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NASA NTRS2016

Correlation of Normal Gravity Mixed Convection Blowoff Limits with Microgravity Forced Flow Blowoff Limits

The axisymmetric rod geometry in upward axial stagnation flow provides a simple way to measure normal gravity blowoff limits to compare with microgravity Burning and Suppression of Solids - II (BASS-II) results recently obtained aboard the International Space Station. This testing utilized the same BASS-II concurrent rod geometry, but with the addition of normal gravity buoyant flow. Cast polymethylmethacrylate (PMMA) rods of diameters ranging from 0.635 cm to 3.81 cm were burned at oxygen concentrations ranging from 14 to 18% by volume. The forced flow velocity where blowoff occurred was determined for each rod size and oxygen concentration. These blowoff limits compare favorably with the BASS-II results when the buoyant stretch is included and the flow is corrected by considering the blockage factor of the fuel. From these results, the normal gravity blowoff boundary for this axisymmetric rod geometry is determined to be linear, with oxygen concentration directly proportional to flow speed. We describe a new normal gravity 'upward flame spread test' method which extrapolates the linear blowoff boundary to the zero stretch limit in order to resolve microgravity flammability limits-something current methods cannot do. This new test method can improve spacecraft fire safety for future exploration missions by providing a tractable way to obtain good estimates of material flammability in low gravity.

microgravity↗

NASA NTRS2017

Mixed Convection Blowoff Limits as a Function of Oxygen Concentration and Upward Forced Stretch Rate for Burning Pmma Rods of Various Sizes

Normal gravity flame blowoff limits in an axisymmetric pmma rod geometry in upward axial stagnation flow are compared with microgravity Burning and Suppression of Solids II (BASS-II) results recently obtained aboard the International Space Station. This testing utilized the same BASS-II concurrent rod geometry, but with the addition of normal gravity buoyant flow. Cast polymethylmethacrylate (pmma) rods of diameters ranging from 0.635 cm to 3.81 cm were burned at oxygen concentrations ranging from 14 to 18 by volume. The forced flow velocity where blowoff occurred was determined for each rod size and oxygen concentration. These blowoff limits compare favorably with the BASS-II results when the buoyant stretch is included and the flow is corrected by considering the blockage factor of the fuel. From these results, the normal gravity blowoff boundary for this axisymmetric rod geometry is determined to be linear, with oxygen concentration directly proportional to flow speed. We describe a new normal gravity upward flame spread test method which extrapolates the linear blowoff boundary to the zero stretch limit to resolve microgravity flammability limits, something current methods cannot do. This new test method can improve spacecraft fire safety for future exploration missions by providing a tractable way to obtain good estimates of material flammability in low gravity.

blowoff↗

NASA NTRS1944

Study of Jet-Propulsion System Comprising Blower, Burner, and Nozzle

A study was made of the performance of a jet-propulsion system composed of an engine-driven blower, a combustion chamber, and a discharge nozzle. A simplified analysis is made of this system for the purpose of showing in concise form the effect of the important design variables and operating conditions on jet thrust, thrust horsepower, and fuel consumption. Curves are presented that permit a rapid evaluation of the performance of this system for a range of operating conditions. The performance for an illustrative case of a power plant of the type under consideration id discussed in detail. It is shown that for a given airplane velocity the jet thrust horsepower depends mainly on the blower power and the amount of fuel burned in the jet; the higher the thrust horsepower is for a given blower power, the higher the fuel consumption per thrust horsepower. Within limits the amount of air pumped has only a secondary effect on the thrust horsepower and efficiency. A lower limit on air flow for a given fuel flow occurs where the combustion-chamber temperature becomes excessive on the basis of the strength of the structure. As the air-flow rate is increased, an upper limit is reached where, for a given blower power, fuel-flow rate, and combustion-chamber size, further increase in air flow causes a decrease in power and efficiency. This decrease in power is caused by excessive velocity through the combustion chamber, attended by an excessive pressure drop caused by momentum changes occurring during combustion.

FUEL CONSUMPTION↗

NASA NTRS1988

LANDER program manual: A lunar ascent and descent simulation

LANDER is a computer program used to predict the trajectory and flight performance of a spacecraft ascending or descending between a low lunar orbit of 15 to 500 nautical miles (nm) and the lunar surface. It is a three degree-of-freedom simulation which is used to analyze the translational motion of the vehicle during descent. Attitude dynamics and rotational motion are not considered. The program can be used to simulate either an ascent from the Moon or a descent to the Moon. For an ascent, the spacecraft is initialized at the lunar surface and accelerates vertically away from the ground at full thrust. When the local velocity becomes 30 ft/s, the vehicle turns downrange with a pitch-over maneuver and proceeds to fly a gravity turn until Main Engine Cutoff (MECO). The spacecraft then coasts until it reaches the requested holding orbit where it performs an orbital insertion burn. During a descent simulation, the lander begins in the holding orbit and performs a deorbit burn. It then coasts to pericynthion, where it reignites its engines and begins a gravity turn descent. When the local horizontal velocity becomes zero, the lander pitches up to a vertical orientation and begins to hover in search of a landing site. The lander hovers for a period of time specified by the user, and then lands.

