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NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of these samples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system. The results of these studies will also be examined in this paper.

Darius Yaghoubi↗

NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of thesesamples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system.

Mars↗

NASA NTRS1992

Critical effects of downstream boundary conditions on vortex breakdown

The unsteady, compressible, full Navier-Stokes (NS) equations are used to study the critical effects of the downstream boundary conditions on the supersonic vortex breakdown. The present study is applied to two supersonic vortex breakdown cases. In the first case, quasi-axisymmetric supersonic swirling flow is considered in a configured circular duct, and in the second case, quasi-axisymmetric supersonic swirling jet, that is issued from a nozzle into a supersonic jet of lower Mach number, is considered. For the configured duct flow, four different types of downstream boundary conditions are used, and for the swirling jet flow from the nozzle, two types of downstream boundary conditions are used. The solutions are time accurate which are obtained using an implicit, upwind, flux-difference splitting, finite-volume scheme.

Kandil, Osama↗

NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Four deflected thrust nozzle concepts, designed to operate at the low pressure ratio typical of high bypass-ratio turbofan engines for medium speed (subsonic) V/STOL aircraft, were studied. Maps of overall performance characteristics and exit velocity distributions are used to highlight similarities and differences between the four concepts. Analytically determined secondary flows at the exit of a 90 deg circular pipe bend are compared with the experimental results from the more complex three dimensional geometries. The relative impact of total-pressure losses and secondary flows on nozzle thrust coefficient is addressed by numerical integration of exit velocity measurements.

Burstadt, P. L.↗

NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Burstadt, P. L.↗

NASA NTRS1973

Effect of fineness ratio on boattail drag of circular-arc afterbodies having closure ratios of 0.50 with jet exhaust at Mach numbers up to 1.30

An investigation was conducted to determine the effect of fineness ratio on the drag of circular-arc boattails at subsonic and low supersonic speeds. The boattails had closure ratios of 0.50 and incorporated convergent nozzles. The investigation was conducted statically and at Mach numbers from 0.40 to 1.30 at 0 deg angle of attack with jet total-pressure ratios varying from jet off to about 6, depending on Mach number. Low-fineness-ratio boattails had large separated-flow regions and the highest drag at all Mach numbers. Subsonic pressure-plus-friction drag levels were generally similar for boattails which did not have large separated regions. Drag-rise Mach number increased as boattail fineness ratio increased.

Reubush, D. E.↗

NASA NTRS1976

STAN5: A program for numerical computation of two-dimensional internal and external boundary layer flows

A large variety of two dimensional flows can be accommodated by the program, including boundary layers on a flat plate, flow inside nozzles and diffusers (for a prescribed potential flow distribution), flow over axisymmetric bodies, and developing and fully developed flow inside circular pipes and flat ducts. The flows may be laminar or turbulent, and provision is made to handle transition.

Crawford, M. E.↗

NASA NTRS1992

Nozzle/Diffuser For Test Section Of Wind Tunnel

Nozzle-diffuser improves distribution of flow in test section and increases recovery of pressure in first diffuser. Has rectangular or circular cross section like contraction duct just upstream of test section and is, in effect, extension of contraction duct. Because flow returning to fan at higher pressure, less fan power needed to maintain same rate of flow.

Barna, P. Stephen↗

NASA NTRS1986

Effects of initial condition on circular and elliptic jet noise

Noise produced by cold elliptic jets was studied experimentally using a 4 cm air jet in an anechoic chamber by varying the aspect ratio, nozzle exit boundary layer, and Mach number. Elliptic geometry is found to accentuate the sound produced by pairing of the shear layer vortices in the laminar boundary jet. Calculation of the pressure field produced by the pairing of elliptic vortex rings shows that the ratio of sound intensities produced by the pairing of inviscid elliptic vortex filament rings should be equal to their aspect ratio, the minor axis producing greater sound. Tripping the nozzle boundary layer changes the instability of the shear layer and removes the laminar vortices. If the sound produced by pairing and/or breakdown is considered to be the difference between laminar and turbulent jet noise, the pairing mechanism is seen to be very significant in laminar boundary layer jets.

Bridges, J. E.↗

NASA NTRS1979

Asymmetry of a circular jet observed in near and far fields

Turbulence intensity measurements were made in three different-size jets at various Mach numbers and the aerodynamic jet axis, defined as the locus of points in the flow at which the turbulence intensity is minimum, was determined. The deviations from the geometrical jet axis were on the order of 2-4% of the nozzle diameter, and the pure jet axis had the form of a helical wave. A similar asymmetry in the radiated noise was also noted. The measured circumferential distribution of rms fluctuating velocity signals confirmed the asymmetry of even a circular jet.

Yamamoto, K.↗

NASA NTRS1972

Normal impingement loads due to small air jets issuing from a base plate and reflecting off a platform for various jet Mach numbers, separation distances, and ambient pressures

An investigation was conducted in a 12.5-meter-diameter vacuum sphere to determine the impingement loads due to air jets issuing from and perpendicular to a circular base and reflecting off a square platform, that is, a simulation of rendezvous maneuvering, docking, launch, impact dampers etc. The nozzles had exit Mach numbers of 1, 3, 5, and 7. The ambient pressures were 0.0006, 5, 225, and 760 torr. Under near-field separation distances and at 0.0006 torr, reflections were significant; and ratios of the impingement force to thrust on both plates in the biplane arrangement varied from about 750 for exit Mach number 1 to 120 for exit Mach number 7. The far-field force ratios were near unity for the platform and zero for the base and indicated few, if any, reflections. Some reversals and rapid changes in loads were obtained at transition distances between the near and far fields. In general, increasing the exit Mach number or ambient pressure reduced the impingement loads.

