NASA Technical Memorandum 107287

AIAA-96--2863

HAN-Based Monopropellant

Assessment for Spacecraft

Robert S. Jankovsky

Lewis Research Center

Cleveland, Ohio

Prepared for the

32nd Joint Propulsion Conference

cosponsored by AIAA, ASME, SAE, ASEE

Lake Buena Vista, Florida, July 1-3, 1996

National Aeronautics and

Space Administration

HAN-BASED MONOPROPELLANT

ASSESSMENT FOR SPACECRAFT

Robert S. Jankovsky

Aerospace Engineer

NASA Lewis Research Center

Cleveland, Ohio 44135

Abstract

The growing cost of space missions, the need for increased mission performance, and concerns associated with

environmental issues are changing rocket design and propellant selection criteria. Whereas a propellant's

performance was once defined solely in terms of specific impulse and density, now environmental safety, operability,

and cost are considered key drivers. Present emphasis on these considerations has heightened government and

commercial launch sector interest in Hydroxylammonium Nitrate (HAN)-based liquid propellants as options to

provide simple, safe, reliable, low cost, and high performance monopropellant systems.

Introduction

Monopropellant system development has been pursued

for decades. The simplicity of monopropellant feed and

control systems make them very attractive for missions

where the high performance (Isp) of bipropellants can be

traded for high reliability and low cost. Many

monopropellants have been investigated, but only a few

have found continued applications. Hydrazine (N2H4) is

by far the most widely used monopropellant and is

applied in many types of attitude control thrusters,

insertion stages, and gas generators. It has relatively

high performance (compared to cold gas systems), an

extensive flight heritage, and is commonly referred to as

the state-of-the-art. Toxicity & flammability hazards,

however, are N2H4's major drawbacks. These hazards

necessitate expensive ground handling procedures that

limit N2H4's utility for small cheap spacecraft. This

has led NASA to establish a program to develop an

advanced monopropellant rocket system for future

applications.

Specifically, the NASA Advanced Monopropellant

Program is emphasizing the requirements of several

mission scenarios. These include, for example, orbital

insertions and attitude control for small Earth-science

spacecraft and injection retro for planetary spacecraft. In

general, the advanced monopropellant thruster being

developed will be of high value in missions where

simple, cost-effective, relatively high thrust systems are

desired or where insufficient power is available for

electric propulsion options. The primary goal of the

program is to demonstrate a flight-type monopropellant

thruster that is environmentally benign and has a

specific impulse greater than that of NEH4, nominally

220 seconds.

Based on the search for monopropellants that could meet

these goals Hydroxylammonium Nitrate-based (I-IANbased) liquid monopropellants were identified. HANbased monopropellants have been pursued by the Army

as liquid gun propellants (LGP) for many years.

Through the Army liquid gun program, HAN-based

propellants have shown promise in the areas of

environmental/health and safety, performance, density,

and thermal management. The impact on satellite design

of the anticipated propellant enhancements were assessed

in a number of studies. One such study compared a

HAN-based monopropellant and a hydrazine

monopropellant propulsion subsystem for three NASA

missions, Total Ozone Mapping Spectrometer-Earth

Probe (TOMS-EP), Tropical Rainfall Measuring
Mission (TRMM), and Microwave Anisotropy Probe

(MAP). The comparison indicated a savings in fuel

mass, fuel volume, tankage volume, and tankage mass.

The average reduction of fuel mass was 17.5%, fuel

volume was reduced by 41.8%, the propellant tankage

volume was reduced by 38% and the tankage mass was

reduced by 35%. 1

Potential payoffs of an advanced monopropellant system

for both commercial and military applications has led to

the coordination of NASA and DOD programs. The

NASA/DOD programs are being coordinated through

the Integrated High Payoff Rocket Propulsion

Technology (IHPRPT) program. The program is

taking the approach of achieving a first generation

design that balances maximum performance with

highest probability of successfully providing a flighttype thruster by the year 2000.

Baseline Propellant Formulation

The Army has developed a number of HAN-based

monopropellants for use in artillery guns. Three of

these formulations are LP1846, LP1845 and LP1898.

Table I shows the formulations of LP 1846 and LP1898

with 20 percent water by weight along with LP1845, a

variation of LP1846 with reduced water content. They

are all HAN-based and differ only in the carbon

containing component. LP1846 and LP1845 use

triethanolaramonium nitrate (TEAN) and LP1898 uses

diethylhydroxylammonium nitrate (DEHAN).2 These

formulations are salts dissolved in water. HAN is

oxygen rich, and is commonly refered to as the oxidizer,

the other salt is fuel rich and is refered to as the fuel.