Source record↗

NASA NTRS2000

Burning Plastics Investigated in Space for Unique US/Russian Cooperative Project

It is well known that fires in the low-gravity environment of Earth-orbiting spacecraft are different from fires on Earth. The flames lack the familiar upward plume, which is the result of gravitational buoyancy. These flames, however, are strongly influenced by minor airflow currents. A recent study conducted in low gravity (microgravity) on the Russian orbital station Mir used burning plastic rods mounted in a small chamber with a controllable fan to expose the flame to airflows of different velocities. In this unique project, a Russian scientific agency, the Keldysh Research Center, furnished the apparatus and directed the Mir tests, while the NASA Glenn Research Center at Lewis Field provided the test materials and the project management. Reference testing and calibrations in ground laboratories were conducted jointly by researchers at Keldysh and at the NASA Johnson Space Center's White Sands Test Facility. Multiple samples of three different plastics were burned in the tests: Delrin, a common material for valve bodies; PMMA, a plastic "glass"; and polyethylene, a familiar material for containers and films. Each burned with a unique spherical or egg-shaped flame that spread over the rod. The effect of varying the airflow was dramatic. At the highest airflow attainable in the combustion chamber, nearly 10 cm/sec (a typical ventilation breeze), the flames were bright and strong. As airflow velocity decreased, the flames became shorter but wider. In addition, the flames became less bright, and for PMMA and polyethylene, they showed two colors, a bright part decreasing in volume and a nearly invisible remainder (see the photographs). Finally, at a very low velocity, the flames extinguished. For the plastics tested, this minimum velocity was very low, around 0.3 to 0.5 cm/sec. This finding confirms that at least a slight airflow is required to maintain a flame in microgravity for these types of materials.

Friedman, Robert↗

NASA NTRS2003

Radiative Heat Loss Measurements During Microgravity Droplet Combustion in a Slow Convective Flow

Radiative heat loss from burning droplets in a slow convective flow under microgravity conditions is measured using a broad-band (0.6 to 40 microns) radiometer. In addition, backlit images of the droplet as well as color images of the flame were obtained using CCD cameras to estimate the burning rates and the flame dimensions, respectively. Tests were carried out in air at atmospheric pressure using n-heptane and methanol fuels with imposed forced flow velocities varied from 0 to 10 centimeters per second and initial droplet diameters varied from 1 to 3 millimeters. Slow convective flows were generated using three different experimental configurations in three different facilities in preparation for the proposed International Space Station droplet experiments. In the 2.2 Second Drop-Tower Facility a droplet supported on the leading edge of a quartz fiber is placed within a flow tunnel supplied by compressed air. In the Zero-Gravity Facility (five-second drop tower) a tethered droplet is translated in a quiescent ambient atmosphere to establish a uniform flow field around the droplet. In the KC 135 aircraft an electric fan was used to draw a uniform flow past a tethered droplet. Experimental results show that the burn rate increases and the overall flame size decreases with increases in forced-flow velocities over the range of flow velocities and droplet sizes tested. The total radiative heat loss rate, Q(sub r), decreases as the imposed flow velocity increases with the spherically symmetric combustion having the highest values. These observations are in contrast to the trends observed for gas-jet flames in microgravity, but consistent with the observations during flame spread over solid fuels where the burning rate is coupled to the forced flow as here.

Hicks, Michael C.↗

NASA NTRS2018

Preliminary Assessment of the Use of Heavy Gases in Two-Stage Light Gas Guns

This paper discusses ballistic range testing at muzzle velocities of 0.7 to 2.7 km/s. Herein, we review several techniques for achieving these velocities - starting with the single stage gas gun and the powder gun. These techniques can have velocity limitations, both on the high end and on the low end, very high powder breech pressures, incomplete and inconsistent powder burn and can send unburned powder grains downrange to confound impact data. To try to resolve these issues, it was decided to study the use of the two stage light gas gun operated with a heavier working gas than is normally used. This would have the effect of lowering the usual muzzle velocity range of the two-stage gun using hydrogen to cover the desired low velocity range. Preliminary results are presented from firings with two NASA guns. Twenty-five shots were made with helium and nine shots with argon. Muzzle velocities of 1.1 to 4 km/s were obtained with helium and velocities of 0.7 to 2.7 km/s were obtained with argon. Overall, the heavy gas technique appears quite promising, but more firings are needed to fill out the database.

