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A dual bell nozzle (DBN) consists of two bell nozzles of different geometric area ratios attached at the region called inflection. This distinct geometry modification allows the nozzle to adapt to ...

This wok focuses to develop a numerical computation program allowing to design new contours of a supersonic nozzle having several bells, adapted to several levels of different altitudes, going from sea level and progressively with the altitude up to space, giving a supersonic uniform and parallel flow to the exit section and a maximum possible thrust without loss, aiming to reduce considerably the side loads caused in the conventional and the dual bell nozzles when the ambient pressure decrease with the altitude. The first bell has a sonic, uniform and parallel inlet to the throat, and a sea level adaptation with a reduced supersonic Mach number, while the other bell have a supersonic, parallel and uniform inlet and an adaptation to a given altitude with a progressed increase of supersonic Mach number. The transition from one level of adaptation to another adjacent is done without mechanical activation. The purpose of this type of nozzle is to have the possibility of flying in several supersonic regimes adapted to several different altitudes with a reduced side loads. This type of nozzle is named by Several Bell Nozzle and has inflection points between bell and adjacent other. The design is done in the context of a calorically and thermally perfect gas. The problem is to design a typical bell having a uniform and parallel supersonic inlet and exit with two given Mach numbers. The design is made by the use of the Method of Characteristics. The resolution of the equations is done numerically by the finite difference corrector predictor algorithm. The validation of the results is controlled by the convergence of the ratio of the critical sections, calculated numerically, to that given by the theory. In this case, all the design parameters converge automatically to the desired solution. The application is made for axisymmetric MLN having 3, 4, 5 and 10 bells.

The typical afterbody flow of a space launcher is characterized by a strong interaction of the engine’s exhaust jet and the separated shear layer emerging from the main body. This interaction is further complicated by strong changes in the spatial and temporal behavior of the afterbody flow during the atmospheric ascent of a launcher. Theoretically, a dual-bell nozzle not only allows for a gain in payload compared to standard single-bell nozzles, but also it alters the wake flow topology due to the two nozzle modes. To predict the benefits as well as the additional risks, the afterbody flow of a generic space launcher model equipped with a cold-flow dual-bell nozzle is investigated in detail. The flow was analyzed for sub-, trans- and supersonic Mach numbers ranging from 0.3 to 2.9 for a variety of nozzle pressure ratios. Particle image velocimetry measurements and schlieren measurements with high repetition rate were performed to determine the dynamics of the separated shear layer, the nozzle jet and their interaction. It is shown that the reattachment length of the base flow decreases with increasing nozzle pressure ratio. Furthermore, the nozzle pressure ratio at which the dual-bell nozzle switches from sea-level mode to altitude mode is reduced by $$15\%$$ 15 % with high subsonic outer flow and by as much as $$65\%$$ 65 % for an outer flow at a Mach number of 1.6. Even for a constant nozzle pressure ratio, the nozzle flow topology depends on the Mach number of the outer flow.

A numerical study is conducted to investigate the impact of a film-cooled dual-bell nozzle extension on its operation mode transition behavior. Therefore, unsteady Reynolds-averaged Navier–Stokes s...

The dual-bell nozzle is a promising altitude adaptive nozzle concept for improving the performance of existing space launchers as well as future reusable launch vehicles. It is characterized by an inflection in the wall geometry which links a base nozzle with a nozzle extension. Two working modes are therefore attained: a low-altitude mode in which the expansion is limited to the first bell and a high-altitude mode in which full expansion is achieved. Even though this capability makes it possible to reduce the non-adaptation losses in engines working from sea level to almost vacuum conditions, the natural transition between the two modes usually takes place prematurely during the ascent trajectory and destructive side-loads may arise. This results in a limitation on the performance gain. In this work, a dual-bell nozzle design capable of upgrading the parallel-staged Ariane 5 launcher is studied. Fluidic control is investigated as a possible solution to limit premature transition and reduce side-loads. The control is obtained by means of a secondary injection in the proximity of the inflection point: different injection strategies are investigated. Reynolds-averaged Navier–Stokes equations at steady state conditions are solved with an in-house CFD tool, to investigate the flow field and determine the required secondary mass flow rate and the range of nozzle pressure ratio for which the control is effective.

