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Tests were made to determine the effect of fillets between wings and fuselage on the drag and propulsive efficiency of a high-wing cabin monoplane. These tests were made in the 20-foot Propeller Research Tunnel of the National Advisory Committee for Aeronautics.
An investigation was conducted in the NACA 20-foot wind tunnel of the ramming effect of three general types of carburetor intake rams for radial engines, namely, the internal constant area type, the external constant area type, and the external expanding type. The rams were installed on a radial air- cooled engine nacelle, and tests were made with and without the propeller operating. The results indicated that the external types having entrances near the front of the engine cowling gave the greatest ramming effect. The propeller increased the ramming effect for the external types. From considerations of the ramming effect, the best entrance location for the external types was close to the nose of the engine cowling. For the internal type, the best location was in a plane perpendicular to the propeller shaft and immediately forward of the engine cylinders.
This report presents the results of tests made in the 7-by 10-foot wind tunnel and in the 20-foot tunnel of the National Advisory Committee for Aeronautics to determine the drag of a number of airplane wheels, wheel fairings, and landing gears designed or selected for an airplane of 3,000 pounds gross weight. All tests were made on full-size models; those in the 7-by 10-foot tunnel were made at air speeds up to 80 miles per hour and those in the 20-foot tunnel were made at air speeds up to 100 miles per hour. Although most of the landing-gear tests were made in conjunction with a fuselage and at 0 degree pitch angle, some of the tests were made in conjunction with fuselage plus wings and a radial air-cooled engine and at pitch angles from -5 degrees to 6 degrees to obtain an indication of the general effect of these various items on landing-gear drag. All tests were made in the absence of propeller slipstream.
While NASA's current next-generation launch vehicle research has largely focused on advanced all-rocket single-stage-to-orbit vehicles (i.e. the X-33 and it's RLV operational follow-on), some attention is being given to advanced propulsion concepts suitable for 'next-generation-and-a-half' vehicles. Rocket-based combined-cycle (RBCC) engines combining rocket and airbreathing elements are one candidate concept. Preliminary RBCC engine development was undertaken by the United States in the 1960's. However, additional ground and flight research is required to bring the engine to technological maturity. This paper presents two options for flight testing early versions of the RBCC ejector scramjet engine. The first option mounts a single RBCC engine module to the X-34 air-launched technology testbed for test flights up to about Mach 6.4. The second option links RBCC engine testing to the simultaneous development of a small-payload (220 lb.) two-stage-to-orbit operational vehicle in the Bantam payload class. This launcher/testbed concept has been dubbed the W vehicle. The W vehicle can also serve as an early ejector ramjet RBCC launcher (albeit at a lower payload). To complement current RBCC ground testing efforts, both flight test engines will use earth-storable propellants for their RBCC rocket primaries and hydrocarbon fuel for their airbreathing modes. Performance and vehicle sizing results are presented for both options.
A status of MEMS-based micropropulsion activities conducted at JPL will be given. These activities include work conducted on the so called Vaporizing Liquid Micro-Thruster (VLM) which recently underwent proof-of-concept testing, demonstrating the ability to vaporize water propellant at 2 W and 2 V. Micro-ion engine technologies, such m field emitter arrays and micro-grids are being studied. Focus in the field emitter area is on arrays able to survive in thruster plumes and micro-ion engine plasmas to serve as neutralizers aW engine cathodes. Integrated, batch-fabricated Ion repeller grid structures are being studied as well as different emitter tip materials are being investigated to meet these goals. A micro-isolation valve is being studied to isolate microspacecraft feed system during long interplanetary cruises, avoiding leakage and prolonging lifetime and reliability of such systems. This concept relies on the melting of a thin silicon barrier. Burst pressure values as high as 2,900 psig were obtained for these valves and power requirements to melt barriers ranging between 10 - 50 microns in thickness, as determined through thermal finite element calculations, varied between 10 - 30 W to be applied over a duration of merely 0.5 ms.
