Real gas effects in the flow of methane and natural gas through critical flow nozzles
Equations for mass flow of methane and natural gas mixtures through critical flow nozzles, including real gas effects
Johnson, R. C.↗
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Real-gas effects in flow metering
Real gas mass flow rate computation through sonic nozzle
Johnson, R. C.↗
Calibration of sonic valves for the laminar flow control, leading-edge flight test
Sonic needle valves were calibrated to measure and control airflow in the suction system for the leading-edge flight test. The procedure and results for the calibration flow test of 4:41 flight valves are given. Mass-flow rates, which ranged from 0.001 to 0.012 lbm/sec, and maximum back pressure were measured for total temperatures from -30 F to 75 F and total pressures from 120 to 540 psf. Correlating equations are obtained for mass-flow rate as a function of total pressure, total temperature, and valve opening length. The most important aspect of flow measurement and control is found to be the measurement of valve opening length.
Petley, D. H.↗
Propellant-Flow-Actuated Rocket Engine Igniter
A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0.001 lb/s) and pressure drops on the order of 30 to 50 kilopascal (a few psi). An analytical model has been created and tested in conjunction with a precisely calibrated reference model. The analytical model accurately captures the overall behavior of this innovation. The model is a simple "volume-orifice" concept, with each chamber considered a single temperature and pressure "node" connected to adjacent nodes, or to vent paths through flow control orifices. Mass and energy balances are applied to each node, with gas flow predicted using simple compressible flow equations.
Wollen, Mark↗
Boundary-layer-ingesting inlet flow control system
A system for reducing distortion at the aerodynamic interface plane of a boundary-layer-ingesting inlet using a combination of active and passive flow control devices is disclosed. Active flow control jets and vortex generating vanes are used in combination to reduce distortion across a range of inlet operating conditions. Together, the vortex generating vanes can reduce most of the inlet distortion and the active flow control jets can be used at a significantly reduced control jet mass flow rate to make sure the inlet distortion stays low as the inlet mass flow rate varies. Overall inlet distortion, measured and described as average SAE circumferential distortion descriptor, was maintained at a value of 0.02 or less. Advantageous arrangements and orientations of the active flow control jets and the vortex generating vanes were developed using computational fluid dynamics simulations and wind tunnel experimentations.
Owens, Lewis R.↗
Aerodynamic Evaluation of Two Compact Radial-Inflow Turbine Rotors
The aerodynamic evaluation of two highly loaded compact radial turbine rotors was conducted at the NASA Lewis Research Center Small Engine Component Test Facility (SECTF). The experimental results were used for proof-of-concept, for modeling radial inflow turbine rotors, and for providing data for code verification. Two rotors were designed to have a shorter axial length, up to a 10-percent reduced diameter, a lighter weight, and equal or higher efficiencies with those of conventional radial inflow turbine rotors. Three configurations were tested: rotor 1, having a 40-percent shorter axial length, with the design stator (stator 1); rotor 1 with the design stator vanes closed down (stator 2); and rotor 2, slightly shorter axially and having higher loading, with stator 2. The stator had 36 vanes and the rotors each had 14 solid blades. Although presently uncooled, the rotor blades were designed for thicknesses which would allow cooling passages to be added. The overall stage performance measurements and the rotor and stator exit flow field surveys were obtained. Measurements of steady state temperatures, pressures, mass flow rates, flow angles, and output power were made at various operating conditions. Data were obtained at corrected speeds of 80, 90, 100, 110, and 120 percent of design over a range of equivalent inlet-to-exit pressure ratios of 3.5, 4.0, 4.5, and 5.0, the maximum pressure ratio achieved. The test showed that the configuration of rotor 1 with stator 1 running at the design pressure ratio produced a flow rate which was 5.6 percent higher than expected. This result indicated the need to close down the stator flow area to reduce the flow. The flow area reduction was accomplished by restaggering the vanes. Rotor 1 was retested with the closed-down stator vanes and achieved the correct mass flow. Rotor 2 was tested only with the restaggered vanes. The test results of the three turbine configurations were nearly identical. Although the measured efficiencies of the compact designs fell 2 to 3 points below the predicted efficiency of 91 percent, they did meet and exceed by up to 2.5 percentage points the efficiences of state-of-the-art turbines found in the literature.