The turbulent wake of a planar generic space launcher equipped with a dual-bell nozzle is numerically investigated to examine the interaction of the dual-bell nozzle jet and the wake flow. The simulation is performed at transonic freestream condition, i.e., freestream Mach number $Ma_{\infty }= 0.8$ and freestream Reynolds number based on the launcher thickness ReD = 4.3 ⋅ 105, with the dual-bell nozzle operating at sea-level mode. A zonal RANS/LES approach is used and the time-resolved flow field data is analyzed by classical spectral analysis and modal decomposition techniques, i.e., proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD). The overall flow topology of the recirculation region downstream of the base and the pressure loads on the outer nozzle fairing are only slightly affected by the modified nozzle shape. However, the changed nozzle flow topology characterized by the flow separation at the nozzle contour inflection leads to a backflow region and an entrainment of the outer flow into the nozzle extension which results in increased pressure loads on the inner nozzle wall. Using spectral, POD, and DMD analyses, the outer wake flow is investigated, revealing a growing and contracting of the separation bubble and an undulating motion of the shear layer similar to the “cross-pumping” and “cross-flapping” motion detected in previous investigations of a configuration with a classical nozzle and a jetless backward facing step setup. The spectral and modal analysis of the nozzle flow shows that the increased pressure loads detected at the inner wall of the nozzle extension are caused by an interaction of the separated shear layer inside the nozzle extension with the shock pattern that leads to a streamwise oscillation of the shock and a pumping or wave-like motion of the shear layer.

Conical aerospike nozzles are advanced rocket nozzles suitable for altitude compensation. A spike nozzle has geometry that resembles a conventional bell nozzle turned inside out. Lack of a solid ou...

In the field of space science and aerospace technology, altitude adaptive nozzles recently have got more importance. Dual bell nozzle presents a solution for the altitude adaption. Dual bell nozzle has dual operating modes (at low and high altitudes) without any mechanical activation. For future reusable launch vehicles, it is necessary to understand the active control of the working mode of the dual bell nozzle. Film cooling with injection of the secondary fluid provides a solution for the active control of the nozzle. The effectiveness of this secondary injection is investigated here by injecting the secondary fluid at three different locations viz. at the throat, at the base nozzle, and at the inflection point for the low-altitude as well as high-altitude operation mode. A two-dimensional steady RANS simulation has been carried out on an axisymmetric parabolic contour dual bell nozzle. Turbulence is modelled with k-Omega SST model with the updated coefficients. The present study highlights the thermodynamic aspects of the film cooling in a dual bell nozzle.

A dual-bell nozzle comes under the classification of an altitude compensating nozzle. The absence of any moving parts makes the dual-bell nozzle a better substitute for the existing bell-shaped nozzles. In this study, a numerical simulation is carried out on a 2D axisymmetric dual-bell nozzle model to find the nozzle wall pressure distribution at different nozzle pressure ratio (NPR) which is validated with the experimental results from previous literature. The main flow is modeled as LO2/LH2 combustion mixture and secondary flow is modeled as GH2. Film cooling is introduced at the inflection point of the dual-bell nozzle to study the influence of the GH2 cooling film to the flow characteristics inside the nozzle. The influence of secondary coolant injection to the movement of separation location and film cooling effectiveness is studied by varying the coolant flow rate.

