Results for “NOZZLE THRUST COEFFICIENT” — Page 4

NASA NTRSDate not supplied

Computational Modeling of Two Mars Powered Descent Vehicle Concepts Tested in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not suitable for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flowfield predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at lower HIAD CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Supersonic Retropropulsion↗

NASA NTRSDate not supplied

Computational Modeling of Two Mars Powered Descent Vehicle Concepts Tested in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not suitable for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flowfield predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at lower HIAD CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Supersonic Retropropulsion↗

NASA NTRS1959

Theoretical Performance of Liquid Hydrogen with Liquid Oxygen as a Rocket Propellant

Theoretical rocket performance for both equilibrium and frozen composition during expansion was calculated for the propellant combination liquid hydrogen and liquid oxygen at four chamber pressures (60, 150, 300, and 600 lb/sq in. abs) and a wide range of pressure ratios (1 to 4000) and oxidant-fuel ratios (1.190 to 39.683). Data are given to estimate performance parameters at chamber pressures other than those for which data are tabulated. The parameters included are specific impulse, specific impulse in vacuum, combustion-chamber temperature, nozzle-exit temperature, molecular weight, molecular-weight derivatives, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, thermal conductivity, Mach number, and equilibrium gas compositions.

Gordon, Sanford↗

NASA NTRS1996

Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure

An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.

Jankovsky, Robert S.↗

NASA NTRS1978

Aerodynamic characteristics induced on a supercritical wing due to vectoring twin nozzles at Mach numbers from 0.40 to 0.95

The induced lift characteristics are studied for a vectored-thrust concept in which jet-exhaust nozzles were located in the fuselage at the wing trailing edge. The wing had a supercritical airfoil section. The investigation was conducted at Mach numbers from 0.4 to 0.95, angles of attack up to 14 deg, and thrust coefficients up to 0.35, and nozzle deflection angles of 0 deg and 30 deg. Separate force balances were used to determine both total aerodynamic and thrust forces alone which allowed for a direct measurement of jet turning angle at forward speeds. The Reynolds number per meter varied from 8.20 x 1 million to 12.80 x 1 million. The results show that the configuration with the supercritical wing generally had a better performance with respect to both lift augmentation and drag reduction than the same configuration with a 64 series airfoil.

Capone, F. J.↗

NASA NTRS1951

Theoretical performance of lithium and fluorine as a rocket propellant

Theoretical performance for liquid lithium and liquid fluorine as a rocket propellant was calculated with assumptions both of equilibrium and frozen composition during expansion. Parameters included were specific impulse, combustion-chamber temperature, nozzle-exit temperature, composition, mean molecular weight, characteristic velocity, coefficient of thrust, and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, the maximum equilibrium specific impulse calculated was 335.5 pound-seconds per pound. The effect of ionization on calculated performance was shown to be negligible by comparison of values of various parameters calculated both with and without ionized products of combustion.

FUELS -ROCKETS (INCLUDES FUELS AND OXIDANT)↗

NASA NTRS1952

Theoretical performance of liquid ammonia, hydrazine and mixture of liquid ammonia and hydrazine as fuels with liquid oxygen biflouride as oxidant for rocket engines : I-mixture of liquid ammonia and hydrazine

Theoretical performance for mixture of 36.3 percent liquid ammonia and 63.7 percent hydrazine with liquid oxygen bifluoride as rocket propellant was calculated on assumption of equilibrium composition during expansion for a wide range of fuel-oxidant and expansios ratios. Parameters included were specific impulse, combustion-chamber temperature, nozzle exit temperature, composition mean molecular weight, characteristic velocity, coefficient of thrust and ratio of nozzle-exit area to throat area. For chamber pressure of 300 pounds per square inch absolute and expansion to 1 atmosphere, maximum specific impulse was 295.8 pound-seconds per pound. Five percent by weight of water in the hydrazine lowered specific impulse from about one to three units over a wide range of weight-percent fuel.

Huff, Vearl N↗

NASA NTRS1956

Internal performance characteristics of short convergent-divergent exhaust nozzles designed by the method of characteristics

Internal performance data on a short exhaust nozzle designed by the method of characteristics were obtained over a range of pressure ratios from 1.5 to 22. The peak thrust coefficient was not affected by a shortened divergent section, but it occurred at lower pressure ratios due to reduction in expansion ratio. This nozzle contour based on characteristics solution gave higher thrust coefficients than a conical convergent-divergent nozzle of equivalent length. Abrupt-inlet sections permitted a reduction in nozzle length without a thrust-coefficient reduction.