We will report on the test apparatus, calibrations of the system, the test procedure used and results for 13, 20 and 30-segment isolators.
The Fluid Acquisition and Resupply Experiment (FARE) program, managed by the Marshall Space Flight Center Space Propulsion Branch with Martin Marietta Civil Space and Communications as the contractor, consisted of two flights designated FARE I and FARE II. FARE I flew in December 1992 on STS-53 with a screen channel liquid acquisition device (LAD) and FARE II flew in June 1993 on STS-57 with a vane-type LAD. Thus, the FARE I and II flights represent the two basic LAD categories usually considered for in-space fluid management. Although both LAD types have been used extensively, the usefulness of the on-orbit data has been constrained by the lack of experimentation beyond predicted performance limits, including both propellant fill and expulsion. Therefore, the FARE tests were designed to obtain data that would satisfy two primary objectives: (1) Demonstrate the performance of the two types of LADs, screen channel and vane, and (2) support the anchoring of analytical models. Both flights were considered highly successful in meeting these two primary objectives.
Data are presented indicating the rates and magnitudes of energy released by the liquid-phase reactions of various propellant combinations. The data show that this energy release can contribute significantly to the rate of vaporization of the incoming propellants and thus aid the combustion process. Nevertheless, very low performances were obtained in rocket motors with conventional impinging-jet injectors when highly reactive systems such as N104-N2H4, were employed. A possible explanation for this low performance is that the initial reactions of such systems are so rapid that liquid-phase mixing is inhibited. Evidence for such an effect is presented in a series of color photographs of open flames using various injector elements. Based on these studies, some requirements are suggested for injector elements using highly reactive propellants. Experimental results are presented of motor tests using injector elements in which some of these requirements are met through the use of a set of concentric tubes. These tests, carried out at thrust levels of 40 to 800 lb per element, demonstrated combustion efficiencies of up to 98% based on equilibrium characteristic velocity values. Results are also presented for tests made with impinging-jet and splash-plate injectors for comparison.
This paper addresses the problem of building trust in online predictions of a battery powered aircraft's remaining available flying time. A set of ground tests is described that make use of a small unmanned aerial vehicle to verify the performance of remaining flying time predictions. The algorithm verification procedure described here uses a fully functional vehicle that is restrained to a platform for repeated run-to-functional-failure experiments. The vehicle under test is commanded to follow a predefined propeller RPM profile in order to create battery demand profiles similar to those expected in flight. The fully integrated aircraft is repeatedly operated until the charge stored in powertrain batteries falls below a specified lower-limit. The time at which the lower-limit on battery charge is crossed is then used to measure the accuracy of remaining flying time predictions. Accuracy requirements are considered in this paper for an alarm that warns operators when remaining flying time is estimated to fall below a specified threshold.
Traditional probabilistic risk assessment approaches often require failure scenarios to be explicitly defined through event sequences that are then quantified as part of the integrated analysis. This approach becomes difficult when failure propagation paths change as a function of the system operation. Additionally, if the propagation paths represent interactions among even a modest number of components, the scenario count becomes combinatorially intractable. This paper presents an alternate approach for quantifying the probability of failure propagation in such a case. Rather than explicitly defining scenario sequences, simple physical models are created for each of the components. In this way, only the physical states and rules of component interactions must be defined, rather than event sequences for each individual scenario. Initiating failures are introduced into the system, either randomly or as defined by relative likelihood, and the failures cascade through the system via the interaction rules. This process is repeated using Monte Carlo methods and, as a result, the most probable scenarios “self-evolve” in terms of both sequence path and frequency. This approach was applied to failures occurring in the engine compartment of a space launch vehicle with four liquid rocket engines and four high-pressure helium tanks. Each engine was modeled with key components, such as turbomachinery, combustion chamber, propellant lines, and additional support systems. Three test cases were conducted with different high-energy engine failures. End results of interest included an additional engine-out failure and tank burst, which represent the loss-of-mission (LOM) and loss-of-crew (LOC) failure environments, respectively. Observations show that almost every scenario outcome is unique and that many scenarios involve complex chain reactions that are difficult to predict. This validates the usefulness of the modeling approach in assessing the overall risks to the crew during a launch vehicle abort.