The Formula Student rulebook specifies that a 20 mm circular restrictor is to be placed in the intake manifold. Conventionally, a straight profile converging–diverging nozzle, i.e. venturi, is used in place of a restrictor. The authors propose an alternative design, i.e. a curved profile converging–diverging nozzle, whose design has been adapted from that of a bell nozzle used in rocket engines. Firstly, the effect of change of various parameters on the pressure drop across the nozzle using CFD simulations is studied. It was found that a bell-shaped nozzle provided 32.4% decrease in pressure drop as opposed to straight profile nozzle. Thereafter, to analyse the effect of change of nozzle geometries on the engine’s performance, Ricardo WAVE 1-D simulations were done. WaveMesher was used to model the bell nozzle geometries in the WAVE model. It was found that a bell nozzle provided a maximum of 2.6% increase in brake power and 2.6% increase in torque at 7000 RPM as opposed to the straight profile nozzle. The ultimate motive behind the whole exercise was to improve the lap times of our car on-track in the competition. In order to get an idea about the increase or decrease in the lap times of the vehicle, point mass lap time simulation tool—Optimum Lap—was used. The car was simulated for the competition’s Autocross lap due to variety of fast and slow corners with long straights.

International audience; Abstract: With the ever rising demands for cheaper payload delivery to orbit, dual bell nozzle with a potential theoretical performance gain of up to 10% can represents one of the major ways of advance to achieve the goal. The well-known transition unsteadiness in a dual-bell rocket nozzle represents one of the major concerns for the development of such altitude compensating type nozzle (ACN). The present study proposes and investigates the possibilities of secondary radial injection for flow regime transition control. Present experimental results demonstrate potential effects of secondary injection on transition and re-transition control even with a relatively low secondary injection mass flow rates. Doing so, the transition and retransition process were significantly delayed towards the ideal transition point. In addition, side loads have been found to be greatly decreased or even eliminated.

Previous research on Dual-Bell nozzle flow always neglected the influence of the outer flow on the nozzle flow and its transition from sea level to altitude mode. Therefore, experimental measurements on a Dual-Bell nozzle with trans- and supersonic external flows about a launcher-like forebody were carried out in the Trisonic Wind Tunnel Munich with particle image velocimetry, static pressure measurements and the schlieren technique. A strongly correlated interaction exists between a transonic external flow with the nozzle flow in its sea level mode. At supersonic external flow conditions, a Prandtl–Meyer expansion about the nozzle’s lip decreases the pressure in the vicinity of the nozzle exit by about 55%. Therefore a new definition for the important design criterion of the nozzle pressure ratio was suggested, which considers this drastic pressure drop. Experiments during transitioning of the nozzle from sea level to altitude mode show that an interaction about the nozzle’s lip causes an inherently unstable nozzle state at supersonic free-stream conditions. This instability causes the nozzle to transition and retransition, or flip-flop, between its two modes. This instability can be eliminated by designing a Dual-Bell nozzle to transition during sub-/transonic external flow conditions.

The turbulent wake of an axisymmetric generic space launcher equipped with a dual-bell nozzle is simulated at transonic ($$Ma_\infty = 0.8$$ and $$Re_D = 4.3\cdot 10^5$$) and supersonic ($$Ma_\infty = 3$$ and $$Re_D = 1.2\cdot 10^6$$) freestream conditions, to investigate the influence of the dual-bell nozzle jet onto the wake flow and vice versa. In addition, flow control by means of four in circumferential direction equally distributed jets injecting air encountering the backflow in the recirculation region is utilized to determine if the coherence of the wake and consequently, the buffet loads can be reduced by flow control. The simulations are performed using a zonal RANS/LES approach. The time-resolved flow field data are analyzed by classical spectral analysis, two-point correlation analysis, and dynamic mode decomposition (DMD). At supersonic freestream conditions, the nozzle counter pressure is reduced by the expansion of the outer flow around the nozzle lip leading to a decreased transition nozzle pressure ratio. In the transonic configuration a spatio-temporal mode with an eigenvalue matching the characteristic buffet frequency of $$Sr_D=0.2$$ is extracted by the spectral and DMD analysis. The spatial shape of the detected mode describes an antisymmetric wave-like undulating motion of the shear layer inducing the low frequency dynamic buffet loads. By flow control this antisymmetric coherent motion is weakened leading to a reduction of the buffet loads on the nozzle fairing.

