Rocket Nozzle Research Articles

Hybrid Rocket Nozzle Erosion with Microaluminum-Added Fuel

Because of its relatively low cost and wide availability, carbon graphite is one of the most widely used ablative nozzle materials in hybrid rocket propulsion. The erosion characteristics of the nozzle material have paramount importance because it directly influences the performance. This is especially the case for upper-stage or in-space rocket motors operating with very long burn times. In this study, the effect of aluminum-added fuel on the graphite nozzle erosion is studied. In the experimental studies, a high regression rate paraffin-based fuel is loaded with micrometer-size aluminum powder for nozzle erosion reduction. In our approach, aluminum is added at high concentrations as a fuel ring in front of the main paraffin-based fuel which contains no aluminum. Based on the motor tests conducted with gaseous oxygen as an oxidizer, it is shown that aluminum addition decreased the nozzle erosion rate up to 45% and increased the nozzle erosion onset time by 1 to . The new method of introducing an energetic powder in a fuel ring positioned at the fore end of the motor offers an easy and scalable way of reducing the nozzle erosion and improving the performance of the rocket motor.

Comparative appraisal of nanofluid flows in a vertical channel with constant wall temperatures: an application to the rocket engine nozzle

ABSTRACT This study aims to investigate the entropy generation of fully developed nanofluid flow in a spongy medium between two fixed absorbent vertical plates with a transpiration cooling mechanism. For heat transfer analysis exponential space-dependent heat source and thermal dependent heat source along with linear radiation are employed. Three distinct scenarios for thermal wall boundary conditions are investigated: i) Constant walls temperature ii) Convective boundary condition on the left wall of the channel and iii) convective boundary conditions on both walls of the channel. Two nanofluids combinations i.e. copper (cylindrical shape) nanoparticles immersed in water and silver (platelet shape) nanoparticles in water are used. It is also noted that studies on the transpiration of nanofluid in rocket engines for cooling purposes are scrutinized when walls are stretchy. But the walls of the rocket engine are fixed. Keeping this in mind, we analyzed the cooling process of the rocket nozzle with a fixed wall channel. The numerical solution of the model equations is attained using the bvp4c method in MATLAB. Findings indicate that the copper/water nanofluid combination performs better than silver/water as for as cooling the nozzle of a rocket engine is concerned. Moreover, convective boundary conditions lower the fluid temperature.

Effect of cure temperature on the thermal degradation, mechanical, microstructural and moisture absorption behavior of vacuum‐only, carbon‐fiber reinforced phenolic composites

AbstractPhenolic resin is the material of choice used for composite liners in solid rocket motor nozzles. We investigate the effect of the cure conditions (94, 116, and 155°C) on vacuum only (VO) processed composite mechanical, thermal, and microstructural changes. Although the Tg's correlated with the processing conditions, mechanical properties showed unexpected non‐linear degradation after 94°C processing. This difference is primarily due to the lower cure state of the resin, which results in poor fiber/matrix coupling and is further magnified by the reduced degree of consolidation observed in VO parts. Elevated thermal exposure was shown to generate larger tensile strains in the out of‐plane direction for the 94°C composites when compared to the other conditions. This resulted in significant increases in crack density, porosity, and delamination due to an increased volume of degradation by‐products, while the higher temperature cured composites maintained good consolidation. Even though all three cure states appeared to display equivalent moisture absorption saturation levels based on weight measurements, the 94°C cured sample was shown to partially dissolve in moisture during humidity exposure, which translated to even higher saturation levels. This larger net absorption resulted in a larger drop in Tg for the 94°C cured material and thus resulted in a more severe thermal mechanical response than the other saturated specimens. This emphasizes how a reduction in cure state will not only affect thermal and mechanical performance but will also translate to a greater degree of degradation than more highly cured specimens, when exposed to humid environments.

Ablative properties, thermal stability, and compressive behaviour of hybrid silica phenolic ablative composites.

Silica phenolic composites are used as thermal protection systems for high temperatures applications like atmospheric re-entry modules, rocket nozzles etc. Lightweight silica phenolic composites (SPGB) were developed using chopped silica fibre, phenolic matrix, and hollow glass microspheres through a compression moulding technique to achieve a cost advantage in thermal protection systems. Ablative properties were evaluated using an Argon plasma jet at a stagnation heat flux of 500 W/cm2. The mass-loss and erosion rates under plasma exposure reduced up to 48% and 32%, respectively, with the addition of hollow glass microspheres. The heat of ablation of composites increased up to 33%. The thermal degradation behaviour of raw materials and SPGB composites was investigated through thermogravimetric analysis, and the scientific know-how was established. Compressive behaviour under uniaxial loading was evaluated, and the theoretical model for the developed three-component composite was established.