Hoffman, S.↗

NASA NTRS1982

The self-excited axisymmetric jet

The results of a number of investigations suggest that large-scale coherent structures and their interactions play key roles in the transport of heat, mass, and momentum, and in the generation of aerodynamic noise. Investigations related to the study of a large-scale coherent structure in the circular jet are considered, taking into account the advantages of inducing controlled perturbation through self-sustained excitation with the whistler nozzle. The primary objective of the present study was to document the jet response to self-excitation and the sensitivity of this response to the initial condition and the Reynolds number R(D). It was also felt that the results would find use in the control of mixing and aerodynamic noise in a jet.

Hasan, M. A. Z.↗

NASA NTRS1991

Computation of a circular-to-rectangular transition duct flow field

This paper presents the results of a Computational Fluid Dynamics (CFD) calibration study of flow through a circular-to-rectangular transition duct. The design of these ducts is critical to the optimum performance of aircraft with rectangular exhaust nozzles, since these ducts transfer the flow from the gas turbine engine to the exhaust nozzle. Two duct inflow conditions are considered, the first with straight inflow and the second with swirling inflow. Both flows contain realistic wall boundary layers. The first case permits examination of the effects of the geometric transition on the flowfield, while the second case adds in the rotational flow effect that can be present from the gas turbine engine. The flowfields associated with the two cases have been shown, both with CFD and experiment, to have significant differences. The Navier-Stokes CFD code PARC with the Baldwin and Lomax turbulence model was used in this study. The turbulence model had to be modified for both cases in order to achieve accurate determination of the edge of the wall bounded vorticity layers and thus turbulent viscosity. The results of this calibration study will be valuable to aircraft designers who rely on CFD to assist in the design and evaluation of propulsion systems.

Sirbaugh, J. R.↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad/Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Panel test articles included a metallic separation bolt imbedded in the compression-pad and heat shield materials, resulting in a circular protuberance over a flat plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the non-equilibrium flow field in the facility nozzle, test box, and flow field over test articles, and comparisons with the measured calibration data.

TPS↗

NASA NTRS1994

Simultaneous three-dimensional velocity and mixing measurements by use of laser Doppler velocimetry and fluorescence probes in a water tunnel

A water tunnel investigation was conducted to demonstrate the capabilities of a laser-based instrument that can measure velocity and fluorescence intensity simultaneously. Fluorescence intensity of an excited fluorescent dye is directly related to concentration level and is used to indicate the extent of mixing in flow. This instrument is a three-dimensional laser Doppler velocimeter (LDV) in combination with a fluorometer for measuring fluorescence intensity variations. This capability allows simultaneous flow measurements of the three orthogonal velocity components and mixing within the same region. Two different flows which were generated by two models were studied: a generic nonaxisymmetric nozzle propulsion simulation model with an auxiliary internal water source that generated a jet flow and an axisymmetric forebody model with a circular sector strake that generated a vortex flow. The off-body flow fields around these models were investigated in the Langley 16- by 24-Inch Water Tunnel. The experimental results were used to calculate 17 quantities that included mean and fluctuating velocities, Reynolds stresses, mean and fluctuating dye fluorescence intensities (proportional to concentration), and fluctuating velocity and dye concentration correlations. An uncertainty analysis was performed to establish confidence levels in the experimental results. In general, uncertainties in mean velocities varied between 1 and 7 percent of free-stream velocity; uncertainties in fluctuating velocities varied between 1 and 5 percent of reference values. The results show characteristics that are unique to each type of flow.

Neuhart, Dan H.↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad/Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Each panel test article included a metallic separation bolt imbedded in Orion compression-pad and heatshield materials, resulting in a circular protuberance over a flat plate. The protuberances produce complex model flowfields, containing shock-shock and shock-boundary layer interactions, and multiple augmented heating regions on the test plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the nonequilibrium flowfield in the facility nozzle, test box, and flowfield over test articles, and comparisons with the measured calibration data.

CFD↗

NASA NTRS2017

CFD Simulations of the IHF Arc-Jet Flow: Compression-Pad Separation Bolt Wedge Tests

This paper reports computational analyses in support of two wedge tests in a high enthalpy arc-jet facility at NASA Ames Research Center. These tests were conducted using two different wedge models, each placed in a free jet downstream of a corresponding different conical nozzle in the Ames 60-MW Interaction Heating Facility. Each panel test article included a metallic separation bolt imbedded in Orion compression-pad and heatshield materials, resulting in a circular protuberance over a flat plate. The protuberances produce complex model flowfields, containing shock-shock and shock-boundary layer interactions, and multiple augmented heating regions on the test plate. As part of the test calibration runs, surface pressure and heat flux measurements on water-cooled calibration plates integrated with the wedge models were also obtained. Surface heating distributions on the test articles as well as arc-jet test environment parameters for each test configuration are obtained through computational fluid dynamics simulations, consistent with the facility and calibration measurements. The present analysis comprises simulations of the non-equilibrium flow field in the facility nozzle, test box, and flow field over test articles, and comparisons with the measured calibration data.

CFD↗

NASA NTRS1974

Solid rocket booster performance evaluation model. Volume 3: Sample case

The solid rocket booster performance evaluation model (SRB-11) is used to predict internal ballistics in a sample motor. This motor contains a five segmented grain. The first segment has a 14 pointed star configuration with a web which wraps partially around the forward dome. The other segments are circular in cross-section and are tapered along the interior burning surface. Two of the segments are inhibited on the forward face. The nozzle is not assumed to be submerged. The performance prediction is broken into two simulation parts: the delivered end item specific impulse and the propellant properties which are required as inputs for the internal ballistics module are determined; and the internal ballistics for the entire burn duration of the motor are simulated.

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