Overview of Propulsion Systems for a Mars Aircraft — 20010091338 (2001)

APPENDIX A: MARS FLYER ROCKET PROPULSIO RISK ASSESSMENT

APPENDIX A: MARS FLYER ROCKET PROPULSIO RISK ASSESSMENT Atlantic Re seaTc h Corporation This report is being reprinted in its entirety as originally plinted in Aplil 2001. The read er should note that the Oliginal page numbers have been retained.

NASAffM-2001-210575 33 .--- .. -~.~ - - -------,

NASAl CR-2001-210709

Mars Flyer Rocket Propulsion Risk Asses s me nt

ARC Testing

Atlantic Research Corporation Niagara Falls, New York Prepared under Contract NAS3- 99197 National Aeronautics and Space Administration Glenn Research Center

April 2001

Trade names or manufacturers' names are used in this report for identification only. This usage does not constitute an official endorsement, either expressed or implied , by the National Aeronautics and Space Administration .

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SECTION PAGE

TABLE OF CONTENTS SECTION PAGE 1.0 INTRODUCTION .......... ...... ....... .. ............ .... ... ................ ................ .

2.0 THR USTE R DESCRIPTION ............... .. ... ... .. .................... .. .. .. .. ... . .. . 4 3.0 PRODUCTION OF MON-25 .. .. ..... ... . .. ......... ........ ...... ..... .. ........ . .. ... . 16 4.0 PROPELLANT CONDITIONING SYSTEM .......................... ........ . 19 5.0 TEST PLAN ........................... ..... ... ....... ....... . .. .... ...... .... .. ..... ............ .. 26 6.0 TEST RESULTS ............................................................................... . 31 7.0 CONCLUSIONS ...... ....... . .. . .. ....... ...... .... . .. .. .... .. .... .. .. .... ... ................. . 45 8.0 RECOMMENDATIONS ........ ..... .... ........ ......... ... ............... ............. ..

LIST OF TABLES Qualification Test Summary ................ ....... .... .......... .... ....... ....... .... .. . 6 2 MON-25 / MMH Baseline Test Matrix ........................................... .. 27 MON-251MMH Pressure and Temperature Mapping Tests ............ .. 28 4 MON-251MMH Test Matrix at -40C .... .... ... .. .................................. . 29 Test Data Summary ........ .. ................. ........ .......................... ... ........ .. .. 32 NASA/CR-2001-210709 III T ABLE OF CONTENTS (Continued) LIST OF FIGURES FIGURE NO. PAGE 1 ARC ION (2 Ibf) Thruster) ........... ........ ..... .. . ... ......... ..... ............... .. . .. 2 2 Thruster Feed Pressure Qualification Box ...... . .. ...... . .. ....... ... .. ... ....... . 7 Pulse Mode Operation Box ...... .. ...... ..... .. .... ........ ...... .... . .. .. . .. . .. .. .... . .. .

3 8 Variation of Specific Impulse with Mixture Ratio .. ........... ..... .... .. ... .

4 9 5 MON-3 , MON-25, and MMH Density Variation with Temperature 6 Fuel/ Oxidizer Density Ratios ... ..... .. ... .. . ... .. .. .... .. ..... ..... ..... .. ..... ..... .. .. . 12 7 Variation ofMMH Viscosity with Temperature .... ... .... ...... .......... .. .. 14 8 Variation ofMON-3 Viscosity with Temperature ........... .. .. .. .. ....... .. . 15 MON-25 Mixing Operation Schematic .... ........ ... ........ .. .... ...... .. ..... .. . 17 10 Propellant Cooling System In Test Cell .... .... .. ... .. ........ .. ...... .. .......... . 20 11 Thruster Mounted in Test Cell .. .. .. .. ...... ...................... .. .................... . 21 Propellant Temperature Time History, Test 35625 .. .. ........ .. ............ .

12 23 13 Heat Exchanger Inlet and Outlet Temperatures, Test 35625 ........ .. .. . 24 14 Thermocouple Locations ............ . .. :. .. ..... .. .. .. .... ............ . .. ... ............ . ..

15 Temperature Time History, Test 35625 .. ..... .. .... ........... ...... .. .. ......... .. 34 16 Thrust, Isp and MR Time History, Test 35625 ....... .. .......... .. ........... .. 35 17 Mixture Ratio Time History, Test 35625 . .. .. .. .... ... .... .. .. .... ........ .. ..... .. 36 Propellant Flowrate Time History, Test 35625 .............. .. .. .. ............ ..

18 37 19 Thruster Mixture Ratio Variation With Propellant Temperature .... .. 38 Specific Impulse Variation with Propellant Temperature .. .. .. .... ...... . 40 21 Effect of Propellant Feed Pressure and Temperature on Mixture Ratio .... .... ... ...... ..... ..... . .. .. ....... .. ... ..... .. .... ...... .. ................ ....... ...... . .. .

22 Effect of Propellant Feed Pressure and Temperature On Specific Impulse .. ... ........ .. ... ... .... ....... .. .. ..... ... .. .. .... ... ....... ..... .... ... .. ...... ..... . .. .. ... 42 23 Effect of Propellant Feed Pressure and Temperature on Injector Temperature .. . .. .... .. ........ .. . .. .... .. .. .... .... .... ...... .... . .. ..... . .. .. . .. ... ........ .. ... .

N AS N CR- 2 001 -2 10709 l V ____ . --------l 1.0 INTRODUCTION The purpose of this report is to describe the results from tests conducted at the Atlantic Research CorporationlLiquid Prolusion Division (ARCILP) located in Niagara Falls, NY under the NASA Glen Mars Flyer Rocket Propulsion Risk Assessment Program. The teclmical objectives of this program were to provide test data on the operational characteristics of a 2 lbf thruster operating with MMH and MON-25 propellants cooled to -40C to simulate conditions expected to be encountered during the Mars Flyer mission.

The thruster used in this program was an ARC ION (2.25 Ibf) thruster which had been used in previous ARC development activities. The thruster was made available to this program at no cost and is described in Section 2.0. Since MON-25 is not available off-the-shelf, ARC manufactured a supply of MON-25 for this program and verified the NO content by assay using an independent company, Vicksburg Chemical. The procedure used to manufacture the MON- 25 and the assay results are discussed in Section 3.0. For this program, ARC developed a propellant conditioning system which was capable of delivering -40C propellant to the thruster for tests of any duration. A description of this system is given in Section 4.0. The test plan is discussed in Section 5.0 and the results from the tests are discussed in Section 6.0. Conclusions are given in Section 7.0 and Recommendations in Section 8.0 This program can be summarized by indicating the manufacture of MON-25 was successful with two assays indicating the NO content of the propellant was 25.0 and 25 .3%.

After a series of development and calibration tests, the propellant conditioning system demonstrated the capability of delivering propellant to the thruster at the target temperatures of 21 C (70F), -IC (30F), -18C (OF), -29C (-20F) and -40C (-40F) for tests of any run length, including three successive tests of 1200s duration of -40C. All but three of the tests in the original plan were conducted with testing being terminated with the depletion of the MON-25 .

Most tests were successfully completed including two where the test cell pressure was increased to 10 torr (0.2 psia) to simulate the Martian atmospheric pressure. The thruster ran well at NASN CR- 2001-2 10 709 Figure 1. ARC 2 Ibf Thruster Design Performance Propellants: MON-3/MMH Thrust: 9.4N

• •

Materials Feed Pressure: 13.8 bar

• •

Disilicide Coated C-l 03 Chamber

•

Mixture Ra ti o: 1.65 Titanium Injector & Cone

•

Area Ratio: 336/1 Specific Impulse: 274s

•

Mass: 0.62 kg Chamber Temperature: 900C

•

Status Demonstrated Life

•

Flight Qualified 151 ,000s

•

210 Delivered 587 kg throughput

•

50 On-Orbit 1278 Cold Starts

•

1,120,000 Pulses

•

NASNC R- 200 1-2 10709 2

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- most conditions with the most notable result being that the cold propellant caused the nominal mixture ratio to shift from 1 .6 5 to 1.90 as the propellant temperatures were reduced from 21 C to -40C with a consequent small decrease in specific impulse. Details of the test results are gi ven in Section 6.0 and Volume 2. These results show that there should be no significant problems operating this thruster with MMHlMON-25 propellants in the cold -40C environment.

NASAfCR-200 1-2 10709 2.0 THRUSTER DESCRIPTION The ION thruster, which was designed for MON-3/MMH propellants, is shown in Figure l . The thruster is radiation cooled and uses fuel-barrier cooling to maintain chamber wall temperatures at levels sufficiently low to ensure the thruster meets propellant throughput requirements. The thruster was designed for stationkeeping on geostationary satellites and has been fully qualified for this mission using solenoid valves.

The thruster was designed for operation with propellants at a nominal temperature of 21 C (70F). Orifices and flow passages were designed to provide a nominal mixture ratio of 1. 65 with MMH and MON-31'ropellants at 21 C. Since propellant densities and viscosities are a function of propellant temperature, it was expected that operation with -40C propellants would result in changes in the thruster operational characteristics. These changes will be discussed shortly. It is important to note that the purpose of the tests described herein was to obtain an understanding of thruster operation with -40C propellants such that one could design a thruster optimized for these conditions at a later time.

The thruster is comprised of three major sub-assemblies: the injector assembly, thrust chamber and expansion cone and the propellant valves. The injector assembly consisting of the injector, distribution ring, thermal stand-off and oxidizer inlet tube are machined from 6Al14V Titanium and electron beam welded together. The injector h~s three unlike doublets in the core region and six fuel film coolant holes equispaced on the periphery. Gold plated Inconel 'C' seals provide the sealing between the valves and the injector assembly ..

The thrust chamber is manufactured from Columbium C103 alloy and is coated with the Hitemco R512E disilicide coating. The 336:1 area ratio expansion cone is machined from 6Al/4V Titanium and is electron beam welded to the thrust chamber.

N ASNCR-200 1- 210709 4 ;

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This thruster has no Pc top so chamber pressures could not be measured. Chamber pressure was estimated by using measured thrust and an assumed value of the thrust coefficient Cfofl.77 .

The thruster uses two Moog Model 51-178 series redundant normally closed solenoid valves. The valve is a fail safe design that remains closed through spring pre-load until opened by electrical energisation. The valve features a teflon seat and no sliding fits. An outlet orifice is fitted into the valve flange for pressure drop control and there is a provision for an orifice in the inlet tube for final thruster trimming. The valve also incorporates a 25 micron absolute filter at the inlet.

2.1 Thruster Heritage The thruster has been subjected to a very extensive qualification program using two thrusters. One thruster was used primarily for steady-state testing, while the second was used primarily for pulse mode testing. The scope of the qualification program included: Two thrusters Hot restarts

• •

Variable pulse width pulse trains 1400 burns

• •

Propellant temperatures: -7 to 54C Thermal stability tests

• •

Feed pressures: 8.3 - 19.8 bar Gas Ingestion Tests

•

Thermal Soakback Tests Pulse trains of over 2700 pulses

•

Some results on the qualification program scope are included in Table 1. The thruster has successfully performed long burns of 1000s and it was expected that the thruster would be able to perform the 20 minute (1200s) burns required during this program without any problems.

NASNCR-200 I -2 10709 Table 1. Qualification Test Summary Parameter Results Qual Thruster 2 Qual Thruster 1 Total nwnber of pulses 1,123,977 216,863 Thermal cycles (cold restarts) 1,278 426 Propellant throughput Steady State

•

Pulse Mode

•

Propellant temperature range (C) -10 to 59 -9 to 57 Injector pre-fire temp range (C) -20 to 96 -11 to 97 Mixture Ratio 1.08 to 1.96 1.31 to 1.97 Thrust (N) 5.8 to 12.5 5.8 to 12.0 Longest SSF duration (s) 1,000 Longest PMF duration (s) 3,600 812 Total on time (s) L51,362 99,354 Total Impulse (N-s) 1,428,192 1,030,620 Figure 2 shows the fuel and oxidizer feed pressure envelope for which the thruster is qualified and the points at which tests were conducted. The qualification range was 8.3 - 19 .7 bar (120-286 psia). The nominal operating conditions for the thruster are feed pressures for the fuel and oxidizer of 13.8 bar (200 psia) and a mixture ratio of l.65. ARC plans to put trim orifices into the thruster which will result in a nominal feed pressure of 15.2 bar (220 psia).

NASNCR-2001-210709 ---~-- . -.- - -- .- ~ ._ . .

. .. _ -- -----, Figure 2. Thruster Feed Pressure Qualification Box Fuel Feed Pressure vs. Oxidant Feed Pressure 21 degrees Centigrade ~

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NASNCR-200 1-2 10709 Figure 3 shows a map of the pulse mode tests which were conducted on the thruster in terms of the on-time and off-time for the pulses and the propellant temperatures at which these tests were conducted. The pulse mapping on this thruster was quite extensive and included long pulse trains (up to 4000 pulses) to demonstrate thermal stability. The thruster demonstrated successful operation at all conditions and no thermal limitations were discovered during the qualification tests.

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t=- ~ • To. 55 C II !J ., " -i 10.00 Figure 1 provides a brief summary of the operating characteristics of the thruster at the nominal operating condition. The specific impulse is 274s and the maximum chamber temperature is only 900C (1650F). This thruster was designed with a high fraction of the fuel in the barrier in order to produce a thermally robust thruster which could satisfy all high propellant throughput requirements for very demanding satellite missions. The thruster was designed to operate .at a mixture ratio of 1.65 with MON-3/MMH propellants.