Base Thermal Environment on Multinozzle Rocket Configurations

Because of lack of quantitative research on rocket base thermal environment under different configurations, three launch vehicles with different numbers of nozzle configurations are studied to analyze and establish criteria for effect of a multinozzle configuration on base heating. To simulate base convective heating during ascents, the three-dimensional compressible Navier–Stokes equations and the shear stress transport turbulence model are used to establish the reverse-flow model. Hence, the base thermal environment at different flight altitudes () was simulated. Results showed that peak heating occurred at lower altitude, and base thermal load was more severe with a compact rocket nozzle layout. Impinging jet caused by recovery shock affected the base flowfield through the formation of stagnation bubbles on the base plate, and led to the ring peak heating. The closure of reverse flow depends on the nozzle pressure ratio, mass flow rate of reverse flow, and effective exhaust boundary length. Consequently, a factor was defined to assess the base convective heating generated by reverse flow in various configurations, and the standard for evaluating rocket base heating was established by introducing the factor , which yielded results within 16% deviation from test data.

The Investigation of Plume-Regolith Interaction and Dust Dispersal during Chang’E-5 Descent Stage

The plume-surface interaction that occurs as a result of a variable-thrust engine exhaust plume impinging on soil during landings is critical for future lunar mission design. Unique lunar environmental properties, such as low gravity, high vacuum, and the regolith layer, make this study complex and challenging. In this paper, we build a reliable simulation model, with constraints based on landing photos, to characterize the erosion properties induced by a low-thrust engine plume. We focus on the low-thrust plume-surface erosion process and erosion properties during the Chang’E-5 mission, aiming to determine the erosion difference between high- and low-thrust conditions; this is a major concern, as the erosion process for a low-thrust lunar mission is rarely studied. First, to identify the entire erosion process and its relative effect on the flat lunar surface, a one-to-one rocket nozzle simulation model is built; ground experimental results are utilized to verify the simulated inlet parameters of the vacuum plume flow field. Following that, plume flow is considered using the finite volume method, and the Roberts erosion model, based on excess shear stress, is adopted to describe plume-surface interaction properties. Finally, a Lagrangian framework using the discrete phase model is selected to investigate the dynamic properties of lunar dust particles. Results show that erosion depth, total ejected mass, and the maximum particle incline angle during the Chang’E-5 landing period are approximately 0.2 cm, 335.95 kg, and 4.16°, respectively. These results are not only useful for the Chang’E-5 lunar sample analysis, but also for future lunar mission design.

Development and Characterization of Silicon Carbide Coating on Graphite Substrate

The development of materials with unique and improved properties using low cost processes is essential to increase performance and reduce cost of the solid rocket motors. Specifically, advancements are needed for boost phase nozzle. As these motors operate at very high pressure and temperatures, the nozzle must survive high thermal stresses with minimal erosion to maintain performance. Currently three material choices are being exploited; which are refractory metals, graphite and carbon-carbon composites. Of these three materials graphite is the most attractive choice because of its low cost, light weight, and easy forming. However, graphite is prone to erosion, both chemical and mechanical, which may affect the ballistic conditions and mechanical properties of the nozzle. To minimize this erosion Pyrolytic Graphite (PG) coating inside the nozzle is used. However, PG coating is prone to cracking and spallation along with very cumbersome deposition process. Another possible methodology to avoid this erosion is to convert the inside surface of the rocket nozzle to Silicon Carbide (SiC), which is very erosion resistant and have much better thermal stability compared to graphite and even PG. Due to its functionally gradient nature such a layer will be very adherent and resistant to spallation. Despite its very good adhesion due to its functionally gradient nature, this layer due to its porous nature exhibit poor oxidation performance compared to a dense SiC layer. The current research is focused on synthesizing, characterizing and oxidation testing of a bi-layer; a functionally gradient inner layer and dense outer layer, SiC coating on graphite.

