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The work focuses on the characterization of the flowfield over a truncated linear aerospike by combining theoretical grounds, numerical simulations and experimental tests. The experimental investigations are carried out on a test rig designed at Politecnico di Torino for advanced nozzle testing. Fully three-dimensional CFD analyses are performed on the actual geometry of the experimental nozzle model. At low nozzle pressure ratios (nprs) the analysis combines numerical simulations and experimental testing, which are also used for validating the CFD results. At higher nprs, the flowfield characterization is performed only by three-dimensional CFD analyses. In addition to the validation of the numerical method, the edge effects at different nprs have been observed. [ABSTRACT FROM AUTHOR]

The work focuses on the characterization of the flowfield over a truncated linear aerospike by combining theoretical grounds, numerical simulations and experimental tests. The experimental investigations are carried out on a test rig designed at Politecnico di Torino for advanced nozzle testing. Fully three-dimensional CFD analyses are performed on the actual geometry of the experimental nozzle model. At low nozzle pressure ratios (nprs) the analysis combines numerical simulations and experimental testing, which are also used for validating the CFD results. At higher nprs, the flowfield characterization is performed only by three-dimensional CFD analyses. In addition to the validation of the numerical method, the edge effects at different nprs have been observed.

Turbine blades and rocket nozzles can be efficiently protected from combustion chamber exhaust hot gases by film cooling technic. A crescent-shaped block placed downstream of the injection hole can significantly improve cooling effectiveness. The main objective of this research work is to investigate the influence of the crescent-shaped block length and width on film cooling effectiveness. ANSYS CFX was used to conduct analysis of a flat plate configuration with cylindrical holes. Nine configurations of the crescent shaped blocks were considered. For each case, the effect of blowing ratios (0.5 and 1) was investigated. The turbulence model shear stress transport (SST) is used for approximating turbulence. Good agreement was obtained by comparing the analysis results with the experimental data. The result indicated that block width variation has a considerable impact on film cooling performance. However, slight effect of block length on cooling effectiveness was obtained. Comparing all analyzed configurations, the best cooling effectiveness was reached for the model 9 (W = 3d, B = 1.5d). [ABSTRACT FROM AUTHOR]

To enhance the thermal life of rocket exhaust nozzles, the hot side of copper liners is coated with thermal barrier coatings (TBCs) to provide thermal insulation and oxidation resistance. However, interface failures often occur between M-CrAlY bond coats and nozzle liners due to significant differences in their thermal expansion coefficients (CTE). This study explores the use of Laser Powder-Directed Energy Deposition (LP-DED) to clad NiCrAlY onto a CuCrZr substrate, as the process offers localized heating which can offer better bond strength. Optimization trials were conducted using single and multi-track studies to identify optimal parameters. Due to the low energy absorption of the CuCrZr substrate to 1070 nm laser sources, cladding was performed at a high energy density of 135 J/mm2 with a 1.2 g/min feed rate to achieve defect-free clads with sufficient diffusion. The bulk of the NiCrAlY clads showed γ′-Ni 3 Al, β-NiAl, and γ-Ni phases, while Y 4 Al 2 O 9 and Y 2 O 3 oxides formed on the top surface due to aluminum and yttrium depletion at high temperatures. The clads exhibited cellular dendritic microstructures at the bulk region, and planar microstructures were observed at the dilution zone. EBSD-KAM maps showed higher dislocation density near the interface due to CTE mismatch across substrate and clad. Scratch tests confirmed strong adhesion with no interface cracks, though crack propagation was observed from the edges after 50 isothermal cycles, driven by copper erosion. With Cu diffusion, interface region exhibited a graded microstructure which could enhance CTE, improving compatibility compared to standard NiCrAlY alloys. • The bulk clad region possessed cellular dendritic structures as a result of rapid solidification. • Conversion of β-NiAl to γ′-Ni3Al and Y 4 Al 2 O 9 was observed at the surface of clad. • Diffusion of Cu resulted in the formation of planar microstructure near the interface. • Bimetallic clad exhibited superior adhesion as a result of graded microstructure. [ABSTRACT FROM AUTHOR]

Thrust augmentation in rocket engines using secondary injection in the diverging part of a nozzle is an innovative extension of after burners. This technique finds application in single stage to orbit propulsion devices, where the nozzle has to work at varying ambient pressures. Experimental and numerical studies have been conducted with varying cross flow and injection conditions to analyse the performance augmentation in a 2D nozzle. Schlieren images and wall pressure data are obtained from the experiment. Simulations are con-ducted using a HLLC scheme based finite volume solver. A detailed description of flow physics resulting due to the introduction of sonic angled jet into expanding supersonic flow is presented. It is found that the injection angle, pressure and main flow pressure have notable influence in the performance of the nozzle.