Simonyi, P. Susan↗
Liquid Oxygen/Liquid Methane Test Summary of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility
Tests were conducted with the RS18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA s Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propellant systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to approx.120,000 ft (approx.37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. Propellant flow rate was measured using a coriolis-style mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup. LO2 flow ranged from 5.9-9.5 lbm/sec (2.7-4.3 kg/sec), and LCH4 flow varied from 3.0-4.4 lbm/sec (1.4-2.0 kg/sec) during the RS-18 hot-fire test series. Thrust was measured using three load cells in parallel. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Data was obtained at multiple chamber pressures, and calculations were performed for specific impulse, C* combustion efficiency, and thrust vector alignment. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes.
Melcher, John C., IV↗
Development of sensor to measure hot gas leaks from flanges and seals in plumbing of saturn vehicle summary report, 30 jun. 1963 - 30 jun. 1964
Transducer to measure hot gas leak-mass flow leakage
LEAKAGE↗
Experimental study of cross-stream mixing in a cylindrical duct
An experimental investigation of cross stream injection and mixing was conducted with application to a low NO sub x combustor for the High Speed Civil Transport (HSCT). Mixing in a cylindrical chamber was studied for transverse injection from slanted slot and round orifice injectors. Momentum ratio, density ratio, and number were studied. Quantitative measurement of injectant concentration distributions were obtained by planar digital imaging of the Mie scattered light from an aerosol seed uniformly mixed with the injectant. The unmixedness, defined as the ratio of the r.m.s. concentration fluctuation to mean concentration in a plane perpendicular to the main flow direction, was found to be primarily a function of momentum ratio and injector spacing. An optimum spacing is indicated. Unmixedness is also a function of orifice size, or mass flow ratio, but the mass flow dependence can be accounted for by normalizing the unmixedness with its maximum theoretical value. The data indicate that a density ratio greater than unity retards mixing. It was found that above a certain momentum flux ratio, mixing with slanted slot injectors was better than with round hole injectors.
Vranos, A.↗
Experimental study of cross-stream mixing in a cylindrical duct
An experimental investigation of cross stream injection and mixing has been conducted with application to a low NO(x) combustor for the HSCT. Mixing in a cylindrical chamber has been studied for transverse injection from slanted slot and round orifice injectors. Momentum ratio, density ratio, and injector geometry were the primary variables. Slanted slots of various size, aspect ratio, and number were studied. Quantitative measurement of injectant concentration distributions were obtained by planar digital imaging of the Mie-scattered light from an aerosol seed uniformly mixed with the injectant. The unmixedness, defined as the ratio of the rms concentration fluctuation to mean concentration in a plane perpendicular to the main flow direction, was found to be primarily a function of momentum ratio and injector spacing. An optimum spacing is indicated. Unmixedness is also a function of orifice size, or mass flow ratio, but the mass flow dependence can be accounted for by normalizing the unmixedness with its maximum theoretical value. The data indicate that density ratio greater than unity retards mixing. It was found that above a certain momentum flux ratio, mixing with slanted slot injectors was better than with round hole injectors.
Vranos, A.↗
Tables of critical-flow functions and thermodynamic properties for methane and computational procedures for both methane and natural gas
Procedures for calculating the mass flow rate of methane and natural gas through nozzles are given, along with the FORTRAN 4 subroutines used to make these calculations. Three sets of independent variables are permitted in these routines. In addition to the plenum pressure and temperature, the third independent variable is either nozzle exit pressure, Mach number, or temperature. A critical-flow factor that becomes a convenient means for determining the mass flow rate of methane through critical-flow nozzles is tabulated. Other tables are included for nozzle throat velocity and critical pressure, density, and temperature ratios, along with some thermodynamic properties of methane, including compressibility factor, enthalpy, entropy, specific heat, specific-heat ratio, and speed of sound. These tabulations cover a temperature range from 120 to 600 K and pressures to 3 million N/sq m.