Thermomechanical response of out-of-autoclave infused carbon-phenolic laminates for rocket engine applications subjected to surface ablation

Cost-affordable and reliable manufacturing technologies to produce Carbon/Phenolic Composites (CPCs) are essential for competitive access to space. These materials are used as ablative Thermal Protection System (TPS) for space applications, such as rocket nozzles or the heat shields of space vehicles. Out-Of-Autoclave (OOA) infusion processes can reduce the manufacturing cost of composites: in this work, OOA produced CPC laminates were compared to Carbon/Epoxy Composites (CECs). The CPCs produced in this work were mechanically characterized to assess their quality and compared with state-of-the-art counterparts through the data available in the open literature. CPC laminates instrumented with ultra-small thermocouples were subjected to Oxy-Acetylene Torch (OAT) ablation testing and a numerical approach was assessed to predict the temperature reached in the laminate layers. Post-burning mechanical tests performed on specimens subjected to the ablation test showed that when the temperature does not exceed a given threshold, the material retains adequate properties to perform a structural role.

CFD Analysis of Convergent-Divergent Nozzle for Low Altitude Rocket

The convergent-divergent nozzle in rocket engines is used to turn the chemical energy developed via the combustion chamber into kinetic energy. The wall of the nozzle is properly contoured to efficiently accelerate the combustion product towards the nozzle exit. However, for low-to medium-altitude rockets (typically within the range of 10 to 145 km), nozzle contouring increases the overall production cost due to a more complicated fabrication process. The development of the dual conical nozzle seeks to compromise between performance and ease of fabrication. The present investigation seeks to analyze small rocket nozzles involving a typical conical nozzle, Truncated Optimized Contour (TOC) and Truncated Idealized Contour nozzles and newly designed dual conical nozzles. Rocket nozzles with various divergence angles and exit area ratios are numerically analyzed. It was discovered that the best dual conical nozzle can produce an average thrust comparable to that of a standard conical nozzle. In some cases, the dual conical nozzle outperforms its conical and contoured counterparts. Keywords— Dual conical nozzle, TOC nozzle, TIC nozzle, shockwaves, thrust.

Experimental and Computational Investigation of Divergent Thrust Losses In Rocket Nozzles

The nozzle is a critical component of any rocket engine because it transforms high pressure and high temperature (with sluggish net velocity) within the combustion chamber into high velocity but low pressure. Individually, the molecules in the mixed propellants have rapid velocity in the combustion chamber, but directions are randomly pointing in all directions. A convergent-divergent nozzle, also known as a CD nozzle, is fashioned like a tube with two bulging ends and a thin part in the center. This structure/design provides the nozzle with a balanced shape. By turning the flow's heat energy into kinetic energy, the nozzle accelerates hot, pressured gas traveling past it in the axial (push) direction. Whereas the nozzle provides the acquired incremental thrust to the rocket engine, it also results in unwanted thrust losses. Out of which divergent losses in the divergent section are the primary foci of our research. Moreover, in this research paper, we produced verified results to reduce the thrust losses at the divergent sections of the nozzle by optimizing the nozzle with different half divergent angles.

Two Dimensional CFD Analysis on Different Rocket Nozzles

The reduction of Earth-to-orbit launch costs in conjunction with an increase in launcher reliability and operational Efficiency is the key demands on future space transportation systems, like single-stage-to-orbit vehicles (SSTO). The realization of these vehicles strongly depends on the performance of the engines, which should deliver high performance with low system complexity. Performance data for rocket engines are practically always lower than the theoretically attainable values because of imperfections in the mixing, combustion, and expansion of the propellants. The main part of the project addresses different nozzle concepts with improvements in performance as compared to conventional nozzles achieved by Different Mach numbers, thus, by minimizing losses caused by over- or under expansion. The design of different nozzle shapes and flow simulation is done in gambit and fluent software’s respectively for various parameters

Investigation on the Mechanical Properties of the Post-Static Fire Testing Effect of Graphite Material used in the RX320 Rocket Motor Nozzle

The rocket nozzle is a crucial and complicated part of a rocket, primarily for rocket RX320. Its function is to control the exhaust flow, so the rocket gets the maximum possible forward thrust. However, graphite materials are used inside rockets due to good mechanical properties and high melting temperature resistance. This study discussed the effect of post-static rocket fire testing on the mechanical properties of graphite materials and compared them with the technical data sheet (TDS) of its graphite. Mechanical properties testing has been done to determine the average flexural strength value with the flexural test method and the average value of the compressive test. The graphite has an average flexural test value of 40.185 MPa, with an average elastic modulus of 0.656 GPa. Furthermore, the average compressive test of graphite has a value of 67.71 MPa, with an average load of 1,560 N. The mechanical properties trends are slightly decreasing due to post static fire testing compared to TDS. It was around 41.2% for flexural strength and 50.7% for compressive strength.