The thruster was designed to provide moderate performance in terms of Isp, but, more importantly, to operate without any thermal limitations over the entire range of duty cycles which could be encountered during satellite operation. The nominal mixture ratio for the thruster NASNCR-2001-210709 --- ----- ------ is 1.65 with MON-3/MMH and the design has a high fraction of the fuel in the barrier. At nominal conditions, the Rupe Number for the doublet element core is about 1.75, rather than the more optimum value of 1.00; this is by design to obtain a wide operational envelope rather than a narrow envelope with higher performance. Rupe Number decreases to ward a more optimum value of 1.0 as mixture ratio decreases below 1.65 and Isp increases slowly as MR decreases as shown by the data in Figure 4.

Figure 4. Variation of Specific Impulse with Mixture Ratio ~ 270 .

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235 +-------;-- f-- ----+----i----+-------i -- - -- i 1.3 1.4 1.S 1.6 1.7 1.8 1.9 2.0 2.2 MARS FLYER TEST CONDITIONS As mentioned earlier, the thruster was designed for operation with 21 C propellants and some changes in operational characteristics were expected at the Mars Flyer conditions due to the change in propellant properties. The propellant properties which have the largest influence on thruster operation are density and viscosity.

Figure 5 shows how the density of MON-3, MON-25 and MMH vary with temperature over the -46C (-50F) to 49C (120F) range. All three propellants show the same trend of density increasing slowly as propellant temperature decreases. Note that MON-25 is somewhat less dense than MON -3 due to the higher NO content.

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In assessing performance ofbipropellant thrusters, one of the most common design parameters is the Rupe Number given by R = (p u d) ox (1) (p u d) fuel In Eq. (1 ), P is the propellant density, u is the orifice injection velocity and d is the orifice diameter; these parameters are for the injector core. Optimum bipropellant combustion is obtained for R = 1.0. Eq. (1) can be rewritten as (2) where Pr andpo are the fuel and oxidizer densities, .dr and do are the diameters of the fuel and oxidizer orifices in the injector core, MR is the mixture ratio, and B is the fraction of the barrier fuel injected into the thrust chamber wall. The parameter B is fixed by the ratio of the fuel core and barrier orifice diameters. Thus, for a given thruster design, R is affected by test conditions only through the density ratio and mixture ratio.

Figure 6 shows the oxidizer/fuel density ratio for MON-31MMH and MON-251MMH as a function of temperature. For MON-3/MMH at 2IC, the density ratio is approximately 1.64 while for MON-25IMMH at -40C, the value is about 1.625. The difference between these two values in terms of their effect on Rupe Number is negligible. In fact, the variation of the density ratio for both propellant combinations over the range of temperatures shown in Figure 6 is sufficiently small that no significant effect on Rupe Number or thruster performance would be expected. The conclusion then, is that the thruster performance will not be affected to any significant degree by the propellant density changes associated with operating at -40c.

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Figures 7 and 8 show the variation of viscosity o~MMH and MON-3 with temperature.

Note that the MON-3 data extends downward only to 2C (35F). No data is available for the viscosity ofMON-25 so we will use the MON-3 data as an estimate. The data in Figure 7 show that in going from 2lC to -40C , the viscosity ofMMH increases by about a factor of7 . Using a straight line extrapolation of the data in Figure 8, in going from 21 C to -40C , the viscosity of MON-25 would be estimated to increase by about a factor of2.

For a thruster operating at a given feed pressure, an increase in propellant viscosity will result in an increase in the frictional pressure drops losses through the thruster flow passages and consequently, a reduction in propellant flowrate. Since the increase in the fuel viscosity as propellant temperatures are reduced to -40C is much larger than the increase in the oxidizer viscosity, the fuel flowrate will decrease more than the oxidizer flowrate. This will cause the nominal thruster mixture ratio to increase as propellant temperature decreases and as will be shown in Section 6, this is the trend that is observed in the test data.

The results in Figure 4 show that thruster Isp decreases slowly as mixture ratio increases and therefore, we would expect Isp to decrease as propellant temperature decreases because of the mixture ratio shift. This trend is also observed in the test data.

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NASNCR-2001-210709 3.0 MON-25 PRODUCTION PROCESS MON-25 is not available either commercially or from Kelly AFB stockpiles. Therefore, the MON-25 used in this program was manufactured by ARCILP at our Niagara Falls facility.

ARC /LP had experience in the manufacturing ofMON-3 and MON-l 0 from previous and ongoing programs. The technique is straightforward as long as proper handling and mixing procedures are followed.

MON-25 consists of a mixture of 75% nitrogen tetroxide (N204) and 25% nitric oxide (NO) by weight. The starting constituents for the mixture are MON-3, which is readily available at our facility from~ther programs, and commercially available bottles of NO . The mixing process involves an exothermic reaction of the NO gas with N204 to create nitrogen trioxide (N203), which gives MON-25 its characteristic bluish-green color. ARC employed a cooling chamber during the mixing operation to both limit the oxidizer temperature increase from the reaction and to lower the vapor pressure of the resulting mixture to better utilize the available NO from its storage bottle. The initial temperature of the mixing bottle and MON-3 was approximately -1 C (30°F).

Figure 9 shows a schematic ofllie mixing operation. All hardware associated with the production ofMON-25 was passivated to ensure the minimization of MON-25 leaching excess iron into solution. This included all mixing vessels, transfer/~eed lines and samples bottles. To minimize the transfer operations required for this test series, a common mixing/storage vessels was used. The mixing/storage vessel consisted of a standard cylindrical 9-gallon tank with a stand tube inside along its center axis. The gaseous NO is introduced into the tank through this perforated stand tube at the bottom of the tank.

NASNCR-2001-210709 , - - - - - - ••• - _. - - - - - 0 - I Figure 9. MON-25 Mixing Operation Schematic The flow of NO into the tank was controlled by a NO gas regulator which limited the pressure of the NO to about 1.6 bar (23 psia), which is approximately the vapor pressure of the final mixture ofMON·25 at -I C (30F). Both pressure and temperature were monitored during the mixing operation. Periodically, the NO flow was turned off to allow the reaction to reach completion and for the mixing/storage vessel to be recooled. The pressure and temperature of the NO bottles were monitored during this operation. The NO bottle was allowed to return to ambient temperature to calculate the mass of NO which has been introduced into the mixing/storage vessel through the use of the perfect gas law.

After it has been determined that sufficient NO has been added to the oxidizer, the oxidizer was sampled using an evacuated standard Hoke bottle. The sample was shipped to Vicksburg Chemical for chemical analysis per MIL-P·26539D to verify the NO content of the oxidizer. The two samples analyzed showed an NO content of25 .0% and 25.3%.

Approximately 22.7 kg (50 Ibm) ofMON-25 was manufactured.

NASAlCR-2001 ·2 10709 After completion of the mixing operation, gaseous helium was introduced into the mixing/storage vessel. This gas pressure blanket insures that the NO does not come out of solution preferentially at ambient temperatures. The pressure was maintained at about 8.6 bar (125 psia), which is higher than the vapor pressure ofMON-25 at 32C (90F) (a maximum summer time temperature for the Niagara Falls facility).

NASNCR-2001-210709

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4.0 PROPELLANT AND HARDW ARE TEMPERATURE CONDITIONING SYSTEMS To support the Mars Flyer program, ARC developed a propellant conditioning system with the capability to cool the propellants to -40C for tests of any run duration. A schematic of the propellant cooling system is shown in Figure 10 .

In this system, liquid carbon dioxide (C0 ) is used to cool a bath filled with a mixture of approximately 60/40 ethylene glycol and· water. The freezing point for this mixture is about - SOC (-S8F). This mixture was chosen as a tradeoff between the required temperature capability of the fluid and the ability to pump it to the propellant heat exchangers. For example, by reducing the water in the fluid, the freezing point could be reduced further, however, since the viscosity of the fluid increases rapidly as the ethylene glycol content is increased and temperature is reduced, it becomes difficult to pump an adequate supply of the fluid to the heat exchangers. After some experimentation, the 60/40 mixture was found to be satisfactory for these tests. The CO circulates through the bath inside a set of copper coil tubing. It enters the coils as a liquid and exits as a gas; most of the cooling capability is, therefore, in the heat of vaporization of the CO A mechanical agitator was also used in the bath to in crease the heat 2 • transfer rate between the bath fluid and the liquid CO 2 .

The glycol bath is located outside of the test cell. The cold bath fluid is pumped through an insulated line to the propellant heat exchanges inside the test cell. The cooling fluid is split inside the cell and di re cted to separate fuel and oxidizer heat exchangers. After leaving the propellant heat exchangers, the fluid is returned to the bath via a single line.

As shown in Figure 10, the propellant heat exchangers are located in the test cell downstream of the propellant tanks and tlowmeters. Figure 11 shows the thruster mounted in the test cell with the propellant heat exchangers located just behind the thruster. The distance between the propellant heat exchangers and the thruster valve was approximately 10 cm.

Thermocouples (FIT, orT; See Figure 10) were located in the fuel and oxidizer lines just NASNCR-2001-210709

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Pump Heat Exchan er Fluid .-------------1f----+----------- __ Test Cell Fuel Ox Tank Tank MMH MON-25 Fuel Ox Flowmeters Flowmeters Fuel Ox Heat Heat Hx out Exchanger Exchanger Fuel Ox FIT Figure 10. Propellant Cooling System in Test Cell NASAlCR-2001-210709 20 L r - --- Propellant Heat Exchanger Figure 11. Thruster Mounted in Test Cell NASAlCR-2001-210709 upstream of the thruster to measure propellants temperatures as close as possible to the thruster inlet. The propellant heat exchangers contain copper coils through which the propellants flow .

The cold glycoVwater mixture flows around the outside of the coils and cools the propellant.

The heat exchanger is a counterflow design with the cold glycol entering the heat exchanger at the location where the propellant exits to the thruster. Typically, the propellant enters the heat exchanger at about 21 C (70F) and leaves at a temperature within about 1C (2F) of the entering glycoVwater mixture.

Figure 12 shows a high resolution time history of the propellant temperatures for Test 35625 which was a 1200s run at the nominal 15.2 bar (220 psia) feed pressure with nominal- 40C propellants. Except for a short period during the start transient, the propellant conditioning system maintained the propellants within ± 1C (±2F) of 40C for the entire 1200s run. The start transient is due to the cleaning of the small amount of propellant in the short line between the propellant heat exchangers and the thruster. Before each test, a short bleed flow is conducted into a catch tank to clear the propellant in this line, however, bleed flow times on these tests was kept short due to the limited amount of MON-25 available for these tests.

As noted earlier, the mixture of ethylene glycol and water used as a heat exchanger fluid becomes very viscous at low temperatures. When conditioning propellants to --40e, it was found that the heavy load imposed on the pump by this viscous fluid would cause the pump to over heat and cycle off during long tests. The two valleys in the propellant temperatures shown in Figure 12 are due to the pump cycling off and then being restarted. Figure 13 shows the time history of temperatures of the heat exchanger fluid at the heat exchanger inlet and outlet for this test. These temperatures are quite constant over most of the test except for the two points where the pump cycles off and is restarted. Note that in cooling the propellant, the temperature increase of the fluid is only Ie (2F) from heat exchanger inlet to exit. The system was designed with a good deal of margin in terms of the thermal cooling capability of the propellant heat exchangers and because of this margin, even when the pump shuts down and is restarted, the propellant temperatures were maintained within 1e of the target --40e valve for the entire 1200s test.

NASNCR-200 1-2 10709

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Before each test, the thruster and valves were cooled to a temperature which was below the propellant temperature. It was difficult to cool the hardware with any degree of precision, however, the objective was hardware temperatures equal to or lower than the propellant temperatures, and this was achieved in all cases. Hardware temperatures were measured by thermocouples located on the back of the inj ector. The hardware was cooled by throttling liquid CO through an orifice. In going through the orifice, the liquid flashed to vapor at the test cell pressure of about 1.5 torr (0.03 psia). This cold CO gas was directed at the hardware through four lines, one at each valve and two at the injector from opposing sides. This cold gas injection system had independent valves for the valve and injector cooling circuits and they were cycled onloffuntil the desifed temperatures were obtained. As noted earlier, it was difficult to achieve precise control of the hardware temperatures using this approach. The data shows that for all tests, hardware temperatures were below the propellant temperatures and for many tests, the hardware start temperatures were welI below the propellant temperatures which represents a very severe environment. In any future testing, a system which provides more precise control of hardware temperatures should be developed.

NASAlCR-2001- 210709 5.0 TEST PLAN The test procedures and test plan for this activity are documented in ARC PP-80704 (Test Procedure) and ARC PP-80704, Addendum 1 (Hot Fire Test Matrix). These docume nt s we re reviewed and approved by NASA before testing was initiated.

Ta bles 2, 3 and 4 show the planned tests. Table 2 shows the baseline tests with 21 C (70F) propellants. The test matrix consists of tests a different feed pressures from 6.9 - 20.7 bar (100 -3 00 psia), three tests to measure the effect of mixture ratio, and one test where the cell pr essure was increased to 10 torr (0.2 psia) to simulate the Mars atmospheric environment.

Shown in the last ce-l.umn is the ARC test number for the planned tes t. Table 3 shows the matrix for the tests to map thruster performance as a function of propellant temperature and feed pressure. Table 4 shows the matrix for testing with -40C propellants. All tests were completed except for Tests 6, 8, and 9 in Table 4. These tests were not conducted since of the MON -25 supply was exhausted.

Figure 14 shows the thermocouple instrumentation on the thruster. Two thermocouples (lCBT! , ICBT2) wee located on the backside of the injector. One thermocouple was located on each valve and valve mount and four thermocouples were located at the nozzle joint. These thermocouples were used to set pre-test hardware temperature and to monitor thruster behavior during hot fire testing.

NASN CR- 2001- 21 0709 26

I I i I I I

l

I 35617 35611 35612 35613 35614 35615 35616 35606 35601 35608 Number ARC Test 35607, 35610 - Survey Survey - - Survey Baseline Pc Survey Pc Survey Pc Pc Survey Pc Survey MRSurvey MR MR - Comments Baseline Pulse Mars Simulation - -- - Path Flow - Bellows Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter - - - - 104 104 104 Inj.