It is well known that the dynamic performance of a rocket or launch vehicle is enhanced when the length of the divergent section of its nozzle is reduced or the nozzle exit area ratio is increased. However, there exists a significant performance trade-off in such rocket nozzle designs due to the presence of random side loads under overexpanded nozzle operating conditions. Flow separation and the associated side-load phenomena have been extensively investigated over the past five decades; however, not much has been reported on the effect of side loads on the attitude dynamics of rocket or launch vehicle. This paper presents a quantitative investigation on the influence of in-nozzle random side loads on the attitude dynamics of a launch vehicle. The attitude dynamics of launch vehicle motion is captured using variable-mass control-volume formulation on a cylindrical rigid sounding rocket model. A novel physics-based stochastic model of nozzle side-load force is developed and embedded in the rigid-body model of rocket. The mathematical model, computational scheme, and results corresponding to side loading scenario are subsequently discussed. The results highlight the influence of in-nozzle random side loads on the roll, pitch, yaw, and translational dynamics of a rigid-body rocket model. [ABSTRACT FROM AUTHOR]

Abstract A numerical investigation is conducted to study an H 2 − O 2 rocket nozzle flow in chemical and vibrational nonequilibrium. Therefore, a 9 reactions kinetic model was implemented in our two-temperature house code. The vibrational relaxation times taken from Skrebkov's theoretical model and utilized in our simulations are found to be better suited for the H 2 − O 2 mixture than those evaluated using Millikan & White semi-empiric formula. The utilized 9 reactions kinetic model demonstrates a good modeling of the chemical nonequilibrium. The results show the presence of three regions in propulsive nozzles: an equilibrium region, a nonequilibrium region followed by a return to equilibrium region. Vibrational nonequilibrium effects on flowfield parameters and nozzle performances in fuel-rich flows of H 2 − O 2 rocket are investigated, by comparing this baseline simulation to vibrational equilibrium simulation. Vibrational nonequilibrium effects on flowfield parameters and on nozzle performances are computed and shown to be minor. A reduction of 90% of computation time is observed when using the vibrational equilibrium configuration instead of the vibrational nonequilibrium configuration. Highlights • Vibrational relaxation times from Skrebkov's model are found to be better suited. • Vibrational equilibrium, nonequilibrium and return to equilibrium regions are detected. • Vibrational nonequilibrium effects are computed and shown to be minor. • Vibrational equilibrium configuration reduces the computation time by 90%. [ABSTRACT FROM AUTHOR]

Single stage to orbit propulsion devices are being developed as part of low cost access to space endeavors. Sea level operation of high area ratio rocket nozzle used in rocket engines leads to an overexpanded flow condition resulting in high side loads. Secondary injection of propellants in high area ratio nozzle is an attractive option to overcome the inefficiency of operation of such nozzles in sea level conditions in addition to the augmentation of thrust. A numerical study on thrust augmentation in high area ratio nozzle by secondary injection of propellants is presented here. The turbulent compressible reacting flow in rocket nozzle with auxiliary injection is simulated using conservation equations for chemical species based on finite rate chemistry model and compressible Navier-Stokes equations with AUSM+-up upwind scheme based unstuctured finite volume solver. An optimized eight step, six species reduced H2-O2finite chemistry reaction model is used to model the supersonic combustion. The indigenously developed solver has an efficient rescaling algorithm to alleviate the effect of stiffness in conventional explicit algorithm for simultaneous solution of reacting flow. The code is validated using the wall pressure and hydrogen concentration values reported for the similar high area ratio rocket nozzle. Accurate prediction of nozzle performance is possible with present turbulent reacting flow simulation as it take care of all losses in nozzle flow. Extensive computations have been performed for the performance estimation of high area ratio rocket nozzle for various prospective auxiliary injection options.

