPa. Only leak-free assemblies 

continued to be tested.

Figure 2: Schematic of thruster holder

Test Facility/Setup

Experiments were performed inside one of NASA 

GRC’s vacuum facilities, a 200-cubic foot test tank, 

equipped with a 2-foot and a 3-foot diameter port, 

both isolated from the main test tank through 

hydraulic valves. This allows interchange of test 

hardware without the need to fully recycle the tank 

vacuum. Dual diffusion pumps are able to maintain a 

vacuum down to 8 x 10-7 torr with low flow rate. At 

increasing flow rate, tank pressure will rise, up to 

approximately 7.5 x 10-4 torr with 15000 sccm of 

nitrogen, at which point the diffusion pumps are set 

to shut off, and mechanical pumps will take over. 

Two ion gages at either side of the tank monitor tank 

pressure. 

A schematic of the test setup is shown in Figure 3. 

Chamber pressure was measured in the feed system 

slightly upstream of the chamber with a 200 psi, 

high-frequency, piezoelectric pressure transducer. 

Temperature was monitored with a type-K 

thermocouple. A Labview digital data acquisition 

system was used to record thrust, pressure and 

temperature, at a 50-Hz data rate. A strip chart 

recorder was also used to provide real time thrust 

measurement. Vacuum tank pressure and barometric 

pressure were manually recorded from facility 

monitors. A bank of five flow controllers of varying 

ranges (10, 100, 1000, 10000, and 22000 sccm 

nitrogen), mounted in parallel, were used to set the 

volumetric flowrate. 

An inverted pendulum thrust stand was used for 

testing with nitrogen. The thrust stand had been used 

NASA/TM—2003-212225 3
previously in the testing of electric propulsion 

thrusters, such as arcjets, Hall thrusters, and MPD 

thrusters.15 Figure 4 show the thruster assembly 

mounted on a bracket, which in turn is mounted on 

the thrust stand. Thrust stand capability ranged from 

less than a milliNewton to over a Newton. A 

torsional-type thrust stand was used in the helium test 

series. This is a more sensitive thrust stand used in 

the testing of pulsed plasma thrusters (PPT’s).16 In 

both thrust stands, thrust was determined by the 

deflection of the stand as measured by a LVDT. In 

all testing, thrust calibrations were conducted directly 

before each test. 

Figure 3: Schematic of Test Setup

Figure 4: Thruster Assembly Mounted on Thrust 

Stand, View of Nozzle Exit (Not 

Discernable in Holder)

Test Procedure/Data Reduction

For all testing, the desired flowrate was set by the 

mass flow controllers. Test runs were typically 160 

seconds in duration, sufficiently long to perform a 

calibration on the thrust stand, start the gas flow, and 

allow the thrust and pressure to reach steady state. A 

steady-state thrust signal was usually achieved within 

5 seconds from the start of the gas flow. Digital data 

were recorded in 0.02-second intervals. In the thrust 

calibration, three weights (whose mass was measured 

precisely) stringed on a filament wire were 

successively applied to the thrust stand and then 

removed.

In the spreadsheet data file, an average of 100 points 

on each "step" (including the zeroes) was used to 

represent a data point on the calibration curve. A 

least squares fit was done of the calibration data (both 

the upward and downward curves). The curve fit 

parameters were used to determine thrust from the 

steady-state portion of the curve. An average of 1000 

points, taken well into the steady-state region, was 

used to represent the thrust reading. Some 

fluctuations in the thrust level existed, but it was less 

then 3% of the nominal thrust level. The measured 

thrust was corrected to vacuum thrust by adding the 

product of the vacuum pressure and the nozzle exit 

area.

Gage chamber pressure was corrected to absolute 

pressure using a barometric pressure reading taken in 

the facility. An average of 1000 points in the steadystate region were used (simultaneous with the thrust 

measurement) to represent the chamber pressure 

reading.