Johnson, R. C.↗
Wind tunnel blockage tests at Mach 5 of vacuum duct models for two sound radiation shields
Two sound shield models with dummy vacuum exhaust ducts were tested in a Mach 5 pilot quiet tunnel. The first model simulates a new sound shield of 3 in. (7.62 cm) inside diameter and the second model is a shield of 4 in. (10.16 cm) inside diameter. The dummy vacuum exhaust ducts were attached to the external housing of the models. The flow in the first model, which had a by pass mass flow ratio of about 0.6, could not be started except at the two highest test Reynolds numbers where only the central core flow region was started. The flow in the second model with a mass ratio of approximately 0.3 was fully started except at the lowest unit Reynolds number where some unsteadiness and partial flow separation at the wall was observed. Since the external housing and dummy vacuum ducts were the same for both models, these results indicate that the ratio of by pass mass flow to total mass flow for a wind tunnel sound shield of this particular design must be less than about 0.3. Hence, a lower limit is imposed on the inlet diameter of the sound shield in relation to the exit diameter of the wind tunnel nozzle. This lower limit on the inlet diameter may possibly be reduced by improvements in streamlining of the external housing and ducts, by reductions in blockage area, or by the use of external ducting shrouds.
Beckwith, I. E.↗
Application of Sweeping Jet Actuators on the NASA Hump Model and Comparison with CFDVAL2004 Experiments
Flow separation control over a wall-mounted hump model was studied experimentally to assess the performance of sweeping jet actuators. Results were compared to that of the 2004 CFD validation experiment (CFDVAL2004), which examined flow separation control with steady suction and unsteady zero-net-mass-flow actuators. Comparisons were carried out at low and high amplitude excitations. In addition to the active flow control methods, a passive flow control method (i.e., vortex generator) was used to complement the dataset. Steady/unsteady surface pressure measurements and surface oilflow visualization were used in the performance assessment of the actuators. The results indicated that the sweeping jet actuators are more effective than the steady suction and unsteady zero-net-mass-flow actuators. For the same momentum coefficient, the sweeping jet actuators produced more flow acceleration upstream of separation, more pressure recovery downstream, and consistently a smaller separation bubble.
Koklu, Mehti↗
Cold-air investigation of a turbine for high-temperature-engine application. 4: Two-stage turbine performance
The efficiency of a two-stage turbine is discussed. The turbine efficiency was 0.932 for equivalent design operating conditions (speed) and specific work, which compares closely to the value of 0.929 that would be estimated using the first-stage efficiency. The mass flow obtained with the two-stage configuration indicated that the mass flow characteristics of the two stages were closely matched at design operating conditions. The stage work split at these conditions was 0.505-0.495, which was close to the design work split of 0.515-0.485.
Whitney, W. J.↗
Modifications to the nozzle test chamber to extend nozzle static-test capability
The nozzle test chamber was modified to provide a high-pressure-ratio nozzle static-test capability. Experiments were conducted to determine the range of the ratio of nozzle total pressure to chamber pressure and to make direct nozzle thrust measurements using a three-component strain-gage force balance. Pressure ratios from 3 to 285 were measured with several axisymmetric nozzles at a nozzle total pressure of 15 to 190 psia. Devices for measuring system mass flow were calibrated using standard axisymmetric convergent choked nozzles. System mass-flow rates up to 10 lbm/sec are measured. The measured thrust results of these nozzles are in good agreement with one-dimensional theoretical predictions for convergent nozzles.
