A I M ELECTRIC PROPULSION CONFERENCE BROADMOOR HOTEL, COLORADO SPRINGS, COLO. $9 MARCH 11-13, 1963

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NASA RESEARCH ON RESISTANCE-IIEATD EYDROGEN 3EII’S

BY John R. Jack md. E r n i e W. Spisz Lewis Research Center National. Aerooauti-s a d Space Admb-istratior clewatma, Ohio 63023

i r -

F..- publication rights reserved by AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS, 500 Fifth Ave.. New York 36, N. Y.

Abstracts may be published without permission if credit is given to the author and to AIAA.

NASA RESEARCH ON RESISTANCE-HEATED HYDROGEN JETS

By John R. Jack and Ernie W. Spisz

Lewis Research Center National Aeronautics and Space Administration

Cleveland, Ohio

INTRODUCTION

This paper, the third in a series (refs. 1 and a ) , is a status report r' W 0 N I w

on a continuing research effort at the Lewis Research Center on resistance- heated hydrogen jets. The design concept, which has been under investigation since early 1959, utilizes an electrically heated tungsten heat exchanger to achieve propellant temperatures of the order of 5500' R. used exclusively as the propellant because of its capability of achieving specific impulses of 1000 seconds or higher at these temperatures.

Hydrogen has been

The resistance heating approach for generating high-temperature gases is competitive with arc-jet devices for missions that require specific im- pulses of the order of 1000 seconds. In addition, this approach is receiving increased attention because of the following potentially attractive character- istics: (1) operating flexibility, (2) starting and restarting ease, (3) ability to utilize an a-c or d-c power supply, (4) high reliability, (5) long operating life, and (6) high efficiency. A major disadvantage to this heating approach is that it is inherently temperature limited by the heater element material to maximum vacuum specific impulses of the order of 1100 seconds.

J

The current status and the planned research on this propulsion device at the N M A Lewis Research Center is discussed herein.

THRUSTOR DESIGN AND PERFORMANCE CONSIDERATIONS

The performance characteristics desired for an electrothermal thrustor are (1) a vacuum specific impulse of 1000 seconds or higher, (2) high over- all efficiency, and (3) long operating life. stagnation temperatures required for a given frozen specific impulse for various pressure levels. To achieve a specific impulse of 1000 seconds, the gas stagnation temperature must be of the order of 5000° R and, consequently, the heat-exchanger temperature must exceed this value. Thus, for satisfactory operation at this temperature level, the heat-exchanger material must be tungsten or perhaps one of the refractory carbides. Of course, specific im- pulses greater than 1000 seconds (or temperatures greater than 5000' R) could be considered with these materials. It should be pointed out, however, that operation at temperatures approaching the melting point of the heat-exchanger

Figure 1 shows the hydrogen

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material will reduce the operational lifetime of the unit because of the incyeased sublimation rate.

Figure 1 indicates also that the influence of operating pressure on specific impulse is small at this temperature level. The operating pressure level, however, does have a large influence on frozen flow losses and a corresponding effect on overall thrustor efficiency. This effect is indicated in figure 2 for a stagnation temperature of 5040° R (ref. 3). Note that at a pressure of 1 atmosphere the vacuum specific impulse is about 950 seconds and the frozen flow efficiency is approximately 0.85. If no recombination occurs in the nozzle, the frozen flow efficiency represents the maximum overall thrustor efficiency attainable.

Another factor which affects overall thrustor efficiency is expansion efficiency, which is defined as the ratio of jet power to the power avail- able for thrust (unfrozen portion of gas power). Both the frozen flow and expansion efficiencies have been combined in figure 3 to obtain an overall nozzle efficiency for hydrogen at a stagnation pressure of 1 atmosphere,

v

With the aid of figure 3, some general conclusions regarding the maximum efficiency of an electrothermal thrustor may be made. For example, if it is assumed that losses other than frozen flow and expansion are small, the maximum engine efficiency that can be obtained in a tank (when a nozzle pressure ratio of 100 is assumed) is about 52 percent and the corresponding specific impulse is 875 seconds for a temperature of 5500° R. this operating point to space conditions, where the nozzle pressure ratio

and the specific impulse to approximately 1025 seconds.

Extrapolating

i s very large, increases the overall efficiency to approximately 71 percent L

The stagnation conditions chosen for the design of the experimental thrustors were a pressure of 1 atmosphere and a maximum temperature level of 5000' to 5500' R.

level of the order of 1 lb. figure 3 for typical thrustors operating in a tank and in space at the assumed conditions is presented in table I.

In addition the input power level and hydrogen flow rate were chosen as 30 kw and 10- 3 lb/sec, respectively, to provide a thrust

A summary of the performance estimated from

MPERIMXNTAL RESFARCH AT NASA

Thrus tors

The current research program on the resistance-heated hydrogen rocket at the Lewis Research Center utilizes two basic thrustor designs. One of these is a low-resistance (5x10-3 ohm) device employing a tungsten-tube heat exchanger whereas the other is a high resistance (3x10-l ohm) device utilizing a tungsten-wire-coil heat exchanger.

In order to measure heat losses conveniently and to eliminate design problems, such as sealing and electrical connections, the tube heat-exchanger

model was water cooled and was tested with two nozzles, one water cooled and one radiation cooled. The specific design used is shown in figure 4. The heat-exchanger element is composed of a 1-inch-diameter tungsten tube with a wall thickness of 0.010 inch and an effective length of 6g inches. design point of 30 kw, the electrical chara,c-teristics of the unit are 2300 amperes and 13 vol.ts.

1 At the

Because of the large number of cooling leads and the size of the elec- trical leads required to carry currents of the order of 2300 amperes, the unit was mounted on the outside of the facuum facility (fig. 5). This ap- proach makes it convenient to take detailed measurements but requires that the thrust be determined indirectly. The me:thod c.hosen was to measure the total pressure at the center of the nozz,le exit with. a tungsten probe. A single centerline measurement is justified since it has been found that the jet is fairly uniform [ref. 2) ~ This particular approach is not precise, but is a. convenient and expedient t,echnique to obtain fairly accurate and reliable thrust data, Complete de.t,ails of this method are presented in reference 2.

The .tube heat-exchanger model, although very u s e f u l for obtaining data, is incapable of yielding high gas heating efficiencies or high overall thrustor efficiencies. in an attempt. to achieve the desired higher efficiencies, the wire’-coil heat-exchanger uni.t was designed. The particular design now being used is shown in figure 6. This design is not water cooled and, consequently, should increase the gas heating efficiency considerably.

The coil heat exchanger consists of 63 mil tungsten wire approximately _/

SO inches long. At t:he 30 kw design point, the unit operates at about 300 amperes at 100 volts, getting power to the engine and, in addition, reduces ma.rkedly the ohmic heat losses of the electrical leads. Furthermore, since the water-cooling leads have been elimina.ted and the size of the electrical leads reduced, the thrust can be determined directly by mounting the thrustor on a thrust stand.

The high voltage simplifies greatly the problem of

Experimental Performance

Experimental results a,re presen-ted for both the tubular heat-exchanger model and the wire-coi.1 model. Most of the data. presented are for the tube model, which has been investigated more extensively. Both uni.ts were designed for operation at identiml conditions, that is, a, power level of SO kw and a hydrogen flow rate of lom3 lb/sec. The units are operated in a vacuum tank at an ambient pressure of 10 to 20 mm Hg. The nozzles used have a throa:t diameter of 0”214 inch, an area ratio of 5.33, and a pressure ratio of approximately 50,

Tubular heat-exchanger model. - Experimental results were obta.ined or1 --.- l____l_

the tube design for input powers up to 38 kw, and gas and heat-exchanger temperatures as high as 4500° and 5000° R, respectively. The basic heat

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exchanger has been operated with a water-cooled graphi-te nozzle and a

t o i n v e s t i g a t e t h e increased performance t h a t could he expected because of t he r e s u l t i n g higher heating e f f i c i ency .