The dual-bell nozzle is an altitude adaptive nozzle concept that offers two operation modes. In the framework of the German Research Foundation Special Research Field SFB TRR40, the last twelve years have been dedicated to study the dual-bell nozzle characteristics, both experimentally and numerically. The obtained understanding on nozzle contour and inflection design, transition behavior and transition prediction enabled various follow-ups like a wind tunnel study on the dual-bell wake flow, a shock generator study on a film cooled wall inflection or, in higher scale, the hot firing test of a thrust chamber featuring a film cooled dual-bell nozzle. A parametrical system study revealed the influence of the nozzle geometry on the flow behavior and the resulting launcher performance increase.

International audience; Dual-bell axisymmetric propulsive nozzle, as the most prominent altitude compensating nozzle concept has been investigated for transition modes and effects. Experimental and numerical investigations show that dual-bell nozzle can be very effective for desired large envelope launcher stage trajectories. Transition regimes are studied in the high altitude simulation wind-tunnel and supported by the numerical simulations. Analysis of the employed diagnostics shows an existing particular hysteresis between the transitioning modes and pinpoints to possibilities to control the early transition. Further numerical investigation on active flow control indicates the promising potentials of the proposed fluidic injection system

In order to gain the supersonic and hypersonic speed, the nozzle plays a major role in the aviation industry. Based on the nozzle design, the required mach number can be calculated. In this paper, the convergent divergent bell nozzle which is basically used for the supersonic flow is analysed and designed using CATIA Software. The mesh of the designed nozzle is carried out in ANSA and then analysed using CFD. Different nozzle designs are assessed through of CFD analysis to choose the best performing nozzle that can be manufactured for experiments. For the experimental test Raspberry Pi, pressure sensor and Python coding was developed to test bell nozzle pressure.

Numerical analysis is performed on a dual bell nozzle designed for an experimental semi-cryogenic rocket engine application to predict the effect of film injection on nozzle flow characteristics and transitional behavior. Both start-up and high-altitude operation regimes are simulated to predict actual flight conditions of a dual bell nozzle. Coolant is injected at two locations and at different operating conditions to study the effect of various parameters on flow behavior. Changes in shock interaction and shock patterns during startup flow with and without the injection of coolant are analyzed. An empirical relation for conventional nozzles has been applied to determine the transition pressure ratio. Actual transition flow from base to extension nozzle is found to be different from the calculated transition pressure ratio and flow transition occurs early in case of secondary injection. A large reduction in wall temperature is observed with the film flowing along the nozzle wall and it does not have any adverse effect on flow transition. Instead of efficiency, a film cooling effectiveness for high temperature flow is used to understand the effect of coolant temperature on reduction of heat flux and mixing characteristics. Mass flow rate of coolant is seen to have significant effect on mixing and flow separation.

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.

Previous research on Dual-Bell nozzle flow always neglected the influence of the outer flow on the nozzle flow and its transition from sea level to altitude mode. Therefore, experimental measurements on a Dual-Bell nozzle with trans- (\(Ma_\infty = 0.8\)) and a supersonic (\(Ma_\infty \) = 1.6 & 2.0) external flows about a launcher-like forebody were carried out in the Trisonic Wind Tunnel Munich with particle image velocimetry and the schlieren technique. The sea level mode was investigated in transonic conditions, whereas transition and the altitude mode took place in supersonic conditions. The results show that there is a strong interaction between the nozzle flow and the outer flow in sea level mode, highly dominated by screeching. In contrast, there is no apparent correlation between the nozzle flow and the outer flow in the altitude mode. Transition from sea level to altitude mode shows multiple retransitions over a wide range of nozzle pressure ratios. This is due to an interaction of the nozzle flow with a supersonic expansion about the nozzle’s lip. For the feasibility of the Dual-Bell concept, future research should investigate if a transition in transonic free-stream conditions is possible without the flip-flop effect.

