Small hydrogen/oxygen rocket flowfield behavior from heat flux measurements

The mixing and heat transfer phenomena in small rocket flow fields with fuel film cooling is not well understood. An instrumented, water-cooled chamber with a gaseous hydrogen/gaseous oxygen injector was used to gather steady-state inner and outer wall temperature profiles. The chamber was tested at 414 kPa (60 psia) chamber pressure, from mixture ratios of 3.41 to 8.36. Sixty percent of the fuel was used for film cooling. These temperature profiles were used as boundary conditions in a finite element analysis program, MSC/NASTRAN, to calculate the local radial and axial heat fluxes in the chamber wall. The normal heat fluxes were then calculated and used as a diagnostic of the rocket's flow field behavior. The normal heat fluxes determined were on the order of 1.0 to 3.0 MW/meters squared (0.6 to 1.8 Btu/sec-inches squared). In the cases where mixture ratio was 5 or above, there was a sharp local heat flux maximum in the barrel section of the chamber. This local maximum seems to indicate a reduction or breakdown of the fuel film cooling layer, possibly due to increased mixing in the shear layer between the film and core flows. However, the flow was thought to be completely laminar, as the throat Reynolds numbers were below 50,000 for all the cases. The increased mixing in the shear layer in the higher mixture ratio cases appeared not to be due to the transition of the flow from laminar to turbulent, but rather due to increased reactions between the hydrogen film and oxidizer-rich core flows

Weight savings in aerospace vehicles through propellant scavenging

Vehicle payload benefits of scavenging hydrogen and oxygen propellants are addressed. The approach used is to select a vehicle and a mission and then select a scavenging system for detailed weight analysis. The Shuttle 2 vehicle on a Space Station rendezvous mission was chosen for study. The propellant scavenging system scavenges liquid hydrogen and liquid oxygen from the launch propulsion tankage during orbital maneuvers and stores them in well insulated liquid accumulators for use in a cryogenic auxiliary propulsion system. The fraction of auxiliary propulsion propellant which may be scavenged for propulsive purposes is estimated to be 45.1 percent. The auxiliary propulsion subsystem dry mass, including the proposed scavenging system, an additional 20 percent for secondary structure, an additional 5 percent for electrical service, a 10 percent weight growth margin, and 15.4 percent propellant reserves and residuals is estimated to be 6331 kg. This study shows that the fraction of the on-orbit vehicle mass required by the auxiliary propulsion system of this Shuttle 2 vehicle using this technology is estimated to be 12.0 percent compared to 19.9 percent for a vehicle with an earth-storable bipropellant system. This results in a vehicle with the capability of delivering an additional 7820 kg to the Space Station

Advanced APS Impacts on Vehicle Payloads

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethyl hydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination and scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible

Hydrogen/oxygen auxiliary propulsion technology

A survey is provided of hydrogen/oxygen (H/O) auxiliary propulsion system (APS) concepts and low thrust H/O rocket technology. A review of H/O APS studies performed for the Space Shuttle, Space Tug, Space Station Freedom, and Advanced Manned Launch System programs is given. The survey also includes a review of low thrust H/O rocket technology programs, covering liquid H/O and gaseous H/O thrusters, ranging from 6600 N (1500 lbf) to 440 mN (0.1 lbf) thrust. Ignition concepts for H/O thrusters and high temperature, oxidation resistant chamber materials are also reviewed

High-Performance Monopropellants and Catalysts Evaluated

The NASA Glenn Research Center is sponsoring efforts to develop advanced monopropellant technology. The focus has been on monopropellant formulations composed of an aqueous solution of hydroxylammonium nitrate (HAN) and a fuel component. HAN-based monopropellants do not have a toxic vapor and do not need the extraordinary procedures for storage, handling, and disposal required of hydrazine (N2H4). Generically, HAN-based monopropellants are denser and have lower freezing points than N2H4. The performance of HAN-based monopropellants depends on the selection of fuel, the HAN-to-fuel ratio, and the amount of water in the formulation. HAN-based monopropellants are not seen as a replacement for N2H4 per se, but rather as a propulsion option in their own right. For example, HAN-based monopropellants would prove beneficial to the orbit insertion of small, power-limited satellites because of this propellant's high performance (reduced system mass), high density (reduced system volume), and low freezing point (elimination of tank and line heaters). Under a Glenn-contracted effort, Aerojet Redmond Rocket Center conducted testing to provide the foundation for the development of monopropellant thrusters with an I(sub sp) goal of 250 sec. A modular, workhorse reactor (representative of a 1-lbf thruster) was used to evaluate HAN formulations with catalyst materials. Stoichiometric, oxygen-rich, and fuelrich formulations of HAN-methanol and HAN-tris(aminoethyl)amine trinitrate were tested to investigate the effects of stoichiometry on combustion behavior. Aerojet found that fuelrich formulations degrade the catalyst and reactor faster than oxygen-rich and stoichiometric formulations do. A HAN-methanol formulation with a theoretical Isp of 269 sec (designated HAN269MEO) was selected as the baseline. With a combustion efficiency of at least 93 percent demonstrated for HAN-based monopropellants, HAN269MEO will meet the I(sub sp) 250 sec goal