S104 S104 S 104 S I04 S104 ~104 ~ ~ ~ ~104 S104 - - ("F) Start _ 104 104 104 ___ ~104 ~104 SI04 SI04 SI04 SI04 ~104 ~ ~ ~ ~104 Valve L......

_ Table 2 _ Temperatures 70 70 70 70 70 70 70 70 70 70 70 ~ __ Propellant MON-25/MMH Baseline Test Matrix L- -- of 1 I 1 1 1 1 1 1 1 No.

Pulses '---- ~ - - _ - 60/- 601- 60/ 60/- 60/- 60/ (sec) 60/- 601- 601- 601 Run Duration 0_110 On/Off --~ - (psia) 1901220 220/190 2201220 2201220 100/100 150/150 2001200 2501250 300/300 220/250 220/220 FFP/OFP 8 9 I 5 6 7 2 3 4 10 11* No.

Test , IV -..J

z ;J> en ~ f IV o o ..... IV ....- o -..J o \0

I - - 35642 35643 35634 35635 35636 35363 35638 35629 35630 35630 35632 35633 35619 35620 35621 35622 35623 36639 35640 35641 - Number ARC Te'st · __ ___ _ L- -- ----- -- Tests -- - --- Mars - Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Pc Survey Simulation Comments --- - Mapping -- -_. -- -- ~ --- Path Flow Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter .Flowmeter - -_.

- .

._ Temperature -' - - - ..

0 -- 0 0 0 0 0 O O O O O -20 -40 ~ ~ ~ ~ ~ 90 ~3 ~3 ~3 9 ~ Inj. ~-40 ~-20 ~-20 ~-2 ~ ~-20 ~-4 ~-40 ~ -~ ---- Table 3.

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Te~eratures~F) 0 0 0 0 30 30 30 30 30 -40 -40 -40 -40 -40 -20 -20 -20 -20 -20 ---- MON-25/MMH Pressure and Propellant of I I I I I I 1 I 1 I 1 I . I 1 1 I 1 1 1 1 Pulses "-----~ 'No.

' - Duration - I- I- I- I- I- I- 30/- 301 - 30 60/ 30/- 30 60/- 301 - 30 301 - 30 60 30/- (sec) 601- 601 - 60/- 60/- 60/- 30/- 30 Run On/Off - -- - ---- (psia) 100/100 150/150 100/100 1501150 150/150 100/100 ISO/ISO 220/220 250/250 300/300 2201220 250/250 300/300 100/100 220/220 250/250 300/300 220/220 250/250 300/300 FFP/OFP 1 4 5 6 7 9 17 19 2 10 11 12 13 14 15 16 18 20 --~~~~- 3* No.

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I Test 5644 5646 3 35645 3 35618 35647 3 35649 3 35626 35 35624 35628 Number ARC Survey Pulse Test Pul se Test Pulse Test MR Survey MR MR Survey Comments Duration Test Duration Test Duration Test Duration Test Duration Test Duration Test Duration Test Duration Test Duration Test ----.---~- Path Flow Bellows Bellows Bellows Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter Flowmeter -40C -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 Inj. 5 5 5 5 5 5 5-40 5 5-40 5 5 5 5 5 ("F) Start -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 5 5 5 5 5 5 5 5 5 5 5 5 5 5 Valve - - Table 4.

-- Temperatures - -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 -40 - Propellant MON-2SIMMH Test Matrix at - - of I 1 1 1 - I 1 I 1 1 I I 1 20 20 20 No.

Pulses I- /- I - I- I- 0.5 / - 0.1 / 2/0.2 60 60/- 60/- (sec) 300 300 0.1 0. 0.51 Run Duration 6001- 600 6001- 300/- 1200/- 1200 OnlOff 220 / / (psia) 2201190 220/220 2201220 220 220/220 220/220 220/220 220/220 220/250 2201220 220 2201220 220/220 190/220 2201220 FFP/OFP 5 6 7 8 14 15 1 2 3 4 10 11 12 13 No.

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, t- eo , t- eo NASAlC R- 2 001 -2 10 709 - ----- -- - - -'- - -' ~----.-- . -~- -~ - ~~- - --- - -- 6.0 TEST RESULTS Table 5 provides a summary of all the tests conducted during this program. The detailed test data is given in Volume 2. Most of the tests were of30s or 60s duration and for these tests,

I

the results shown in Table 5 are at the end of the run. For the long dur ation 600s and 1200s runs, a data slice at 60s and at the end of the run are given in Table 5. Given in Table 5 are results for 1 thrust, Isp, mixture ra ti o(MR), propellant flowrates(total, fuel, oxidizer), propellant temperatures and valve and injector temperatures before the start of the test. Also shown is a value fo r Pc which is calculated from the measured thrust, assuming a value for the thrust coefficient(Cf) of 1.770. C* is calculated from this estimated value for Pc. The maximum chamber temperature measured during t he test using the Agema infrared camera is also given in Table 5.

Test 35607 was a 30s test which was terminated at 28s when an excessive amount of fuel was observed coming from the nozzle. This test was successfully rerun as Test 35610. It was suspected that there was gas in the oxidizer line during Test 35607 which caused poor combustion. This thruster has a high fraction of the fuel in the barrier so the chamber runs relatively cold, however, a consequence of this is that it is difficult to bum all the fuel, particularly when the feed pressure is low and the propellant is cold. Typically, at the low feed pressure conditions fuel was observed collecting around and vaporizing from the rim of the nozzle.

Test 35464 was a pulse test with a 0. SOO /0.500s duty cycle. This test was terminated before completion when fuel was observed coming from the nozzle and the strip chart da ta showed no eviden ce that combustion was occurring in the chamber. The previous two pulse tests with duty cycle of0.10010.100s and 0.2001 0.200s ran successfully.

Test 35649 was terminated early since the MON-25 oxidizer was exhausted early in the test. During the test, fuel was observed running out of the nozzle. The test data shows the mixture ratio decreasing during the test, indicating the oxidizer was being exhausted.

ASAfCR-200 1-2 10709 ru peIa peIa 0.20 0.

m • oombuilion no depleted during ox prequre preMO'" cell Test cell COMMENTS Test terminated Test Terminated, Slopped: , C -2 -a -1 -4 -3 34 31 28 31 30 33 29 22 18 22 23 26 ~ -<41 -38 -52 -43 ~ -29 -43 -39 -51 -22 - -19 -21 -21 -33 -30 -29 -28 -43

T ... -31 -43 -42 -41

Injector C -5 -7 22 23 24 23 19 19 24 20 17 19 16 12 23 24 25 33 -43 -43 -51 -58 $ -48 -51 -48 -49 -48 -48 -12 -12 - - - -28 - -32 -32 - -38 -36 -38 -4<4 -49 -47 -47 Fuel Hardware Va"'" 6 4 C -3 -9 20 -4 22 25 25 24 21 19 24 22 19 16 11 10 22 31 Ox -38 -53 ~2 -39 -43 -53 -51 -43 -47 -4<4 -4<4 - - - -23 -2 -22 -30 -29 - -35 -3 -37 -43 -48 -45 -46 Pre-te.t Va"'" 2 9 C ·2 -3 -2 -1 22 22 21 21 21 21 19 21 39 19 25 21 21 38 -39 - -39 -<41 -<41 -<41 -<41 -41 -39 -41 -39 ... -39 -42 - -39 -38 -17 - -19 -19 -1 -29 -30 -30 -31 -31 -39 -<40 -<40 -41 Fuel 2 9 C -2 -2 -1 -1 25 21 21 21 21 21 21 21 21 20 18 31 Ox -38 -39 -36 ~ -<41 -<41 -<40 -<41 -39 -<40 -41 -39 -37 -<40 -17 - -18 -2 -30 -<40 -41 ... -19 -18 -29 - -29 -38 -39 -39 -<40 PropejIani emperaturea 'T ' C 46 49 50 46 48 26 47 19 18 11 17 17 23 18 20 19 18 18 28 32 32 17 19 24 11 14 18 15 51 46 48 50 26 34 24 26 19 21 23 29 -34 -22 -42 - Temp InjeCtor C 623 293 359 591 668 662 673 678 893 679 262 629 712 878 258 679 732 912 264 665 734 891 615 462 Max Temp Cham note-4 -~~.--- Ox 3H 150 166 806 304 361 523 511 - gmll 2.301 2. 1.742 1. 2,470 2. 2.532 2.046 2.371 0.266 2. 2. 2.304 1.606 1.946 2.380 2.569 2.846 2.417 2.352 2.377 2. 2.400 2.367 2.359 2.345 2.356 2.318 1. 1.874 2.323 2. 2.809 1.532 1.947 2.367 2.508 2.842 1.557 1.883 2.366 2.559 2.858 0.261 0.503 2.640 2.165 -- 540 ~7 165 440 264 893 533 232 857 -424 - Fuel .

gmll Test Data Summary 1. 1.357 1.127 0.947 1. 1.728 1.406 1 1.291 0. 1. 1.273 1. 0. 1.014 1.253 1.341 1. 1.232 1.278 1.263 1.275 1. 1.281 1.214 1.302 1.264 1.286 0.916 1.126 1.408 1.527 1.709 0. 1.078 1.353 1.484 1.659 0.847 1.050 1.328 1.432 1.587 0.136 0.233 1.241 1.264 ~ PropeIIani FlooIr8les 726 507 135 534 514 662 745 568 960 633 649 640 636 632 646 573 622 605 038 518 390 025 740 992 404 934 694 397 737 ToCal gmt.

3. 2.869 2 . 4.010 4. 3 .938 3. 3. 0.433 3 . 3.647 3. 2.499 2. 3. 3.929 4.381 3. 3.631 3. 3. 3. 3. 3. 3.647 3. 3. 2.438 3.001 3.731 4. 4. 2. 3. 3. 3. 4.501 2. 3. 3.991 4.444 0. 0. 3.881 3.429 3. 2.

Program 59 --- - MR 616 585 546 254 604 624 396 837 622 600 865 794 920 900 858 963 840 682 852 946 846 942 865 802 663 650 644 644 787 806 764 690 713 839 793 .1 1. 1. 1. 1. 1. 1.801 1. 1. 1. 1. 1. 1. 1.798 1. 1. 1.931 1. 1. 1. 1. 1. 1. 1. 1. 1.801 1. 1. 1. 1.664 1. 1. 1. 1. 1. 1. 1. 1. 1. 1. 1.762 1.787 1.601 1.915 2 2.127 1.712 1.

mil 1504 1384 1355 1171 1522 1549 1440 1435 1324 1561 1438 1457 1259 1564 1441 1514 1217 1458 1444 1463 1357 1464 1446 1442 1461 1474 1 1283 1456 1492 1522 1155 1330 1480 1509 1539 1096 1344 1477 1505 1526 1404 1441 Cslar Mars Flyer 5.

Pc 17 94 18 77 93 21 41 17 43 67 65 26 74 23 20 43 85 73 66 79 57 69 82 76 91 66 51 04 66 01 bar 4. 3. 7. 8. 7. 6. 6. 7. 6. 6. 2. 4. 7. 7. 8. 5. 6. 6. 6.77 6.27 6. 6. 6. 6. 6. 4.90 6. 3. 7. 7. 3.35 7.12 6. 3 .31 7. 8.74 5.12 8.81 5. 6.94 7.64 8.63 6.93 6.28 -.-.--- Table .1 .9 .7 .8 .8 .2 .6 .2 lap oec 254.8 249.8 216.7 280.5 285.1 266 264.6 244.0 201.1 266.3 264 266.0 130 231 .4 266.4 265.5 278.5 225.0 266.6 266.7 269.5 249 269.8 266.6 266.2 273.0 271.8 197.2 268.7 280.1 272.9 278.4 277.6 235 275 212 245.9 283 201 .9 247.4 272.6 277.3 281 .7 208.4 2.24.1 258.3 265.9 Instrumentation error.

_-- ..

an N 15 54 04 28 20 28 24 97 05 57 52 61 90 66 35 52 70 61 72 94 83 90 98 30 01 90 50 76 12 86 61 62 94 . .58 . . 28 83 . . . . to 8.77 7.03 4. 9.12 8.77 0.85 9.46 9.43 3. 6.72 8. 9. 9. 9. 8. 9. 9. 9. 9. 9. 6. 9. 4. 7. 4. 1.

4 . 7. 9.88 0. 9. 8.

11 12 10 10.59 11 10 10 10 10 12.42 10. 10. 12. 10. 12.

ThrueI

-_

due lost 60 60 60 80 60 60 60 nil 60 60 60 60 60 60 80 60 60 60 60 60 60 60 30 30 60 30 30 30 30 60 30 30 30 30 60 30 30 na na na oec 300 600 600 Data Slice 1200 1200 1200 been 1 1 1 1 1 1 1 1 1 1 20 20 20 11 Pu_ --- -35617 has 28 60 60 60 60 60 60 60 60 60 60 60 60 80 80 30 30 60 30 30 30 60 30 30 30 -~ 60 30 30 30 30 60 60 60 10 .20 .50 oec 300 600 -~--~--.- Tlme 1200 1200 1200 0nI0ff 201 50/ .10/.10 .10/.10 . .

1.77 35606 0.07865 cm2 . Run .2 .2 .2 .2 .2 .9 .8 .9 .2 .9 .2 .2 .2 .2 .2 .2 .9 .3 .2 .2 .2 .2 .9 .2 .2 .2 Cf '" bar Testa ------ 915 916 915 Pres 915 915 ing 2115 2115 2115 for 2115 8113 2117 2111 2117 2115 2115 2115 2115 2115 3110 2115 2117 2115 2117 2115 2115 2117 : . .7f20.7 . .1/15.2 . . 8. .3110.3 .2115.2 .7f20.7 .