Analytical investigation over the Four Inlet C & D Nozzle Using Ansys

The engine's interpretative thrust production, such as nozzles, has been restructured for higher staging. Modern expansions of combustion systems, such as rocket nozzles, will be adjusted to meet the needs of today's applications. In this paper an analytical study of fluid flow in a convergent-divergent nozzle with 4 inlets the nozzle is design with a convergent angle of 45 and a divergent angle of 15. we have observed the CFD of the C-D nozzle by swerving the inlet Mach parameter from 2,2.5,3,3.5,4,4.5,5 This study observed the usage of the usage of the four inlets in the convergent and the divergent nozzle rather that a single inlet the study is caried out in the in the 2-D Simulation the Design of the nozzle is performed in the Solid-Works software and the simulation is performed in the Ansys fluent Software. The counter of the flow is been observed. The results concluded that in the analytic study the usage of the 4- Inlet is better than the single inlet. Key Words: Ansys, convergent-divergent nozzle, nozzle, CFD, Mach number

Effect of lid height and blowing ratio on film cooling effectiveness of a novel lidded hole configuration

Film cooling is one of the promising technologies used for protecting rocket nozzles and turbine blades from combustion chamber hot gases. This paper proposes a novel shape of film cooling injection hole, called lidded hole, that can offer significant enhancement of cooling performance. ANSYS CFX is used to perform 3D numerical simulations of a flat plate with a single row of lidded holes, in which the k–ε model approximates turbulence effects. Four cases are investigated to highlight the influence of the hole's lid height (H/d = 0, 0.25, 0.5, 0.75). The effect of blowing ratios (M = 0.5, 1, 1.5) is also analyzed for each configuration. The numerical results of this study are compared with available experimental data, and, generally, a good agreement is achieved. The results obtained show that the lidded hole configuration reduces the coolant flow separation which improves significantly the film cooling effectiveness. In addition, increasing the blowing ratio leads to an increase in lateral and centerline cooling effectiveness. Comparing all studied cases, the optimum coolant coverage was obtained for the lidded hole configuration with H/d = 0.25 at M = 1.5.

A Numerical Simulation Study for A Dual Thrust Solid Propellant Rocket Motor Nozzle

CFD simulation through the rocket motor nozzle is an important issue as it gives a better understanding of the flow behavior and shows the evolution of the parameters along with the nozzle. CFD simulations using ANSYS FLUENT were conducted to model the flow through an experimental dual thrust solid rocket motor nozzle at two pressure values representing the two pressure levels. A two-dimensional axisymmetric numerical simulation was carried out using a pressure-based solver to solve the governing equation and the standard k-ε model to model the turbulence through the motor nozzle. The hot gas properties used in the simulation model were calculated using the CEA software (chemical equilibrium with applications). The parameters marched out from simulations were used to calculate the thrust values for each case. The results calculated through the numerical models compared with the experimental measurements and showed an acceptable level of accuracy.

Design and numerical analysis on supersonic rocket nozzle

Rocket nozzle is part of a rocket in which the burnt propellant from the combustion chamber can exhaust through the nozzle to get the desire amount of thrust and required Mach number. The nozzle turns the static high pressure high temperature gas to low pressure, low temperature, and high velocity gases. To analyse this we have used CFD. Where CFD is the fluid mechanics which is integral part of the design, this is based on numerical methods and algorithm method. To analyse the fluid flow in supersonic rocket nozzle the CFD analysis software is used. Mainly in this work had focused on the pressure, temperature, velocity and Mach number of the mass flow in nozzle by changing the divergent angle. The oblique shock is a main conception and the presence of oblique shock which will create the instabilities. And it is visualized that the shock wave is creating in the nozzle at the 5° of divergent angle and it slowly started eliminating from the nozzle after 5° of divergent angle. To overcome this problem some of the trials are made for different divergent angle such as 5°, 10° 11°, 12°, 15°, and 20°. After 12o of divergent angle the flow parameters will start get reducing. At 20o the sudden drop of flow parameter will occur. By considering all the results at different divergent angle, the 11° angle is more efficient and found all the flow parameters such as velocity, temperature, pressure and Mach number of the mass flow are at required condition.