Keyes, J. W.↗
Medium power hydrogen arcjet performance
An experimental investigation was performed to evaluate hydrogen arcjet operating characteristics in the range of 1 to 4 kW. A series of nozzles were operated in modular laboratory thrusters to examine the effects of geometric parameters such as constrictor diameter and nozzle divergence angle. Each nozzle was tested over a range of current and mass flow rates to explore stability and performance. In the range of mass flow rates and power levels tested, specific impulse values between 650 and 1250 sec were obtained at efficiencies between 30 and 40 percent. The performance of the two larger half angle (20, 15 deg) nozzles was similar for each of the two constrictor diameters tested. The nozzles with the smallest half angle (10 deg) were difiicult to operate. A restrike mode of operation was identified and described. Damage in the form of melting was observed in the constrictor region of all the nozzle inserts tested. Arcjet ignition was also difficult in many tests and a glow discharge mode that prevents starting was identified.
Curran, Francis M.↗
Medium power hydrogen arcjet performance
An experimental investigation was performed to evaluate hydrogen arcjet operating characteristics in the range of 1 to 4 kW. A series of nozzles were operated in modular laboratory thrusters to examine the effects of geometric parameters such as constrictor diameter and nozzle divergence angle. Each nozzle was tested over a range of current and mass flow rates to explore stability and performance. In the range of mass flow rates and power levels tested, specific impulse values between 650 and 1250 sec were obtained at efficiencies between 30 and 40 percent. The performance of the two larger half angle (20, 15 deg) nozzles was similar for each of the two constrictor diameters tested. The nozzles with the smallest half angle (10 deg) were difficult to operate. A restrike mode of operation was identified and described. Damage in the form of melting was observed in the constrictor region of all the nozzle inserts tested. Arcjet ignition was also difficult in many tests and a glow discharge mode that prevents starting was identified.
Curran, Francis M.↗
Compressor Stall Recovery Through Tip Injection Assessed
Aerodynamic stability is a fundamental limit in the compressor design process. The development of robust techniques for increasing stability has several benefits: enabling higher loading and fewer blades, increasing safety throughout a mission, increasing tolerance to stage mismatch during part-speed operation and speed transients, and providing an opportunity to match stages at the compressor maximum efficiency point, thus reducing fuel burn. Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors if the injection is activated before stall occurs. This approach to stall suppression requires that a reliable stall warning system be available. Tests have recently been performed to assess whether steady injection can also be used to recover from fully developed stall. If mass injection is effective in recovering from stall quickly enough to avoid structural damage or loss of engine power, then a stall warning system may not be required. The stall recovery tests were performed on a transonic compressor rotor at its design tip speed of 1475 ft/sec using four injectors evenly spaced around the compressor case upstream of the rotor. The injectors were connected to an external air source. In an actual engine application, the injected air would be supplied with compressor bleed air. The injectors were isolated from the air source by a fast-acting butterfly valve. With the injectors turned off, the compressor was throttled into stall. Air injection was then activated with no change in throttle setting by opening the butterfly valve. The compressor recovered from stall at a fixed throttle setting with the aid of tip injection. The unsteady operating characteristic of the rotor was measured during these tests using high-response pressure sensors located upstream and downstream of the rotor. The figure shows the results, where the unsteady pressure and mass flow are superimposed on the steady operating characteristic. The total injected mass flow was equal to 1.3 percent of the compressor flow. The solid line with no solid squares on it denotes the operating point during the beginning of throttle closure and the initial drop into stall. The gray traces denote the operating point during an additional throttle closure that occurred over the next 1200 rotor revolutions (4 sec). The dashed line denotes the recovery from stall that occurred during 90 rotor revolutions (0.3 sec) after the injectors were activated with no change in throttle setting. Tip injection not only recovers the compressor from stall, but also restores the compressor to its pre-stall level of pressure rise. In contrast, standard stall recovery schemes such as compressor bleed, stator vane actuation, or engine throttle modulation result in a loss of pressure rise across the compressor, which results in a loss of engine power.
Suder, Ken L.↗