6. . 6. . 6. .7f20.7 6.9/6.9 .7f20.7 6. .7(20.7 .

Is 13 15 15. 15 OFPIFFP 15 13 10.3/10.3 17. 2O 15 15.2113.1 10 15 17. 2O 15 15. 15. 15 15. 10. 15. 17. 20 10.3/10.3 15. 17. 2O 10.3/10.3 15.2/15.2 17.2/17.2 20 15.2/15.2 15. 15. 15 15.2113.1 13.1/15.2 Feed -----.- naum thrust coIum in Oct Oct Nov Nov Nov Nov Nov Date 4 5 thMil 28 29 calculation 10 11 12 C· from temperature data .

for No Test 35615 35606 35607 35608 35609 35610 35611 35612 35613 3561<4 35616 35817 35618 35619 35620 35621 35622 35623 35624 35625 35626 35627 35628 35629 35630 35631 35632 35633 35634 35635 35636 35637 35636 35639 35640 35641 35642 35643 35644 35645 35646 35647 35646 35649 chamber NOTES calculated .

1 2 3 .. 5 6 7 8 9

10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 36 39 40 41 42 43 44 No Pc The .

2. Throat area 3. For pulse runs, values <I TEST 1.

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- Figure 15 shows the time history for the propellant, fuel valve, oxidizer valve and injector temperatures for Test 35625, which was a 1200s test at the nominal 15.2 bar (220 psia) feed pressure and with -40C propellant. Figure 15 shows that the injector and valve temperatures were below -40C at the start of the run. Pre-test hardware temperatures for each test are given in Table 5. The results in Figure 15 show the valve and injector temperatures rise smoothly during the run with steady-state valve temperatures being reached after 500s and the injector temperature after about 300s .

Figure 16 shows the time history of thrust, specific impulse and mixture ratio for Test 35625. Thrust and Isp are relatively constant except for two bumps which appear to be correlated with the dips in the propellant temperatures shown in Figure 12. Figure 17 shows a higher resolution chart of the mixture ratio time history. As the injector temperature increases, the mixtur~ ratio shifts downward reaching a steady-state value of about 1.84. The nominal mixture ratio for the thruster with 21 C propellants is 1.65; the changes in propellant viscosity caused by operation with -40C-propel1ants cause the mixture ratio to shift from a nominal 1.65 to 1.84.

Figure 18 shows the time history of the propellant flowrates for Test 35625. The flows are quite steady throughout the run. The oxidizer flow shows two small oscillations associated with the pump cycling off and being restarted. The fuel flow shows a small increase during the start transient as the injector temperature increases as this is responsible for the decrease in mixture ratio during the start transient.

Figure 19 shows the effect of propellant temperature on thruster mixture ratio for tests at the nominal 15 .2 bar (220 psia) feed pressure. The data covers the propellant temperature range of25C to -42C (77F to -44F). The data trend shows that as propellant temperature decreases, the mixture ratio increases. As discussed in Section 2, this is caused by the large increase in the fuel viscosity as the propellant temperature decreases. This thruster was designed for a nominal mixture ratio of 1.65 with 21 C (70F) propellants and as the results show, the mixture ratio shift is NASAl CR- 2001-210709 33 -- --- -- - -- --- ' -- - - -- --- - --- -

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- -- - - -- - - - - - - - - - - modest for temperatures above about -4C (25F), but increases rapidly for temperature below this level. It should be noted that it would be relatively easy to redesign the thruster by modifying the orifices and flow passages to provide a mixture ratio of 1.65 with --40C (-40F) propellants such that optimum performance could be obtained for the Mars environment.

Figure 20 shows how thruster specific impulse is affected by propellant temperature. The effect is quite modest with Isp decreasing by about lOs as the propellant temperature is decreased from 25C (77F) to --42C (-44F). Figure 4 shows how Isp varies with MR for 21 C (1 OF) propellants and these results show that as MR increases from 1.65 to 1. 85 for the 16 bar tests, Isp decreases by about 5s. This suggests that the decrease in Isp due to operation with cold propellants is caused about equally by the induced shift in mixture ratio and by the propellant temperature, itselr.

Figure 21 shows the effect of both feed pressure and propellant temperature on the thruster mixture ratio. The results show that at all propellant temperatures, except the 21C (70F) case, mixture ratio is relatively insensitive to feed pressure. With 21 C (70F) propellants, the mixture ratio decreases as feed pressure decreases and falls sharply for feed pressures below about 10.3 bar (150 psia). This occurs because at low feed pressures, the oxidizer te nds to boil in the injector causing a decrease in the oxidizer flowrate and a downward shift in mixture ratio.

This does not occur with cold propellants since they can absorb the injector heating without reaching the saturation temperature of the oxidizer.

Figure 22 shows how thruster specific impulse varies with feed pressure and propellant temperatures. The results show that Isp decreases as feed pressure and propellant temperatures decrease. The effect of propellant temperature has been discussed. The decrease with feed pressure is typical f or a thruster of this type and is due to a lower combustion efficiency as mass flowrate and chamber pressure decrease.

ARC's 2 lbf thruster uses a high level of fuel barrier cooling to maintain low chamber and injector temperatures when the thruster is firing. The chamber temperatures are measured NASNCR-2001-210709 ---- - ~ - ---- --

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Figure 23 shows how the injector temperatures are affected by the propellant feed pressure and temperature. Due to the high level of barrier cooling in this thruster, the injector 21 C (70F) propellant, the temperatures are quite low. At the nonnal operating condition with injector temperature is about 49C (120F) over the entire range of feed pressures. As shown in Figure 23, the injector temperature decreases as the propellant temperature is reduced_ As propellant temperature is reduced, the injector temperatures show a greater sensitivity to feed pressure. The strong influence of propellant temperature on the injector temperature is expected since the propellant provides the primary cooling for the injector.

NASNCR-2001-210709 43 _ 0 _ -- - - _ -- - 60 Propellant T emperature ....

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6 8 10 12 14 16 18 20 22 Feed Pressure bar Figure 23 . Effect of Propellant Feed Pressure and Temperature On Injector Temperature NASAlCR-2001- 210709 ---- ---- ------ --- ---- -- -- -- - - -----~-- 7.0 CONCLUSIONS ARC's ION thruster was successfully tested with propellant cooled to -40C and the thruster and valves cooled below this level. The thruster was operated successfully over the entire propellant temperature range of 21 C to -40C (70F to -40F) and feed pressure range of 6.9 - 20.7 bar (100-300 psia) required by the NASA-approved test matrix. While the thruster was not designed to operate at these conditions, it did demonstrate the capability to operate successfully at the temperature conditions expected to be encountered in the Mars environment.

When operating with -40C propellants, the thruster experienced a shift in mixture ratio from the nominal value of 1. 65 with 21 C propellants to about 1.85. This shift is caused by the increase in the MMH viscosity as propellant temperatures are reduced with a consequent reduction in the MMH flowrate and increase in mixture ratio. The increase in mixture ratio and the lower energy content of the -40C propellants cause a slight reduction of specific impulse of about lOs compared to the baseline value with 21 C propellants.

During this program, ARC successfully demonstrated the ability to manufacture MON-25 as required to conduct these tests. The procedures are now in place to make this propellant for any future activities. Further, ARC developed and demonstrated propellant conditioning systems for the fuel and oxidizer with the capability to deliver propellants to the thruster at -40C with a tolerance of ±1 C for tests of any duration. This ability to make propellants and to precisely control propellant temperatures provides NASA the demonstrated capability to conduct such Mars environment tests in the future.

The testing conducted on this program was exploratory in nature. The purpose was to use an existing thruster design to determine if there were any unexpected problems which would be encountered when operating a thruster with MON-251MMH propellants at -40c. The conclusion to be derived from these tests is that no such problems were uncovered and that a properly designed thruster should have no difficulty operating in the Mars environment.

NASN CR- 2 001 -2 10709 - ________ J

-- l

8.0 RECOMMENDATIONS Propellant Properties There is a lack of data on the properties ofMON-25, particularly at the low temperature conditions of interest for the Mars environment. Data on the thermodynamic and transport properties of this propellant over the 100C to -SOC range is needed to support thruster design and performance evaluation activities. It is recommended that NASA compile a database of MON-2S properties, identify deficiencies in the database, and initiate a program to obtain the required data.

Thruster Design The results in this report show that a thruster designed for 21 C propellants will operate somewhat differently when tested with -40C propellants. The reasons for the mixture ratio shift and specific impulse decrease observed herein are understood. The mixture ratio shift can be corrected by either modifying the injection orifices in the thruster or possibly, by using a · different set of trim orifices upstream of the valve. Modifying the design will allow one to recover the Isp lost to the mixture ratio shift. It is recommended that in any future activity to examine propulsion for the Mars environment, NASA provide a set of design objectives for the thruster and consider modifying an existing design to meet these requirements before initiating a test evaluation.

NASNCR-200 1-210709 46 L.~ . _._ Form Approved REPORT D OCUMENTATION PAGE OMB N o. 0704-0188 Public reporting burden for this collection of information is estimated to average 1 hour per response , including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reduci ng th is burden , to Washington Headquarters Services , Directorate for Informa tion Operations and Rep orts, 1215 Jefferson Davis Highway , Suite 1204 , Arlington , VA 22202-4302 , and to the Office of Management and Budget , P aperwo rk Reduclion Project (0704-0188) , Washington , DC 20503 .

1. AGEN C Y USE ONLY (Leave blank) R E PORT DATE 3 REPORT TYPE AND DATES COVERED 1 .

Apli l2001 Final Contractor Report

r

4. TITLE A ND SUBTITLE 5 . FU N DING N UMBERS Mars Flyer Rocket Prop ul sion Risk Assessment ARC Testing VVU - 755- B4-01 -00 6. AUTH O R(S) NAS3- 99197 Atlantic Research Corporation 7. PERFORMING ORGANIZATION NAME( S ) AN D ADDRESS ( ES ) 8 . PERFORMI NG ORGA N IZ ATI ON REPORT N UMBER Atlantic Research Corporation 6686 Walmore Road E-12642 Niagara Falls, New York 14304-1697 9 . SP ONSORING/MONITORING AG ENC Y NAM E(S ) AND A DDRESS ( ES ) 10. SPO N SORING /M ONITO RING AGENCY REPORT N U M B ER National Aeronautics and Space Admini stra ti on Washington, DC 20546-000 1 NASA CR-2001-210709 11 . SUPPLEMENTARY NOTES Project Managers, Brian Reed, 216 -977 -7489, and James Biaglo w, 216 -977- 7480, Power and On-Board Prop ul sion Technology Division, NASA Glenn Research Center, organization code 5430.

12a . DISTRIBUTION/AVAILABILITY STA T EM ENT 1 2b . DISTRIBUTION CODE Unclassified - Unlimited Subject Category: 20 Distribution: Nonstandard Available electronically at hrrp:l/gltrs.grc.nasa.gov/GLTRS Thi s publication i s available from the NASA Cent er for AeroSpace Infoml a ti on, 3 01 -621-0390.

13 . ABST R ACT (Maximum 200 wo rds) This report describes the investigation of a lO-N, bipropellant thruster, operating at _40 ° C, wi th monome th ylhydrazine (M MH) and 25% nitric oxide in nitrogen tetroxide (MON-25). The thruster testing was conducted as part of a risk reduction activity for the Mars Flyer, a proposed mi ss ion to fl y a miniature airplane in the Martian atmosphere. Testing was conducted using an existing thruster, designed for MMH and MON-3 propellant s. MON-25 oxidizer was success- fully manufactured from MON-3 by the addition of nitric oxide. The thruster was operated successfully over a range of propellant temperatures ( -40 to 21 °C) and feed pressures (6 .9 to 20.7 kP a). The thruster hardware was always equal or lower than the propellant temperature. Most tests were 30- and 60-second duration s, with 600- and 1200-second duration and pulse testing also conducted. When operating at _40 °C, the mi xture ratio of the thruster shifted from th e nominal value of 1.65 to about 1.85, probably caused by an increase in MMH viscos it y, with a corresponding reduction in MMH flowrate. Specific in1pulse at -40 °C (at nominal feed pressures) was 267 sec, while petiormance was 277 sec at 21 ° C. This difference in petiormance was due , in part, to th e mixture ratio shift.

15 . NUMBER O F PA GES 14. SUBJ E CT TERMS Liquid rocket propellants; Rocket testing; Mixed oxides of nitrogen; Mars fl yer 16. PRIC E CODE A04 17 . S ECU R ITY CLASSIFICATI ON 18 . SECUR I TY CLASSIFICATION 19 . SECURITY CLASSIFICATION 20. LIMI T ATION OF ABSTRA CT OFREPOAT OF TH IS PAGE OFABSTAACT Unclassified Unclassified Unclassified Standard Form 298 ( Rev. 2-89) NSN 7540-01-280-5500 Prescribed by ANSI Std. Z39-18 298-102

Next: APPENDIX B: MARS FLYER ROCKET PROPULSION RISK ASSESSMENT →

APPENDIX B: MARS FLYER ROCKET PROPULSION RISK ASSESSMENT

APPENDIX B: MARS FLYER ROCKET PROPULSION RISK ASSESSMENT Kaiser Marquardt Final Report This report is being reprinted in its entirety as originally printed in Aplil 200 1. The reader should note that the original page numbers have been retained.

NASAffM-2001-210575

NASAl CR-2001-21 0710

Mars Flyer Rocket Propulsion Risk Assessment

Kaiser Marquardt Testi ng

Kaiser Marquardt Van Nuys, California Prepared under Contract NAS3- 99198 National Aeronautics and Space Administration Glenn Research Center

April 2001

Trade names or manufacturers ' names are used in this report for identification only. This usage does not constitute an official endorsement, either expressed or implied, by the National Aeronautics and Space Administration .

