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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Supersonic Retropropulsion Experimental Results from the NASA Langley Unitary Plan Wind Tunnel

A new supersonic retropropulsion experimental effort, intended to provide code validation data, was recently completed in the Langley Research Center Unitary Plan Wind Tunnel Test Section 2 over the Mach number range from 2.4 to 4.6. The experimental model was designed using insights gained from pre-test computations, which were instrumental for sizing and refining the model to minimize tunnel wall interference and internal flow separation concerns. A 5-in diameter 70-deg sphere-cone forebody with a roughly 10-in long cylindrical aftbody was the baseline configuration selected for this study. The forebody was designed to accommodate up to four 4:1 area ratio supersonic nozzles. Primary measurements for this model were a large number of surface pressures on the forebody and aftbody. Supplemental data included high-speed Schlieren video and internal pressures and temperatures. The run matrix was developed to allow for the quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and bias errors due to flow field or model misalignments. Preliminary results and observations from the test are presented, while detailed data and uncertainty analyses are ongoing.

Berry, Scott A.↗
NASA NTRS2012

Validation of OVERFLOW for Supersonic Retropropulsion

The goal is to softly land high mass vehicles (10s of metric tons) on Mars. Supersonic Retropropulsion (SRP) is a potential method of deceleration. Current method of supersonic parachutes does not scale past ~1 metric ton. CFD is of increasing importance since flight and experimental data at these conditions is difficult to obtain. CFD must first be validated at these conditions.

Schauerhamer, Guy↗
NASA NTRS2014

Supersonic Retropropulsion Test 1853 in NASA LaRC Unitary Plan Wind Tunnel Test Section 2

A supersonic retropropulsion experiment was conducted in the Langley Research Center Unitary Plan Wind Tunnel Test Section 2 at Mach numbers of 2.4, 3.5, and 4.6. Intended as a code validation effort, this study used pretest computations to size and refine the model such that tunnel blockage and internal flow separations were minimized. A 5-in diameter 70 degree sphere-cone forebody, which can accommodate up to four 4:1 area ratio nozzles, followed by a 9.55 inches long cylindrical aft body was selected for this test after computational maturation. The primary measurements for this experiment were high spatial-density surface pressures. In addition, high speed schlieren video and internal pressures and temperatures were acquired. The test included parametric variations in the number of nozzles utilized, thrust coefficients (roughly 0 to 4), and angles of attack (-8 to 20 degrees). The run matrix was developed to also allow quantification of various sources of experimental uncertainty, such as random errors due to run-to-run variations and systematic errors due to flowfield or model misalignments. To accommodate the uncertainty assessment, many runs and replicates were conducted with the model at various locations within the tunnel and with model roll angles of 0, 60, 120, and 180 degrees. This test report provides operational details of the experiment, contains a review of trends, and provides all schlieren and pressure results within appendices.

Berry, Scott A.↗
NASA NTRS2020

Model Design and Pre-Test CFD Analysis for a Supersonic Retropropulsion Wind Tunnel Test

Future Mars human lander missions will require using powered descent beginning at supersonic conditions, something which has never been done before at Mars. Significant aerosciences challenges exist due to interactions between the retrorocket exhaust plumes and the freestream flow that alter the aerodynamic behavior of the powered descent vehicle. Historically, wind tunnel tests have been used to study supersonic retropropulsion with inert gas exhaust simulants. Also, SpaceX has successfully decelerated the Falcon 9 first stage numerous times by using powered flight at supersonic conditions. On the computational side, Reynolds-Averaged Navier-Stokes flowfield simulations are regularly completed at full-scale conditions. However, the available ground and flight data do not provide a basis for calibrating the computational uncertainties for aerodynamic interference forces and moments on proposed Mars descent vehicles, either because of insufficient data or dissimilar vehicles geometries and/or conditions. Thus, additional ground testing is needed to continue addressing the aerosciences risks for large-scale human Mars landers. To that end, a retropropulsion wind tunnel test will be conducted in the NASA Langley Unitary Plan Wind Tunnel in 2020. The test is designed to improve upon similar past tests, both in terms of model design and measured data. The test campaign will use subscale model geometries derived from the two current NASA powered descent reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. This paper covers the reference vehicles, test objectives, model design, scaling parameters, test matrix, and computational analysis to date.

Karl T Edquist↗
NASA NTRS2011

Toward Supersonic Retropropulsion CFD Validation

This paper begins the process of verifying and validating computational fluid dynamics (CFD) codes for supersonic retropropulsive flows. Four CFD codes (DPLR, FUN3D, OVERFLOW, and US3D) are used to perform various numerical and physical modeling studies toward the goal of comparing predictions with a wind tunnel experiment specifically designed to support CFD validation. Numerical studies run the gamut in rigor from code-to-code comparisons to observed order-of-accuracy tests. Results indicate that this complex flowfield, involving time-dependent shocks and vortex shedding, design order of accuracy is not clearly evident. Also explored is the extent of physical modeling necessary to predict the salient flowfield features found in high-speed Schlieren images and surface pressure measurements taken during the validation experiment. Physical modeling studies include geometric items such as wind tunnel wall and sting mount interference, as well as turbulence modeling that ranges from a RANS (Reynolds-Averaged Navier-Stokes) 2-equation model to DES (Detached Eddy Simulation) models. These studies indicate that tunnel wall interference is minimal for the cases investigated; model mounting hardware effects are confined to the aft end of the model; and sparse grid resolution and turbulence modeling can damp or entirely dissipate the unsteadiness of this self-excited flow.

Kleb, Bil↗
NASA NTRSDate not supplied

OVERFLOW Analysis of Supersonic Retropropulsion Testingon the CobraMRV Mars Entry Vehicle Concept

The CFD solver OVERFLOW was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a pre-test study. Aerodynamics of the CobraMRV are summarized and the sensitivity of vehicle loads to CFD parameters at a subset of operating conditions are investigated. Specific numerical methods, boundary conditions, and turbulence modeling options have been selected after considering the complex phenomena in SRP flows. Proper shock-capturing methods and dynamic grid adaption are necessary to correctly capture vehicle loads and dynamics. Present data indicates that at particular conditions, upstream asymmetries originating from the tunnel inflow plane contribute to measurable asymmetries on the CobraMRV heatshield.

OVERFLOW↗
NASA NTRS2013

Supersonic Retropropulsion CFD Validation with Ames Unitary Plan Wind Tunnel Test Data

A validation study of Computational Fluid Dynamics (CFD) for Supersonic Retropropulsion (SRP) was conducted using three Navier-Stokes flow solvers (DPLR, FUN3D, and OVERFLOW). The study compared results from the CFD codes to each other and also to wind tunnel test data obtained in the NASA Ames Research Center 90 70 Unitary PlanWind Tunnel. Comparisons include surface pressure coefficient as well as unsteady plume effects, and cover a range of Mach numbers, levels of thrust, and angles of orientation. The comparisons show promising capability of CFD to simulate SRP, and best agreement with the tunnel data exists for the steadier cases of the 1-nozzle and high thrust 3-nozzle configurations.

Schauerhamer, Daniel G.↗
NASA NTRS2012

Supersonic Retropropulsion Technology Development in NASA's Entry, Descent, and Landing Project

NASA's Entry, Descent, and Landing (EDL) space technology roadmap calls for new technologies to achieve human exploration of Mars in the coming decades [1]. One of those technologies, termed Supersonic Retropropulsion (SRP), involves initiation of propulsive deceleration at supersonic Mach numbers. The potential benefits afforded by SRP to improve payload mass and landing precision make the technology attractive for future EDL missions. NASA's EDL project spent two years advancing the technological maturity of SRP for Mars exploration [2-15]. This paper summarizes the technical accomplishments from the project and highlights challenges and recommendations for future SRP technology development programs. These challenges include: developing sufficiently large SRP engines for use on human-scale entry systems; testing and computationally modelling complex and unsteady SRP fluid dynamics; understanding the effects of SRP on entry vehicle stability and controllability; and demonstrating sub-scale SRP entry systems in Earth's atmosphere.

Edquist, Karl T.↗
NASA NTRSDate not supplied

Seeding Method for Velocimetry and Visualization of Supersonic Retropropulsion Nozzle Plumes

In the current work, the “Venturi seeder” method for nozzle plumes is improved and studied to determine suitability for particle image velocimetry. This seeding method involves independently pressurizing the primary flow path and a separate seed liquid reservoir, connecting them at the throat of a Venturi contraction upstream of the nozzle. The pressure differential at the throat causes the liquid to enter the nozzle channel where it is atomized into particles by strong shearing forces. This new seeding system was characterized by conducting a series of diagnostics using different exit nozzle pressure ratios and seed reservoir pressures. Shadowgraph imaging confirmed the nozzle plumes to be underexpanded jets. A particle sizing device determined that generated DEHS oil particles had median aerodynamic diameters of 0.67, 0.69, and 0.73 μm for seed liquid reservoir pressure supplies of 0.67, 0.60, and 0.93 MPa and primary flow path pressures of 0.67, 0.76, and 0.93 MPa measured downstream of the Venturi throat, respectively. These diameters were within the threshold for acceptable response time in typical gas flows (≤ 1 μm). PIV experiments were conducted on the nozzle plume for all presented cases; mean axial and transverse velocities appeared as expected for the underexpanded jet structure, including the Mach disk and re-acceleration regions. Additional sizing analysis based on the particle response times across the normal shock again showed particle diameters to be well less than 1 μm for all cases. The results suggest that this method is an inexpensive and relatively simple solution to the problem of seeding nozzle plumes. The method will be used in testing of supersonic retropropulsion models in the Unitary Plan Wind Tunnel at The NASA Langley Research Center in 2022.

particle image velocimetry↗
NASA NTRS2017

Advancing Supersonic Retropropulsion Using Mars-Relevant Flight Data: An Overview

Advanced robotic and human missions to Mars require landed masses well in excess of current capabilities. One approach to safely land these large payloads on the Martian surface is to extend the propulsive capability currently required during subsonic descent to supersonic initiation velocities. However, until recently, no rocket engine had ever been fired into an opposing supersonic freestream. In September 2013, SpaceX performed the first supersonic retropropulsion (SRP) maneuver to decelerate the entry of the first stage of their Falcon 9 rocket. Since that flight, SpaceX has continued to perform SRP for the reentry of their vehicle first stage, having completed multiple SRP events in Mars-relevant conditions in July 2017. In FY 2014, NASA and SpaceX formed a three-year public-private partnership centered upon SRP data analysis. These activities focused on flight reconstruction, CFD analysis, a visual and infrared imagery campaign, and Mars EDL design analysis. This paper provides an overview of these activities undertaken to advance the technology readiness of Mars SRP.

Braun, Robert D.↗
NASA NTRS2017

Advancing Supersonic Retropropulsion Using Mars-Relevant Flight Data: An Overview

Advanced robotic and human missions to Mars require landed masses well in excess of current capabilities. One approach to safely land these large payloads on the Martian surface is to extend the propulsive capability currently required during subsonic descent to supersonic initiation velocities. However, until recently, no rocket engine had ever been fired into an opposing supersonic freestream. In September 2013, SpaceX performed the first supersonic retropropulsion (SRP) maneuver to decelerate the entry of the first stage of their Falcon 9 rocket. Since that flight, SpaceX has continued to perform SRP for the reentry of their vehicle first stage, having completed multiple SRP events in Mars-relevant conditions in July 2017. In FY 2014, NASA and SpaceX formed a three-year public-private partnership centered upon SRP data analysis. These activities focused on flight reconstruction, CFD analysis, a visual and infrared imagery campaign, and Mars EDL design analysis. This paper provides an overview of these activities undertaken to advance the technology readiness of Mars SRP.