In the nitrogen series, testing was conducted at 

flowrates of 10, 20, 50, 100, 200, 500, 1000, and 

2900 sccm of nitrogen. For higher area ratio 

thrusters, testing was also done at 4350, 5800, 7250, 

and 8700 sccm of nitrogen. Testing was started at 10 

sccm, increased to the maximum allowable by the 

capacity of the diffusion pumps, then decreased back 

to 10 sccm. In this way, two sets of measurements 

were obtained at each flow rate (except for the 

maximum). A similar procedure was used in the 

helium testing, with the volumetric flowrates set the 

same as in the nitrogen testing, but with the mass 

flowrate corrected for helium.

Bias Error in Test Data

Thrust coefficient performance data for nozzles 

tested with ambient-temperature nitrogen were 

previously reported in Reference 12. The modeling 

effort on these nozzles at Penn State lead to a 

determination that the measured flow rates and 

chamber pressures were in disagreement. In most 

cases, characteristic velocity and specific impulse 

values based on pressure and thrust measurements 

were much higher than maximum theoretical for the 

mass flowrate measurements. This indicated a 

NASA/TM—2003-212225 4
significant error in one or more of the measured 

parameters. The calculation of thrust coefficient 

(from the measured thrust, chamber pressure, and 

throat area) as the nozzle performance parameter did 

not reveal this discrepancy.

The experimental setup was disassembled before the 

discrepancy was discovered, however, an 

examination of the test procedures for determining 

each of the measured parameters suggests the source 

of the error. The measured parameters were,

• Throat Area, determined from measurements of 

the throat width and depth, using either a 

scanning electron microscope (SEM) or a

measuring microscope

• Chamber Temperature, measured at the inlet tube 

of the holder (just upstream of the micronozzle 

chamber), using a chromel-alumel thermocouple

• Chamber Pressure, measured at the inlet tube of 

the holder (just upstream of the micronozzle 

chamber), using a 0 to 200 psig dynamic 

pressure transducer and corrected to psia, using a 

barometric pressure gauge in the vacuum test 

facility

• Mass Flowrate, determined from the volumetric 

flowrate set by a bank of five volumteric flow 

controllers (10, 100, 1000, 10000, and 22000 

sccm nitrogen), corrected to mass flowrate 

according to the gas used

• Vacuum Thrust, determined from the site thrust 

measured directly on the thrust stand and 

corrected to vacuum thrust. The product of the 

tank pressure (assumed to be the nozzle exit 

pressure) measured from two ion gauges in the 

tank and the nozzle exit area, determined from 

measurements of the nozzle width and throat 

depth (assumed to be the same as the exit depth) 

using an measuring microscope, is added to the 

measured thrust

The source of error cannot be determined definitely, 

since the experimental setup had been long since 

disassembled when the error was discovered. All of 

the flow controllers were calibrated before testing. A 

small leak in the large flow controllers feeding the 

same manifold as the smaller flow controllers could 

have lead to erroneously low mass flowrate 

measurements at the small flowrates used in the 

testing. Each flow controller was isolated with a 

hand valve, which had to leak as well. The pressure 

transducer was also calibrated before testing and 

there was no indication of improper operation during 

testing. An error in the thermocouple used for 

stagnation temperature would have been obvious 

during testing as an open circuit. The dimensional 

measurements (throat and nozzle exit) were made in 

a procedure described in Reference 12. The errors in 

those measurements are given in Table II (the 

measurements made with SEM are noted in the 

table). They are not large enough to account for the 

large errors indicated in the performance parameters.

The rocket performance parameters that can be 

calculated from the basic measurements are,

• Characteristic Velocity (C*) = 

Chamber Pressure * Throat Area/ Mass Flowrate

• Thrust Coefficient (Cf) = 

Vacuum Force/(Chamber Pressure* Throat Area)

• Specific Impulse (Isp) = 

Vacuum Force/Mass Flowrate

Generically, C* measures chamber performance 

(upstream of the throat), Cf measures nozzle 

performance (downstream of the throat), and Isp is an 

overall measure of rocket performance. C* and Isp 

were higher than theoretically possible, in most cases, 

significantly higher. The error increased as the 

flowrate decreased and was substantial for flowrates 

below 100 sccm. Cf, however, was very low, which 

seemed to be reasonable for miniature, rectangular 

nozzles, dominated by viscous flows.

The evidence, then, points to an error in the mass 

flowrate, since they are in denominator of both 

equations for C* and Isp, while it is absent from the 

Cf equation. If the actual mass flowrate was much 

higher than the flowrate thought to be set by the flow 

controllers, then both C* and Isp would be inflated, 

while Cf would be unaffected. If an inflated vacuum 

force was responsible for the high Isp values, then it 

would not be reflected in the C* numbers. In the 

same line of reasoning, if an inflated chamber 

pressure was responsible for the high C* numbers, 

then Isp would not be affected. Of course, there is 

the possibility that both vacuum force and chamber 

pressure were inflated, canceling each other in the Cf 

equation. There is no definite way of disproving this 

assertion, although, two different thrust stands were 

used for the nitrogen and helium testing. Since the 

error occurred in both test series, it seems unlikely 

that a bias existed in both thrust stands. 

NASA/TM—2003-212225 5
Nozzle 

ID

(Series)

Area 

Ratio

Throat 

Width 

(μm)

Error 

(%)

Exit 

Width 

(μm)

Error 

(%)

Depth 

(μm)

Error 

(%)

146 

(N2)

1.0 372

0.2%

372

0.2 %

295

5.8 %

153 

(N2 & 

He)

1.5 230

(SEM)

8.9 %

353 

(SEM)

5.8 %

271 

(Est.)

7.0 %

169 

(N2)

4.8 367

0.3 %

1756

0.2 %

198

3.4 %

170 

(He)

4.8 366

0.2 %

1761

0.2 %

246

4.8 %

178 

(N2)

9.7 362

0.3 %

3511

0.04 %

260

1.8 %

182 

(He)

9.7 362

0.1 %

3510

0.01 %

385

2.5 %

195 

(He)

14.5 362

0.3 %

5260

0.01 %

238

4.0 %

208 

(N2)

14.9 353

0.8 %

5246

0.1 %

220

8.6 %

226 

(N2 & 

He)

18.3 380 

(SEM)

4.2 %

6933 

(SEM)

2.7 %

243 

(Est.)

5.8 %

230 

(N2)

19.5 360

0.2 %

7008

0.1 %

243

3.2 %

243 

(N2 & 

He)

23.1 367 

(SEM)

3.3 %

8467 

(SEM)

1.8 %

214 

(Est.)

3.1 %

237 

(N2)

24.2 361

0.3 %

8759

0.01 %

221

15.7 %

Table II: Dimensional Measurements of Micronozzles

All of the testing was conducted using a bank of mass 

flow controllers, ranging from 10 to 22,000 sccm, 

arranged in parallel. When the test apparatus was put 

together, it was thought that testing would be 

conducted well above 10,000 sccm. Testing always 

started at 10 sccm and most of the data was gathered 

at points below 2900 sccm (usually above that value, 

the vacuum in the test tank was compromised). Any 

small addition of gas flow to the system could 

seriously bias the volumetric flowrate values. This 

rationale seems to be supported by the fact that the 

larger errors occurred at the small flowrate values. 

Figures 5 and 6 plot the C* values versus volumetric 

flowrate for nitrogen and helium, respectively. There 

clearly is an error in most, if not all, of the values, but 

the error increases greatly as flowrate decreases.

The above arguments seem to suggest an undetected 

bias in the volumetric flowrate measurements in both 

test series. The bias had the affect of creating a 

larger flowrate than indicated by the instrumentation. 