Purpose The purpose of this paper is to design a double parabolic nozzle and to compare the performance with conventional nozzle designs.Design/methodology/approach The throat diameter and divergent length for Conical, Bell and Double Parabolic nozzles were kept same for the sake of comparison. The double parabolic nozzle has been designed in such a way that the maximum slope of the divergent curve is taken as one-third of the Prandtl Meyer (PM) angle. The studies were carried out at Nozzle Pressure Ratio (NPR) of 5 and also at design conditions (NPR = 3.7). Experimental measurements were carried out for all the three nozzle configurations and the performance parameters compared. Numerical simulations were also carried out in a two-dimensional computational domain incorporating density-based solver with RANS equations and SST k-ω turbulence model.Findings The numerical predictions were found to be in reasonable agreement with the measured experimental values. An enhancement in thrust was observed for double parabolic nozzle when compared with that of conical and bell nozzles.Research limitations/implications Even though the present numerical simulations were capable of predicting shock cell parameters reasonably well, shock oscillations were not captured.Practical implications The double parabolic nozzle design has enormous practical importance as a small increase in thrust can result in a significant gain in pay load.Social implications The thrust developed by the double parabolic nozzle is seen to be on the higher side than that of conventional nozzles with better fuel economy.Originality/value The overall performance of the double parabolic nozzle is better than conical and bell nozzles for the same throat diameter and length. [ABSTRACT FROM AUTHOR]

This work deals with experimental characterisation of ultra-high-temperature ceramic materials for aerospace applications. Current research activities are focused on development and testing of new-class ultra-high-temperature ceramic matrix composites for near-zero ablation Thermal Protection Systems and near-zero erosion rocket nozzles. To investigate materials behaviour in atmospheric re-entry conditions, relevant tests are carried out on samples with different shapes with a supersonic arc-jet facility, with specific total enthalpies higher than 20 MJ kg−1 and temperatures over 2000 K in a gas atmosphere with high concentration of atomic oxygen. The Aerospace Propulsion Laboratory allows investigating rocket components or subsystems manufactured in innovative materials, such as nozzles and nozzle inserts, but also small specimens exposed to the rocket exhaust plume, in highly relevant operating conditions. Computational models for numerical simulations of high-enthalpy flows, in both arc-jet and propulsion environments, and thermal analyses of the material samples are developed to support the experimental activities. [ABSTRACT FROM AUTHOR]

A gas generating composition (GGC) was designed with Guanidinium azotetrazolate (GZT) as fuel and Ammonium perchlorate (AP) as an oxidizer for burst testing of nozzle closure at rig level, simulating solid rocket motor ignition transient and propellant flame temperature. EXPLO5 software code was used for formulation studies. The selected GGC formulation was subjected to thermokinetic analysis for understanding the thermal decomposition behaviour as well as evaluating kinetic and thermodynamic parameters of decomposition.The non‐isothermal kinetic parameters of the composition over the extent of conversion were determined using two isocoversional methods viz; Flynn‐Wall‐Ozawa (FWO) and Kissinger–Akahira–Sunose (KAS). GGC in pellet form was fired using a pyrotechnic pressure generating device (PPG) in test rig level with nozzle closure. The nozzle closure was burst opened at a pressure of 0.985 MPa with a pressurization rate of 10.3 MPa s−1. [ABSTRACT FROM AUTHOR]

Abstract The efficiency of an axially symmetric convergent-divergent micronozzle expander at zero ambient pressure is studied. Its contribution to total thrust is considered by analyzing the relative thrust determined as thrust of the expander computed to nozzle thrust without it. An inviscid one-dimensional flow was considered as the starting point for the subsequent analysis. For that inviscid one-dimensional flow model, expander of infinite length obtains finite relative thrust. Theoretical analysis of shear stress influence shows the existence of an optimal length for the expander. The value of the optimal length evidently depends on the shape of the expander wall. This conclusion was numerically confirmed by results from numerical simulations made for various axisymmetric nozzles. Navier-Stokes equations were used for the numerical simulations. Detailed analysis of flow properties shows that the expander having a wall partially inclined to the axis of symmetry has sometimes higher relative thrust than a similar traditional one due to the appearance of local circulation zones with non-zero bottom pressure resulting in additional thrust force. Reliable numerical analysis of micronozzle flows allows avoiding nozzle "blockage" with corresponding crucial decrease in thrust force and loss of spacecraft maneuverability. In addition, discussed circulation zones can negatively influence optical equipment and solar batteries of a space vehicle due to the low speed of exhaust gas flows near their surface and, consequently, deposition of fuel particles. So the problem of micronozzle optimization evidently belongs to space flight safety issues. Highlights • The thrust of convergent-divergent micronozzle is studied numerically. • The expander with a wall partially inclined to the axis of symmetry was analyzed. • That expander obtains sometimes higher relative thrust than a traditional one. • Additional thrust was obtained due to circulation zones with non-zero bottom pressure. • Achieved micronozzle optimization can be applied to problems of space flight safety. [ABSTRACT FROM AUTHOR]