Advanced Monopropellant Thruster Technology Tested

A new family of environmentally friendly, low-freezing-point, high-density monopropellants is being developed under a NASA Glenn technology program. New monopropellant technology would greatly benefit a range of small (<100 kg) satellites and spacecraft missions. These monopropellants are mixtures of hydroxylammonium nitrate (HAN), fuel, and water. Primex Aerospace Company, under contract to the NASA Glenn Research Center at Lewis Field, tested a 1-lbf thruster using a HAN-based monopropellant formulation. Over 8000 sec of total test time was accumulated on a single thruster using the blowdown duty cycle typical of state-of-the-art monopropellant systems

Long Life Testing of Oxide-Coated Iridium/Rhenium Rockets

22-N class rockets, composed of a rhenium (Re) substrate, an iridium (Ir) coating, and an additional composite coating consisting of Ir and a ceramic oxide, were tested on gaseous oxygen/gaseous hydrogen (GO2/GH2) propellants. Two rockets were tested, one for nearly 39 hours at a nominal mixture ratio (MR) of 4.6 and chamber pressure (Pc) of 469 kPa, and the other for over 13 hours at a nominal MR of 5.8 and 621 kPa Pc. Four additional Ir/Re rockets, with a composite Ir-oxide coating fabricated using a modified process, were also tested, including one for 1.3 hours at a nominal MR of 16.7 and Pc of 503 kPa. The long lifetimes demonstrated on low MR GO2/GH2 suggest greatly extended chamber lifetimes (tens of hours) in the relatively low oxidizing combustion environments of Earth storable propellants. The oxide coatings could also serve as a protective coating in the near injector region, where a still-mixing flowfield may cause degradation of the Ir layer. Operation at MR close to 17 suggests that oxide-coated Ir/Re rockets could be used in severely oxidizing combustion environments, such as high MR GO2/GH2, oxygen/hydrocarbon, and liquid gun propellants

Evaluation of oxide-coated iridium-rhenium chambers

Iridium-coated rhenium (Ir-Re) provides long life operation of radiation-cooled rockets at temperatures up to 2200 C. Ceramic oxide coatings could be used to increase Ir-Re rocket lifetimes and allow operation in highly oxidizing environments. Ceramic oxide coatings promise to serve as both thermal and diffusion barriers for the iridium layer. Seven ceramic oxide-coated Ir-Re, 22-N rocket chambers were tested with gaseous hydrogen/gaseous oxygen (GHz/G02) propellants. Five chambers had thick (over 10 mils), monolithic coatings of either hafnia (HfO2) or zirconia (ZrO2). Two chambers had coatings with thicknesses less than 5 mils. One of these chambers had a thin-walled coating of ZrO2 infiltrated with sol gel HfO2. The other chamber had a coating composed of an Ir-oxide composite. The purpose of this test program was to assess the ability of the oxide coatings to withstand the thermal shock of combustion initiation, adhere under repeated thermal cycling, and operate in aggressively oxidizing environments. All of the coatings survived the thermal shock of combustion and demonstrated operation at mixture ratios up to 11. Testing the Ir-oxide composite-coated chamber included over 29 min at mixture ratio 16. The thicker walled coatings provided the larger temperature drops across the oxide layer (up to 570 C), but were susceptible to macrocracking and eventual chipping at a stress concentrator. The cracks apparently resealed during firing, under compression of the oxide layer. The thinner walled coatings did not experience the macrocracking and chipping of the chambers that was seen with the thick, monolithic coatings. However, burn-throughs in the throat region did occur in both of the thin-walled chambers at mixture ratios well above stoichiometric. The burn-throughs were probably the result of oxygen diffusion through the oxide coating that allowed the underlying Ir and Re layers to be oxidized. The results of this test program indicated that the thin-walled oxide coatings are better suited for repeated thermal cycling than the thick-walled coating, while thicker coatings may be required for operation in aggressively oxidizing environments

High-Temperature Oxidation Behavior of Iridium-Rhenium Alloys

The life-limiting mechanism for radiation-cooled rockets made from iridium-coated rhenium (Ir/Re) is the diffusion of Re into the Ir layer and the subsequent oxidation of the resulting Ir-Re alloy from the inner surface. In a previous study, a life model for Ir/Re rockets was developed. It incorporated Ir-Re diffusion and oxidation data to predict chamber lifetimes as a function of temperature and oxygen partial pressure. Oxidation testing at 1540 deg C suggested that a 20-wt percent Re concentration at the inner wall surface should be established as the failure criterion. The present study was performed to better define Ir-oxidation behavior as a function of Re concentration and to supplement the data base for the life model. Samples ranging from pure Ir to Ir-40 wt percent Re (Ir-40Re) were tested at 1500 deg C, in two different oxygen environments. There were indications that the oxidation rate of the Ir-Re alloy increased significantly when it went from a single-phase solid solution to a two-phase mixture, as was suggested in previous work. However, because of testing anomalies in this study, there were not enough dependable oxidation data to definitively raise the Ir/Re rocket failure criterion from 20-wt percent Re to a Re concentration corresponding to entry into the two-phase region

Experimental and analytical comparison of flowfields in a 110 N (25 lbf) H2/O2 rocket

A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket was examined through the RPLUS code using the full Navier-Stokes equations with finite rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses were performed for a range of mixture ratios and fuel film cooling pcts. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four pct. Thrust coefficients computed by the code are found to be within two pct. of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty