Rocket Exhaust Research Articles

Reactions of Halomethanes with γ-Alumina Surfaces. 2. X-ray Photoelectron and Temperature-Programmed Reaction Spectroscopic Studies

The reactions of CF3Cl, CF2Cl2, CFCl3, and CCl4 with γ-alumina powders have been studied using temperature-programmed reaction (TPR) and X-ray photoelectron (XPS) spectroscopies. In the TPR experiments, powder samples that had been dehydroxylated by heating to 1000 K in vacuum were dosed at 100 K with halomethanes and then gradually heated. Desorbing species were monitored as a function of substrate temperature using a line-of-sight quadrupole mass spectrometer. Hydrogen chloride and halomethyl fragments, which are indicative of halomethane dissociative chemisorption, were observed to desorb at temperatures below 150 K. Carbon dioxide began to desorb between 240 and 320 K; the onset temperature of CO2 desorption correlated inversely with the degree of chlorination of the compound. The CO2 most likely arises from COn (carbonate and/or formate) species which are formed via the low-temperature dissociative chemisorption of the halomethanes. Repeated heating and dosing of the alumina sample resulted in a drop in CO2 production, implying that surface active sites become depleted. In situ XPS analysis of heat-treated powders that had been dosed at 150 K with chlorofluoromethanes revealed the presence of both organic and inorganic forms of fluorine. Powders dosed at 200 K had much lower organic halogen coverages. Halogen uptake probabilities, which are estimated to be ∼10-5 from the data, increased as the degree of chlorination of the halomethane increased. These results indicate that chlorofluoromethane compounds will probably decompose on alumina solid-propellant rocket motor exhaust particles in the stratosphere, forming adsorbed Al−X (X = F, Cl) and COn species and releasing gas phase HCl and CFxCly fragments. However, the impact of these processes on global stratospheric halomethane and ozone concentrations is likely to be minimal.

A physico-mathematical model of rocket exhaust plumes

An axisymmetric nonisobaric supersonic jet mixing with the atmosphere or expanding into a vacuum is considered as a model of a rocket engine plume. The expanding caused by the nonisobarity, turbulent mixing and afterburning are main processes forming a plume structure at different altitudes. Optical characteristics of the plume are determined by thermal radiation of condense-phase particles and vibration-rotation transitions in molecules. In the visible and ultraviolet spectral ranges chemiluminescent mechanisms make significant contribution to plume radiation. The Navier-Stokes equations with additional terms describing chemical kinetics and vibrational relaxation have been used for numerical simulation of plumes. An original technique has been developed to calculate selective radiation transfer in nonuniform nonequilibriam medium. Calculated plume spectra of typical liquid and solid fuel rocket engines at different altitudes are presented for infrared, visible and ultraviolet ranges.

Observations of the Character of a Water Jet

High-velocity jets of fluid are encountered in both natural and man-made settings. The geyser, the fireman's hose, and the rocket exhaust are notable examples. Study of the fluid mechanics literature reveals the axisymmetric fluid jet as a fundamental fluid configuration that is rich in physical features. It is the purpose of this paper to discuss a few basic, but important, aspects of the behaviour of a water jet and ways that they may be directly observed by the senses, without recourse to sophisticated measuring devices, to the benefit of science and engineering students of all ages. These are observations which may be readily performed in undergraduate laboratories or even as homework assignments.

Extremally steered singular thrust-arcs for space flight in an atmosphere

This paper advances prior results on the problem of time-free, Mayer-optimal control of a blunt space vehicle's flight in an atmosphere. The extremally steered, first order, singular arc is partially singular and the Lie-bracket formalism employed previously for the totally singular case is not quite applicable. However, by differentiating the switching function, a relationship is obtained between the totally and partially singular arcs. For the special case of coplanar flight, the singular thrust control is expressed in a nonlinear state feedback form. This form indicates that the singular thrust magnitude is separable as the sum of five types of forces: one due to gravity alone, another due to centrifugal effects, a third due to drag alone, a fourth due to atmospheric density gradient, and finally a fifth due to a coupling between these forces. The extremal steering angle along the singular arc is obtained as an implicit function of the states. An explicit solution is obtained for the special case of a low-density atmosphere. For proper combinations of steering and rocket exhaust speed, the singular arc satisfies the generalized Legendre-Clebsch condition and the Goh-Robbins condition making it a strong candidate for Mayer optimality.