Bogdanoff, David W.↗

NASA NTRS2018

Preliminary Assessment of the Use of Heavy Gases in Two-Stage Light Gas Guns

Bogdanoff, David W.↗

NASA NTRS1968

Discussion of paper by Antonio Alfonso-Faus, ''Solar wind and planetary rotation.''

Solar wind mechanism for origin of planetary rotation, discussing previous calculations of angular velocity and couple acting on planet

Alfonso-Faus, A.↗

NASA NTRS1964

Alumina Size Distributions from High-Pressure Composite Solid-Propellant Combustion

Alumina size distributions were obtained for both a coarse and a fine oxidizer composite propellant burning in a nitrogen atmosphere over the pressure range from atmospheric to 500 pounds per square inch. The amount of additive agglomeration was found to be significantly higher for the coarse oxidizer propellant and decreased with increasing pressure to the 0.3 power over the range from atmospheric pressure to 250 pounds per square inch. High-speed photographs of the burning propellant surface revealed that the additives moved on the surface with the average particle velocity decreasing with pressure to approximately the 0.3 power over the pressure range from atmospheric to 50 pounds per square inch. The empirical relation between particle velocity and pressure was used to modify an agglomeration criterion presented previously. Evaluation of the critical aluminum diameter required for agglomeration indicated that both propellant types used in this study should experience some agglomeration over the pressure range studied. The aluminum size required for agglomeration was found to increase with increasing pressure. The experimental findings of the critical aluminum diameter required for agglomeration were in reasonable agreement with the calculated data. The volume mean diameter of the alumina was found to decrease with increasing pressure.

Povinelli, Louis A.↗

NASA NTRS1985

Combustor flame flashback

A stainless steel, two-dimensional (rectangular), center-dump, premixed-prevaporized combustor with quartz window sidewalls for visual access was designed, built, and used to study flashback. A parametric study revealed that the flashback equivalence ratio decreased slightly as the inlet air temperature increased. It also indicated that the average premixer velocity and premixer wall temperature were not governing parameters of flashback. The steady-state velocity balance concept as the flashback mechanism was not supported. From visual observation several stages of burning were identified. High speed photography verified upstream flame propagation with the leading edge of the flame front near the premixer wall. Combustion instabilities (spontaneous pressure oscillations) were discovered during combustion at the dump plane and during flashback. The pressure oscillation frequency ranged from 40 to 80 Hz. The peak-to-peak amplitude (up to 1.4 psi) increased as the fuel/air equivalence ratio was increased attaining a maximum value just before flashback. The amplitude suddenly decreased when the flame stabilized in the premixer. The pressure oscillations were large enough to cause a local flow reversal. A simple test using ceramic fiber tufts indicated flow reversals existed at the premixer exit during flickering. It is suspected that flashback occurs through the premixer wall boundary layer flow reversal caused by combustion instability. A theoretical analysis of periodic flow in the premixing channel has been made. The theory supports the flow reversal mechanism.

Proctor, M. P.↗

NASA NTRS1990

Control/structure interaction study of a Space Station Freedom first flight concept during reboost

In the current design, the Space Station first-flight configuration will be assembled in a LEO and reboosted to a higher altitude such that the orbital decay due to atmospheric drag forces will lower the spacecraft to the proper altitude at the appropriate time to rendezvous with the next-assembly flight of the orbiter. The reboost maneuver will be performed by firing the on/off reaction control system jets to increase the forward velocity. This paper describes the design of a basic orbital reboost maneuver based on a burn-coast-burn scenario and a closed-loop attitude control system which will control the firing sequences of the jets to maintain the spacecraft attitude within required bounds. The design is then used to investigate the infuence of the elastic component response on the reboost jet firing logic.

Lim, Tae W.↗

NASA NTRS1972

Solid-propellant motors for high-incremental-velocity low-acceleration maneuvers in space

The applicability of solid-propellant rockets into a regime of high-performance long-burning tasks beyond the capability of existing motors is discussed. Successful static test firings have demonstrated the feasibility of: (1) utilizing fully case-bonded end-burning propellant charges without mechanical stress relief; (2) using an all-carbon radiative nozzle markedly lighter than the flight-weight ablative nozzle it replaces, and (3) producing low spacecraft acceleration rates during the thrust transient through a controlled-flow igniter that promotes operation below the previous combustion limit.

Shafer, J. I.↗

NASA NTRS1928

Combustion time in the engine cylinder and its effect on engine performance

As part of a general program to study combustion in the engine cylinder and to correlate the phenomena of combustion with the observed performance of actual engines, this paper presents a sketchy outline of what may happen in the engine cylinder during the burning of a charge. It also suggests the type of information needed to supply the details of the picture and points out how combustion time and rate affect the performance of the engine. A theoretical concept of a flame front which is assumed to advance radially from the point of ignition is presented, and calculations based on the area and velocity of this flame and the density of the unburned gases are made to determine the mass rate of combustion. From this rate the mass which has been burned and the pressure at any instant during combustion are computed. This process is then reversed in an effort to determine actual rates of combustion and flame velocities from the pressures as recorded on indicator diagrams. The effects of different rates of combustion on engine performance are then discussed and the importance of proper spark advance is emphasized.