Available from NASA Center for Aerospace Information National Technical Information Service 7121 Standard Drive 5285 Port Ro yal Road Springfield, VA 22100 Hanover, MD 21076 Price Code: A03 Price Code: A03 Available electronically at http : // gltrs.grc.nasa.gov I GLTRS TABLE OF CONTENTS 1 INTRODUCTION ••.•••.•••.•.••• ••• ••••• ••• •.•.••..•.•...••..•.....•..•••• ••••• ••.•.•. •• .•...•.••.• ••••••••••••• .. •.•...•••••••••••.••••• 1 2.4 MON-25 PRODUCTION ... ... ...... ......... .. .... ...... .... . ....... ... .. .. .. .. .. .......... ..... ...... . ............ ... .......... ..... .. 8 3.1 IGNITION CHARACTERISTICS ..... .... ....... ... ...... . .................. ... ...... ....... ...... .. .... ... ... ................... .... . 12 3.3 STEADy-STATE CHARACTERISTICS .. .................. ... ........ ..... .......... ..... ....... .... ....... ..... ........ ... .... ... 18 3.5 PERFORMA CE AT 1.4-KPA CELL PRESSURE ... .. .. .. ..... ........ ...... ..... ..... .... ..... .... ... ... ............. ..... ... 20 NASA/CR-2001-2 107 10 III LIST OF FIGURES FIGURE 2 R-S3 PROPELLANT INJECTION AND FUEL-FILM COOLING SCHEME .. ... ........... ........ ....... ..... .... .. 2 FIGURE4 R-53 THRUSTER MOUNTED 0 CELL 9 SMALLROCKETTHRUSTSTAND ............................... .. 3 FIGURE 5 ROOTS BLOWERS AND STOKES PUMPS FOR CELL EVACUA TIO .' ............................. . ..... ...... .. .4 FIGURE 6 CELL 9 DATA ACQUISITIO AND FACILITY CO TROL CE TER .... ... ....... .. ... ... ... .......... .. .. ........ .4 FIGURE 8 NITROGE PURGES FOR E GINE THERMAL CONDITIONING ...... ... . ...................................... .. .. 6 FIGURE 14 TYPICAL POWER SPECTRAL DENSITY VERSUS FREQUENCY OF A RAW THRUST SIG AL . ... ... 14 FIGURE 15 TYPICAL THERMAL CHARACTERISTICS OF A lO-SEC HOT FIRE; F = 8.9N, OfF = 2.2 .... . ... . 14 FIGURE 19 SPECIFIC IMPULSE VERSUS PROPELLANT MIXTURE RATIO .. .. .. .. ................ ...... .. ... ' . .... .... ... .. 17 FIGURE22 E GINECONTROLVALVEANDMA IFOLDTEMPERATURESFORALO G-DURATIO BURN.19 FIGURE 24 ACCELEROMETER TRACE DURING FIRST 20.48 MS OF HOT FIRE .. ... .... .. .. .... . .... .......... . ....... . 21 FIGURE 25 POWER SPECTRAL DE SITY FOR THE FIRST 20.48 MSEC OF HOT FIRE . .. ..... .... ..... ... . .... ... .... . 22 LIST OF TABLES TABLE 1 ENGINE TEST SETUP-SPECIFIC INSTRUMENTATION ... .. ............ ....... ...... . ..... .... ... .. .. . ... ... .. ... ... .. 9 TABLE 2 ACTUAL TEST MATRIX PERFORMED ...... ........ ..... ... ... .... .. ... ..... .... . ..... .... ... .... ....... .... . .... .... .. .. 11 TABLE 3 SUMMARY OF PERFORMANCE AT ELEVATED CELL PRESSURE . .... .... .... ..... .. .... ............ ....... ... 20 ASA/CR-2001-2 107 10 iv ~~~-.- . - _.- ------_ .- LIST OF NOME CLATURE Symbols characteristic velocity

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ignition delay F , Ff thrust (vacuum, unless otherwise specified) C* efficiency 17c * Is p, ISP specific impul e (vacuum , unles otherwise specified) N 0 dinitrogen tetroxide 2 4 NO nitric oxide oxidizer-to-fuel mixture ratio OfF P ee )) , PCELL test cell pres ure fuel inlet manifold pressure P mf, PMF oxidizer inlet manifold pressure P mo ,PMO t time T el> TC1 injector-chamber interface temperature #1 T , TC2 injector-chamber interface temperature #2 e2 T , TCH chamber temperature eh Tm TMF fuel inlet manifold temperature f, oxidizer inlet manifold temperature T mo, TMO valve time off t off valve time on t on T f] ,TF1 oxidizer-side valve flange temperature fuel- side valve flange temperature T f2, TF2 T , TVF fuel-side valve temperature vf T vo , TVO oxidizer-side valve temperature valve current Iv, VI Vv, VV valve voltage fuel flow rate W fu , WFU W , WOX oxidizer flow rate ox Acronyms ACC1 accelerometer #1 C-103 columbium-metal alloy CD-ROM compact disc read-only memory gaseous nitrogen GN2 LN2 liquid nitrogen MMH monomethylhydrazine mixed oxides of nitrogen, 75 % N204 / 25 % NO by weight MON-25 MON-3 mixed oxides of nitrogen, 97 % N204 / 3% NO by weight NASA National Aeronautics and Space Administration RAM random-access memory National Instruments signal conditioning system SCX!

v NASA/CR- 2001-21071O 1 INTRODUCTION The Mars Flyer mission proposed by the National Aeronautics and Space Administration will fly a miniature airplane in the Martian atmosphere on the centennial anniversary of the Wright broth er 's firs t powered flight. At the time of proposal, both, an electric motor-driven propeller and a chemical rocket engine were und er consideration for the Mars Flyer propulsion sy tem. As part of a lisk reduction investigation, Kaiser Marquardt was contracted to validate, by te t demonstration, the use of a chemical thruster utilizing mixed oxides of nitrogen and monomethylhydrazine. Th e objective was to provide NASA wi th as much data as possible for the flight system decision.

Preliminary studies performed by NASA showed that a thrust level of 8.9 N would be appropriate for a rocket-powered Mars Flyer. The Mar s Flyer will be power-limited and volume-limited, therefore, the rocket propulsion system will have to employ storable propellants and be capable of operating in the Martian environment without thermal management. Without thermal management, the entire propulsion system, including propellants, will be subjected to extremely low temperature as the average diurnal temperature of the surface-level Martian atmosphere is -40 degrees Celsius.

An experimental investigation was conducted to assess the performance of an 8.9 -N , bipropellant thruster operating at -40 °C with monomethylhydrazine ( MMH ) and mixed oxides of nitrogen (MON). To facilitate engine operation at low temperatures, dinitrogen tetroxide, N 0 was saturated 2 4 , with nitric oxide, NO, to lower the freezing point. The freezing point of the industry-standard, 3% nitric oxide in dinitrogen tetroxide (MON-3) is -15 °C . By increa ing the nitric oxide content to 25 % (MON -2S), the freez in g point was lowered to -55 °C , thus enabling safe operation of the thruster a t- 40 °C with sufficient margin for error. The thru ter was tested in a near-vacuum environment and conditioned, along with the propellants, to -40 °C prior to hot firing. Thruster operating parameters included oxidizer-to-fuel mixture ratios of 1.6 to 2.7 and inlet pressure ranging from 689 to 2070 kPa.

The test matrix consisted of many lO -second firings and several 60, 300, 600, and 1200-seco nd firings . Measurements included thrust, propellant flow rates, propellant inlet pressures and temperatures, engine temperatures, system vibrations, and valve control- ignal characteristics. Data obtained from testing were analyzed to determine engine performance characteristics such as ignition delay, specific impulse, I versus oxidizer-to-mixture ratio, Off , I p versus thrust, F, and the sp, vibration frequency spectrum. Preliminary results indicate that the additional nitric oxide not only permits lower propellant temperatures, but also compensates for the loss of performance associated with the lower propellant temperatures by introducing additional chemical energy to the combustion process. Results exhibited comparable, if not superior, performance as compared to those obtained with the same thruster fired at normal operating temperatures ( ~22 D C) using MON-3 and MMH.

2 TEST PROGRAM 2.1 Test Article The test article used in this investigation was the Kaiser Marquardt Model R-S3 8.9-N bipropellant thruster shown in Figure 1. The thruster consists of three primary components, the control valve, the propellant injector, and the chamber-nozzle assembly. The engine control valve is a Moog torque- motor bipropellant type designed to simultaneously control the flow of both oxidizer and fuel. The injector consists of a single unlike-doublet with two separate injection ports for fuel-film cooling of the combustion chamber wall. The portion of fuel used for fuel-film cooling was, by design, to be of the total flow. As a result of manufacturing difficulties, however, the cooling fuel was only 30% NASA/CR - 2001-210710 20% of the total flow , which resulted in significantly higher chamber temperatures during hot firing.

The single unlike-doublet is a pair of injection ports designed to impinge a stream of oxidizer with a stream of fuel inside the combustion chamber and very near the external face of the injector (see schematic shown in Figure 2). The chamber-nozzle assembly consists of a single piece of forged C- 103 columbium (niobium) alloy with a silicide coating to prevent oxidation. In these tests, the chamber was operated at temperatures of up to 1500 °C for a total of nearly 10,000 seconds.

Figure 1 Kaiser Marquardt Model R-53 8.9-N bipropellant thruster Injector Unlike Fuel-Film Chamber Doublet Cooling Jets Figure 2 R-53 propellant injection and fuel-film cooling scheme.

2.2 Test Facilities All testing was conducted in Cell 9 of the Kaiser Marquardt Rocket Test Facility located in Van Nuys, California and shown in Figure 3. The Cell 9 Rocket Test Facility includes a small thrust stand (s ee Figure 4) located within a very la rg e vacuum sphere (see Figure 3). The vacuum sphere may be evacuated using a large steam ejection system or a series of Stokes pumps and Roots blowers shown in Figure 5. The facility is fully automated and operated from a remote data acquisition and control center shown in Figure 6. Data acquisition was performed using a National Instruments SCX! signal conditioning system controlled by an Intel Pentium II 500 MHz computer having 1 gigabyte of RAM , NASA/CR-2001-210710 2 an 18 gigabyte hard drive, a 53-cm monitor, and a CD-ROM writer. The MON-25 propellant was produced on-site and analyzed in the Kaiser Marquardt chemistry lab.

Figure 3 Kaiser Marquardt Cell 9 rocket test facility.

Figure 4 R-53 thruster mounted on Cell 9 small rocket thrust stand.

NASA/CR-2001-210710 3

I

Figure 5 Roots blowers and Stokes pumps for cell evacuation.

Figure 6 Cell 9 data acquisition and facility control center.

2.3 Facility Modifications To accommodate testing of the Model R-53 8.9-N thruster at -40 °C, it was necessary to modify the existing propellant conditioning system and to make provisions for thermal conditioning of the thruster hardware. Normally, propellants are conditioned using an Environmental Industries e El ) Model ME2110 refrigeration system. This system, however, is only capable of conditioning propellants to -7 ° C. Therefore, a Government-supplied thermal conditioning system consisting of two FrS Model RClOBO, shown in Figure 7, was used in conjunction with the existing system . This combination system enabled conditioning of the propellants to temperatures below -46 °C.

Conditioning of the propellants was achieved using a system of co-annular plumbing, where the propellants were passed through an interior passage encased in an outer tube of flowing refrigerant.

This system of co-annular plumbing began at the main propellant run tanks, extended to the engine valve inlets, and was interrupted only by tho e transition joints and valves where the system could not be praCtically implemented. Conditioning was divided into two stages. The first stage utilized the NASA/CR-200 1-2107 10 4 L- ___ ~ normal EI refrigeration system with ethylene glycol as the refrigerant and the second stage utilized the FrS refrigeration system with a silicon-based refrigerant. The first stage conditioned the propellants in the storage tanks and in the lines up to the cell vacuum sphere. The second stage conditioned the propellants from ju st inside the cell wall up to the engine inlets. The first stage temperature was controlled at -7 °C, while the second stage temperature was set at the desired run temperature of down to -40 DC.

To provide a stable propellant temperature throughout the duration of hot firings, lasting up to 20 minutes, one-gallon, chilled accumulator tanks were installed serially in-line with the propellant supply lines. These tanks were located inside the cell vacuum sphere and controlled with the second- stage thermal conditioning system. This arrangement provided a buffer between the propellant chilled to the desired run temperature and the higher-temperature propellant entering from outside the cell vacuum sphere.

Figure 7 FTS Model RCIOBO refrigeration units in protective shed.

The engine hardware was conditioned to -40 °C on the thrust stand through the use of a nitrogen gas purge. Liquid nitrogen was circulated through a reservoir, which contained a helical heat exchanger tube through which gaseous nitrogen was passed. The cold gaseous nitrogen was introduced to the engine hardware through a system of normal impingement jets shown in Figure 8. A ring containing several jet orifices cooled the engine valve, injector, and chamber. A bar, also containing several jet orifices, cooled the engine inlets and the short sections of unconditioned, flexible propellant lines.

Figure Q shows a schematic of the entire setup including the propellant and engine hardware conditioning systems.

NASA/CR-2001-21071O Fig ure 8 Nitrogen pur ges for engine thermal conditioning.