Ultra-High-Temperature Ceramic (UHTC) materials, because of their high temperature resistance, are suitable as thermal protection systems for re-entry vehicles or components for space propulsion. Massive UHTC materials are characterized by poor thermal shock resistance, which may be overcome using C or SiC fibers in a UHTC matrix (UHTCMC). The University of Naples “Federico II” has a proven experience in the field of material characterization in high-enthalpy environments. A hypersonic arc-jet facility allows performing tests in simulated atmospheric re-entry conditions. The Aerospace Propulsion Laboratory is employed for testing rocket components in a representative combustion environment. Ad-hoc computational models are developed to characterize the flow field in both facilities and perform thermal analysis of solid samples. Current research programs are related to a new-class of UHTCMC materials, for rocket nozzles and thermal protection systems. The activities include design of the prototypes for the test campaign, numerical simulations and materials characterizations. [ABSTRACT FROM AUTHOR]

This study experimentally examines the heat flux to objects outside of a firing solid propellant rocket motor plume by measuring the heat flux to gages located at various locations with respect to the rocket nozzle. Because the application of interest may involve multiple motors firing simultaneously, the heat flux from multiple motors is projected based on data collected for a single motor test, and compared to the data from two, three-motor configuration test. Data showing the enhancement from three motors firing can be substantially higher than a single motor firing when the three motors are arranged in a triangular bundle, but this was not found to be the case when the three motors were arranged in a linear bundle (linear to the instrumentation). Based on results of this study, it was concluded that a material of concern which is exposed to as many as 14 motors firing simultaneously, should survive. [ABSTRACT FROM AUTHOR]

W-ZrC composites with different W contents from 48 to 73 vol.% have been synthesized by reactive melt infiltration of Zr2Cu melt into porous preforms from partially carburized W powders at 1300 °C for 1 h in vacuum. The influences of carbon content and porosity in the preforms on microstructure and mechanical properties of W-ZrC composites are investigated. Cold isostatic pressing followed by pre-sintering process is used to produce porous preforms with suitable porosities of 53.6-47% under a pressure of 100 MPa to allow sufficient penetration of Zr2Cu melt into the preforms. Small amounts of Cu-rich phases form in the synthesized W-ZrC composites after a complete reaction of y/2xZr2Cu(l) + WCy(s) = y/xZrCx(s) + W(s) + y/2xCu(l). These Cu-rich phases are distributed not only at the phase boundaries of W matrix and ZrC grains, but also in the interior of ZrCx grains. With decreasing W content from 73 to 48 vol.% in the W-ZrC composites, the flexural strength and fracture toughness increase from 519 to 657 MPa and from 9.1 to 10.6 MPa m1/2, respectively. [ABSTRACT FROM AUTHOR]

An improved constant volume cycle (CVC) model is developed to analyze the nozzle effects on the thrust and specific impulse of pulse detonation rocket engine (PDRE). Theoretically, this model shows that the thrust coefficient/specific impulse of PDRE is a function of the nozzle contraction/expansion ratio and the operating frequency. The relationship between the nozzle contraction ratio and the operation frequency is obtained by introducing the duty ratio, by which the key problem in the theoretical design can be solved. Therefore, the performance of PDRE can be accessed to guide the preliminary shape design of nozzle conveniently and quickly. The higher the operating frequency of PDRE is, the smaller the nozzle contraction ratio should be. Besides, the lower the ambient pressure is, the larger the expansion ratio of the nozzle should be. When the ambient pressure is 1.013 × 10 Pa, the optimal expansion ratio will be less than 2.26. When the ambient pressure is reduced to vacuum, the extremum of the optimal thrust coefficient is 2.236 9, and the extremum of the specific impulse is 321.01 s. The results of the improved model are verified by numerical simulation. [ABSTRACT FROM AUTHOR]

An experimental investigation was conducted in a cold flow test facility to identify the origin of various flow conditions that lead to side-load generation in a truncated ideal contour nozzle (of area-ratio 20.66) especially at moderate nozzle pressure ratio range of 20 to 42 during the start-up and shutdown sequences. The major contributors seem to be the transition in flow conditions namely, the change in the circumferential shape of re-circulation region inside the nozzle from a cylindrically dominated regime to a conical one and the end-effect regime that initiate highly unsteady flow conditions in the separation region preceding these transitions. Other flow transitions such as those initiated by the onset of test gas condensation and vice-versa result in a downstream or an upstream jump in separation, respectively, that causes the overall side-load signal to increase. During this flow regime, an increase in the length of the upstream influence region accompanied by a rise in the peak standard deviation value is also observed. [ABSTRACT FROM AUTHOR]