Variations on these formulations are being developed for

rocket monopropellant applications. They are being

derived from the aforementioned Army formulations as

aqueous mixtures of HAN and one or more nitrate salts.

Issues specific to rocket monopropellants, such as low

pressure ignition and combustion with a clean exhaust

plume, are being considered.

To date, though, LP1846, LP1845 and LP1898 are the

most closely studied HAN-based formulations and have

the largest body of data available. Therefore, the

discussion herein will use LP1846, LP1845 and

LP1898 as examples to illustrate the advantages of

HAN-based monopropellants.

Benefits Over State-of-Art

Based on the considerable amount of work done on the

Army's HAN-based LGPs, benefits of a HA_N-based

monopropellant propulsion system are anticipated in the

areas of safety, performance, density, and thermal

management as discussed below.

HAN-based propellants, mostly LP1846, have been the

subject of numerous studies concerning the health and

safety risks associated with them. To date all data

collected is favorable. Both the generant and the

exhaust are benign. No extrordinary clothing is required

for handling. Water repellent materials and elastomeric

gloves are sufficient. Utility clothing is acceptable if it

is removed and the skin is washed after a spill.

Animal studies have shown that toxicity is not a major

concern. LP1846 has proven negative as both a

carcinogen and mutagen. 4 No inhalation hazards are

associated with these propellants (unless aerosolizedS),

or their exhaust products (CO2, N2 and H20). In

addition, the propellant has proven not to be flammable

or sensitive at atmospheric pressure.

The health and safety characteristics of N2H4 are a stark

contrast to the ones just described for HA/q-based

monopropellants. N2H4 poses a threat both in terms of

toxicity and flammability. N21-14, both as a liquid and

vapor, is a confirmed carcinogen, mutagen in animals

and is flammable at atmospheric pressure. It requires

specialized clothing, facilities, and equipment. This all

translates into large infrastructure and high costs

associated with handling N21-14.4,6

Performance

A comprehensive investigation of potential

formulations and a careful balance of the critical issues

of performance, ignition, material limitations, and

contamination is underway. Variables being considered

for formulation of a HAN-based rocket monopropellant

are the fuel ingredient and amount of water. Performance

of HAN-based liquid propellants is highly dependent on

the amount of water in the mixture. The more water,

the lower the exhaust temperature and, to first order, the

lower the specific impulse. The fuel ingredient trades

are maximizing the heat of formation or fuel value

2
while minimizing the molecular weight of the exhaust.

The fuel ingredient trades are still continuing, therefore,

for purposes of benchmarking the performance of HANbased monopropellants, LP 1846, LP1845, and LP 1898

were used.

Figure 1 is a graph of theoretical specific impulse (Isp)

vs. area ratio for LP1846 (XM46) and LP1898 with

variations in water content. These predictions were

obtained using a one-dimensional equilibrium code. 7

Further refinement of the theoretical Isp predictions to

account for decomposition inefficiencies and heat losses

can be found in Table II. Two estimates of specific

impulse efficiency (rllsp) are made, a conservative rllsp

estimate of 85% along with an anticipated rllsp of

92%.g It can be seen that a delivered Isp of between

220 and 240 seconds is achieveable with the Army

LGP's, and even higher Isp, in the range of 270 seconds

is achievable, with a reduced water version of these

baseline formulations.

Comparitively, N2H4 monopropellant thruster systems

have delivered Isp of between 220 and 230 seconds.9

The density of a HAN-based rocket monopropellant can

be estimated by looking at the Army LGP formulations

in Table HI. The similarity between LP1846, LP1845,

LP1898 and the rocket monopropellant under

developement is in both the ingredients and their

quantities (-60 wt % HAN and -20 wt % H20).

Therefore, the storage density of the advanced rocket

monopropellant can be estimated as -1.4 g/co. This is

a 40% improvement over the 1.0 g/cc of N2H 4 .6

Thermal Management

Thermal management of a HAN-based monopropellant

system is driven by the viscosity. Figure 2 has a graph

of the viscosity of the Army monopropellants as a

function of temperature.2 From this, it can be seen that

at approximately 240 K the propellants transition into a

region of dramatic viscosity variation. This sets the

lower bound on the usable temperature range of the

propellant. Increased viscosity below this temperature

makes the propellant incompatible with typical

propellant feed systems. Therefore, the practical limits

of these types of monopropellants appears to be

approximately 240 K. The exact temperature limit of

the propellant and amount of thermal management

ultimately will depend on the exact formulation,

propulsion system design, mission, and satellite design.

In contrast, a N2I-I4 propulsion system is limited not by

viscosity variations, but rather the freezing point of the

propellant (see figure 2). N2H4 freezes at 273K and is

generally maintained at a minimum temperature of

280K. This 40K difference in propulsion system

temperature requirements translates into possible

reduction or elimination of thermal management power

requirements for HAN-based systems.