If there is error in the flowrate measurements, then 

the Reynolds number calculations are in error (lower 

than the actual value). The same is true of the 

performance parameters calculated using mass 

flowrate (C* and Isp), as well as theoretical thrust 

calculated from isentropic relations. However, there 

is no reason to suspect the force measurements, 

particularly since two different thrust stands were 

used in the nitrogen and helium testing. Thus, the 

data collected can be accurately presented in terms of 

thrust coefficient versus stagnation pressure for a 

given set of area ratios, but the calculated Reynolds 

number is low by orders of magnitude at the lowest 

flowrates.

Figure 5: Nitrogen C* Performance Based on 

Suspected Bias Flowrate Data

NASA/TM—2003-212225 6
Figure 6: Helium C* Performance Based on 

Suspected Bias Flowrate Data

Test Data

Despite the issues with the flowrate data, the 

conclusions drawn in the previous report did not 

change. Figure 7 plots the thrust coefficient versus 

area ratio for four sets of chamber pressures. In 

Figure 8, thrust coefficient is plotted against chamber 

pressure for each nozzle (set of area ratios). 

Appendix A lists the same data in tabular format. 

The data shows that nozzle performance increases to 

a maximum around area ratio 5, then drops off 

sharply to a minimum between area ratio 10 and 15. 

Above area ratio 15, however, there is another 

performance increase until a secondary maximum 

around area ratio 20. The data are consistent with 

computational results using Direct Simulation Monte 

Carlo (DSMC) in a three-dimensional contour, out to 

area ratio 10, which show a maximum thrust at an 

area ratio of 6.13

Figure 9 plots thrust coefficient versus area ratio over 

three sets of chamber pressures for the thrusters 

tested with cold helium gas. Figure 10 shows thrust 

coefficient versus chamber pressure for each nozzle 

(set of area ratios). Appendix B lists the same data in 

tabular format. There are fewer data points than the 

nitrogen test series, but the maximum performance 

was achieved at the lowest area ratio tested (1.5). 

Performance decreased after this point, with a

minimum occurring around area ratio 10. As with 

the nitrogen test series there was an increase in 

performance above area ratio 15. Unlike the nitrogen 

flow, this performance increase continued beyond 

area ratio 20, although there were few data points to 

draw conclusions from.

Figure 7: Thrust Coefficient vs. Area Ratio for Cold 

Gas Nitrogen

Figure 8: Thrust Coefficient vs. Chamber Pressure 

for Cold Gas Nitrogen

Figure 9: Thrust Coefficient vs. Area Ratio for Cold 

Gas Helium

NASA/TM—2003-212225 7
Figure 10: Thrust Coefficient vs. Chamber Pressure

for Cold Gas Helium

Figures 11, 12, and 13 compare nitrogen and helium 

thrust coefficients for nozzles with area ratio nozzles 

1.5, 18.3, and 23.1, respectively. The nitrogen flows 

always provide better nozzle performance than 

helium at chamber pressures below 1.6 atm.

However, the slope of the helium curve was steeper 

than the nitrogen curve, suggesting that at higher 

chamber pressures, helium flows would provide 

better performance than nitrogen. 

Thrusters in a decomposing solid micropropulsion 

system could be sized for a range of chamber 

pressures, even within the same array, depending on 

the propulsion function desired. The propellants used 

in the concept are likely to produce primarily 

nitrogen,7 so there may be a performance advantage 

to operating at 1 to 2 atm chamber pressure.