Single stage to orbit propulsion devices are being developed as part of low cost access to space endeavors. Sea level operation of high area ratio rocket nozzle used in rocket engines leads to an overexpanded flow condition resulting in high side loads. Secondary injection of propellants in high area ratio nozzle is an attractive option to overcome the inefficiency of operation of such nozzles in sea level conditions in addition to the augmentation of thrust. A numerical study on thrust augmentation in high area ratio nozzle by secondary injection of propellants is presented here. The turbulent compressible reacting flow in rocket nozzle with auxiliary injection is simulated using conservation equations for chemical species based on finite rate chemistry model and compressible Navier-Stokes equations with AUSM+-up upwind scheme based unstuctured finite volume solver. An optimized eight step, six species reduced H2-O2finite chemistry reaction model is used to model the supersonic combustion. The indigenously developed solver has an efficient rescaling algorithm to alleviate the effect of stiffness in conventional explicit algorithm for simultaneous solution of reacting flow. The code is validated using the wall pressure and hydrogen concentration values reported for the similar high area ratio rocket nozzle. Accurate prediction of nozzle performance is possible with present turbulent reacting flow simulation as it take care of all losses in nozzle flow. Extensive computations have been performed for the performance estimation of high area ratio rocket nozzle for various prospective auxiliary injection options. [ABSTRACT FROM AUTHOR]

The evolution of the incident shock in the plane overexpanded jet flow or in the axisymmetric one is analyzed theoretically and compared at the whole range of governing flow parameters. Analytical results can be applied to avoid jet flow instability and self-oscillation effects at rocket launch, to improve launch safety and to suppress shock-wave induced noise harmful to environment and personnel. The mathematical model of “differential conditions of dynamic compatibility” was applied to the curved shock in non-uniform plane or axisymmetrical flow. It allowed us to study such features of the curved incident shock and flow downstream it as shock geometrical curvature, jet boundary curvature, local increase or decrease of the shock strength, flow vorticity rate (local pressure gradient) in the vicinity of the nozzle lip, static pressure gradient in the compressed layer downstream the shock, and many others. All these quantities sufficiently depend on the flow parameters (flow Mach number, jet overexpansion rate, nozzle throat angle, and ration of gas specific heats). These dependencies are sometimes unusual, especially at small Mach numbers. It was also surprising that there is no great difference among all these flowfield features in the plane jet and in the axisymmetrical jet flow out of a nozzle with large throat angle, but all these parameters behave in a quite different way in an axisymmetrical jet at small and moderate nozzle throat angles. [ABSTRACT FROM AUTHOR]

Physical details regarding base pressure low-frequency oscillations between rocket nozzles, their excitation and maintenance, are considered. Amplitude – frequency characteristics of these oscillations, as well as sequence of their type change, are studied. A single nozzle, a two-nozzle unit and a ring nozzle imitating multi-nozzle unit, are investigated in the present study. [ABSTRACT FROM AUTHOR]

Rocket motors that are designed to operate at high altitudes need a nozzle with a large expansion ratio to maximize the thrust at much lower atmospheric pressure than that of the sea level pressure. Accurate performance of these nozzles cannot be obtained when static tested on ground. Computational fluid dynamics (CFD) analyses have been performed to characterise the supersonic exhaust diffuser (SED). The results obtained from the CFD analyses have been found to be in good agreement with experimental and numerical values reported in the published literature. Started and un-started regions of the SED have been identified with the CFD results. [ABSTRACT FROM AUTHOR]

The objective of this paper is to numerically investigate the three-dimensional side loads and associated flow physics of a conical nozzle during the fire-in-the-hole staging event of the multistage rocket. The transient flow in nozzle is examined by a three-dimensional time-accurate numerical method with the start-up process of the upper-stage motor and the prescribed nominal separating motion of the lower stage. The impacts of asymmetric lateral motion due to non-nominal staging movement on the nozzle side loads are also numerically analyzed. These motion deviations include the relative angle between axes of two stages, relative lateral displacement between two stages, superposition of relative angle with relative lateral displacement, and angle of thrust vector control. As presented in the simulation results, the flow separation in nozzle is significantly prolonged and intensified by the obstruction of the lower stage. With the location of shockwave reflection approaching the dome of lower stage, the nozzle flow structure experiences two different, stable free shock separation (FSS) patterns. The transition from the first pattern to the second pattern results in the flow oscillation and the side loads peak. The staging motion deviations result in not only the enlarged nozzle side loads, but also a sustained unidirectional effect that may influence the flight stability of the upper stage. An essential ingredient driving the flow unsteadiness and asymmetry is related to the asymmetric reverse supersonic flow deflected by the dome and skirt corner of lower stage. Lastly, the flight data of maximum lateral force during staging of Scout's stage I and stage II is discussed and compared with the computed results in the article. [ABSTRACT FROM AUTHOR]