Braun, Robert D.↗
NASA NTRS2014

Entry, Descent, and Landing with Propulsive Deceleration: Supersonic Retropropulsion Wind Tunnel Testing and Shock Phenomena

The future exploration of the Solar System will require innovations in transportation and the use of entry, descent, and landing (EDL) systems at many planetary landing sites. The cost of space missions has always been prohibitive, and using the natural planetary and planet's moon atmospheres for entry, and descent can reduce the cost, mass, and complexity of these missions. This paper will describe some of the EDL ideas for planetary entry and survey the overall technologies for EDL that may be attractive for future Solar System missions. Future EDL systems may include an inflatable decelerator for the initial atmospheric entry and an additional supersonic retropropulsion (SRP) rocket system for the final soft landing. A three engine retropropulsion configuration with a 2.5 in. diameter sphere-cone aeroshell model was tested in the NASA Glenn Research Center's 1- by 1-ft (1×1) Supersonic Wind Tunnel (SWT). The testing was conducted to identify potential blockage issues in the tunnel, and visualize the rocket flow and shock interactions during supersonic and hypersonic entry conditions. Earlier experimental testing of a 70deg Viking-like (sphere-cone) aeroshell was conducted as a baseline for testing of a SRP system. This baseline testing defined the flow field around the aeroshell and from this comparative baseline data, retropropulsion options will be assessed. Images and analyses from the SWT testing with 300- and 500-psia rocket engine chamber pressures are presented here. In addition, special topics of electromagnetic interference with retropropulsion induced shock waves and retropropulsion for Earth launched booster recovery are also addressed.

Palaszewski, Bryan↗
NASA NTRS2017

High Ballistic Coefficient Mars EDL with Supersonic Retropropulsion

Future Mars missions will require the ability to land increasingly heavy systems with high precision, possibly at high altitude landing sites. A key obstacle in landing systems with high ballistic coefficient is Mars’ very thin atmosphere, approximately one hundred times less dense than Earth’s atmosphere. As a result, hypersonic deceleration occurs slowly throughout atmospheric entry and additional deceleration methods are required before landing. Previous landers up to and including MSL relied on Viking-heritage technologies including a Disk-Gap-Band (DGB) supersonic parachute and a blunt body aeroshell. Safe and effective deployment of a DGB parachute can be accomplished only within prescribed bounds on dynamic pressure and Mach number.2 In addition to the restrictive constraint this imposes on the entry trajectory, there is a concern that for a parachute to effectively slow an entry vehicle with a high ballistic coefficient, its diameter must be large enough to impose certification costs, modeling uncertainty, or mission risk that could be unacceptable. This has prompted investigation into alternative mission concepts based on use of supersonic retropropulsion (SRP) without a parachute. This work investigates SRP as an enabling technology for Mars EDL of high ballistic coefficient vehicles.

Noyes, Connor↗
NASA NTRS2018

Supersonic Retropropulsion on Robotic Mars Landers: Selected Design Trades

Many concepts for future robotic Mars lander missions require landing heavier payloads than those landed to date. Mars lander architectures to date have relied on a parachute to help slow the lander; however the effectiveness of a parachute in the thin Martian atmosphere is diminished with heavier payloads unless the diameter of the parachute is increased or it is deployed at a higher Mach number, both of which are significant technical challenges. In addition the parachute can be successfully deployed only within a specific Mach number and dynamic pressure range. Targeting the entry trajectory to hit this “Mach-Q box” imposes constraints on the entry ballistic coefficient, limiting it to ~ 150-200 kg/m^2. Eliminating the parachute from the design requires descent engine ignition at supersonic speeds (Supersonic Retropropulsion, or SRP). SRP increases the propellant requirement, but also allows entry ballistic coefficients of ~600 kg/m^2 or more, with the consequence of significantly increased entry mass and landed payload mass.

Lorenz, Christopher G.↗