Figure 11: Nitrogen and Helium Nozzle Performance 

for Area Ratio = 1.5 (Micronozzle 153)

Figure 12: Nitrogen and Helium Nozzle Performance 

for Area Ratio = 18.3 (Micronozzle 226)

Figure 13: Nitrogen and Helium Nozzle Performance 

for Area Ratio = 23.1 (Micronozzle 243)

Summary

A micropropulsion concept that utilizes decomposing 

solid propellants offers many advantages for small 

and miniature spacecraft. However, there are many 

technical challenges facing the development of this 

concept, including performance of the rectangular 

micronozzles. Testing has been conducted of flat 

micronozzles under ambient-temperature nitrogen 

and helium flows. This testing, in conjunction with 

numerical modeling, has indicated that a short area 

ratio will provide the best performance for 

micronozzles. Testing with nitrogen showed a 

performance peak around area ratio 5, which is 

consistent with computational modeling that has been 

conducted. Testing with helium flows showed

performance peaking at area ratio 1.5, suggesting that 

no nozzles might be best for these flows. Testing 

with nitrogen and helium flows both indicated 

performance minimums around area ratio 10 and 

secondary performance peaks above area ratio 15. 

Testing with ambient-temperature flows in flat, 

rectangular micronozzles are the first step to 

understanding nozzle optimization at this scale. 

More testing needs to be conducted to understand the 

influence of throat area (small throat widths), throat 

aspect ratio (larger depths), and gas temperature 

NASA/TM—2003-212225 8
(heated flows) on the performance of these nonoptimized nozzles.

References

1. Collins, D., Kukkonen, C., and Venneri, S., 

“Miniature, Low-Cost, Highly Autonomous 

Spacecraft – A Focus for the New Millennium,” 

IAF 95-U.2.06, 46th International Astronautical 

Congress, Oslo, Norway, October 1995.

2. Proceedings of the AFOSR/DARPA/AFRL/VSD 

Micro/Nanotechnology for Micro/Nanosatellites 

workshop, Edited by A. Das, Alburqueque, NM, 

April 1998.

3. Ketsdever, A. and Mueller, J., “Systems 

Considerations and Design Options for 

Microspacecraft Propulsion Systems,” AIAA 99-

2723, 35th Joint Propulsion Conference, Los 

Angeles, CA, June 1999.

4. Janson, S., “Chemical and Electric 

Micropropulsion Concepts for Nanosatellites,” 

AIAA Paper 94-2998, 30th Joint Propulsion 

Conference, Indianapolis, IN, July 1994.

5. de Groot W. and Oleson, S., “Chemical 

Microthruster Options,” AIAA Paper 96-2868, 

32nd Joint Propulsion Conference, Lake Buena 

Vista, FL, July 1996.

6. Mueller, J., “Thruster Options for 

Microspacecraft: A Review and Evaluation of 

Existing Hardware and Emerging Technologies,” 

AIAA Paper 97-3058, 33nd Joint Propulsion 

Conference, Seattle, WA, July 1997.

7. de Groot, W., Reed, B., and Brenizer, M., 

“Preliminary Results of Solid Gas Generator 

Micropropulsion,” NASA TM-1999-208842, 

AIAA 98-3225, 34nd Joint Propulsion 

Conference, Cleveland, OH, July 1998.

8. Bayt, R., Ayon, A., and Breuer, K., “A 

Performance Evaluation of MEMS-based 

Micronozzles,” AIAA 97-3169, 33nd Joint 

Propulsion Conference, Seattle, WA, July 1997.

9. Alexeenko, A., Levin, D., Gimelshein, S., 

Collins, R., and Markelov, G., “Numerical 

Simulation of High-Temperature Gas Flows in a 

Millimeter-Scale Thruster,” AIAA 2001-1011, 

39th Aerospace Sciences Meeting, Reno, NV, 

January 2001.

10. Grisnik, S., Smith, T., and Saltz, L., 

“Experimental Study of Low Reynolds Number 

Nozzles,” AIAA Paper 87-0992, 19th

International Electric Propulsion Conference, 

Colorado Springs, CO, 1987.

11. Arrington, L., Reed, B., and Rivera, A., “A 

Performance Comparison of Two Small Rocket 

Nozzles,” AIAA 96-2582, 32nd Joint Propulsion 

Conference, Lake Buena Vista, FL, July 1996.