Experimental and numerical investigations are carried out to analyze the effect of coolant injector configuration on overall film cooling performance in a divergent section of a rocket nozzle. Two different injector orientations are investigated: (1) shaped slots with a divergence angle of 15° (semi-divergent injector) (2) fully divergent slot (fully divergent injector). A 2-dimensional, axis-symmetric, multispecies computational model using finite volume formulation has been developed and validated against the experimental data. The experiments provided a consistent set of measurements for cooling effectiveness for different blowing ratios ranging from 3.7 to 6. Results show that the semi divergent configuration leads to higher effectiveness compared to fully divergent slot at all blowing ratios. The spatially averaged effectiveness results show that the difference between the two configurations is significant at higher blowing ratios. The increase in effectiveness was around 2 % at BR = 3.7 whereas it was around 12 % in the case of BR = 6. Numerical results show the presence of secondary flow recirculation zones near the jet exit for both the injectors. An additional recirculation zone present in the case of fully divergent injector caused an increase in mixing of the coolant and mainstream, and a reduction in film cooling performance. [ABSTRACT FROM AUTHOR]

The objective of this effort is to numerically investigate the effect of the fire-in-the-hole staging event on the start-up flow separation and side loads of a conical nozzle during flight. The axisymmetric transient flow structure in nozzle was examined by time-accurate numerical method with time-varying chamber pressure and separation gap width. Three-dimensional time-accurate computations to examine the side loads due to non-nominal staging movement were conducted for two particular cases: misalignment between two stages and the angle of the nozzle thrust vector direction. Each asymmetry case comprised several specific staging state points which determined the corresponding chamber pressure and staging gap width. As presented by the simulation results, the flow separation in nozzle is significantly prolonged and intensified by the obstruction of the lower stage dome. When the flow separation location moves close to the nozzle exit, a strenuously axial oscillation of the separation shock appears and lasts for a period of time. In the three-dimensional asymmetric computations, in spite of strenuous oscillation, the time-average of nozzle side loads indicate that the asymmetric staging situation could lead to large sustained lateral disturbance forces. An essential ingredient driving the flow unsteadiness and asymmetry is related to the deflected and reversed supersonic flow by the dome and interstage wall corner of lower stage. [ABSTRACT FROM AUTHOR]

Numerical simulation of separated flows in rocket nozzles is challenging because existing turbulence models are unable to predict it correctly. This paper addresses this issue with the Spalart–Allmaras and Shear Stress Transport (SST) eddy-viscosity models, which predict flow separation with moderate success. Their performances have been compared against experimental data for a conical and two contoured subscale nozzles. It is found that they fail to predict the separation location correctly, exhibiting sensitivity to the nozzle pressure ratio (NPR) and nozzle type. A careful assessment indicated how the model had to be tuned for better, consistent prediction. It is learnt that SST model's failure is caused by limiting of the shear stress inside boundary layer according to Bradshaw's assumption, and by over-prediction of jet spreading rate. Accordingly, SST's coefficients were empirically modified to match the experimental wall pressure data. Results confirm that accurate RANS prediction of separation depends on the correct capture of the jet spreading rate, and that it is feasible over a wide range of NPRs by modified values of the diffusion coefficients in the turbulence model. [ABSTRACT FROM AUTHOR]

Near-field characterization of the acoustical environment near rockets has often involved extrapolating far-field measurements. However, because far-field amplitude data reveals only limited information about source characteristics, a vector intensity measurement system and analysis package has been developed to examine source features more directly. This paper describes the development of the measurement and analysis capability and its application to a horizontal firing of a GEM-60 solid propellant rocket motor firing conducted at ATK Space Systems near Promontory, Utah. An analysis of near-field intensity data provides insight both into the spatial extent and principal radiation lobe as a function of frequency. For 50 Hz, the far-field spectral peak frequency in the maximum radiation direction, the dominant source region derived from tracing the near-field intensity vectors spans 17-32 nozzle diameters, with peak radiation at ~68o . At high frequencies, the radiation results from a more contracted region that occurs farther upstream and is directed at about ~85°. These results point to the potential utility of near-field vector intensity measurements, in part because the near-field environments represented do not agree with historical far-field data-based models. [ABSTRACT FROM AUTHOR]

