Solid Rocket Motor Exhaust Research Articles

NUMERICAL SIMULATION OF WATER COOLING OF THE PLUME OF AN EXPERIMENTAL SOLID ROCKET MOTOR

Typical solid rocket motor exhaust emissions consist of high temperature gaseous pollutants including hydrochloric acid, chlorine, and alumina particles. A scrubber is developed to treat the gaseous and particle pollutants present-reducing the detrimental impact of such testing on the environment. A series of water spray configurations are studied to cool the plume without alumina particle modelling. A system consisting of two jets oriented at 60 degrees offers the best performance without compromising safety; the exhaust gases are cooled from over 3000 K to ambient over a distance of 1 meter from the nozzle exit. This paper is published with the permission of the authors granted to 3AF - Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International Conference.

Numerical Simulation of Water Cooling of the Plume of an Experimental Solid Rocket Motor

The main driver for the present study is the development of a scrubbing system to treat the gaseous and particle pollutants present in the plume of an experimental solid rocket motor (SRM) and thereby reduce the detrimental impact of such testing on the environment as described in the European Union-FESR, MUR PON Research and Innovation 2014-2020, Generazione_E project. Typical SRM exhaust emissions consist of both high temperature gaseous pollutants including hydrochloric acid and chlorine and alumina particles resulting from the combustion of aluminium impregnated in the original solid grain of propellant. The present work studied a series of water spray configurations to cool the plume of an experimental SRM, proposed by AVIO Spa, without alumina particle modelling. A system consisting of two jets oriented at 60 degrees appeared to offer the best performance without compromising safety, where the exhaust gases would be cooled from over 3000 K to 300 K in a distance of 1 meter from the nozzle exit.

Large-Eddy Simulations of a Reactive Solid Rocket Motor Plume

Large-eddy simulation (LES) of a solid rocket motor plume is performed with a reduced chemical scheme that is able to describe the postcombustion occurring in the multispecies supersonic plume between the unburnt exhaust and ambient air. The reduced chemical scheme is first evaluated on laminar flame cases in conditions representative of those encountered in the LES of the reactive plume. The reduced mechanism is then applied to three-dimensional LES, which provides a more detailed representation of the structure of the solid rocket motor exhaust plume compared with previous studies based on one-dimensional box models or three-dimensional Reynolds-averaged Navier–Stokes simulations. The present chemical model does not include alumina particles, whose effects at the plume scale are still uncertain, yet the simulation offers a prediction of the chlorine species distribution, which is known to produce strong interactions with stratospheric ozone, eventually leading to ozone depletion.

Fabrication and erosion studies of C–SiC composite Jet Vanes in solid rocket motor exhaust

C–SiC composite Jet-Vanes were fabricated using liquid silicon infiltration and tested in a plume of aluminum based solid propellant to study erosion resistance. The Jet-Vanes showed excellent resistance to thermo-oxidative erosion; average linear and mass erosion rates were 1 mm/s and 5 g/s, respectively. Morphology of the eroded surfaces suggests that alumina particles in the plume hit the leading edge of the Jet-Vane, damage it and some material is eroded away. Residual silicon melts and reacts with oxygen to form silica which in turn reacts with SiC matrix. The matrix of SiC, silicon and un-reacted carbon is loosened and erode by high shear forces. Once carbon fibers get exposed directly to the plume, these may be both eroded and oxidized.

TEMPERATURE MEASUREMENT OF SOLID ROCKET MOTOR EXHAUST PLUME BY ABSORPTION-EMISSION SPECTROSCOPY1*

A method, which measures the temperature of solid rocket motor exhaust plume, was developed by employing an improved sodium line reversal process. The formula for calculating the temperature was improved and simplified. The temporal temperature-time distributions of the exhaust plume of double base propellant rocket motors were given by the established method. The maximum time resolution and accuracy for the temperature measurement are 40 μs and 8%, respectively. This kind of temperature measurement does not disturb the combustion flow field of exhaust plume and is simpler and more sensitive. The maximum temperatures of SQ-2 propellant exhaust plumes are 2234 K and 2202 K for 240 mm and 500 mm apart from the motor nozzle, respectively. More combustion phenomena of the exhaust plume can be measured by using an existing instrument.

