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Computer program can be used to predict gas dynamic and chemical properties of underexpanded rocket plumes from sea level to altitude above which viscous continuum-flow assumption with distinct shocks is no longer valid.
Optical spectrometry can provide means to characterize rocket engine exhaust plume impurities due to eroded materials, as well as combustion mixture ratio without any interference with plume. Fiberoptic probes and cables were designed, fabricated and installed on Space Shuttle Main Engines (SSME), allowing monitoring of the plume spectra in real time with a Commercial of the Shelf (COTS) fiberoptic spectrometer, located in a test-stand control room. The probes and the cables survived the harsh engine environments for numerous hot-fire tests. When the plume was seeded with a nickel alloy powder, the spectrometer was able to successfully detect all the metallic and OH radical spectra from 300 to 800 nanometers.
For the first time, a suite of three lidar sensors have been used in flight to scan a lunar-like hazard field, identify a safe landing site, and, in concert with an experimental Guidance, Navigation, and Control (GN&C) system, guide the Morpheus autonomous, rocket-propelled, free-flying test bed to a safe landing on the hazard field. The lidar sensors and GN&C system are part of the Autonomous Precision Landing and Hazard Detection and Avoidance Technology (ALHAT) project which has been seeking to develop a system capable of enabling safe, precise crewed or robotic landings in challenging terrain on planetary bodies under any ambient lighting conditions. The 3-D imaging flash lidar is a second generation, compact, real-time, air-cooled instrument developed from a number of cutting-edge components from industry and NASA and is used as part of the ALHAT Hazard Detection System (HDS) to scan the hazard field and build a 3-D Digital Elevation Map (DEM) in near-real time for identifying safe sites. The flash lidar is capable of identifying a 30 cm hazard from a slant range of 1 km with its 8 cm range precision at 1 sigma. The flash lidar is also used in Hazard Relative Navigation (HRN) to provide position updates down to a 250m slant range to the ALHAT navigation filter as it guides Morpheus to the safe site. The Doppler Lidar system has been developed within NASA to provide velocity measurements with an accuracy of 0.2 cm/sec and range measurements with an accuracy of 17 cm both from a maximum range of 2,200 m to a minimum range of several meters above the ground. The Doppler Lidar's measurements are fed into the ALHAT navigation filter to provide lander guidance to the safe site. The Laser Altimeter, also developed within NASA, provides range measurements with an accuracy of 5 cm from a maximum operational range of 30 km down to 1 m and, being a separate sensor from the flash lidar, can provide range along a separate vector. The Laser Altimeter measurements are also fed into the ALHAT navigation filter to provide lander guidance to the safe site. The flight tests served as the culmination of the TRL 6 journey for the lidar suite and included launch from a pad situated at the NASA-Kennedy Space Center Shuttle Landing Facility (SLF) runway, a lunar-like descent trajectory from an altitude of 250m, and landing on a lunar-like hazard field of rocks, craters, hazardous slopes, and safe sites 400m down-range just off the North end of the runway. The tests both confirmed the expected performance and also revealed several challenges present in the flight-like environment which will feed into future TRL advancement of the sensors. The flash lidar identified hazards as small as 30 cm from the maximum slant range of 450 m which Morpheus could provide, however, it was occasionally susceptible to an increase in range noise due to heated air from the Morpheus rocket plume which entered its Field-of-View (FOV). The flash lidar was also susceptible to pre-triggering on dust during the HRN phase which was created during launch and transported by the wind. The Doppler Lidar provided velocity and range measurements to the expected accuracy levels yet it was also susceptible to signal degradation due to air heated by the rocket engine. The Laser Altimeter, operating with a degraded transmitter laser, also showed signal attenuation over a few seconds at a specific phase of the flight due to the heat plume generated by the rocket engine.
Data from absorption coefficients and fine-structure parameters measured for water vapor have been incorporated in an analytic program useful in evaluating heating by radiation from the exhaust plume of a large rocket.