NASA/CR- 200 1-2 107 10 6 Main Main Oxidizer Fuel GN2 Tank Tank Supply (MON-25) (MMH) LN2 Coriolis Coriolis Heat Flowmeter Flowmeter Exchanger Cell Vacuum Fuel Oxidizer Sphere Accumulator Accumulator GN2 Purge Bar GN2 Purge R-53 Ring Thruster Conditioned from Ambient to -6°C __ II'iIl""!!' Conditioned from Ambient to -50 °C Figure 9 Facility Flow Schematic for Mars Flyer Testing NASA/C R- 2001-2 1 071 0 7 2.4 MON-25 Production The required mixed oxides of nitrogen composition, 75 % dinitrogen tetroxide (N 0 ) and 25 % nitric 2 4 oxide ( NO) by weight, was produced on-site at Kai ser Marquardt. Production was ac hieved by introducing gaseous NO to commercially available MON-3 (3% NO in N 0 in a controlled 2 4) reaction. Production began with a known amount of MON-3 in the run-storage tank. Nitric oxide was then allowed to flow into the run-storage tank from an external storage cylinder mounted on a weight balance. Because the saturation of NO into N 0 involves an exothern1ic reaction, the flow of 2 4 the nitric oxide was carefully controlled so as to preclude thermal run-away. The storage tank was conditioned to -7 °C to lower the vapor pre ss ure of the mixture and to accelerate the process. For reference, Figure 10 shows the vapor pressure of the mixed oxides of nitrogen. A tank stllTing mechanism was also used to further accelerate the process. After the required amount of NO had been transferred and the saturation process was complete, a sample was taken for composition analysis. The species composition analysis was pelformed in the Kaiser Marquardt chemistry laboratory and according to specification MIL-PRF-26539E. After completion of production, the storage tank was held under pressure (> 700 kPa ) to ensure composition continuity. This also permitted shutdown of the thern1al conditioning system during storage.

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Temperature, oK Figure 10 Vapor pressure of the mixed oxides of nitrogen.

2.5 Instrumentation and Data Acquisition In addition to the standard facility instrumentation, the engine test setup was equipped with instrumentation to measure thrust, propellant flow rates , propellant inlet pressures, propellant inlet temperatures , valve control signal characteristics, various engine temperatures, and system vibrations.

All measurements were calibrated in standard (English) units and later converted to metric (System International) units. Table I shows a summary of the engine test setup-specific instrumentation along NASA/CR-2001-21071O 8 ~------ - - - - with the ranges of interest. Figure 11 sho ws the locations of all engine thennocouple s. Data acquisition was perfonned using a National In truments SCXI signal conditioning system controlled by an Intel Pentium II 500 MHz computer having 1 gigabyte of RAM, an 18 gigabyte hard drive , a 53-cm monitor, and a CD-ROM writer.

Table 1 Engine Test Setup-Specific Instrumentation MEASUREMEN T ACRONYM UNITS MIN MA X CUT Fuel Flow Rate WFU Ibm/sec 0.002 0.01 Oxidizer Flow Rate WOX Ibm/sec 0.002 0.01 Thrust FT Ibf 0 5 Fuel Inlet Manifold Pressure PMF psia 0 300 PMO psia 0 300 Oxidizer Inlet Manifold Pr essure Cell Pressure PCELL psia 0 1.0 Valve Voltage VV volts 0 50 Valve Current VI amps 0 1 O F Fuel Inlet Manifold Temperature TMF -60 120 of Oxidizer Inlet Manifold Temperature TMO -60 120 O F Chamber Temperature TCH 1500 3000 2800 max O F Injector/Chamber Temperature TC1 -60 1000 800 max of TC2 -60 1000 800 max Injector/Chamber Temperature of Valve Temperature - Fuel Side TVF -60 500 250 max of Valve Temperature - Oxidizer Side TVO -60 500 250 max O F -60 1000 250 max Flange Temperature TF1 of Flange Temperature TF 2 -60 1000 250 max CT Accelerometer ACC1 -100 +100 C> NASA/CR-2001-2 1071 0 9 I

. I

Figure 11 Engine thermocouple locations.

NASA/C R- 200 1- 210710 I

_______________ - _~ ____ J

3 RESULTS AND DISCUSSI ON The thruster was tested wi th propellant inlet pre ures ranging fr om 7 23 to 23 15 kP a, yielding vacuum thrust levels of 4 .5 to 10.6 N with mixture ratios from 1.6 to 2.7. Table 2 sho ws a summary of the tests performed.

Table 2 Actual Test Matrix Performed R UN pm o pmr t ON t OFF PULSES t )r op p" " kPa kPa sec sec kPa °C 688 1571 1565 10 --- I <0.15 -40 689 1570 1555 10 --- I -40 <0.15 690 1567 1586 10 --- I -40 <0 .15 691 15 66 165 8 10 --- I -40 <0. 15 692 1564 1795 10 --- I -40 <0 .15 693 1543 1836 10 --- I -40 <0. 15 694 154 9 1 828 10 --- I -40 <0.15 695 1550 1904 10 --- I -40 <0.15 2022 10 I 696 1576 --- -40 <0. 15 697 1522 2076 10 --- I -40 <0.15 698 1479 21 08 --- I -40 <0. 15 699 1447 2134 JO --- I -40 <0.15 700 1449 2177 10 I -40 <0 . 15 --- 701 1621 1766 10 --- I -40 <0. 15 702 1657 1729 10 --- I -40 <0 . 15 703 1657 1693 10 --- I -40 <0.15 704 10 I <0.15 1657 1624 --- -40 705 16 95 155 9 10 --- I - 40 < 0.15 706 1727 1591 10 --- I -40 <0. 15 707 1824 162 9 10 --- I -40 < 0.15 708 1553 1835 10 --- I -40 <0. 15 709 1696 1971 10 --- I -40 < 0. 15 710 1 86 1 21 08 10 --- I -40 <0. 15 711 1967 2248 10 --- I -40 <0.15 713 2028 2315 10 --- I -40 <0. 15 714 1407 1697 10 --- I -40 <0. 15 715 1555 10 I -40 <0.15 1306 --- 716 1202 1416 10 --- I -40 <0.15 717 1072 127 8 10 --- I -40 <0.15 718 929 1105 10 - -- I -40 < 0.15 719 792 963 10 --- I -40 <0. 15 720 723 848 10 --- I -40 <0. 15 I < 0. 15 721 1551 1831 10 --- -40 I <0. 15 722 1544 1 828 300 --- -40 726 1564 1839 10 - -- I < 0. 15 -7 727 1555 183 1 10 --- I < 0. 15 -7 728 1607 175 8 10 --- I -7 <0 . 15 729 1556 1836 10 --- I -18 <0.15 730 1553 1827 10 --- I -18 <0. 15 <0. 15 733 1566 1838 10 --- I -18 735 1552 1835 0.1 0.1 20 -40 <0.15 736 1551 1828 0.1 0.1 20 -40 <0.15 737 1551 1825 0.1 0.1 20 -40 <0. 15 739 1551 1827 1 200 --- I -40 <0. 15 740 1551 1 827 1 200 --- I -40 <0.15 741 1547 1 826 60 --- I -40 1. 5 <0 .1 5 742 1 551 1 827 1 200 --- I -40 743 1200 I <0. 15 1519 1 853 - -- -40 744 1528 1827 600 --- I -40 < 0.15 745 1520 18 19 600 --- I -40 <0 . 15 746 1533 1820 600 - -- I -40 <0.15 747 1536 1820 300 --- I -40 < 0. 15 748 1540 1826 300 --- I -40 <0.15 <0.15 749 1530 1828 300 --- I -40 750 1551 1827 300 --- I 40 <0. 15 NA SA/C R- 200 1- 2107 JO 11 Initial tests were performed to as certain the propellant inlet pressure producing a vacuum thrust level, F of 8.9 N and an oxidizer-to-fuel mixture ratio, OfF, of 2. 1. The engine was designed to produce vac , these nominal values for equal inlet pressures of 1517 kP a. Improper propellant-flow trim orifices, however, required oxidizer and fuel inlet pre ss ures of 1551 and 1827 kPa , respectively, to produce the nominal perfom1ance.

Mea surements included thrust, propellant flow rate s, propellant inlet pressures and temperatures, engine temperatures, system vibrations, and valve control-signal characteristics. Data obtained from testing were anal yzed to determine engine performance characteristics such as ignition delay, specific impulse, Isp, versus oxidizer-to-mixture ratio , O/F , Isp versus thrust, F, and the vibration frequency spectrum. In each firing, 1 second of pre-fire data was taken to obtain an accurate pre-run thrust tare and to ensure that all instrumentation was operating properly prior to firing. In each case, 10 seconds of data was taken after the hot firing to obtain a post -run thrust tare and to observe thermal soak back trends.

The total firing time for the matrix described in Ta ble 2 is nearly 8600 second s. Although the amount of fuel-film cooling was much le ss than the design value, resulting in ve ry high chan1ber temperatures, the engine did survive the entire test matrix . In fact, an additional 2400 seconds of firing time were put on the engine to dispose of the excess propellant s. Record of these final runs was not possible, because the remaining propellant was downstream of the flow meters.

3.1 Ig ni ti on Characteristics Ignition delay was measured as the time between control signal " ON " and the first instance of measurable thrust minus the valve response time. Figure 12 shows the ignition delay plotted versus vacuum thrust level for a constant mixture ratio of 2 . 1±0 . OS . As the thrust level was decreased from approxin1ately 10.6 N to 4.S N, the ignition delay increased from 7 .S ms to 11 .S ms . Ignition delay versus mixture ratio , on the other hand, display ed no significant trend. Overall, ignition delay is very similar to that exhibited by the sa me thru ster fired at normal operating temperatures (~22 °C) using MON -3 andMMH .

I Mode l R ·53 at · 40 · C MON·25 / MMH

\

O/F = 2.1 ± 0 .0 5

~

E 10

-

:>: ~ (1) C 9 c: o

~

'"

~

E> 8 ~ .-..

I 6,," = 0 .08 03 F' ·1. 8554 F +1 8.154 ms I 4 5 6 7 8 9 10 11 12 Thrust, N Figure 12 Ig ni ti on delay versus th rus t.

NASA/CR-2001-210710 12 - - - -- - - At low thru t levels, roughly less than 6 N, the startup proces exhibited a two- tage behavior.

Immediately after ignition, the engine produced very little thrust, le ss than 1 N, for about 10 ms after which, the thrust abruptly reached its full steady state value. Thi delay in full combustion results in an accumulation of propellants followed by a detonation and a momentary spike in chamber pressure and thrust. The engine, however, was not damaged as a result of this phenomena.

3.2 Performance Mapping Thirty-three lO-second hot firings were made at various thrust levels and mixture ratios to map the domain of performance for a range of inlet pressure with the engine and propellants conditioned to - 40 0 c. Initially, several trim runs were made to identify the nominal inlet pre ssures yielding an oxidizer-to-fuel mixture ratio, OfF , of 2.1 at a vacuum thrust level , F , of 8.9 N. The nominal inlet vac pressures, pm o and Pmf, were determined to be 1550 and 1830 kPa, respectively. Excursions from the nominal inlet pressures were such that, either the thrust was held constant at nominal while varying the mixture ratio or, vice ve rsa. Performance data shown in the following plots and listed in Table 4 of the Appendix are averaged over the next-to-last second of each run .

Figure l3 shows a typical thrust trace for a 10-second hot firing. The thrust trace shown in Figure 13 was passed through a digital band-reject filter designed to eliminate both 60-Hz noise and thrust- stand ringing which would otherwise obfuscate the true engine thrust response. Figure 14 shows the power spectral density of the raw, unfiltered thrust signal, clearly identifying the 160-Hz natural frequency of the thrust-stand load cell and the 60-Hz line interference.

16.0 Model R·53 at ·40 ' C MON·25 I MM H 14.0 O/F = 2.1; I, p = 290 sec 12.0 10.0 Z "lif 8.0 ::J ....

.s::.

I- 6.0 4.0 2.0 0. 00 2 4 6 8 10 12 Time, sec Figure 13 Typical filtered thrust trace for a lO-second run ( Run #693).

NASA/CR-2001-210710 13 90 r- N 80 r- ~ Load Cell Z Natural Frequency 70 r- "'. ~ 160 Hz ~ 'en 60 r- c: II> 50 r- C1l ...

-

II> 40 r- a.

If) ...

30 r- II> :;: 20 r- D..

60-Hz Line 10 __ Interference

I I I I I

°0 ~~~~ 1 ~ 00 ~"~ 2 ~ 0 ~ 0 ~~~ 3 ~ 0 ~ 0 ~~~ 4~0~0 ~~~ 500 Frequency, Hz Figure 14 Typical power spectral density versus frequency of a raw thrust signal.

Figure 15 shows traces of the chamber and the injector-chamber interface temperatures, T , and T el eh and T , respectively_ During hot fire, the chamber temperature increases monotonically and reaches e2 a maximum of 1360 d c. Immediately after engine shutdown, T eh decreases exponentially. Chamber temperatures below 900 °C could not be ascertained using the optical pyrometer and are, therefore, not plotted. Both injector-chamber interface temperatures increase very quickly over the first two seconds of hot firing then roll over and increase slowly throughout the entire run, reaching a maxin1um of 160 °C. After engine shutdown, both T el and T e2 increase and reach a maxin1um soak- back temperature of 290 °C after 7 seco nd s. Other engine temperature were not plotted because they showed no significant change over the 10-second firing duration_ Engine Hot Fire U 1!!

u :l

200 -

ai ~ ....

II> :l a.

E ~ II> II> ~ a.

....

100 1100 E II> II> ,Q ~ E C1l .s::.

U 0 1000 20 900 5 10 15 Time , sec Figure 15 Typical thermal characteristics of a 10-sec hot fire; F = 8.9N, OfF = 2.2.

NASA/CR-2001-210710 14

_J

Figure 16 shows a plot of C* efficiency versus vacu um thrust level for a constant mixture ratio of 2.1tO.OS. The C* efficiency is defined as the ratio of actual-to-theoretical characteristic velocities.