Krull, H George↗

NASA NTRS1972

Internal performance of a 10 deg conical plug nozzle with a multispoke primary and translating external shroud

An experimental investigation was conducted in a nozzle static test facility to determine the performance characteristics of a cold-flow, 21.59-centimeter-diameter plug nozzle with a multispoke primary. Two multispoke primary nozzles, a 12-spoke and a 24-spoke, were tested and compared with an annular plug nozzle. The supersonic cruise configurations for both spoke primaries performed about the same, with a gross thrust coefficient of 0.974, a decrease of approximately 1.5 percent from the reference nozzle. The takeoff configuration for the 12-spoke primary had a gross thrust coefficient of 0.957, a decrease of 1.5 percent from the reference nozzle, and the 24-spoke primary had a gross thrust coefficient of 0.95.

Bresnahan, D. L.↗

NASA NTRS1983

Ejector nozzle test results at simulated flight conditions for an advanced supersonic transport propulsion system

Results are presented of wind tunnel tests conducted to verify the performance improvements of a refined ejector nozzle design for advanced supersonic transport propulsion systems. The analysis of results obtained at simulated engine operating conditions is emphasized. Tests were conducted with models of approximately 1/10th scale which were configured to simulate nozzle operation at takeoff, subsonic cruise, transonic cruise, and supersonic cruise. Transonic cruise operation was not a consideration during the nozzle design phase, although an evaluation at this condition was later conducted. Test results, characterized by thrust and flow coefficients, are given for a range of nozzle pressure ratios, emphasizing the thrust performance at the engine operating conditions predicted for each flight Mach number. The results indicate that nozzle performance goals were met or closely approximated at takeoff and supersonic cruise, while subsonic cruise performance was within 2.3 percent of the goal with further improvement possible.

Nelson, D. P.↗

NASA NTRS2014

Computational Simulations of Convergent Nozzles for the AIAA 1st Propulsion Aerodynamics Workshop

Computational Fluid Dynamics (CFD) simulations were completed for a series of convergent nozzles in participation of the American Institute of Aeronautics and Astronautics (AIAA) 1st Propulsion Aerodynamics Workshop. The simulations were performed using the Wind-US flow solver. Discharge and thrust coefficients were computed for four axisymmetric nozzles with nozzle pressure ratios (NPR) ranging from 1.4 to 7.0. The computed discharge coefficients showed excellent agreement with available experimental data; the computed thrust coefficients captured trends observed in the experimental data, but over-predicted the thrust coefficient by 0.25 to 1.0 percent. Sonic lines were computed for cases with NPR >= 2.0 and agreed well with experimental data for NPR >= 2.5. Simulations were also performed for a 25 deg. conic nozzle bifurcated by a flat plate at NPR = 4.0. The jet plume shock structure was compared with and without the splitter plate to the experimental data. The Wind-US simulations predicted the shock structure well, though lack of grid resolution in the plume reduced the sharpness of the shock waves. Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations and Detached Eddy Simulations (DES) were performed at NPR = 1.6 for the 25 deg conic nozzle with splitter plate. The simulations predicted vortex shedding from the trailing edge of the splitter plate. However, the vortices of URANS and DES solutions appeared to dissipate earlier than observed experimentally. It is believed that a lack of grid resolution in the region of the vortex shedding may have caused the vortices to break down too soon

DES↗

NASA NTRS1995

Internal performance characteristics of thrust-vectored axisymmetric ejector nozzles

A series of thrust-vectored axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at the Langley research center. This study indicated that discontinuities in the performance occurred at low primary nozzle pressure ratios and that these discontinuities were mitigated by decreasing expansion area ratio. The addition of secondary flow increased the performance of the nozzles. The mid-to-high range of secondary flow provided the most overall improvements, and the greatest improvements were seen for the largest ejector area ratio. Thrust vectoring the ejector nozzles caused a reduction in performance and discharge coefficient. With or without secondary flow, the vectored ejector nozzles produced thrust vector angles that were equivalent to or greater than the geometric turning angle. With or without secondary flow, spacing ratio (ejector passage symmetry) had little effect on performance (gross thrust ratio), discharge coefficient, or thrust vector angle. For the unvectored ejectors, a small amount of secondary flow was sufficient to reduce the pressure levels on the shroud to provide cooling, but for the vectored ejector nozzles, a larger amount of secondary air was required to reduce the pressure levels to provide cooling.

Lamb, Milton↗

NASA NTRS1993

Experimental investigation of spherical-convergent-flap thrust-vectoring two-dimensional plug nozzles

An experimental investigation of multiaxis thrust vectoring nozzles with spherical convergent flaps (SCF) and a convertible (in-flight deployable) centerline plug were tested in the NASA-Langley 16-Foot Transonic Tunnel Static Test Facility. Parameters tested during the static test included plug length, external shroud length, shroud internal angle, and yaw vector angle. Results indicated that the SCF convertible-plug nozzle flow is highly three dimensional and characterized by internal flow separation and shock effects. The resultant thrust and discharge coefficient levels were lower than previous wedge nozzle results. In addition, increasing shroud internal angle increased resultant thrust ratio but decreased discharge coefficient. Increasing shroud length increased resultant thrust ratio performance for NPR's greater than 6.0 but for NPR's less than 6.0, resultant thrust ratio decreased as shroud length increased. Increasing plug length caused an increase in resultant thrust ratio and an increase in discharge coefficient. Increasing geometric yaw vector angle had no effect on resultant thrust ratio and little effect on discharge coefficient.