12. Reed, B., de Groot, W., and Dang, L., 

“Experimental Evaluation of Cold Flow 

Micronozzles,” AIAA 2001-3521, 37th Joint 

Propulsion Conference, Salt Lake City, UT, July 

2001.

13. Alexeenko, A., Levin, D., Gimelshein, S., 

Collins, R., and Reed, B., “Numerical Modeling 

of Axisymmetric and Three-Dimensional Flows 

in Microelectromechanical Systems Nozzles,” 

AIAA Journal, Vol. 40, No. 5, pp. 897-904.

14. Alexeenko, A., Levin, D., Gimelshein, S., 

Collins, R., and Reed, B., “Numerical Study of 

Flowfield Structure and Thrust Performance for 

3-D MEMS-Based Nozzles,” 32nd AIAA Fluid 

Dynamics Conferences, St. Louis, MO, June 

2002.

15. Haag, T., “Thrust Stand for High-Power Electric 

Propulsion Devices,” Rev. Sci. Instrum., Vol. 62, 

No. 5, pp. 1186-1191, May 1991.

16. Haag, T., “Thrust Stand for Pulsed Plasma 

Thrusters,” Rev. Sci. Instrum., Vol. 68, No. 5, 

pp. 2060-2067, May 1997.

NASA/TM—2003-212225 9

Appendix A: Cold Gas Nitrogen Micronozzle 

Performance Data

Nozzle 146, Area Ratio = 1.0

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.722 0.067

1.00 0.713 0.066

1.00 0.761 0.070

1.00 0.893 0.083

1.03 0.649 0.058

1.03 1.16 0.103

1.07 1.72 0.149

1.07 1.66 0.143

1.12 2.65 0.219

1.12 2.68 0.222

1.29 5.77 0.415

1.29 5.75 0.411

1.56 10.9 0.644

1.57 10.9 0.645

2.13 22.0 0.954

Nozzle 153, Area Ratio = 1.5

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.771 0.127

1.00 0.837 0.137

1.03 0.924 0.145

1.00 0.860 0.141

1.11 1.19 0.174

1.09 1.06 0.158

1.22 1.82 0.243

1.12 1.36 0.198

1.18 2.93 0.403

1.29 2.42 0.306

1.80 6.10 0.392

1.73 5.56 0.522

2.53 11.8 0.761

2.40 10.9 0.743

4.03 24.3 0.983

4.05 24.5 0.985

5.87 39.8 1.10

Nozzle 169, Area Ratio = 4.8

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.01 0.720 0.098

1.00 0.821 0.114

1.00 0.800 0.110

1.00 0.792 0.109

1.05 1.12 0.147

1.05 1.18 0.154

1.14 1.70 0.204

1.11 1.68 0.207

1.22 2.61 0.293

1.21 2.61 0.295

1.54 5.66 0.503

1.55 5.71 0.506

2.06 11.2 0.742

2.03 11.1 0.746

3.10 26.0 1.15

Nozzle 178, Area Ratio = 9.7

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.880 0.098

1.00 0.785 0.087

1.00 0.829 0.088

1.00 0.937 0.100

1.10 1.07 0.103

1.10 1.14 0.109

1.23 1.58 0.136

1.24 1.66 0.141

1.43 2.68 0.198

1.42 2.72 0.203

1.92 6.00 0.331

1.90 6.08 0.337

2.70 11.8 0.463

2.66 11.9 0.470

4.97 28.7 0.609

4.98 29.7 0.630

Nozzle 208, Area Ratio = 14.9

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.850 0.115

1.00 0.785 0.106

1.00 0.836 0.111

1.06 1.15 0.139

1.05 0.996 0.121

1.11 1.33 0.153

1.11 1.33 0.153

1.31 2.41 0.236

1.37 2.26 0.211

1.86 5.60 0.385