A novel method to measure the velocity of ejecta generated by the impingement of rocket plumes on a lunar or planetary surface is presented. Dust Ejecta RADAR Technology (DERT), is a Millimeter Wave Doppler Radar sensor that provides a unique dataset based on direct measurements of ejecta particle velocities during plume-surface interactions, which will help inform risk assessments associated with impacts by high-speed ejecta particles. DERT has measured velocities of a pendulum (~1 m/s), a rotating strip (60 m/s), projected regolith simulant BP-1 (~250 m/s) and rifle rounds (~820 m/s). Signal-to-noise ratios of up to 60 dB with regolith simulant BP-1 have been demonstrated with a CW 94 GHz unit. Data collected from DERT will inform the development of high-fidelity computational models of plume-surface interaction effects and will help address NASA Strategic Knowledge Gaps related to characterizing entry, descent and landing effects, and the risks associated with high-speed ejecta on the lander and lunar surface assets. Recent efforts to develop DERT as a flight instrument are presented as well as the development of a Lunar Regolith Ejecta Simulator (LuRES) system to calibrate DERT.
Proposed experiments for analyzing rocket plumes are reported. Two groups of experiments were studied: (1) those that would help define some of the parameters that characterize the plume and (2) those that would enable evaluation of some of the contamination effects of the plume environment on various items of interest. The items investigated, the purpose of the investigation, are given in tabular form.
Tests at M = 8.2 show that a simulated rocket plume at the base of a blunted cone can cause large areas of separated flow, with dramatic effects on the heat transfer rate distribution. The plume was simulated by solid discs of varying sizes or by an annular jet of gas. Flow over the cone without a plume is fully laminar and attached. Using a large disc, the boundary layer is laminar at separation at the test Reynolds number. Transition occurs along the separated shear layer and the boundary layer quickly becomes turbulent. The reduction in heat transfer associated with a laminar separated region is followed by rising values as transition occurs and the heat transfer rates towards the rear of the cone substantially exceed the values obtained without a plume. With the annular jet or a small disc, separation occurs much further aft, so that heat transfer rates at the front of the cone are comparable with those found without a plume. Downstream of separation the shear layer now remains laminar and the heat transfer rates to the surface are significantly lower than the attached flow values.
Wind tunnel tests were performed to determine pressures, heat transfer rates, and gas recovery temperatures in the base region of a rocket firing model of the space shuttle integrated vehicle during simulated yawed flight conditions. First and second stage flight of the space shuttle were simulated by firing the main engines in conjunction with the SRB rocket motors or only the SSME's into the continuous tunnel airstream. For the correct rocket plume environment, the simulated altitude pressures were halved to maintain the rocket chamber/altitude pressure ratio. Tunnel freestream Mach numbers from 2.2 to 3.5 were simulated over an altitude range of 60 to 130 thousand feet with varying angle of attack, yaw angle, nozzle gimbal angle and SRB chamber pressure. Gas recovery temperature data derived from nine gas temperature probe runs are presented. The model configuration, instrumentation, test procedures, and data reduction are described.
Experimental aerodynamic investigations were conducted in a 3.5-foot hypersonic wind tunnel. The model used for this test was a 0.010-scale of the Configuration 2 Space Shuttle Orbiter and the External Tank. Six-component aerodynamic force and moment data were recorded over an angle of attack range from -8 deg to +30 deg at 0 deg and 5 deg angles of sideslip. Data was also recorded during beta sweeps of -8 deg to +10 deg at angles of attack of -10 deg, 0 deg, and 30 deg. All testing was done at Mach 7.3. Various elevon, rudder and orbiter to external tank attaching structures and fairings were tested to determine longitudinal and lateral-directional stability characteristics. Non-metric exhaust plumes were installed during a portion of the testing to determine the effects of the main propulsion system rocket plumes.
A series of three tests was conducted using solid rocket propellants to determine the effects a solid rocket plume would have on thermal protective surfaces (TPS). The surfaces tested were those which are baselined for the shuttle vehicle. The propellants used were to simulate the separation solid rocket motors (SSRM) that separate the solid rocket boosters (SRB) from the shuttle launch vehicle. Data cover: (1) the optical effects of the plume environment on spacecraft related surfaces, and (2) the solid particle size, distribution, and composition at TPS sample locations.