11 _ C ;C WGI (1) 'I C· - • C th eo reticol The C* efficiency increased with thrust from 80 .5% at 4.23 N to 93 .0 % at 10.2 N. For thrust level above 8.5 N, the C* efficiency appears to be relatively constant, however, it drops off sharply for thrust levels of less the 7 N. The empirical data was used to develop the following analytical curve fit for 'l7 c* as a function of thrust at an OfF of 2. 1: 3 2 'l7 • = 0.0007630F -0.0220SF +0.2150F +0.2187; 4.2 N < F < 10.2 N; 0/ F = 2.1 (2) C A chemical eqUilibrium computer code wa used to determine the theoretical C * values. The actual C* values were calculated based on measured vacuum specific in1pulse, Is p. vac , calculated thrust coefficient, and an assumed nozzle efficiency of 97 %. Th e thrust coefficient along with the corresponding chamber pressure was determined iteratively using a chemical equilibrium and nozzle expansion code.

1.000 ModeJ R-S3 at -40 ·c 0. 975 MON-2S / MMH OIF = 2.1 ± 0.05 0. 950

•

0.925 ~ ....

--

~ >- () 0. 900 c: Q)

~

'(3 ..

0. 875 == W

/

~ 0.850 U / J 0.825

I

0. 800

0.775 'l e' = 0 . 000 7630 F" - 0. 02205 F' + 0 . 2150 F + 0 .21 87 r---

0.750 10 11 12 4 5 6 789 Thrust, N Figure 16 C* efficiency versus thrust.

Figure 17 shows C* efficiency as a function of the oxidizer-to-fuel mixture ratio, Off, for a constant vacuum thrust level of 8.9±0.22 N . As the mixture ratio is increased, the C* efficiency initially increases from 92.6% at 1.61 to a maximum recorded value of 93.3 % at 1.99 after which it decreases to 87.1 % at 2.75. The experin1ental data was used to construct the following analytical curve fit for a thrust level of 8.9 N: 'l7 =-0.0755(0/F? + 0.2827(0/F)+ 0.6629; 1.61< OIF < 2.75; F =8.9N (3) c- NASA/CR-2001-210710 IS Using Eg. 3, the optimal mixture ratio is l.87 producing an T] c* of 92.8%, however, additional data is necessary to validate these values of regression analysis since the actual data show a maximum closer to 2.0.

1.000 I Model R·53 at -40 ·c 0.975 MON-25 / MMH - Thrust = 8.9 ± 0.22 N 0.950 -II ..

~ 0.925 ~-.- ....

c: Q)

~

lit....

'u 0.900 :E W ..

() 0. 875

"--

""

0. 850

"

11e' = -0. 0755 (O/F)' + 0.2827 (O/F) + 0.6629 0. 825

I

0.80~ .4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 Mixture Ratio Figure 17 C* efficiency v ersus propellant mix ture ratio.

I Model R-53 at -40 ' C MON-25 / MMH 310 r- Ol F = 2.1 ± 0.05 ()

Q) -

- ....

1/1

~

Q) 290 .!!!.

::J

V'

E () ;;::

/

'u 270 Q) Q.

/

"

250 1 .. _ 0.2331 P - 6.930 F' + 69 .13 F + 66 .71 sec 4 5 6 7 8 9 10 11 12 Thrust , N Figure 18 Specific impulse v ersus thrust.

Figure 18 shows a plot of vacuum specific impulse versus vacuum thrust for a constant mixture ratio of 2.1i o.os. Because the thrust coefficient is nearly constant at 1.84, the specific impulse is directly NASA/CR- 2001-210710 16 proportional to the C* efficiency and therefore shows the same trends . That is, as thrust increases, specific irnpul e also increases over the range tested. At 4.46 N of vacuum thrust, the measured I sp was 257.6 sec, while at 10.58 N the I sp was 298.1 sec. For thrust levels above 8.5 N, the specific inlpulse appears to be relatively constant, however, it drops off sharply for thru t levels of Ie s the 7 N. The empirical data was used to construct the following analytical curve fit for an O/P of 2.1: I =0.2331p -6.930P :! +69 . 13P+66.71(sec); 4.2N<P<10.2N; O/P=2.1 (4) sp Figure 19 shows a plot of vacuum specific impulse as a function of the oxidizer-to-fuel mixture ratio, O/P, for a constant vacuum thrust level of 8.9±0.22 N. Again, because the thrust coefficient is nearly constant at 1.84, the specific impulse is directly proportional to the C* efficiency and therefore hows the same trends. As the mixture ratio i increased, I s p. vac initially increases from 294.1 ec at 1.61 to a maxinlum recorded value of 300.7 sec at 1.99 after which it decreases to 271.1 sec at 2.75. The experimental data was used to construct the following analytical curve fit for a thrust level of 8.9 N:

I sp = -41.559(0/ F r + 160.58(0/ F)+ 142.84 (sec); 1.61 < 0/ F < 2.75; F = 8.9N (5)

Using Eq. 3, the optimal mixture ratio is 1.93 producing an I p of 298.0, however, additional data is necessary to validate these values of regression analysis since the actual data show a maximum closer to 2.0.

Model A-53 at -40 · C 305 r- MON-25 / MMH Thnlst = 8.9 ± 0. 22 N

300 •

....

............ .

u ~ 3l 295

1' -

i 290

::J

~

~ 285 u

.~

:E 280 u Q)

i'-

/Ji 275

\

r---ll .. = -41.559 (O/F)' + 160.58 (O /F) + 142 .84 sec

I

26~.4 1.6 1.8 2.2 2.4 2.6 2 .8 2 3 Mixture Ratio Figure 19 Specific impulse ve rsus propellant mixture ratio.

Figure 20 shows as plot of chamber temperature averaged over the next-to-Iast second of each ten- second run versus vacuum thrust for a constant mixture ratio of 2.1±0.05. As thrust was increased from 4.46 N to 10.58 N, the chamber temperature increased from 993°C to 1751 0c. This was to be expected as both the combustion efficiency and amount of total energy increase with thrust over the r~nge of inlet pressures tested. The empirical data was used to construct the following analytical curve fit for an O/P of 2.1: NASA/CR-2001-21 07 10 17 (4)

T e ll = -11 .5 45P 2 +297 .7 P -105.7 (K); 4.2 N < P < 10.2 N; 0/ P = 2.1

Model R·53 at .40 'c MON-25 / MMH O /F = 2.1 ± 0.05 ~

f!r

~

::J JI~ ~ /- (1) c..

V

E (1) ~

/

...

(1) ~ .0 E

/'

(\I ~

/

-

T", _ -11 .545 F' + 297 .7 F -105 .7 (K) 4 7 89 11 12 5 6 10 Thrust, N Figure 20 Chamber temperature versus thrust after 10 seconds of hot fire.

3.3 Steady-State Characteristics The plot shown in Figure 21 displays the quasi-steady-state characteristics of the Model R-53. The term quasi-steady-state is used here, because, as indicated in the plot, the mixture ratio decreases with tin1e and could not be held constant at its initial value of 2.1. The OIP does eventually, however, reach a steady value of approximately 1.82. For this mixture ratio , the injector-chamber interface temperatures, T e l and T , reach steady-state value of 145 and 180 °C, respectively. The chamber e2 temperature reaches a steady-state value of 1420 D C.

NASA/CR- 2001-210710 18 400 r-------------------------------~ 1800 2.25 350 1750 1700 2. 00 () OfF

er

;:, () (\I

0_ 200 -

TC2 1600 1.75 ~ ...

N Q) a: () a.

Q) ~ 150 1550 E ...

TC1 Q) 0/1 ;:, ~ ...

-

1 50 .~ U 100 Q) . ::: .&l ~ E 50 1450 (\I TCH ~ () o 1400 1.25 -50 1.00 Time, sec Figure 21 Quas i- steady- stat e thermal char acteristics fo r a long-du ra tion burn.

Figure 22 shows various temperature traces of the engine control valve and propellants for the same 1200-second bum shown in Figure 21. Note that the engine valve temperatures do not achieve steady state even after a 1200-sec bum. After 1200 seconds, the valve flange temperatures, Tn and T f2, are approximately 0 °e and the valve body temperatures, T and Tvf , are approximately - 12 °e. The vo propellant manifold temperatures, T me and T , are shown for reference. It was not possible to mf maintain Tm e and T mf at --40 °e for the entire duration of the run .

20 r----.----~----_r----_r----,-----,_, 10~--~----_+----~----~----4-----+H Or---~----_+-----r----_r----- -40 ~~~----_+-----r----_r----~----+_; Time, sec Figure 22 Engine control valve and manifold temperatures for a long-duration burn.

NASA/CR-2001-21071O 19 4.0 Model A-53 at _40 °C MON· 25 / MMH O/F = 2.1 3 .0 Z

en

2.0 ::::I ....

.l: t- 1.0 Ti me, sec F igu re 23 Sever a l pulses from ser ies of 20 wi th 50% d uty cycle and a 0.2-sec perio d 3.4 P ul sing Performance Figure 23 shows several typical pulses from a series of 20 with a 50% duty cycle and a 200-ms period (t o n/t off = 0.1/0.1 seconds). After 20 pulses, the vacuum impulse bit reaches an average value of 0.22 N-s.

3.5 Per fo r mance at 1.4- kPa Cell Pr essure Although the engine was designed to operate in a vacuum environment, a single run was made to ascertain the performance of the engine at nominal inlet pressures with a cell pressure of 1.4 kPa.

The idea was to simulate the average atmospheric pressure on the surface of Mars. It was expected that the flow inside the nozzle, designed for a much lower exhaust pressure, would separate and performance would drop off significantly. If the Model R- 53 were to be flown within the atmosphere of Mars, the nozzle would have to be redesigned and optimized for the 1.5-kPa ambient pressure. In any event, a sixty-second run was made and results were obtained. Video recordings did indeed show nozzle flow separation well inside the nozzle. Table 3 shows a summary of the averages over the last second of the run.

Tab le 3 Su mm a r y of Per for m ance at Elevate d Ce ll P ressure PAR AME TER VA L UE UN IT S Oxidizer inlet pressure 1547 kPa Fuel inlet pressure 1826 kPa Oxidizer inlet temperature -31 °C Fuel inlet temperature -37 °C Oxidizer flow rate 2.07 g/sec Fuel flow rate 1.085 g/sec Mixture ratio 1.9 Thrust 7.7 N NASA/CR-2001-210710 20 Specific impulse 249.0 sec Cell pressure 1.5 kPa Chamber temperature 1443 °C 3.6 Vibration Characteristics An accelerometer was attached to the back of the engine control valve to ascertain the system vibration characteristics. Figure 24 shows a typical accelerometer trace for the first 20.48 ms of a nominal lO-sec hot firing. The accelerometer had an accurate response range of up to 50,000 Hz with a resonant frequency of nearly 90,000 Hz. The data was sampled at 200,000 Hz using an IOTech high-speed data acquisition system. The highest-amplitude vibrations were experienced during the first 2 ms of startup after which the vibration was significantly reduced. There are momentary increases in the vibration amplitude, but they rarely exceed ±10 g.

Figure 25 shows the power spectral density versus frequency obtained from the accelerometer response trace shown in Figure 24.

C) C .Q ~ 0 (1) Cii u -10 u oct -20 -40 ~--J.--l--..J....---l--+----l--+----I---+---H -50 0 2 4 6 8 10 12 14 16 18 20 Time , msec Figure 24 Accelerometer trace dur in g first 20.48 ms of hot fire.

NASA/CR-2001-210710 21 ---- -- ... --_ ...

0.5 M ,....

0.4 X N ;;: N Cl :; 0.3

- 'iii

c: Q) C ~ 0.2 ti Q) a.

en ...

Q) 0.1 == Cl.

° 0 ~~~~~ ~~ 2 ~ 0 ~~~~~~~~W-~~ Frequency, kHz Figure 2S Power spectral density for the first 20.48 msec of hot fire.

NASA/CR-2001- 210710 22 4 NEW TECHNOLOGY While no new teclmologies were developed within the scope of this effort, the technology demonstrated may be applied to other mi ss ions besides the Mar Flyer program . Te st results show great promise for any power-limited space application where therm al management of the propulsion ystem may be restricted or even eliminated. Such missions include interplanetary travel, planetary de cent and ascent, and intra-atmospheric planetary navigation.

5 SUMMARY AND CONCLUSIONS An experimental investigation was conducted to assess the perform ance of an 8.9-N, bipropellant thruster operating at -40 °C with monomethylhydrazine ( MMH ) and mixed oxide of nitrogen (MON). To facilitate engine operation at low temperatures , dinitrogen tetroxide, N 0 was saturated 2 4, wi th nitric oxide, NO, to lower the freezing point. The freezing point of the industry-standard, 3% nitric oxide in dinitrogen tetroxide (MON-3 ) is -15 °C. By increasing the nitric oxide content to 25 % (MON-25), the freezing point was lowered to -55 °C, thus enabling safe operation of the thruster at - 40°C with sufficient margin for error. The thruster was tested in a near-vacuum environment and conditioned, along wi th the propellants, to -40 °C prior to hot firing. Thruster operating parameters included oxidizer-to-fuel mixture ratios of 1.6 to 2.7 and inlet pressure ranging from 689 to 2070 kPa.

The test matrix consisted of many 10-second firings and several 60, 300, 600, and 1200- econd firings. Measurements included thrust, propellant flow rate , propellant inlet pressures and temperatures, engine temperature s, system vibrations, and valve control-signal characteristics. Data obtained from testing were analyzed to determine engine perfomlance characteristics such as ignition delay, specific impulse, I sp, versus oxidizer-to-mixture ratio, Off , Isp versus thrust, F, and the vibration frequency spectrum .

Although more data will be necessary to fully characterize the performance of the Model R-53 at -4 '0 °C fired with MMH and MON-25 , the following general statements can be made.

• The engine was successfully operated at a thrust level of 8.9-N and -40 °C when fired with MMH and M ON -25.

• Performance is comparable to operation at normal temperatures (22° C) with MMH and MON-3.

• At nominal inlet pressures, the ignition delay is approximately 8 m s, which is comparable to operation at normal temperatures with MMH and MON-3.