2.03 5.19 0.326

2.78 11.0 0.506

3.06 9.98 0.417

5.15 23.2 0.575

5.27 23.0 0.558

7.88 38.0 0.616

8.01 38.7 0.618

7.88 38.5 0.625

NASA/TM—2003-212225 11
Nozzle 226, Area Ratio = 18.2

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.910 0.102

1.00 0.879 0.099

1.00 0.806 0.090

1.00 0.811 0.091

1.04 1.09 0.113

1.05 1.06 0.109

1.09 1.21 0.119

1.08 1.19 0.119

1.27 2.35 0.198

1.27 2.32 0.196

1.79 5.80 0.242

1.79 5.72 0.344

2.58 11.7 0.488

2.56 11.7 0.492

4.19 24.8 0.637

4.18 24.3 0.626

6.43 42.4 0.709

8.43 58.4 0.745

8.42 59.4 0.758

10.2 74.2 0.780

10.2 74.1 0.780

12.4 97.3 0.842

Nozzle 230, Area Ratio = 19.5

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.843 0.101

1.00 0.770 0.093

1.00 0.838 0.101

1.00 0.881 0.106

1.00 1.23 0.139

1.01 1.14 0.128

1.06 1.58 0.169

1.06 1.51 0.161

1.20 2.62 0.248

1.20 2.61 0.247

1.59 6.06 0.433

1.58 6.20 0.445

2.18 11.9 0.623

2.18 12.0 0.626

3.82 29.2 0.866

3.79 28.9 0.864

5.50 46.3 0.955

Nozzle 243, Area Ratio = 23.1

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

0.96 0.854 0.113

0.96 0.822 0.109

0.95 0.872 0.116

0.97 0.861 0.112

1.03 1.12 0.138

1.03 1.12 0.138

1.10 1.51 0.173

1.14 1.71 0.191

1.32 2.79 0.269

1.32 2.71 0.260

1.84 6.07 0.417

1.82 6.07 0.423

2.66 11.8 0.562

2.66 12.0 0.568

4.88 28.3 0.733

4.97 29.1 0.740

9.33 62.3 0.845

9.31 62.0 0.844

12.8 88.4 0.877

Nozzle 237, Area Ratio = 24.2

Vacuum Thrust 

(mN)

Thrust Coefficient

0.97 0.911 0.117

0.95 0.789 0.103

0.98 0.900 0.114

0.99 0.907 0.114

1.04 1.26 0.151

1.04 1.08 0.129

1.16 1.81 0.194

1.16 1.85 0.198

1.34 2.96 0.275

1.34 2.94 0.274

1.86 6.37 0.425

1.86 6.35 0.425

2.72 12.2 0.559

2.71 12.2 0.561

5.87 33.6 0.710

5.88 34.5 0.729

8.32 50.7 0.757

NASA/TM—2003-212225 12

Chamber Pressure 

(atm)
Appendix B: Cold Gas Helium Micronozzle 

Performance Data

Nozzle 153, Area Ratio = 1.5

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.028 0.005

1.00 0.023 0.004

1.02 0.049 0.008

1.02 0.070 0.011

1.06 0.159 0.024

1.06 0.200 0.030

1.09 0.278 0.041

1.09 0.334 0.049

1.11 0.360 0.052

1.11 0.429 0.062

1.23 1.12 0.145

1.23 1.16 0.151

1.49 2.93 0.314

1.48 3.00 0.323

Nozzle 170, Area Ratio = 4.8

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.06 0.019 0.002

1.08 0.023 0.002

1.08 0.043 0.004

1.12 0.086 0.008

1.12 0.115 0.011

1.15 0.148 0.014

1.15 0.188 0.018

1.17 0.247 0.023

1.17 0.240 0.023

1.32 0.640 0.054

1.31 0.702 0.059

1.41 1.01 0.079

1.41 1.07 0.084

1.61 1.76 0.121

1.61 1.81 0.124

1.70 2.10 0.137

1.69 2.17 0.141

Nozzle 182, Area Ratio = 9.7

Chamber 

Pressure (atm)

Vacuum Thrust 

(mN)