Double Divergent Nozzle is a type of altitude compensating nozzle having less weight and minimum modifications when compared to other altitude compensating nozzle concepts. Planar double divergent nozzle is similar to a dual bell nozzle but, it is having a rectangular cross section. The present study involves numerical analysis of a planar double divergent nozzle and its comparison with a single divergent nozzle using ANSYS Fluent Software. The flow is simulated over various Nozzle Pressure Ratios (NPRs) and corresponding wall static pressure and specific impulse are studied. A sneak transition can be observed when the flow changes from low altitude mode to high altitude mode. The transition occurs early for double divergent nozzle having low NPR and vice versa. The wall static pressure variation shows a strong dependence of inflection angle on pressure variation, separation location and recirculation regions inside the nozzle.

To evaluate the impact of dual-bell nozzles on the payload mass delivered into geostationary transfer orbit by Ariane 5 Evolution Cryotechnique Type A (ECA), detailed studies were conducted. For this purpose, a multitude of Vulcain 2 extension contours were designed. The two variation parameters were the starting point and the inflection angle of the nozzle extension. As the most upstream starting point, the position of the turbine exhaust gas injection was chosen. Geometrical restrictions were imposed by the launch pad ELA 3. Considering these parameters, an analytical and a numerical method were applied to predict the impact of the dual-bell nozzle on the payload mass. The analytical approach yields a correlation between specific impulse, nozzle mass, and payload mass increment. The numerical approach was conducted applying German Aerospace Research Center’s trajectory simulation code Trajectory Optimization and Simulation of Conventional and Advanced Transport Systems. Both calculation procedures yield...

Numerical study was conducted on the LOX-RP1 engine bell nozzle to replace the dual bell nozzle for determining the optimum expansion ratio and greater thrust at sea level. Here base nozzle is a conventional Rao’s TIC nozzle, where throat radius, exit radius, inflection angle and exit angles are 138mm, 826mm, 33° and 8° respectively. The total length of dual bell nozzle and existing bell nozzle are kept same. Numerical analyses were carried out in ANSYS FLUENT software on different dual bell nozzle geometry to evaluate the thrust and expansion ratio. Numerical analysis is performed by two dimensional, axi-symmetric, steady state, pressure based solver with SST k-ω turbulence model at different ambient pressures to recreate the patterns of flow in the nozzle at different nozzle pressure ratio (NPR) and expansion ratio. The profile has made by commercial software SOLID WORKSTM. Numerical simulations and flow separation locations are validated with the experimental data published earlier. It is observed that with area ratio 150 and inflection angle 15° the thrust has increased by 5.46% at sea level and around 10.5% in vacuum, where the increase in exit radius and mass is only 96 mm and 6 kg respectively.

The nozzle is an element of the rocket engine in which the potential energy of gases generated during combustion is converted into the kinetic energy of the gas flow. The design parameters of the nozzle have a decisive influence on the ballistic characteristics of the engine. Designing a nozzle assembly is therefore one of the most responsible stages in developing a rocket engine design. The paper presents the results of the simulation of three types of rocket propulsion nozzles. Calculations were made using CFD (Computational Fluid Dynamics) in ANSYS Fluent software. The analyzed types of nozzles differ in shape. The analysis referred to conical nozzle, a bell type nozzle with a conical supersonic part and a bell type nozzle. Calculation results are presented as pressure, velocity, turbulence kinetic energy and eddy viscosity distributions in the longitudinal section. The results show that the cone nozzle generates strong turbulence in the critical section, which negatively affects the flow of the working gas. In the case of a bell nozzle, the transformation of the wall caused the elimination of flow disturbances in the critical section, which reduced the probability of waves that can form before or after the trailing edge. The most sophisticated construction, the bell-type nozzle, allows for maximizing performance without adding extra weight. The bell type nozzle can be used as a starter and auxiliary engine nozzle due to its advantages.