This paper addresses the problem of building trust in the online prediction of a eUAV’s remaining available flying time powered by lithium-ion polymer batteries. A series of ground tests are described that make use of an electric unmanned aerial vehicle (eUAV) to verify the performance of remaining flying time predictions. The algorithm verification procedure described is implemented on a fully functional vehicle that is restrained to a platform for repeated run-to-functional-failure (charge depletion) experiments. The vehicle under test is commanded to follow a predefined propeller RPM profile in order to create battery demand profiles similar to those expected during flight. The eUAV is repeatedly operated until the charge stored in powertrain batteries falls below a specified limit threshold. The time at which the limit threshold on battery charge is crossed is then used to measure the accuracy of the remaining flying time prediction. In our earlier work battery aging was not included. In this work we take into account aging of the batteries where the parameters were updated to make predictions. Accuracy requirements are considered for an alarm that warns operators when remaining flying time is estimated to fall below the specified limit threshold.
Technology development to mature a hybrid rocket option for a potential Mars Ascent Vehicle has been underway at the Jet Propulsion Laboratory and Marshall Space Flight Center. A conceptual design was completed in 2016 and has been used to determine the motor size for risk reduction development. The task plan for the FY17 testing, actually completed in mid FY18, is discussed. The major achievement of this effort was successful full scale hotfire testing with a representative ambient Earth storable propellant combination: SP7/MON3. A near term plan, through July 2019 is briefly discussed.
Three Hg propellant isolators (two cathode and one main) failed during testing in the Mission Profile Life Test. These failures involved contamination of the surface of the alumina insulating body which resulted in heating of the vaporizer by leakage current from the high voltage supply, with subsequent loss of propellant flow rate control. Failure analysis of the isolators showed the surface resistance was temperature dependent and that the alumina could be restored to its original insulating state by grit blasting the surface. The contaminant was identified as carbon and the most likely sources identified as ambient facility hydrocarbons, directed back-sputtered facility materials, and outgassing from organic insulating materials within the thruster envelope. Methods to eliminate contamination from each of these sources are described.
The product of incomplete reaction of monomethylhydrazine (MMH) and nitrogen tetroxide (NTO) propellants, or fuel-oxidizer reaction product (FORP), has been hypothesized as a contributory cause of an anomaly which occurred in the chamber pressure (PC) transducer tube on the Reaction Control Subsystem (RCS) aft thruster 467 on flight STS-51. A small hole was found in the titanium-alloy PC tube at the first bend below the pressure transducer. It was surmised that the hole may have been caused by heat and pressure resulting from ignition of FORP. The NASA Johnson Space Center (JSC) White Sands Test Facility (WSTF) was requested to define the chemical characteristics of FORP, characterize its reactivity, and simulate the events in a controlled environment which may have lead to the Pc-tube failure. Samples of FORP were obtained from the gas-phase reaction of MMH with NTO under laboratory conditions, the pulsed firings of RCS thrusters with modified PC tubes using varied oxidizer or fuel lead times, and the nominal RCS thruster firings at WSTF and Kaiser-Marquardt. Fourier transform infrared spectroscopy (FTIR), differential scanning calorimetry (DSC), accelerating rate calorimetry (ARC), ion chromatography (IC), inductively coupled plasma (ICP) spectrometry, thermogravimetric analysis (TGA) coupled to FTIR (TGA/FTIR), and mechanical impact testing were used to qualitatively and quantitatively characterize the chemical, thermal, and ignition properties of FORP. These studies showed that the composition of FORP is variable but falls within a limited range of compositions that depends on the fuel loxidizer ratio at the time of formation, composition of the post-formation atmosphere (reducing or oxidizing), and reaction or postreaction temperature. A typical composition contains methylhydrazinium nitrate (MMHN), ammonium nitrate (AN), methylammonium nitrate (MAN), and trace amounts of hydrazinium