Critical Demonstrations

A large body of work has been done under the Army

funded gun program that is directly applicable to the

NASA rocket effort. Work in such areas as safety,

handling, materials compatibility, modelling, ignition

and combustion fundamentals, to mention only a few,

have all been leveraged. The fundamental differences in

the Army gun effort and the NASA rocket effort to be

addressed are operating pressure, duty cycle, and material

limits. Liquid monopropellant artillery guns under

development operate at approximately 200 MPa and fire

a few rounds a minute.lO High pressure operation

enhances the combustion kinetics and the low duty

cycle allows the gun to act as a heat sink to survive the

extreme combustion temperatures. Rocket

monopropellant systems under development must be

designed to operate in the range of 0.5-1.4 MPa, at a

variety of duty cycles in both pulse and steady-state

modes, and at the highest possible performance (i.e.

highest temperature). At low pressures the turbulence

levels will not be present to enhance ignition and

combustion. At multiple duty cycles and with high

performanace (>220 seconds), selection of thruster

materials (specifically catalyst materials) will present a

challenge.

Two NASA contracts have been let via the IHPRPT

program to address these issues and develop a flight-type

thruster operating on a HAN-based monopropellant.

The thruster/propellant development has achieved very

promising results to date. Specifically, over onehundred different ingredients for possible fuel

alternatives have been investigated. This list has been

trimmed to five ingredients through both laboratory and

thruster/reactor testing. Thruster testing used a

pyrotechnic ignition system and heavyweight hardware
to investigate the critical thresholds of the thruster and

propellant designs. In this configuration many HANbased rocket monopropellant formulations were tested at

chamber pressures that varied from 2.5-4.8 MPa. One

of the formulations tested in this configuration was

optimized for performance and demonstrated an Isp of

270 seconds (projected for area ratio of 50:1 at vacuum

conditions) at a TIC* (characteristic exhaust velocity

ef-fieiency) of 95%. The combustion temperature in this

case approached 2500K and had an exhaust that was

largely water.

For the baseline thruster system, ignition studies have

focused on catalytic ignitors. Catalytic ignition and

combustion at pressures of 1.4-2.8 MPa with a nonoptimized HAN-based monopropellant has been

demonstrated. The catalysts primarily being considered

use a Pt-group metal as the active metal and require a

minimal amount of preheat, much the same as in

hydrazine systems. However, as demonstrated in the

pyrotechnic/heavyweight hardware testing, operating at

the maximum potential performance of this type

propellant results in a very high temperature oxidizing

environment. This environment will preclude the use

of conventional catalyst materials. Therefore, additional

ignition concepts will be developed along with a

baseline thruster system that will operate at a reduced

temperature so that a conventional catalyst bed ignition

system can be used. The alternate ignition concepts to

be considered for the next generation advanced

monopropellant thruster system include laser ignition,

electrical ignition, chemical injectant ignition, and an

enhancext catalytic ignition system.

Conclusion

HAN-based monopropellant systems will provide a

step-change in monopropellant technology. Once

available, reduced costs, increased capability, and

decreased complexity will all result. The primary

technical challenges which need to be addressed before

these monopropellants can be applied to satellite

propulsion systems are reliable, repeatable ignition at

low pressure and high temperature oxidation resistant

materials. Low pressure ignition (1.4-2.8 MPa) has

been accomplished with both a pyrotechnic and catalytic

ignition system, but repeatability and durability of the

ignition system is still an issue at this time. Design

changes from the state-of-the-art N2FI4 system such as

combustion chamber material, ignition concept, and

thermal management system may be required, but with

very little affect on overall system mass. A dramatic

impact on mass and volume is anticipated when

compared to N2H4 propulsion systems because of

increased performance and density of this new

propellant. However, the overriding benefit of FIANbased propellants will be reduced toxicity and

flammability hazards and the corresponding reduction in

ground operation costs.

References

I. Davis, G.: Advanced Propulsion for MIDEX

Class Spacecraft, Goddard Spaceflight Center,

September 1, 1995.

. Liquid Propellant 1846 Handbook, Jet

Propulsion Laboratory, U.S. Department of

the Army, ARDEC, Picatinny Arsenal, NJ,

July, 1994.

. Decker, M.M.; Klein N.; Freedman, E.;

Leveritt, C.S.; Wojciechowski, J.Q.: HANBased Liquid Gun Propellants: Physical

Properties, BRL-TR-2864, 1987.

. Thiokol Corporation, Elkton, MD.: X-33

Non-Toxic Liquid Thruster Technology

Demonstration Program Review, 1995.

. Liquid Propellant 1846 Handbook. JPL D8978 DRAFT, U.S. Department of the Army,

ARDEC, Picatinny Arsenal, NJ, March 1992.

. Schmidt, E.W.: Hydrazine and Its Derivifives:

Preparation. Properties. Applications, John

Wiley & Sons, New York, 1984.

. Gordon, S.; and McBride, B.J.: Computer

Pro_tram for Calculation of Complex Chemical

Eauilibrium Compositions. Rocket

Performance. Incident and Reflected Shocks,

and Chapman-Jou_tmet Detonations, NASA

SP-273, 1976.

, Sutton, G. P.: Rocket Propulsion

Fifth Edition, John Wiley & Sons, New York,

1986.

. Hydrazine Handbook, Rocket Research

Company, Aerospace Dvision, Olin Defense

Systems Group

4
10. Klein,N.;Coffee,T.P.;Leveritt,C.S.:

Pressure Oscillations in a Liquid Pro_llant

Gun-Possible De_ndence On Pro_llant

Burning Rate, BRL-TR-3361, 1992.
Propel- [ Component, wt%

lant

HAN TEAN DEHAN

(NH3OH)+ (HOCH2CH2)3 (CH3CH2)

NO3- NH+NOy HNOH+NO3-

LP18463 60.8 19.2 0.0

water

H20

20.0

LPI845a 63.2 20.0 0.0 16.8

LP18982 i 60.7 0.0 19.3 20.0

I

Table I - Composition of Army developed monopropellants for liquid gun applications

[ LP1846 [LPIg45 [LP1898

Combustion Gas

Molecular Weight,

lb/ib tool @ I00 psi

Combustion Gas

Temperature, F

@ 100 psi

C*, Ws_ @

I{30 psi

Theoretical Isp,

@ 100 psi / 50:1

expansion,

vacuum

Estimated Isp

(Theo. x 0.85)

Estimated Isp

(Theo. x 0.92)

22.85

3180

4356

252.9

215.0

23.07

3366

4455

259.4

22.64

3384

4510

262.6

220.5 223.2

232.7 238.7 241.6

Table II - Theoretical Performance of Reference Monopropellants

reference ] 3 3 2

Density, 1.43 1.45 1.39

g/cc @ 25 C

Table II1 - Density of Reference Monopropellants

6
28O

_E 27O

o

¢o

,,' 2so

nO

24o

t.-

_-_0

0

l i i i ....

LPI _ _ X.Ol .\_.-'" I ......

i:::-j..<- ,

)7/I

t !

EXPANSION RATIO

Figure 1 - Theoretical Rocket Performance Predictions

10000

,-,1000

O,.

0

v

>- 100

W

0

w 10

>

,1

HAN-based monopropellants

(LGP1846 & LGP1898)

Freezing

point Hydrazine

i I I

200 300 400 500

Temperature (K)

Figure 2 - Viscosity of HAN-based Monopropellants and Hydrazine
FormApproved

REPORT DOCUMENTATION PAGE OMBNo.0704-0188

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1. AGENCY USE ONLY (LeaveblanlO

4. TITLE AND SUBTITLE

2. REPORT DATE

July 1996

HAN-Based Monopro_llant Assessment for Spacecraft

6. AUTHOR(S)

Robert S. Jardcovsky

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS{ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

3. REPORTTYPE AND DATES COVERED

Technical Memorandum

5. FUNDING NUMBERS

WU-233-1B--1B

8. PERFORMING ORGANIZATION

REPORT NUMBER

E- 10362

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM- 107287

AIAA-96-2863

11. SUPPLEMENTARYNOTES

Prepared for the 32rid Joint Propulsion Conference cosponsored by AIAA, ASME, SAE, and ASEE, Lake Buena Vista,

Florida, July 1-3, 1996. Responsible person, Robert S. Jankovsky, organization code 5330, (216) 977-7515.

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified - Unlimited

Subject Category 20

This publication is available fzom the NASA Center for Aea'oSpaeeln.fommaon, (301) 621--0390.

13. ABSTRACT (Maximum 200 words)

The growing cost of space missions, the need for increased mission performance, and concerns associated with environmental issues are changing rocket design and propellant selection criteria. Whereas a propellant's performance was once

defined solely in terms of specific impulse and density, now environmental safety, operability, and cost are considered key

drivers. Present emphasis on these considerations has heightened government and commercial launch sector interest in

Hydroxylammonium Nitrate (HAN)-based liquid propellants as options to provide simple, safe, reliable, low cost, and

high performance monopropellant systems.

14. SUBJECT TERMS

Monopropellant; Hydroxylammonium nitrate; Thruster

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