Lidar for remote measurement of ozone in the exhaust plumes of launch vehicles

Large quantities of chlorine and alumina particles are injected directly into the stratosphere by the current fleet of launch vehicles. Environmental concerns have been raised over the impact of the rocket exhaust on the ozone layer. Recently differential absorption lidar (DIAL) was selected for remote sensing of ozone density within the plumes of Titan IV launch vehicles. The application of DIAL to this very challenging problem is described, and an implementation of UV-ozone DIAL is discussed that holds promise for this application.

In Situ Measurement of Hydrochloric Acid in Rocket Exhaust Clouds Using a Remotely Piloted Aircraft

Abstract An instrumentation system employing an RPV (remotely piloted vehicle) platform was developed for temporally and spatially resolved air pollution measurements, and was used to measure the evolution of gas-phase HC1 in exhaust clouds from a solid rocket motor firing and fuel pit burns. A thermistor and a sensitive (ppmv-level), rapid-response (<0.1 sec) infrared absorption sensor for HC1 were mounted in a flow channel in the RPV, permitting concentration and temperature measurements to be made in the cloud on a several-meter scale. The RPV system was flown in a series of field tests at Thiokol Corporation’s Elkton, MD division to evaluate the HC1 content of the exhaust products of a new Mg-based fuel formulation. Measurements were made in the clouds from Al-based and Mg-based solid fuel pit burns and a Mg-fueled motor firing over periods of several minutes. Elevated temperatures and HC1 concentrations were found to be temporally correlated with video images of the particulate cloud. Cl originating ...

Investigation of microwave attenuation measurements in a laboratory-scale rocket motor plume

Introduction Overview T HE use of microwave diagnostics in rocket exhaust studies can allow determination of electron density and provide information concerning communication signal reception and radar signature. The classical presentation of microwave measurement techniques is for use in underdense or low-collision plasmas,' such as rocket exhausts at low back pressures. In contrast, this investigation is concerned with electron density determinations in rockets exhausting at atmospheric pressure. The appropriate use of assumptions is reviewed, and a graphical presentation of index-ofattenuation behavior is provided to supplement the assessment of the experimental results. An experimental investigation using multiple-frequency microwave diagnostics was performed to assess the use of microwave attenuation measurements in laboratory-scale solid-rocket exhaust plumes for ranking propellants according to their production of electrons. The challenges of this work are interpreting measurements in a collisional plasma coupled with the small plume diameter of the laboratory-scale rocket exhaust.

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.

Solid rocket exhaust in the stratosphere - Plume diffusion and chemical reactions

A model has been developed to examine, on a local scale, the reactions of rocket exhaust from solid rocket motors with stratospheric ozone. The effects were examined at two different altitudes. Results of the modeling study indicate that afterburning chemistry of reactive exhaust products can cause local but transient (on the order of several minutes) loss of ozone. The modeling study included potential heterogeneous reactions at aluminum oxide surfaces. Results indicate that these potential heterogeneous reactions do not have a major impact on the local plume chemistry. Homogeneous reactions appear to be of more consequence during the early dispersion of the plume. It has also been found that the rate of plume dispersion has a very significant effect on local ozone loss.

Compatibility of infrared band models with scattering

Techniques for the computation of radiative heating from aluminized solid propellant rocket exhaust plumes must account for infrared emission and absorption by hot H2O and CO2 in the presence of strong threedimensional aerosol scattering by micron-sized A12O3 droplets and particles. Radiative heating computations are usually performed over wide spectral intervals using infrared band models. However, no rigorous extension to infrared band models has been proposed for situations with scattering. Indeed, band models and scattering are widely held to be incompatible. Practical engineering calculations of plume radiative heating have therefore proceeded, using various approximations to the transport equation. Although the errors in these approximations are believed to be reasonably small, they have never been quantified because of the lack of rigorous results. In order to help remedy this deficiency, this article develops two different rigorous solutions. Although they are currently restricted to homogeneous media, they nevertheless provide a previously unavailable means of calibrating code performance and assessing the accuracy of various approximation schemes. Some sample applications are provided for the Space Shuttle Solid Rocket Booster.