Measurements of stratospheric plume dispersion by imagery of solid rocket motor exhaust

Plume expansion from nine space shuttle and Titan IV vehicles was measured at altitudes of 18, 24, and 30 km in the stratosphere. The plume diameters, inferred from electronic images of polarized, near‐infrared solar radiation scattered from the exhaust particles, increased linearly with time. The expansion rate was measured for as long as 50 min. Observations made simultaneously at multiple altitudes showed that the expansion rate increased with increasing altitude for six measurements made at Cape Canaveral but decreased between 24 and 30 km for the one measurement made at Vandenberg Air Force Base. The average expansion rates for all measurements are 4.3±1.0 m s−1 at 18 km, 6.8±1.9 m s−1 at 24 km, and 8.7±2.4 m s−1 at 30 km. Expansion rates varied from launch to launch by as much as a factor of 1.6 at 18 km, 2.2 at 24 km, and 2.6 at 30 km. No correlation between the expansion rate and wind speed or wind shear was evident. These data are compared with several models for diffusivity and are used to update a comprehensive particle model of solid rocket motor exhaust in the stratosphere. The expansion rates are required by models to calculate the spatial extent and temporal persistence of the local stratospheric ozone depletion cause by solid rocket exhaust.

In situ measurement of the aerosol size distribution in stratospheric solid rocket motor exhaust plumes

The concentration and size distribution of aerosol in the stratospheric exhaust plumes of two Space Shuttle rockets and one Titan IV rocket were measured using a two component aerosol sampling system carried aboard a WB‐57F aircraft. Aerosol size distribution in the 0.01 µm to 4 µm diameter size range was measured using a two component sampling system. The measured distributions display a trimodal form with modes near 0.005 µm, 0.09 µm, and 2.03 µm and are used to infer the relative mass fractionation among the three modes. While the smallest mode has been estimated to contain as much as 10% of the total mass of SRM exhaust alumina, we find show that the smallest mode contains less than 0.05% of the alumina mass. This fraction is so small so as to significantly reduce the likelihood that heterogeneous reactions on the SRM alumina surfaces could produce a significant global impact on stratospheric chemistry.

In‐situ measurement of Cl2 and O3 in a stratospheric solid rocket motor exhaust plume

The concentration of Cl2 in the stratospheric exhaust plume of a Titan IV launch vehicle was measured with a neutral mass spectrometer carried on a WB‐57F aircraft at 18.9 km altitude. Twenty nine minutes after a twilight Titan IV launch, the mean Cl2 concentration across an 8 km wide plume was 126 ± 44 ppbv, consistent with model predictions that a large fraction of the HCl in solid rocket motor exhaust is converted into Cl2 by afterburning reactions in the hot plume. Co‐incident measurements with ultraviolet absorption photometers also carried on the aircraft show that ozone concentration in the plume was not different from ambient levels. This is consistent with model predictions that nighttime SRM launches will not cause transient ozone loss in the lower stratosphere. The measured Cl2 concentration equals 15% of the ambient ozone concentration suggesting that transient ozone reduction in SRM plume wakes can be expected after daytime launches when solar ultraviolet radiation will photolyze the exhaust plume Cl2.

Predicted Optical Characteristics of Solid Rocket Motor Exhaust in the Stratosphere

Optical characteristics of large solid rocket motor exhaust in the stratosphere are predicted for portions of the near-ultraviolet, visible, and near-infrared spectral regions. Mie scattering calculations show that attenuation by thealumina particlesis caused principally by thelarge-particlemodeof the trimodal particlesizedistribution. The local attenuation in the plume as a function of wavelength and the transplume total attenuation due to particles as a function of time are predicted for several wavelengths. An expression for the total surface area of the particles is derived in terms of known parameters and the Sauter mean diameter. Mie theory is then used to show that this Sauter mean diameter can be measured using a two-color transmissometer. The absorption spectra due to the chemicalconstituents are predicted for the200 ‐400-nm wavelength interval forseveraltimes aftervehiclepassage. It is predicted that these absorptions are comparable to the attenuations due to particle scattering. Nomenclature Ap = cross-sectional area of plume, m 2 a = geometrical cross-sectional area of particle, cm 2 atot = total surface area of all particle per unit volume of air, πm 2 cm i3

Solid Rocket Motor Exhaust Model for Alumina Particles in the Stratosphere

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Ozone decomposition on alumina: Implications for solid rocket motor exhaust