Thrust 

Coefficient

1.07 0.015 0.001

1.07 0.027 0.002

1.10 0.032 0.002

1.09 0.052 0.003

1.15 0.102 0.006

1.14 0.131 0.008

1.19 0.172 0.010

1.18 0.208 0.013

1.21 0.216 0.013

1.20 0.261 0.015

1.35 0.645 0.034

1.34 0.697 0.037

1.62 1.75 0.077

1.65 1.83 0.079

1.70 2.12 0.089

1.71 2.18 0.091

Nozzle 195, Area Ratio = 14.5

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.06 0.007 0.001

1.07 0.026 0.003

1.08 0.019 0.002

1.08 0.049 0.005

1.12 0.081 0.008

1.12 0.122 0.013

1.15 0.146 0.015

1.15 0.194 0.019

1.17 0.186 0.018

1.21 0.337 0.032

1.34 0.633 0.055

1.34 0.674 0.058

1.45 0.982 0.078

1.46 1.00 0.079

1.69 1.55 0.106

1.69 1.63 0.111

1.84 1.89 0.118

1.84 1.93 0.122

Nozzle 226, Area Ratio = 18.2

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.034 0.004

1.00 0.000 0.000

1.00 0.069 0.007

1.00 0.040 0.004

1.04 0.190 0.020

1.03 0.162 0.017

1.06 0.314 0.032

1.06 0.296 0.030

1.08 0.395 0.039

1.08 0.389 0.039

1.21 1.13 0.100

1.21 1.15 0.102

1.51 3.01 0.215

1.51 3.03 0.217

1.60 3.63 0.244

1.60 3.65 0.246

NASA/TM—2003-212225 13
Nozzle 243, Area Ratio = 23.1

Chamber Pressure 

(atm)

Vacuum Thrust 

(mN)

Thrust Coefficient

1.00 0.041 0.005

1.00 0.033 0.004

1.00 0.082 0.010

1.02 0.074 0.009

1.04 0.205 0.025

1.04 0.200 0.024

1.08 0.421 0.049

1.08 0.408 0.048

1.21 1.11 0.117

1.21 1.14 0.119

1.51 3.04 0.255

1.50 3.02 0.254

NASA/TM—2003-212225 14
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19

Decomposing Solid Micropropulsion Nozzle Performance Issues

Brian Reed

Micropropulsion; Nozzle; Spacecraft

Unclassified -Unlimited

Subject Categories: 28 and 20 Distribution: Nonstandard

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

Astronautics, Reno, Nevada, January 6–9, 2003. Responsible person, Brian Reed, organization code 5430, 216–977–7489.

Micropropulsion technology is essential to the success of miniaturized spacecraft and can provide ultra-precise propulsion for small spacecraft. NASA Glenn Research Center has envisioned a micropropulsion concept that utilizes decomposing solid propellants for a valveless, leak-free propulsion system. Among the technical challenges of this decomposing solid micropropulsion concept is optimization of miniature, rectangular nozzles. A number of flat micronozzles were

tested with ambient-temperature nitrogen and helium gas in a vacuum facility. The thrusters were etched out of silicon

and had throat widths on the order of 350 microns and throat depths on the order of 250 microns. While these were halfsections of thrusters (two would be bonded together before firing), testing provided the performance trend for nozzles of

this scale and geometry. Area ratios from 1 to 25 were tested, with thrust measured using an inverted pendulum thrust

stand for nitrogen flows and a torsional thrust stand for helium. In the nitrogen testing, peak nozzle performance was

achieved around area ratio of 5. In the helium series, nozzle performance peaked for the smallest nozzle tested area ratio

1.5. For both gases, there was a secondary performance peak above area ratio 15. At low chamber pressures (< 1.6 atm),

nitrogen provided higher nozzle performance than helium. The performance curve for helium was steeper, however, and

it appeared that helium would provide better performance than nitrogen at higher chamber pressures.