There is a kind of large thin-walled parts in aerospace industry, the machining target surface of which is tightly associated with another specific surface (named correlative surface). And it is always the primary machining objective. However, the preformed correlative surface is significantly different from its original design model due to large profile and thickness errors. Thus, part-referenced machining is necessary to ensure the correlative constraint accuracy. In this article, an integration strategy of OMM and NC machining for the large thin-walled parts with surface correlative constraint is systematically developed. Generally, the integration process consists of correlative constraint analysis, on-machine measurement, machining target surface redesign, and NC machining. Firstly, an isoplanar-based on-machine scanning method is presented for large surface profile information extraction. Then, a unified target surface redesign model is established according to surfaces accompanying relation analysis. Further, to compensate stress-induced monotonic structural deformation, a partitioned measuring and machining approach has been employed. Finally, the liquid rocket engine nozzle as a typical part was employed to verify the validation of the proposed strategy. Coolant channel machining experiments were conducted on a special dual-spindle machine tools. For a nozzle with machining area about 8 m, the correlative accuracy could be controlled in the range of ±0.1 mm. It has been proved that incorporating dimensional metrology feedback to machining process could consistently improve machining quality and efficiency of large thin-walled parts. [ABSTRACT FROM AUTHOR]

The rising demand for reliable wind-tunnel flow data for rocket nozzle design applications led to the development of a new type of facility. Most of the former experimental data in the literature were gathered in conditions far from those encountered in real engines and combustion gas generators. Thus, they are not appropriate for the physical interpretation of complex flow phenomena and computational fluid dynamic tools validation. The new experimental platform is intended to solve this problem by its capability to nearly match most of these conditions. It allows for the investigation of several engine-relevant aerothermodynamic problems such as boundary-layer transition, nozzle flow separation, nozzle exhaust plume interactions, etc. Furthermore, innovative nozzle cooling techniques like film cooling can be studied in this new type of facility. This is demonstrated first by the experimental results given at the end of this paper. High nozzle reservoir pressures and temperatures are achieved by means of detonative combustion of a premixed gas in a confined tube. A detonation wave produces a high-energy flow that expands through a Laval nozzle to the desired experimental flow conditions. The facility simulates enginelike conditions in a combustion environment for various mass ratios of oxidizer-fuel mixtures. The short-duration flow device affords a simple and cost-effective research tool for testing under harsh flow conditions in laboratories. Hence, it will significantly contribute to the understanding of complex rocket nozzle flow phenomena and validation of numerical tools, as well as correlations. [ABSTRACT FROM AUTHOR]

The aim of this paper is to present and compare two different approaches for aeroelastic stability analysis of a flexible over-expanded rocket nozzle. The first approach is based on the aeroelastic stability models developed in a previous work, while the second uses the numerical fluid–structure coupling via the transpiration method. The aeroelastic frequencies of the nozzle obtained by various stability models are compared with those extracted from the numerical coupling by the method of transpiration. Both set of results show an overall good agreement. [ABSTRACT FROM AUTHOR]

The article analyzes the effect of fluid-structure-thermal interaction on mechanical yaw thrust vectoring efficiency of a two-dimensional (2D) rectangular convergent rocket nozzle. The study identifies the boundary conditions for computational fluid dynamics (CFD) domain and structure and examines the temperature and pressure distribution inside the nozzle. It investigates the impact of vane deformation due to pressure loading on velocity component and the effectiveness of thrust vectoring.

While operating in an overexpanded condition, rocket nozzles exhibit dynamic off-axis loads caused by asymmetric internal flow separation. It has been demonstrated that topologies, which only display free shock separation, produce side-load magnitudes lower than those that can transition to restricted shock separation. It is therefore suggested that geometries such as truncated ideal contour nozzles could be used on launchers with large-area-ratio designs providing improved vacuum performance. Part of the challenge of designing nozzles to operate in the overexpanded regime is characterizing and quantifying associated side loads. This paper presents the methodology and results of an analytical model that simulates side loads induced by free shock separation. The pressure distribution within a separated rocket nozzle is produced using free interaction theory, and a spring-damper analogy is used to model shock fluctuations in response to external pressure disturbances. Good agreement has been found between simulated side loads and experimentally measured forces on four conical and one truncated ideal contour nozzle. [ABSTRACT FROM AUTHOR]