Cler, Daniel L.↗

NASA NTRS2004

Generation 1.5 High Speed Civil Transport (HSCT) Exhaust Nozzle Program

The objective of this program was to conduct an experimental and analytical evaluation of low noise exhaust nozzles suitable for future High-Speed Civil Transport (HSCT) aircraft. The experimental portion of the program involved parametric subscale performance model tests of mixer/ejector nozzles in the takeoff mode, and high-speed tests of mixer/ejectors converted to two-dimensional convergent-divergent (2-D/C-D), plug, and single expansion ramp nozzles (SERN) in the cruise mode. Mixer/ejector results show measured static thrust coefficients at secondary flow entrainment levels of 70 percent of primary flow. Results of the high-speed performance tests showed that relatively long, straight-wall, C-D nozzles could meet supersonic cruise thrust coefficient goal of 0.982; but the plug, ramp, and shorter C-D nozzles required isentropic contours to reach the same level of performance. The computational fluid dynamic (CFD) study accurately predicted mixer/ejector pressure distributions and shock locations. Heat transfer studies showed that a combination of insulation and convective cooling was more effective than film cooling for nonafterburning, low-noise nozzles. The thrust augmentation study indicated potential benefits for use of ejector nozzles in the subsonic cruise mode if the ejector inlet contains a sonic throat plane.

Thayer, E. B.↗

NASA NTRS1993

The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

Kacynski, Kenneth J.↗

NASA NTRS1980

The effect of throat contouring on two-dimensional converging-diverging nozzles at static conditions

An experiment was conducted at static conditions to determine the internal performance effects of nozzle throat contouring, the result of increasing the circular-arc throat radius. Five nonaxisymmetric converging-diverging nozzles were tested at nozzle pressure ratios up to 9.0. Data are presented as internal thrust ratios, discharge coefficients, and static-pressure distributions. Comparisons of internal performance data for the five nozzles show that throat contouring increases the value of discharge coefficient but has no significant effect on internal thrust ratio except in cases of internal flow separation. To illustrate the use of the two dimensional converging-diverging (2-D C-D) nozzle data base, a two dimensional inviscid theory was applied to the five configurations. The generally good agreement of data with theoretical results indicates that two-dimensional inviscid theory can be successfully applied to the prediction of 2-D C-D nozzle internal flow.

Mason, M. L.↗

NASA NTRSDate not supplied

Optimization of Plug Nozzles Using CFD Informed Design Space Querying

This paper presents an axisymmetric CFD approach based on Reynolds-averaged Navier-Stokes (RANS) calculations for the shape optimization of plug nozzle contours for both supersonic cruise and landing/takeoff (LTO) conditions. Three key external plug design parameters were defined as independent variables in the optimization scheme, while nozzle performance was evaluated using thrust coefficient (Cfg) and discharge coefficient (Cd) as the primary dependent variables. By varying these parameters, a wide range of axisymmetric conic section plug profiles can be achieved, simplifying the problem. The paper also discusses the supersonic plug nozzle optimization approach, design space seeding, meshing, and CFD simulations for the optimization process. The results demonstrate that the approach is effective in exploring the parameter space and reducing computational expenses, making it a viable option for optimizing nozzle plug contours. Using this optimization toolchain, the study yielded preliminary designs for LTO, cruise, and LTO-cruise balanced plug nozzles.

nozzles↗

NASA NTRSDate not supplied

Optimization of Plug Nozzles Using CFD Informed Design Space Querying

This paper presents an axisymmetric CFD approach based on Reynolds-averaged Navier-Stokes (RANS) calculations for the shape optimization of plug nozzle contours for both supersonic cruise and landing/takeoff (LTO) conditions. Three key external plug design parameters were defined as independent variables in the optimization scheme, while nozzle performance was evaluated using thrust coefficient (Cfg) and discharge coefficient (Cd) as the primary dependent variables. By varying these parameters, a wide range of axisymmetric conic section plug profiles can be achieved, simplifying the problem. The paper also discusses the supersonic plug nozzle optimization approach, design space seeding, meshing, and CFD simulations for the optimization process. The results demonstrate that the approach is effective in exploring the parameter space and reducing computational expenses, making it a viable option for optimizing nozzle plug contours. Using this optimization toolchain, the study yielded preliminary designs for LTO, cruise, and LTO-cruise balanced plug nozzles.

CFD↗