Two tests supporting development of the launch abort system for the Orion MultiPurpose Crew Vehicle were run in the NASA Ames Unitary Plan wind tunnel last year. The first test used a fully metric model to examine the stability and controllability of the Launch Abort Vehicle during potential abort scenarios for Mach numbers ranging from 0.3 to 2.5. The aerodynamic effects of the Abort Motor and Attitude Control Motor plumes were simulated using high-pressure air flowing through independent paths. The aerodynamic effects of the proximity to the launch vehicle during the early moments of an abort were simulated with a remotely actuated Service Module that allowed the position relative to the Crew Module to be varied appropriately. The second test simulated the acoustic environment around the Launch Abort Vehicle caused by the plumes from the 400,000-pound thrust, solid-fueled Abort Motor. To obtain the proper acoustic characteristics of the hot rocket plumes for the flight vehicle, heated Helium was used. A custom Helium supply system was developed for the test consisting of 2 jumbo high-pressure Helium trailers, a twelve-tube accumulator, and a 13MW gas-fired heater borrowed from the Propulsion Simulation Laboratory at NASA Glenn Research Center. The test provided fluctuating surface pressure measurements at over 200 points on the vehicle surface that have now been used to define the ground-testing requirements for the Orion Launch Abort Vehicle.
Here we present a novel use of millimeter wave doppler radar to measure the velocity of ejecta generated by the impingement of rocket plumes on a planetary surface. The Dust Ejecta Radar Technology (DERT) provides a unique dataset based on direct measurements of ejecta particle velocities during plume-surface interactions. DERT is a continuous wave (CW) 94 GHz radar unit that is being developed at Kennedy Space Center. DERT has measured velocities of a pendulum (~ 1 m/s), a rotating strip (60 m/s), projected regolith simulant BP-1 (~ 250 m/s) and rifle rounds (~ 820 m/s). Signal-to-Noise Ratios of up to 60 dB have been demonstrated. Data collected from DERT will inform the development of high-fidelity computational models of plume-surface interaction effects and will help address NASA Strategic Knowledge Gaps related to characterizing entry, descent and landing effects, and the risks associated with high-speed ejecta on the lander and lunar surface assets.
Here we present a novel use of millimeter wave doppler radar to measure the velocity of ejecta generated by the impingement of rocket plumes on a planetary surface. The Dust Ejecta Radar Technology (DERT) provides a unique dataset based on direct measurements of ejecta particle velocities during plume-surface interactions. DERT is a continuous wave (CW) 94 GHz radar unit that is being developed at Kennedy Space Center. DERT has measured velocities of a pendulum (~ 1 m/s), a rotating strip (60 m/s), projected regolith simulant BP-1 (~ 250 m/s) and rifle rounds (~ 820 m/s). Signal-to-Noise Ratios of up to 60 dB have been demonstrated. Data collected from DERT will inform the development of high-fidelity computational models of plume-surface interaction effects and will help address NASA Strategic Knowledge Gaps related to characterizing entry, descent and landing effects, and the risks associated with high-speed ejecta on the lander and lunar surface assets.
Landing vehicle propulsive systems introduce several risks associated with plume/surface interactions (PSI) during the final stages of descent. For example, view obscuration during descent due to the liberation of dust driven by the rocket plume’s interaction with the lunar surface was reported on several Apollo missions. Plume-driven debris impacts on vehicle and nearby assets as well as landing instabilities due to crater formation are also potential risks associated with plume/surface interactions. The importance of studying these plume/surface interactions has been elevated due to NASA’s mission to land humans on the Moon by 2024. Existing and under-development predictive simulation tools are currently being brought to application readiness under a project awarded by the NASA Space Technology Mission Directorate (STMD) Game Changing Development (GCD) program. This tool development is divided into four tasks: (1) simulating plume structure in low-pressure environments, (2) simulating crater development and ejecta, (3) regolith particle phase modeling, and (4) gas-particle interaction modeling. Recent efforts under task 1 include demonstrating existing computational fluid dynamics (CFD) tools for simulating plume structure in low pressure environments similar to those found on Mars. The results of these demonstrative simulations are compared to available experimental data. Emerging capability being developed under task 2 is also presented, demonstrating simulations of crater formation in Lunar conditions.