Rates of ozone decomposition on aluminum oxide (alumina) particles were measured in a flow tube reactor equipped with molecular beam sampling mass spectrometry and ultraviolet absorption spectroscopy, and in a static reaction cell equipped with ultraviolet absorption spectroscopy. Reaction probabilities η are reported for ozone on α‐alumina, γ‐alumina, and Chromatographic alumina (hydroxylated alumina), respectively, over the temperature range −60 to 200°C. This work addresses the potential for stratospheric ozone depletion by launch vehicle solid rocket motor exhaust. Considering best estimates of plume particle size distributions and dispersion rates, we calculate ozone depletion profiles, for direct decomposition on alumina only. The calculated ozone holes are rather narrow. In the worst case, ozone levels are within 5 × 10−5 of ambient in the center of the plume. A simple analysis of the global impact of alumina particles on ozone decomposition indicates a potential steady‐state daytime depletion of < 2.6 × 10−8 at present launch rates.

Local Effects on of Solid Rocket Motor Exhaust on Stratospheric Ozone

Solid rocket motor (SRM) exhaust contains chlorine, an important stratospheric constituent that plays a crucial role in the chemistry of ozone. Models of SRM plume combustion and chemistry suggest that a significant fraction of SRM exhaust chlorine might be in an active form as a result of afterburning reactions and be available for immediate ozone destruction as the plume mixes with the ambient stratosphere. If afterburning does produce free chlorine, the SRMs used by the Titan IV and Space Shuttle have the potential to cause nearly complete ozone loss in regions up to several tens of kilometers in radius, depending on altitude. We present a simulation of SRM exhaust mixing and chemistry that predicts the three-dimensional structure of the prompt local ozone loss during the 8 h following a daytime Titan IV launch. Maps of the loss of the total column ozone abundance show that the maximum loss of about 30% occurs about one hour after launch. Four hours after launch the area of column ozone loss exceeding 8% covers up to 2000 km. The shape and severity of the ozone-depleted region depends only slightly on the interaction between launch trajectory and stratospheric winds. Solar ultraviolet flux on the ground beneath the plume changes in response to decreased ozone absorption, alumina scattering, and absorption by chlorine oxide compounds. The ultraviolet flux increases by 100% at 295 nm, shows no change near 300 nm, and decreases by 30% longward of 310 nm. Existing space instrumentation does not possess the spatial resolution required to measure ozone depletion of the size and magnitude predicted following launches involving large SRMs.

Using a geographical information system for monitoring space shuttle launches: Determining cumulative distribution of deposition and an empirical test of a spatial model

Space shuttle launches produce localized hydrochloric acid deposition. The interaction of solid rocket motor exhaust and deluge water released on the pad at the time of launch results in the formation of an exhaust cloud. The spatial pattern and extent of deposition from the launch cloud are predicted by the rocket exhaust effluent diffusion (REED) model. The actual pattern of deposition has been mapped by field surveys for each shuttle launch since 1981. In this paper we use a geographical information system (GIS) to compare model predictions with ground patterns for 49 shuttle launches. We also compile cumulative maps of deposition patterns needed to consider long-term impacts. The direction of launch cloud movement did not differ significantly from model predictions. The REED model overpredicted both the area that received deposition and the maximum distance from the launch pad that deposition occurred. Severe vegetation damage was restricted to near-field deposition areas within 1980 m north of each launch pad. Total area impacted from launches has been 87.0 ha around pad 39A and 52.9 ha around pad 39B. Far-field deposition has caused leaf spotting from acid droplets or aluminum oxide over a wider and more variable area than near-field. A total of 19,397 ha has received deposition, but 63.6% of this area has received deposition only one time and 92.2% not more than three times. GIS techniques provide means to test spatial models and compile information useful for assessing cumulative impacts.

Ignition/duct overpressure induced by Space Shuttle solid rocket motor ignition

An analytical methodology, with a propagating solid rocket motor (SRM) exhaust front as a perturbation pressure wave generator, is developed to enhance understanding of the ignition/duct overpressure (IOP/DOP) induced during the Space Shuttle liftoff. Waveform, amplitude, and low-frequency responses of IOP/DOP are simulated. For a full-scale configuration, IOP and DOP responses are clearly isolated and distinguishable. However, in the subscale tests, an interaction region of IOP and DOP exists, which obscures interpretation of the test data. The methodology offers a first-order assessment of these phenomena with respect to the launch complex geometry and the SRM ignition transient, and provides guidance for the test data interpretation.