An experimental and computational investigation of supersonic film cooling was conducted on a subscale model of a rocket engine nozzle. A computational model of a convergent-divergent nozzle was generated, incorporating a secondary injection module for film cooling in the divergent section. Computational fluid dynamics simulations were run on the model and different injection configurations were analyzed. The computational fluid dynamics simulations also analyzed the parameters that influence film cooling effectiveness. Subsequent to the computational fluid dynamics analysis and literature survey an angled injection configuration was found to be more effective, therefore the hardware was fabricated for the same. The fabricated nozzle was later fixed to an Air-Kerosene combustor and numerous sets of ex-periments were conducted in order to ascertain the effect on film cooling on the nozzle wall. The film coolant employed was gaseous nitrogen. The results showed substantial cooling along the walls and a considerable reduction in heat transfer from the combustion gas to the wall of the nozzle. Finally the computational model was validated using the experimental results. There was fairly good agreement between the predicted nozzle wall temperature and the value obtained through experiments. [ABSTRACT FROM AUTHOR]

Rhenium is a promising coating material for liquid rocket nozzles because of its extremely superior performance, such as high melting point, high temperature strength and excellent fatigue properties. In this work, Re coating was prepared on fabric-based pierced Carbon/Carbon composite substrate using chemical vapor deposition. Bonding strength, micro hardness and thermal shock resistance of rhenium coating on Carbon/Carbon substrate were investigated respectively. The bonding strength between the coating and the substrate was tested using different methods including tensile tests, shearing tests and micro scratch tests. Thermal shock resistance was investigated by an induction furnace from 2000 o C to the room temperature. The morphologies and characteristics of the coatings were studied by scanning electron microscopy. The results indicated that the highest average value of tensile fracture strength is 20.4 MPa and shearing fracture strength is 21 MPa on Carbon/Carbon composite substrate. The rhenium coating broke up and flaked off near a crack in the scratch test. Micro hardness test suggested that C atoms diffused quickly into the rhenium coating within an hour. The thermal shock lives depended on the number of thermal shock cycles. It is also concluded that rhenium coating had good thermal shock resistant with Carbon/Carbon Composite substrate. [ABSTRACT FROM AUTHOR]

The article discusses research that examines the gas-kinetic density solutions related to the cosine-relation rocket plume model. The researchers assumed that the rocket plume develops along the X axis and that the center for the round nozzle exit is the coordinate origin. The details of the experiment conducted are presented.

Carbon/carbon composites have been typically used to protect a rocket nozzle from high temperature oxidizing gas. Based on the Fourier's law of heat conduction and the oxidizing ablation mechanism, the ablation model with non-linear thermal conductivity for a rocket nozzle is established in order to simulate the one-dimensional thermochemical ablation rate on the surface and the temperature distributions by using a written computer code. As the presented results indicate, the thermochemical ablation rate of a solid rocket nozzle calculated by using actual thermal conductivity, which is a function of temperature, is higher than that by a constant thermal conductivity, so the effect of thermal conductivity on the ablation rate of a solid rocket nozzle made of carbon/carbon composites cannot be neglected. [ABSTRACT FROM AUTHOR]

The objective of this article is to perform detailed analysis of heat transfer in accelerated supersonic nozzle flows with cooled walls. Since most of the heat transfer occurs near the nozzle walls, correct prediction of the boundary layer under strong adverse pressure gradient is therefore required to achieve high fidelity numerical prediction. In this study, a two-equation SST-V turbulence model is used in conjunction with a second-order explicit-implicit method to solve axisymmetric compressible Navier-Stokes equations. First, the effect of the inlet pipe diameter and the associated contraction area on the heat transfer is studied in nozzles having 15° and 30° diverging half-angles. Then, a series of computations are conducted to examine the efficiency of the use of a constant wall temperature as a function of the stagnation temperature in heat transfer calculations. The computations are performed for nominal stagnation pressure of 208 N/cm2and stagnation temperature of 539 K. The computed heat-transfer coefficients are compared to experimental data and a good agreement is found. A pronounced increase in the throat heat transfer coefficient peak is observed accompanied with a reduction in the contraction area ratio. Also, the peak of the heat transfer coefficient for the pipe inlet diameter of 7.8 cm is found to be 70% higher than the one related to the pipe of 16.51 cm diameter. [ABSTRACT FROM PUBLISHER]

Highlights: [•] We designed and tested a breaking wire like recession sensor based on thermocouples. [•] Commercially available raw materials and techniques have been used. [•] Carbon/carbon materials were used to test the recession rate technology. [•] The proposed sensing technique can be applied on TPSs and on rocket nozzle throats. [Copyright &y& Elsevier]