Marvin, Charles F , Jr↗

NASA NTRS1989

Erosive burning threshold conditions in solid rocket motors

Erosive burning has been predicted to be enhanced by factors that increase the level of turbulence close to the propellant surface, such a high cross flow velocity, low surface blowing rate, propellant surface roughness, and adverse pressure gradient. A study is reported which was carried out with the objective of measuring the effects of these parameters on the scaling to larger rocket motor sizes of the transition to, or threshold conditions for, erosive burning rate augmentation. The results are used to develop a scaling criterion for the threshold conditions for erosive burning.

Strand, L. D.↗

NASA NTRS2000

Hydrodynamic Suppression of Soot Formation in Laminar Coflowing Jet Diffusion Flames

Effects of flow (hydrodynamic) properties on limiting conditions for soot-free laminar non-premixed hydrocarbon/air flames (called laminar soot-point conditions) were studied, emphasizing non-buoyant laminar coflowing jet diffusion flames. Effects of air/fuel-stream velocity ratios were of particular interest; therefore, the experiments were carried out at reduced pressures to minimize effects of flow acceleration due to the intrusion of buoyancy. Test conditions included reactant temperatures of 300 K; ambient pressures of 3.7-49 8 kPa; methane-, acetylene-, ethylene-, propane-, and methane-fueled flames burning in coflowing air with fuel-port diameters of 1.7, 3.2, and 6.4 mm, fuel jet Reynolds numbers of 18-121; air coflow velocities of 0-6 m/s; and air/fuel-stream velocity ratios of 0.003-70. Measurements included laminar soot-point flame lengths, laminar soot-point fuel flow rates, and laminar liftoff conditions. The measurements show that laminar soot-point flame lengths and fuel flow rates can be increased, broadening the range of fuel flow rates where the flames remain soot free, by increasing air/fuel-stream velocity ratios. The mechanism of this effect involves the magnitude and direction of flow velocities relative to the flame sheet where increased air/fuel-stream velocity ratios cause progressive reduction of flame residence times in the fuel-rich soot-formation region. The range of soot-free conditions is limited by both liftoff, particularly at low pressures, and the intrusion of effects of buoyancy on effective air/fuel-stream velocity ratios, particularly at high pressures. Effective correlations of laminar soot- and smoke-point flame lengths were also found in terms of a corrected fuel flow rate parameter, based on simplified analysis of laminar jet diffusion flame structure. The results show that laminar smoke-point flame lengths in coflowing air environments are roughly twice as long as soot-free (blue) flames under comparable conditions due to the presence of luminous soot particles under fuel-lean conditions when smoke-point conditions are approached. This is very similar to earlier findings concerning differences between laminar smoke- and sootpoint flame lengths in still environments.

Dai, Z.↗

NASA NTRS1985

Drop-turbulence interactions in a diffusion flame

A monodisperse stream of methanol drops injected along the axis of a turbulent, methane-fueled diffusion flame burning in still air is studied experimentally and theoretically, in order to determine mean and fluctuating phase velocities, mean drop number flux, drop size distributions, and mean gas phase temperature. Measured values were compared with the predictions of two separated flow analyses, namely deterministic separated flow and stochastic separated flow. The stochastic analysis yielded the best agreement with measurements, due to its providing for the turbulent dispersion of drops.

Shuen, J.-S.↗

NASA NTRS1991

Non-isotropic turbulence effects on spray combustion

A numerical model for the prediction of local properties of statistically stationary spray-combusting flows is evaluated by comparison with experimental data. To appraise the relative performance of turbulence models, computations were carried out by the k-epsilon model and the algebraic stress model. The present numerical results show the qualitative agreement with experimental data. In terms of overall local flow properties, the algebraic stress model improves a degree of conformity to the experimental data due to its ability to introduce the nonisotropic turbulence effects. Two swirl numbers are considered to investigate the influence of swirl on the droplet evaporation and trajectories, and the effects of droplet/turbulence interactions in flow properties. It is found that the large swirl produces a higher evaporationn rate, and more intensive turbulent mixing and burning. The discrepancies observed in the results are attributed mainly to uncertainties in the initial spray size and velocity distributions, the droplet/wall impingement interaction, the combustion model with the fast chemistry and the turbulence models dealing with the strong streamline curvature and complex interactions between the dispersed droplets and the continuous gas-phase flows.

Kim, Y. M.↗