Simplified model for calculation of backflow contamination from rocket exhausts in vacuum

A computationally simple procedure for estimating the backflow from a plume expanding into a vacuum has been developed. The continuum flow is modeled as in the Simons method, with straight streamlines radiating from a point on the plume axis of symmetry, constant velocity, and density varying inversely as the square of the radius. The continuous transition from continuum to free-molecular flow is replaced by a suitably defined discontinuity surf ace. Particle number densities outside the discontinuity surface are calculated by assuming that a Maxwellian velocity distribution exists for particles located within small volumes adjacent to that surface. The boundary layer at the nozzle exit is accounted for and may be large compared with the freestream inviscid flow region. Calculated results are compared with existing backflow data for nozzles using CCh as a propellant.

Decomposition of halomethanes on α‐alumina at stratospheric temperatures

We present data that suggest the possibility of a previously unsuspected heterogeneous process involving the decomposition of halocarhon source gases on alumina solid‐propellant rocket motor exhaust particles deposited in the stratosphere. These results, obtained under ultrahigh vacuum conditions, provide evidence that halomethane compounds undergo dissociative chemisorption on α‐alumina surfaces at stratospheric temperatures, leading to the release of more reactive halogen containing species. Simple calculations indicate that halocarbon concentrations will be severely depleted in the immediate wake of the rocket plume; any serious global impact at current projected launch rates requires the existence of one or more processes that regenerate active sites on the surface subsequent to the dissociative chemisorption of halocarbons.

Nucleation, growth, and morphology of dust in plasmas

Abstract

Chemical Microsensors for Satellite Applications

AbstractWe have been investigating several chemical microsensor technologies for the detection of chemicals found in propellant and rocket exhaust plumes and as contaminants for satellite components. In this work we have developed a catalytic metal sensor for the detection of H2contamination in electronic device packages. The sensor is based on the resistance of a thin film Pd/Ni alloy. Data have been obtained on its response at room temperature to various levels of H2as well as the effects of H2O and O2on its performance. Hydrogen levels to 10.5 ppm have been detected and the responsivity is about twice that of similar sensors. The sensor is insensitive to H2O but its response to H2is strongly inhibited by O2.

EXPERIMENTAL STUDIES ON ROCKET EXHAUST PLUMES: SUPPORT FOR COMPUTED SIMULATION

Exhaust plumes of solid propellant missiles are heterogeneous flowfields containing gases and liquids or solid particles. Exhaust species mix with the atmosphere and may produce an afterburning flame. AP composite propellant combustion produces chloric acid which contributes to atmospheric water vapour condensation. A missile exhaust plume has harmful operational consequences. It may disturb the missile guidance system and increase engine and launch platform detectability. Specific computer codes are used in order to simulate operational missile exhaust plume effects. Because these codes have to be verified, experimental techniques have been developed in order to characterize static rocket exhaust plumes. Radiation structure, radiance and temperature are measured by video, infrared and UV camera. Radiative chemical species are identified by spectroradiometry. Flowfield is studied by laser doppler anemometry and laser tomography. Collected smoke particles are analysed by electronic microscopy and X-ray spectrometry. Meteorological thresholds for water vapour condensation are noted for various propellant compositions. Experimental setup, instruments and tested propellant are chosen in order to optimize comparisons between experimental results and numerical computations. Moreover, computer codes are modified to take experimental data into account. Experimental and computed results allow establishment of correlations between exhaust plume characteristics, motor parameters and propellant compositions.

Theory of plume radiance from the bow shock ultraviolet 2 rocket flight

A computational fluid dynamics algorithm is used to simulate the flow field and solid rocket motor plume about the bow shock ultraviolet 2 flight vehicle. A new computationally efficient algorithm to model two-phase gas and particulate flow is developed. The flow over the complete rocket geometry and its interaction with the particle-laden plume is simulated. The computed plume radiance is compared with radiometric and spectroscopic data and good agreement with radiance magnitudes and spectral characteristics is obtained for the intrinsic core rocket exhaust data. The computation underpredicts the observed signal from the far-field photometers by many orders of magnitude. Based on the particulate flow simulations developed here and new experimental data, we hypothesize that the far-field radiation is due to molecular emission. Use of the CO Cameron band spectra obtained on this flight provides an estimate of the governing temperature of such radiation.