The design of a film cooled dual-bell nozzle is presented. The nozzle is part of a thrust chamber assembly that adopts an existing LOX/GH2 thrust chamber. The dual-bell base nozzle, including the gaseous hydrogen cooling film injection, is a downscaled redesign of an already tested film cooled TIC nozzle. Future hot flow tests at the test facility P8 will study the impact of a ROF variation and a cooling film mass flow variation on the operation mode transition of the dual-bell. For this reason, a homogeneous hot flow and cooling film distribution are mandatory. To meet those demands, extensive numerical studies were performed and design optimizations were derived. The test specimen will be operated under sea level conditions.

This Paper has been excavating and investigating two different types of Aerospike nozzles such as external aerospike nozzle and internal-external aerospike nozzle. The main reason for using aerospike nozzle over bell nozzle is automatic altitude compensation. With the help of Method of characteristics and MATLAB codes, geometry of both the nozzles has been created in GAMBIT and flow analysis has been carried out in FLUENT for desired exit Mach number. Analysis has been carried out for both types of nozzles with different truncations. However, in this paper the fuel was taken as Papi 94 and Hydroxyl-terminated polybutadiene (HTPB) with N 2 0 as an oxidant and the combustion products were entered in FLUENT by selecting species transport. The desired exit Mach number resulted in FLUENT is compared to the designed exit Mach number in code for both the two nozzles and their truncations.

Dual bell nozzle is one of the feasible and cost effective techniques for altitude adaptation. Planar double divergent nozzle with a rectangular cross section was designed for two different NPR׳s to simulate and investigate the flow regimes similar to those inside the dual bell nozzle. Measurements involved flow visualization using Schlieren technique and wall static pressure measurements. The flow transition between the two nozzles at the respective inflection points and the formation of recirculation region due to flow separation was analyzed in detail. Cold flow tests were performed on the double divergent nozzle in the over-expanded conditions to study the shock wave characteristics. The results obtained from the two independent double divergent nozzles were compared with those obtained from a single divergent nozzle of the same area ratio. From the experiments it was observed that inflection angle played a key role in defining the type of shock structures existing inside the double divergent nozzles.

Dual-bell nozzle can overcome the performance losses of the conventional bell-shaped nozzles which induced by off-design operations with either over-expanded or under-expanded exhaust flow and minimize the losses of the specific impulse. In United States, Rocketdyne analyzed thrust characteristics according to the shape of the expansion nozzle and NASA conducted hot firing tests with various altitudes. DLR, which is one of the research institute of the Europe, is carrying out research for the different cases of inflection angle, nozzle length and expansion ratio. MAI of Russia applied the slot nozzle to the expansion region in order to reduce the performance losses. In Asia, both the Japan and the India are researching on the dual-bell nozzle and Mitsubishi cooperation of the Japan registered its patent. In this paper, concepts and performance of dual-bell nozzle, which can compensate altitude, are investigated and trends of current research are summarized. It is necessary for Korea to research on the dual-bell nozzle for lucrative space development.

The development of the liquid propellant engine within the frame of the SMILE Project is described in the current paper. The analysis of a generic liquid propelled rocket launcher was performed and resulted in a description of the engines of each stage. The focus of this work is on the performance analysis of the first stage engine. As the liquid engine concept for the first stage aims for a cost efficient solution, three separate concepts are discussed. Thus, a conventional bell nozzle main combustion chamber and two aerospike concepts with multiple combustor chamber configurations geometries are compared and evaluated with respect to each other. The basis for the analytical comparison is a weight driven performance calculation. It is shown that 30 unitary combustion chambers in linear aerospike configuration can compete, regardless of their higher weight, against the other two concepts. An analytical calculation for the maximum change of velocity for each configuration is performed over a variable initial acceleration of the first stage. The analysis indicated only a minor impact on the change of velocity, the difference between best and worst performing configurations is less than 30 m/s, assuming an initial acceleration of 1.4 for the first stage. It can be further stated that a structure, which is fully dependent on the initial acceleration, results in minor differences of the change of velocity for the first stage.