High-fidelity models of plume-regolith interaction are difficult to develop because of the widely disparate flow conditions that exist in this process. The gas in the core of a rocket plume can often be modeled as a time-dependent, high-temperature, turbulent, reacting continuum flow. However, due to the vacuum conditions on the lunar surface, the mean molecular path in the outer parts of the plume is too long for the continuum assumption to remain valid. Molecular methods are better suited to model this region of the flow. Finally, granular and multiphase flow models must be employed to describe the dust and debris that are displaced from the surface, as well as how a crater is formed in the regolith. At present, standard commercial CFD (computational fluid dynamics) software is not capable of coupling each of these flow regimes to provide an accurate representation of this flow process, necessitating the development of custom software. This software solves the fluid-flow-governing equations in an Eulerian framework, coupled with the particle transport equations that are solved in a Lagrangian framework. It uses a fourth-order explicit Runge-Kutta scheme for temporal integration, an eighth-order central finite differencing scheme for spatial discretization. The non-linear terms in the governing equations are recast in cubic skew symmetric form to reduce aliasing error. The second derivative viscous terms are computed using eighth-order narrow stencils that provide better diffusion for the highest resolved wave numbers. A fourth-order Lagrange interpolation procedure is used to obtain gas-phase variable values at the particle locations.
Rocket exhaust research in areas of direct impingement heating, backwash base heating, radio signal attenuation, and interference with line of sight to ground stations
The Low-power Atmospheric Compensation Experiment (LACE) satellite, launched in February 1990 by the Naval Research Laboratory, uses a magnetic damper on a gravity gradient boom and a momentum wheel with its axis perpendicular to the plane of the orbit to stabilize and maintain its attitude. Satellite attitude is determined using three types of sensors: a conical Earth scanner, a set of sun sensors, and a magnetometer. The Ultraviolet Plume Instrument (UVPI), on board LACE, consists of two intensified CCD cameras and a gimbal led pointing mirror. The primary purpose of the UVPI is to image rocket plumes from space in the ultraviolet and visible wavelengths. Secondary objectives include imaging stars, atmospheric phenomena, and ground targets. The problem facing the UVPI experimenters is that the sensitivity of the LACF satellite attitude sensors is not always adequate to correctly point the UVPI cameras. Our solution is to point the UVPI cameras at known targets and use the information thus gained to improve attitude measurements. This paper describes the three methods developed to determine improved attitude values using the UVPI for both real-time operations and post observation analysis.
As a final report for phase 1 of the project, the researchers are submitting to the Tennessee Tech Office of Research the following two papers (reprinted in this report): 'Collision Line Broadening Effects on Spectrometric Data from the Optical Plume Anomaly System (OPAD),' presented at the 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 27-29 June 1994, and 'Calculation of Collision Cross Sections for Atomic Line Broadening in the Plume of the Space Shuttle Main Engine (SSME),' presented at the IEEE Southeastcon '95, 26-29 March 1995. These papers fully state the problem and the progress made up to the end of NASA Fiscal Year 1994. The NASA OPAD system was devised to predict concentrations of anomalous species in the plume of the Space Shuttle Main Engine (SSME) through analysis of spectrometric data. The self absorption of the radiation of these plume anomalies is highly dependent on the line shape of the atomic transition of interest. The Collision Line Broadening paper discusses the methods used to predict line shapes of atomic transitions in the environment of a rocket plume. The Voigt profile is used as the line shape factor since both Doppler and collisional line broadening are significant. Methods used to determine the collisional cross sections are discussed and the results are given and compared with experimental data. These collisional cross sections are then incorporated into the current self absorbing radiative model and the predicted spectrum is compared to actual spectral data collected from the Stennis Space Center Diagnostic Test Facility rocket engine. The second paper included in this report investigates an analytical method for determining the cross sections for collision line broadening by molecular perturbers, using effective central force interaction potentials. These cross sections are determined for several atomic species with H2, one of the principal constituents of the SSME plume environment, and compared with experimental data.