radiation-cooled molybdenum nozzle. The radiation-cooled nozzle was used U

Figure 7 presents t h e s p e c i f i c impulse of t h e u n i t f o r t h e two d i f f e r e n t ~~

nozzles as a function of iiiput power. The spec i f i c impulse i s based upon the measured flow r a t e of lo-' lb/sec and the t h r u s t as evaluated from cen te r l ine to ta l -pressure probe measurement.s. s p e c i f i c impulse of t he u n i t as it operates i n t he tank i s 700 and 790 sec f o r t he water-cooled and radiation-cooled nozzles, respec t ive ly , which corresponding t o e f f i c i e n c i e s of 36 and 45 percent.

A t t h e 30 kw design point t he

A comparison of t he performance da ta of f i gu re 7 f o r t he water-cooled and radiation-cooled nozzle ind ica t e s t he expected b e t t e r performance f o r t he r a d i a t i o n cooled nozzle. t h a t t he d i f fe rence between the da t a f o r t'ne two nozzles may not be as l a r g e as indicated. For example, a t t he 30 kw point, t h e ca l cu la t ed increase i n performance due t o the higher heating e f f ic iency o f t he r a d i a t i o n cooled nozzle i s approximately 50 seconds. This i s about one-half of t he increase ind ica ted by the da ta . The add i t iona l d i f fe rence i s not y e t f u l l y unde.rstood but may be due t o the e r r o r s assoc ia ted with the i n d i r e c t determination of t h r u s t .

Above t,he 15 kw poin t it should be pointed out

Two of t h e most important parameters of a resistance-heated t h r u s t o r are i t s hea t - t ransfer e f fec t iveness and heating e f f ic iency . The hea t - t ransfer

'gas, out - 'gas, i n a Experimental L ef fec t iveness i s defined by the r a t i o

values of t h i s parameter a r e shown i n f igu re 8 . t he capab i l i t y of t he heat-exchanger t o hea t t h e gas t o temperatures ap:proach- ing those of t he heat-exchanger sur faces . The value of i s calcu- l a t e d from a power balance of the u n i t and the propel lan t flow rate, value of temperature, Ttube, i s found from the r e s i s t ance of t h e tube as given by the vol tage and cur ren t measurements, t he tube dimensions, and the r e s i s t i v i t y of tungsten.

'tube - Tgas,in ~~

This parameter i nd ica t e s

~~

The i s spec i f i ed as 530° R. The mean value of t he tube

As noted i n f i gu re 8 , ef fec t iveness values as high as 0.87 were obtained a t the 30-kw design point. This high value and the increasing t r end wi-th s p e c i f i c impulse a r e extremely encouraging. Various design modifications t h a t may produce a more e f f i c i e n t hea t - t ransfer process and achieve higher values of heat-transfer e f fec t iveness a r e present ly being considered.,

The heating e f f ic iency of t,he u n i t which i s defined as the r a t i o of This power i n t o gas t o e l e c t r i c a l power input i s presented i n f i gu re 9.

parameter i s an ind ica t ion of t he a b i l i t y of t he u n i t t o convert e1ectr:Lc energy i n t o thermal energy. The gas power i s obtained from the e l e c t r i c power

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input and the energy removed by the cooling water. cooling, the heating efficiency at an input power of 30 kw is only approxi- mately 50 percent. water-cooled and radiation-cooled nozzle data corresponds closely to the measured nozzle heat losses.

Because of the water -_

The difference of approximately 10 percent between the

Figure 10 compares the performance obtained from probe measurements with the performance calculated for the measured gas power and propellent flow rate. The curve calculated from gas power and propellant flow rate is based upon a one-dimensional, isentropic, frozen flow expansion process. The substantially good agreement between the measured data and the computed curve over the entire input power range indicates a fairly uniform flow. comparison gives a check on the reliability of the technique used to determine thrust.