nitrate and 1,1-dimethylhydrazinium nitrate. The thermal decomposition reactions of FORP compositions used in this study were unremarkable, Neither the various compositions of FORP, the pure major components of FORP, nor mixtures of FORP with propellant-system corrosion products showed any unusual thermal activity when decomposed under laboratory conditions. Off-limit thruster operations were simulated by rapid mixing of liquid MMH and liquid NTO in a confined space. The test hardware was constructed with pressure- and temperature-measurement devices to determine if the expected fuel oxidizer reaction would result in increased energy release when FORP, FORP constituents, or propellant-system corrosion products were present. These tests demonstrated that FORP, MMHN, AN, or Inconel corrosion products can induce a mixture of MMH and NTO to produce component-damaging energies. The simulation-test program was not extensive enough to provide statistical probabilities for these events but did show that such events can occur. Damaging events required FORP or metal salts to be present at the initial mixing of MMH and NTO. Based on the results of these studies, it is suggested that removal or mitigation of a buildup of these materials may decrease the incidence of these high-energy, potentially damaging events.
Report presents the results of tests of a model of a single-engine airplane with two different tilts of the propeller axis. The results indicate that on a typical design a 5 degree downward tilt of the propeller axis will considerably reduce the destabilization effects of power. A comparison of the experimental results with those computed by use of existing theory is included. A comparison of the experimental results with those computed by use of existing theory is included. It is shown that the results can be predicted with an accuracy acceptable for preliminary design purposes, particularly at the higher powers where the effects are of significant magnitude.
It is shown that a time marching Navier-Stokes code called PARC can be utilized to provide a reasonable prediction of the flow field within an inlet for an advanced ducted propeller. The code validation was implemented for a nonseparated flow condition associated with the inlet functioning at angles-of-attack of zero and 25 deg. Comparison of the computational results with the test data shows that the PARC code with the propeller face fixed flow properties boundary conditions (BC) provided a better prediction of the inlet surface static pressures than the prediction when the mass flow BC was employed.
A lightweight hydrogen-oxygen alkaline fuel cell incorporated into the design of a lightweight fuel cell powerplant (LFCP) was analytically and experimentally developed. The powerplant operates with passive water removal which contributes to a lower system weight and extended operating life. A preliminary LFCP specification and design table were developed along with a lightweight power section for the LFCP design, consisting of repeating two-cell modules was designed. Two, four-cell modules were designed incorporating 0.508 sq ft active area space shuttle technology fuel cells. Over 1,200 hours of single-cell and over 8,800 hours of two-cell module testing was completed. The 0.25 sq ft active area lightweight cell design was shown to be capable of operating on propellant purity reactants out to a current density of 600ASF. Endurance testing of the two-cell module configuration exceeded the 2,500-hour LFCP voltage requirements out to 3700-hours. A two-cell module capable of operating at increased reactant pressure completed 1000 hours of operation at a 30 psia reactant pressure. A lightweight power section consisting of fifteen, two-cell modules connected electrically in series was fabricated.
The role of solid phase heterogeneity on the low-pressure L-star instability of nonmetallized AP/PBAN propellants is explored. Four particle size distributions are employed in propellants that are otherwise identical. Over one hundred test firings were conducted in the 21/2 in. diameter L-star burner. Pressure time histories in the chamber and color movies of two firings constitute the raw data. An economical firing program was used which enables the interesting range of L-star values to be covered during a single firing (at a set mean pressure), through the variations in the depleting propellant volume. Time-independent combustion, Helmholtz mode, chuff mode, and the pressure-burst phenomena are revealed as the principal signatures. Of these, the Helmholtz mode is found to be the most ordered form of instability.