The structural body of a rocket nozzle is dynamically exposed to high thermal and mechanical loads, which are caused by the hot gas flow. The flow field is itself significantly influenced by the shape and temperature changes of the nozzle wall. Additionally, for typical hot nozzle walls thermal radiation cannot be neglected. To ensure reliable results the unsteady aerothermoelastic interaction and the radiation have to be taken into account in computational investigations. A coupling method is proposed which is able to transfer unsteady mechanical and thermal loads from a fluid solver to a structural solver and transfer deformations and surface temperatures back. Especially in early design steps a fast but reliable simulation approach is necessary. Therefore, a solver using reduced structural models is applied. To be able to consider the radiation within the aerothermoelastic simulation, the structural solver is extended by appropriate computational methods. Also the ability to consider temperature dependent material behaviour is implemented. The presented computational results show the capability of the current fluid-structure interaction method to simulate the aerothermal behaviour of a dual bell nozzle. The computational predictions of the structural solver of the temperature field considering external and internal radiation and also considering temperature-dependent material properties are validated against computational results from a commercial FE solver and experiments.

Concept of Altitude adaptive rocket nozzles recentl y received greater importance and interest in the sexplorations and other such applications in space and rocket technology. The op erations reliability of rocket launcher and the ear th to orbit rocket launch are the crucial for the space transportation in the future. The performance of the engine components such as t he power plant and the thrust delivery of the engine such as nozzles are in renov ation for the greater performance and applicability for complex space applications. In the recent progress of the combustion expansion system the rocket nozzles are greatly revised from both application and design perspectives. One of such development is the dual b ell nozzle. The publications indicate that the rese arch on the concept of dual bell nozzle is tardy and there is no much progress from the inception of the idea. The specific application purpose designs are tested experimentally and implemented but the large scale development can only be possible if the generalized design parameters can be suggested. In the present paper one of such nozzle is selected and studied using computational fluid d ynamics (CFD) and the results are synthesized for bench marking the general appro ach to study the Dual Bell nozzles. The result shows the variation in the Mach number, pressure, temperature distribution and turbulence i ntensity.

Aerospike nozzle is often described as an inside-out bell shaped nozzle and named for its prominent spike located at centre. This project mainly focuses on design of various types of nozzles such as aerospike, truncated aerospike and spike bell nozzles. In theoretical approach, the efficiency of rocket nozzles is calculated with design parameters such as throat area, exit area, chamber pressure and so on. From theoretical approach it is inferred that the spike bell nozzle has an increased value of thrust when compared to other nozzle types. So meshing and internal flow analysis is carried out for spike bell nozzle alone by using ANSYS CFX 13.0 software. From this flow analysis it can be inferred that the exit Mach number of spike bell nozzle comes close to the theoretical design exit Mach number .Also the flow simulations for the spike bell nozzle is carried out for various parameters such as pressure, temperature and velocity.

The characteristic contour inflection of a dual bell nozzle is the key to altitude adaption. In sea level conditions, it forces the flow to a symmetrical separation, limiting the side load generation and increasing the thrust. After the transition, under high altitude conditions, the nozzle flows full, increasing the vacuum thrust. A hot flow experimental study has been conducted at the German Aerospace Center on a planar dual bell nozzle. The wall temperature distribution has been measured at various depths for the determination of the heat flux through the wall. The region of the inflection is of particular interest for the adjustment of the conventional cooling methods of dual bell nozzles. The contour inflection leads to a local increase of the thermal loads. In addition to the tests, the flow behavior and thermal loads have been calculated with CFD method and compared with the experiment.