In addition, this

The performance data presented on the tubular heat exchanger model indicate that the concept is promising. Sufficient data, understanding, and experience have been accumulated to undertake the design of an efficient thrustor, with regenerative cooling and a large expansion ratio, to approach the performance desired for propulsion.

Coil heat-exchanger model. - Freliminary results were obtained on the coil heat-exchanger model for input powers up to 16 kw and gas temperatures up to 4300' R. input power. power since the heat losses associated with the unit are quite small, i.e., the initial gas power is approximately equal to the input power. The specific impulse is based upon the propellant flow rate and the thrust as measured on a thrust stand. A comparison of measured and calculated per- formance indicates reasonably good agreement.

Figure 11 presents the specific impulse as a function of In this case, the calculated performance is based on the input

-

This particular thrustor has not as yet been operated at the design power level. A comparison can be made, however, at the 15-kw input power level where an overall thrustor efficiency of 73 percent is found for the wire coil unit and 52 percent for the tube heat-exchanger unit. efficiency of the wire-coil device was expected and illustrates the potentially high efficiencies attainable with resistance-heated hydrogen thrustors.

The higher

As a point of interest it is worth mentioning an operating problem that was encountered with this unit. At high gas temperatures and corresponding high voltages, arcing occurred between the wire coil and the nozzle body at the throat section. existed across a gap of more than l/4 inch at pressures of approximately 1 atmosphere. nozzle throat. indicates one of the potential troublesome areas with a high-voltage unit. Design modifications have been incorporated to eliminate this problem.

This arcing was unexpected because less than 100 volts

The arcing resulted in coil destruction and ablation at the The fact that arcing can occur so readily at these conditions

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CONCLUDING REMARKS v

Experimental results have been presented for two basically different design approaches to the resistance-heated hydrogen thrustor. One thrustor employs a tungsten-tube heat exchanger and is water cooled, whereas the other uses a tungsten-wire coil heat exchanger and is radiation cooled. Data were obtained over an input power range of 0 to 38 kilowatts for propellant flow rates of lb/sec. A summary of the experimental results follows:

r- (D 0 N I

W

1. For the water-cooled tube heat-exchanger unit at the design power level (30 kw), a specific impulse of 700 seconds was achieved with an overall thrustor efficiency of 36 percent. Sufficient data and experience have now been obtained on the research unit to design a radiation-cooled thrustor for space application with a potential performance of 1000 seconds specific impulse with an efficiency of 75 percent.

2. For the wire-coil unit with an input power of 15 kw, a specific impulse of 710 seconds with an efficiency of 73 percent was attained. This compares to the values of 600 seconds specific impulse and 52 percent efficiency for the tube heat-exchanger model. The higher efficiency of this device demonstrates the expected improved performance of this design approach and indicates the potentially high efficiencies attainable.

ACKNOWLEDGFNENT

L The authors would like to express their appreciation to their coll.eague Mr. Paul F. Brinich for permitting use of the preliminary data for the wire coil unit, which he designed and tested.

REFERENCES

1. Jack, J. R,: Theoretical Performance of Propellants Suitable for Electrothermal Jet Engines. 15th Annual Meeting ARS, Washington, Dec, 1960, ARS preprint 1506-60.

2. Jack, J, R , : NASA Research on Resistance Heated Hydrogen Jets, AFOSR Symposium on Advanced Propulsion, Cincinnati, Ohio, Oct. 2-4, 1962.

3. King, C, R.: Compilation of Thermodynamic Properties, Transport Properties, and Theoretical Rocket Performance of Gaseous Hydrogen. April 1960.