Loss mechanism and diagnostic elements of rocket exhaust-jet structure from the diffracted field

The small arrival-angle loss mechanism associated with a typical high electron-density rocket exhaust jet has been modelled using an approximate scalar, ‘effective diffracting surface’, description. It contrasts with GTD descriptions using ‘equivalent’ diffracting objects in that its ray-optic basis accounts for changes in the intensity distribution within a diffuse region surrounding the geometrical shadow. Diagnostic information on broad features of jet structure, which can facilitate noninvasive validation of models of the jet structure, have been identified within the diffracted field.

Magnesium-neutralized propellant

Conventional solid rocket motors, which utilize ammonium perchlorate as the propellant oxidizer, produce hydrochloric acid (HC1) as an exhaust product. A low-acid propellant has been developed which substantially reduces the HCl content of the solid rocket exhaust. The predominant mechanism for HC1 exhaust reduction is a neutralization of the acid by the base magnesium hydroxide which is generated in the exhaust plume from the combustion of magnesium (Mg) fuel. Experimental analysis of the Mg-propellant combustion products has shown that HCl levels of less than 1 % are achievable. The Mg-neutralized propellant exhibits properties that are compatible with the reliability, cost, performance, ballistic, and mechanical properties as well as the safety dictates of an advanced launch system (ALS). Predictions for propellant variability and sensitivity suggest that the Mg-neutralized propellant reliability will compare favorably with the current state-of-the-art hydroxy terminated polybutadiene (HTPB) propellants. The range of ballistic tailorability is broad and the low-volumetric solids loading of the Mg propellant results in outstanding mechanical properties and simple processing. The Mg propellant also exhibits DoD zero cards class 1.3 hazards properties which suggests that no compromise in safety will be encountered relative to current propellants used in large booster applications.

The effects of chemical propulsion on the environment

This paper seeks to quantify the effects of chemical propulsion exhaust on both the local launch site and the global environments. Four major areas of concern are discussed: the stratospheric ozone, acid rain, toxicity, and the greenhouse effect. The environmental impacts of both solid and liquid rocket propulsion systems are evaluated. The exhaust species and launch profile of the Space Shuttle, which injects the greatest mass of exhaust products into the atmosphere of any current system, are discussed in some detail. Model calculations predict a global stratospheric ozone reduction of about 0.01% due to chemical propulsion. Acid rain due to the HCl in solid rocket exhaust has a small measurable impact on the local environment, with the mortalities of some plants and small fish very near (<2500 ft) the launch site having been documented. Based on history, the handling of potentially toxic species from the use of chemical propulsion systems is manageable. The relative contribution of chemical propulsion to the global CO 2 burden, the increase of which may lead to global warming, is insignificant. It appears that the perturbation to the natural environment caused by chemical propulsion exhaust is very small and manageable, even for the most optimistic projections of future launch rates.

A Cost-Effective Approach to Evaluate High-Temperature Ablatives for Military Applications

Several methods for evaluating thermal performance of ablative materials under simulated solid rocket exhaust plumes are presented in this paper. A small-scale supersonic rocket motor facility capable of delivering more than 700 Btu/s-ft2 (7,947 kW/m2) was used to simulate the rocket motor environment in the laboratory. A particle feeder with the capability of injecting 40 lb/hr (5.0 gm/s) of aluminum oxide (Al2O3) particles into the flow stream was used to simulate the particle impingement effects of solid rocket exhaust. Application of several gauges for heat flux measurement is discussed in this paper. These gauges include a circular foil calorimeter (Gardon), flat plate calorimeter, and pencil probe surface thermocouple. Several techniques for erosion measurement are also presented. These techniques consist of a linear variable differential transformer (LVDT), eddy current device, ultrasonic device, and laser-based sensor. The operation principle of the LVDT and laser-based sensors is presented, and some typical results are shown. High-speed motion and video analyses were performed to study material surface behavior during testing. Scanning electronic microscope (SEM) analyses were also performed on ablatives to study subsurface structure behavior during ablation. A cost-effective approach was developed to evaluate high-temperature ablatives for military applications.