• Based on regression anal yse, Optinlal performance for the 8.9-N thruster will be obtained at an Off of 1.93 with an anticipated average specific impulse of 298 sec.

• The thruster accumulated nearly 10,000 econds of operation without failure.

• For a duty cycle of 50% with a 200-ms period, the average vacuum impulse bit was 0.22 N-sec • The thruster was successfully operated in an atmosphere of 1.5 kPa.

• The thruster eXhi b.its no detrimental vibrations during hot firing.

NASA/CR-200 1- 21 07 10

Next: APPENDIX →

APPENDIX

Summary Data Listings

I

i I 0 0 0 .0 : 0 0 0 0 0 0 0 0 0 0 0 - 0 ms 8.0 8. 8. 8.0 8. 8.0 8.0 8. 8.0 8. 8. 8. 0 8.0 8.0 8.0 8. 8. 8.0 8. 8. 8. 8. 7.5 7.5 7.5 8.5 8.5 9. 9.5 8.

11.0 11.

9 0 .3 .0 1.

- Pc kPa 630. 0 627.4 639. 1 647.3 645.2 634.2 629.0 632.3 632.5 63 622.3 6 596.4 606.7 6 17.7 629.9 667.9 705.3 746.3 579. 1 545.8 506.8 459.2 360.7 3 15.3 622.9 595 .9 58 583.7 596.0 629.6 734 4 10. 1 f 18 835 8 849 855 854 842 846 844 846 846 84 1 844 844 858 857 847 844 844 809 85 1 84 C .8 1.857 1. 1. 1. 1.839 1. 1 1. 1. 1. 1. 1.853 1. 1.859 1.859 1. 1.846 1.846 1.842 1. 1. 1.845 1. 1. 1.

1. 1.854 1. 1.823 1. 1. 1. 1.

94 87 34 C* 459 1500 1505 1576 1575 15 1595 1590 1594 15 1576 1538 15 17 1502 1447 1457 14 1593 1585 1590 1 1594 1572 1553 15 1457 1370 1580 1 1458 1558 1603 1596 1496 m/sec gs.

11 62 79 32 20 60 50 in

°C 32

T eh 97 1 727 72 1 13 1436 1478 1377 1279 1308 1 1385 1379 1377 1375 13 1343 1313 1279 1273 1352 134 1 13 1306 13 18 13 13 1403 147 1 1299 1246 11 10 1304 .1 .7 .5 vac 1.

Hot Fir . 4.1 0.2 !'

sec Is 284. 0 276.0 293.6 296.3 296 300 298.3 296 294.8 29 292.9 297 .0 29 286.6 284.1 274.4 276.2 27 1.1 299.3 299.3 298.4 299.2 298. 1 299.4 295.6 292.3 285.5 28 273.4 257.6 297 .1 276.4 284.5 nd seco ac .1 N O- F,' 8.78 8.98 8.85 8.7 1 8.5 1 8.66 8.79 8.90 9.45 8 7.18 6.5 1 5.8 1 5. 8.82 8.5 0 8.30 8.32 8.48 8.93 8.92 8.88 9.02 9. 12 9.03 8.85 8.73 8.97 7.73 4.46 10.00 10.40 10.58 1 5 6 ll .9 .1 1. 1.

9.8 5. 6.9 ce Pa P 69.6 88.4 92.7 73. 1 88 .0 76.5 77.6 85.8 97. 1 87.5 99 84. 0 87 .3 1 10 12 11 145.8 12 129.6 107.5 138.9 116.9 147 .1 140.7 lll.l 138.0 12 122.4 125.2 104.7 111.7 108.9 14 12 17 F N 8.99 8.79 8.69 8.5 1 8.79 8.74 8.69 8.46 8.3 1 8.40 8.57 8.7 1 9. 8. 6.92 6.28 5.64 8.64 8.36 8. 8. 8.27 8.79 8.73 8.74 8.82 8.5 1 9.73 7.52 4.87 4.23 10.20 10.41 92 6 1 62 02 13 09 59 99 1. 1. 1.75 1.70 1. 1.

OfF 2.57 2.65 2. 2.46 2. 18 2. 07 2.09 2.22 2 .32 2.4 1 2.50 2.67 2.67 2.75 2. 2. 07 2. 15 2. 2. 15 2. 03 2. 2.13 2. 12 2. 11 2.01 2. 07 2.08 40 34 69 34 01 856 924 893 883 744 Wf g/sec 1.025 1.070 1.129 1.1 1.160 1.1 0. 0. 0.872 0. 1.003 1.046 1.086 1.1 1.150 0.8 0. 0.677 0.632 0.575 0.982 0. 855 0.8 0. 0.878 0. 977 0.998 0.989 0.957 0.949 0.863 0.920 0.863 63 44 07 1 070 035 977 928 32 028 868 803 0 W o .1 Performance Summary Data for l 1.705 glsec 1. 1. 1.868 1.889 2.226 2.17 1 1. 1. 1.580 1.427 1.270 1.1 9 1 2.

2. 196 2.223 2.2 18 2 2. 2. 2. 2. 2.048 2. 125 2.202 2.232 2.301 2. 2.425 2. 2.33 1 2.411 2.468 .1 .7 .2 .5 .6 1.1 1.

°C

40 40 40 40 Tonf -39.9 -38.6 -39.7 -41. 2 - -39.8 -40 -40.8 -42.2 -42.3 -40.7 -40.2 -40.2 -39.6 -41.3 -41.3 -44.2 -46.7 -37.2 -39.3 -39.3 -39.9 -38.9 -39.6 -39.5 - -39 .6 -39.7 -40.5 - - -4 -4 Table 4 2 2 .1 .7 .1 lI10 0. 0.4 0. 1.1

°C

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Form Approved

REPORT DOCUMENTATION PAGE

OMB N o. 0704-0188 Public reporting burden for this collection of informat io n is estimated to average 1 hour per response , including the time for re vi ewing instructions, searching existing data sources, gather! ng an? ma i nt~in i n~ the ?ata neede~ . and comple~ing a~d reviewing the c~lIection of information. S~nd comments regarding this burden estimate or any other aspect of th is collection of Inform~tlon , including suggestions for reducing this burden , to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson DavIs Highway, SUite 1204, Arlington, VA 22202·4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188 ), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) RE PO RT TYPE AND DATES COVERED 12 . REPORT DATE April 2001 Final Contractor Report

r'

TITLE AND SUBTITLE 4.

5. FUNDING NUMBERS Mars Flyer Rocket Propulsion Ri sk Assessment Kai se r Marquardt T es tin g VVU-755-B4-01 -0 0 6. AUTHOR(S ) NAS3- 99198 Kaiser Marquardt PER FORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 7. 8. PERFORMING ORGANIZATION REPORT NUMBER Kaiser Marquardt Rocket Propulsion Systems E-12643 16555 Saticoy Street Van Nuys, California 91406-1739 SPONSOR I NGIMONITORING 9. SPO NS ORING/ MONITORING AGENCY NAME(S) AND ADDRES S(ES) 10.

AGENCY REPORT NUMBER National Aeronautics and Space Administra ti on Washington, DC 20546-0001 NASA CR-2001-21071O 11 . SUPPLEMENTARY NOTES Kaiser Marquardt has been acquired by General Dynami cs, Ordinance and Ta ctical System s, Aerospace Operation s, 11441 Willows Road N.E. , P.O. Box 97009, Redmond, WA 98073-9709. Project Managers, Brian Reed, 216-977-7489 , and James Biaglow, 216-977-7480 , Power and On-Board Pr op ul sion Tec hn ology Di vision, NASA Glenn Research Center, organization code 5430.

DISTRIBUTION/AVA I LABILITY STATEMENT 1 2b. DISTRIBUTION CODE 12 a.

Unclassified - Unlimited Nonstandard Subject Category: 20 Disuibution: Avai l able electronically at http: // gltr s. gr c. na sa. go v/ GLTRS This public atio n is available from th e NASA Ce n te r for AeroS p ace lnfonnation . 301-621-0390.

ABSTRAC T (Maxi mum 200 word s) 13 .

This report describes the investigation of a lO - N, bipropellant thruster, op'erating at -40 ·C, with monomethylhydrazine (MMH) and 25 % nitric oxide in nitrogen tetroxide (MON-25). The thruster testing was conducted as pmt of a risk reduction activity for the Mars Flyer, a proposed mission to fl y a miniature airplane in th e Martian atmosphere. Testing was conducted using an existing thruster, designed for MMH and MON-3 propellant s. The nitric oxide content of MON-3 was increased to 25 %, to lower its freezing point to -5 5 . c. The thruster was conditioned, along with the propellant s, to temperature prior to hot firi n g. Thruster operating parameters included oxidizer-to-fuel mi xture ratios of 1.6 to 2.7 and inlet pressure ranging from 689 to 2070 kPa. The test matrix consisted of many lO-second fIrings and several 60-, 300-, 600-, and 1200-second firi ngs, as well as pulse testin g. The thruster successfully accumulated nem'ly 10,000 seconds of operation without fail ure, at temperatures ranging from -40 ·C to 22 . c. At nominal inlet pressures, the ignition delay was comparable to MMHlMON-3 operation. The optimal performance for the 8.9-N thrust er was determined to be at a mixture ratio of 1.93 with an average specific impulse of 298 sec.

15. NUMBER OF PAGES 14 . SUB JE CT TERMS Liquid rocket propellants; Rocket testing; Mixed oxides of nitrogen; Mars flyer 16 . PRICE C ODE A03 17. SECURI TY CLASSIFICATION 18 . SECURITY CLASSIFICATION 19 . SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OFTHIS PAGE OF A BSTRACT OF RE PORT Unclassified Unclassified U ncJassified Standard F orm 298 (Rev . 2-89) NSN 7540-01-280-5500 Prescribed by ANSI Std . Z39-18 298-102 Form Approved REPORT DOCUMENTATION PAGE OMB N o. 0704-0188 Public reporting burden tor this collection of information is estimated to average 1 hour per response , including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information . Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington , VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188). Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2 REPORT DATE 3 REPORT TYPE AND DATES COVERED 1 .

1 .

August 2001 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUND ING NUMBERS Overview of Propul s ion Systems for a Mars Aircraft VVU-710-70-23-00 AUTHOR (S) 6.

Anthony 1. Colozza, Christopher J. Mill er, Brian D. Reed , Lisa L. Kohout , and Patricia L. Loyselle 7 . PERFORMING ORGANIZATION NAME (S) AND ADDRESS (ES) 8. PERFORMING ORGANIZATION REPORT NUMBER National Aerona utics and Space Administration John H . Glenn Res earc h Center at L ew is Field E-12541 Cleveland, Ohio 44135 - 3 191 9. SPONSORING/MONITORING AGENCY NAME (S) AND ADDRESS (ES ) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administr'ation Washington, DC 20546-0001 NASA TM-2001-210575 11. SUPPLEMENTARY NOTES This repOlt includes NASA CR-2001-210709 and NASA CR-2001-210710 in appendixes. Anthony J . Colozza, D y nac s Engineering Company, Inc ., 2001 Aerospace Parkwa y, Brook Park , Ohio 44142; and Cluistopher J. Miller, Brian D. R ee d, Lisa L. Kohout , and Patricia L. Loy se lle, NASA Glenn Research Center. Responsibl e person , Lisa Kohout , organization code 5440 , 216-433-8004.

12 a. DISTRIBUTION/AVAILABILITY STATE M ENT 1 2b . DISTRIBUT ION CODE Unclassified - Unlimited Subject Category : 07 Distribution: Nonstandard Available electronically at hmJ: // gltrs.grc.na ~ a.go v/ GLTRS Thi s publication is avai lable from th e NASA Center for AeroSpace Information. 30 1 -62 1-0 390.

13 . ABSTRACT ( Ma xi mum 200 words ) The capabilities and performance of an aircraft depends greatly on the ability of the propulsion system to provide thrust. Since the beginning of powered flight, performance has increased in step with advancements in aircraft propulsion systems. The se advances in technology from combustion engines to jet s and rockets have enabled aircraft to exploit our atmospheric environ- ment and fly at altitudes near the Earth 's surface to near orbit at speeds ranging from hovering to several times the speed of sound . One of the main advantages of our atmosphere for these propulsion systems is the availability of oxygen. Getting oxygen basically "free" from the atmosphere dramatically increases the performance and capabilities of an aircraft. This is one of the reasons our present-day aircraft can perform such a wide ran ge of tasks. But this advantage is limited to Earth; if we want to fly an aircraft on another planetary bod y, such as Mars, we will eith er have to carry our own source of oxygen or use a propulsion system that does not require it. The Mars atmosphere, composed mainly of carbon dioxide, is very thin. Because of this low atmospheric density, an aircraft flying on Mar s will most likely be operating, in aerodynamical terms, within a very low Reynolds number regime. Also, the speed of sound within the Martian environment is approximately 20 percent less than it is on Earth. The reduction in the speed of sound plays an important role in the aerodynamic performance of both the aircraft itself and the components of the propulsion system, such as the propeller. This low Reynolds number-high Mach number flight regime is a unique flight environment that is very rarely encountered here on Earth.

14 . SUBJECT TERMS NUMBER OF PAGES 15.

Pl anetary exploration; Propulsion and power; Aircraft 16 . PRICE CODE 17 . SECUR I TY CLASSIFICATION 18. SECURITY CLASSIFICATION 19 . SECURITY CLASSIFICATION 20 . LIM I TATION OF ABSTRACT OF REPORT OFTHIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified S t andard Fo rm 298 (Rev. 2 - 89) NSN 7540-01-280-5500 Prescribed b y ANSI Std. Z39 - 18 298-102

Overview

APPENDIX A: MARS FLYER ROCKET PROPULSIO RISK ASSESSMENT View in PDF

NASAl CR-2001-210709

Mars Flyer Rocket Propulsion Risk Asses s me nt

ARC Testing

April 2001

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