NASA T” 11-275,

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TAEILE I. - ESTIMATED THRUSTOR PERFORMANCE

h

Thrustor parameter

Input power, kw

Mass flow, lb/sec

Nozzle pressure ratio

Heating efficiency

Stagnation temperature, OR

Stagnation pressure, atm

Frozen flow efficiency

Nozzle expansion efficiench

Overall efficiency

Specific impulse, sec

Thrust, lb

Tank operation - ?ater :ooled

30

10-3

53

0.43

3750

0.79

0.99

0.65

0.28

650

1.650

tegeneratively cooled

30

10-3

53

1.0

5500

1.0

0.71

0 . 7 1

0.50

850

0.050

Space operation,

regeneratively cooled

30

10-3

lo5

1.0

5400

1.0

0.75

1 000

0.75

1000

1.0

6N03

0

12

8

4

I

1 E-2067 I \

PRESSURE, a

- LL

F

>- V

w z

LL u LL w

0 _I LL

z W N 0 E LL

3

0 I I I I I , I I I I 1 1 1 / 1 I I I L l , I I I I I I ( .01 .I I I O

L 9 O Z - 3

-JlOOO

W'

a 5 Q V

cn J 9

-500 e - H

m y

0 J L

Z W N 0 CY L

3

I O 0

1.2

I .o

.8 c

0 Z

>-

w - .6 u LL LL w

A N N 0 Z

w .4

.2

0

2000

NOZZLE 1QOO PRESSURE IO00 TEMPERATURE, RATIO, Po 4

IO

200 400 600 800 1000 1200 1400

SPECIFIC IMPULSE, I, SEC Figure 3. - Ferfomance characteristics of hydrogen at a chamber rjressure of 1 atn.

WATER COOLING PROPELLANT ,-WATER COOLING FOR . . . ~ -~ ~~ ~~ ~

INLET 7 / NOZZLE AND NEGATIVE ’ ELECTRICAL LEAD FOR RADIATION

INSULATION

WATER COOLING FOR POSITIVE ELECTRICAL LEA

ELECTRICAL L

Y-RADIATION TUNGSTEN TUBE , HEAT EXCHANGER-

PROPELLANT INLET

Figure 4. . Water-cooled t.Wo~1e.r heat-exchanger rrodel.

L902-3

GRANULAR ZIRCONIA GAS IN I r

I I 0.060 TUNGSTEN

/ / HEATER WIRE

HEATER CHAMBER

Figure 6 . - Radiation-cooled wire-coil heat-exchanger model.

L9oz-x (

I E-2067

800

600

400

200

I O O O r -

-

-

0 WATER COOLED NOZZLE 0 RADIATION COOLED NOZZLE

-

I I I I 1 I

Figure 7 . - Ekperimental performance of tubular heat-exchanger u n i t . Nozzle a rea r a t i o , 5.33; tank pressure, 10 mm Hg; propel lan t , hydrogen; flow r a t e , lb/sec.

1.0 cn cn W Z W > I- -

L L ~ :

I .6

0 WATER COOLED NOZZLE 0 RADIATION COOLED NOZZLE

I I- U w

Figure 8 . - Reat- t ransfer effect iveness of tungsten tube heat-exchanger u n i t . Propel lant , hydrogen; f l o w r a t e , Ih/sec.

L902-8

!-- Q W I

.a

.6

.4

3

OLO

1 E-2067 \

- U -r

0 0 0 -

0 WATER COOLED NOZZLE RADIATION COOLED NOZZLE

5 IO 15 2 0 25

INPUT POWER, P,, KW

Figure 9. - Z q e r i m e n t a l ?.Eating efficiency of tubular i:eat flax rate; io-' ;b/sec.

30

exchanger

35 40

u n i t . Hydrogen

w

0 WATER COOLED NOZZLE

0 RADIATION COOLED N O Z Z L E

GAS POWER, PGAS I KW Figure 10. - Experimental perfomance o f tubular heat-exchanger u n i t .

r a t i o , 5.33; t ank pressure, 10 mm Hg; propel lan t , hydrogen; flow r a t e , Nozzle a rea

lb / sec .

L 9 O Z - 3

IO00

800

600

400

200 0

! E-2067 i

-

PERFORMANCE P ‘-CALCULATED

15 20 25 30 35 5 IO

INPUT POWER, P,, KW Figure 11. - Experimental performance of wire-coil heat-exchanger unit. ratio, 5.33; tank pressure, 10 mm Hg; propellant, hydrogen; flow rate,

Rozzle area lb/sec.