The Effect of Curvature on the Heat Transfer Performance of Regenerative Cooling Passages for a High-Area-Ratio Nozzle

Abstract Numerical studies on the flow and heat transfer characteristics of rectangular regenerative cooling passages with lateral curvature for a high-area-ratio nozzle are presented. Unlike regular geometries of coolant paths used in various engineering applications, high-area-ratio rocket nozzles have steeply curved surfaces over which the cooling passages are provided. Though the inherent inward curvature benefits heat transfer enhancement in the throat region, it is insufficient to circumvent peak heat dissipation demand. Hence, the heat transfer enhancement due to lateral curvature of the coolant fluid passages provided near the throat region is explored in this study. Extensive numerical simulations have been performed to analyze the effect of the geometry of the cooling channel on its flow and heat transfer characteristics. The compressible turbulent flow field inside the nozzle has been resolved to understand the realistic local wall heat transfer characteristics of the typical high-area-ratio rocket nozzle using Advection Upstream Splitting Method (AUSM) scheme-based finite volume solver. Menter’s Shear Stress Transport k–ω turbulence model is used to model the turbulent flow inside the nozzle. Simulations of the incompressible coolant flow and conjugate heat transfer in regenerative cooling passages have been performed with realistic spatially varying local heat flux profiles, resulting due to compressible gas expansion in the convergent–divergent nozzle. Secondary flow structures are formed due to the lateral curvature of the coolant fluid passages and are found to enhance the heat transfer considerably. Further, the effect of coolant flowrate and channel curvature have been examined to explore its suitability to negotiate the peak heat flux dissipation demand at the throat region of the nozzle.

Numerical investigation of surface roughness effects on non-equilibrium flow in expansion section of rocket nozzle

To further understand the non-equilibrium flow characteristics in a rocket nozzle, we have to consider the rough wall surface caused by nozzle ablation, and explore the relative influence of surface roughness on the flow field, composition field, and wall heat transfer in the nozzle. In this study, first, the data of the inner wall morphology of the real nozzle is obtained through measurement experiments. The parameters used to characterize the inner wall morphology of the expansion section of the nozzle are determined. Then, a computational physical model is built based on these parameters. Finally, the simplified chemical reaction mechanism is used to solve the chemical non-equilibrium flow in the nozzle through the eddy dissipation conceptual model. The numerical calculation method of large eddy simulation is used to obtain the fine structure of the flow field with the shape of the inner wall of the expansion section of the nozzle. The results of the numerical study show that the influence of the wall roughness element of the nozzle expansion section caused by ablation cannot be ignored during the working process of the solid rocket motor. For the non-equilibrium flow in the supersonic nozzle, the roughness element is a crucial factor to promote the flow transition and the generation of hairpin vortices, and it also affects the wall heat transfer and gas composition near the roughness element.

Coupled Thermo-Structural Analysis Model of Solid Rocket Motor Nozzle considering the Variation of Friction Coefficient under Operating Conditions

Multi-interface is a typical feature of solid rocket motor nozzles, and the interface evolution law is an important aspect to analyze the thermo-structural response of solid rocket motor nozzles. In this paper, based on the friction coefficient test at different temperatures, a coupled thermo-structural analysis model of solid rocket motor nozzle considering the variation of friction coefficient under operating conditions was established. Firstly, the friction coefficient was tested to model the variation at different temperatures. Then, by adopting structure gap, variable friction coefficient, thermal contact resistance, and friction heat production, a strongly coupled non-linear model was established. Simulations using the non-linear model and traditional model were performed, in which the stress of the throat insert was within the required stress range of the material. The ground firing test result demonstrated the validity of the analysis model, and the non-linear model was in a better agreement with the firing test than the traditional model. Therefore, it can be concluded that with a more specific friction coefficient to represent the friction behaviour of the different parts of the nozzle, the strongly coupled non-linear model established in this paper can reflect the essence of thermo-structural response of solid rocket motor nozzle.

Cooling rocket nozzle extensions with supersonic gas injection

Supersonic helium injected into the rocket exhaust boundary layer protects components from extremely hot gas.