A flow oscillation phenomenon of the dual-bell nozzle at low-altitude operationmode is investigated with cold-flow experiments and numerical simulations. Results indicate the existence of three kinds of periodic flow oscillation modes: a symmetric lowest-frequency oscillation mode, which is an acoustic resonance in the axial direction in an extensionpart of thenozzle, and twohigh-frequencymodes lateral to thenozzle axis originatingdue to vortex shedding andacoustic resonance in the extension nozzle cross section.These periodicflowoscillations are characteristic to dualbell nozzles, and the extension part is the cause of these fluctuations, which are not observed in single-bell nozzles.

Unsteady flow characteristics in a dual-bell altitude compensation nozzle at low altitude operation modes are examined. At the low altitude operation mode, a jet separation point is fixed at the nozzle wall inflection. However, experimental and numerical results show that three modes of periodic jet fluctuations take place at certain range of nozzle pressure ratios. The lowest frequency mode is a symmetric modes and is considered to be caused by the acoustic resonance in longitudinal direction in the extension part of the nozzle. Two kinds of asymmetric modes have higher frequency than the symmetric mode. According to the numerical results, one of them is a jet screech tone caused by the convection of large scale vortices beside the separated jet. The other mode has the same frequency with a resonance mode in the diametrical direction of the cross section of the extension part.

Dual-bell nozzle is one of the altitude compensation nozzle concepts, which consists of two bell-type nozzle with different geometric area ratios. The dual bell nozzle has two operation modes—a low altitude operation mode and a high altitude operation mode. However, the practical transition between these operation modes occurs at lower nozzle pressure ratio (NPR) than the optimum one. In the present study, the effect of introduction of film-cooling flow into the nozzle is examined in order to control the operation mode transition NPR. The experimental and numerical results show that the operation mode transition can be control by injecting the film cooling flow at appropriate inlet pressure and flow rate. By injecting the film cooling flow, the axial momentum near the wall in the extension part becomes small and the separation point movement becomes slower. As a result, the dip of the thrust coefficient during the operation mode transition becomes smaller.

For the past several decades, research dealing with hypersonic flight regimes has been restricted mainly to military applications. Hypersonic transportation could be a possible and affordable solution to travel in the medium term and there is renewed interest from several private organisations for commercial exploitation in this direction. Various combined cycle propulsion configurations have been proposed and the present paper deals with implications for the nozzle component of a ramjet configuration as part of one such combined cycle propulsion configuration. An investigation was undertaken for a method of turbine-based propulsion which enables the hypersonic vehicle to take off under its own power and propel the aircraft under different mission profiles into ramjet operational Mach regimes. The present study details an optimal method of ramjet exhaust expansion to produce sufficient thrust to propel the vehicle into altitudes and Mach regimes where scramjet operation can be initiated. This aspect includes a Computational Fluid Dynamics (CFD)-based geometric study to determine the optimal configuration to provide the best thrust values. The CFD parametric analysis investigated three candidate nozzles and indicated that the dual bell nozzle design produced the highest thrust values when compared to other nozzle geometries. The altitude adaptation study also validated the effectiveness of the nozzle thrust at various altitudes without compromising its thrust-producing capabilities. Computational data were validated against published experimental data, which indicated that the computed values correlated well with the experimental data.

The flow structure of the jet emanating from supersonic nozzle with different profiles at underexpanded condition were investigated systematically. Experiments were carried out on conical, bell and double parabolic nozzle at Nozzle Pressure Ratios (NPR) of 4.5, 5 and 5.5. For the sake of comparison, the divergent length was the same for all the nozzles. Numerical simulations were also conducted in a two dimensional computational domain taking a density-based solver with RANS equations and SST k-ω turbulence model. The experimental and numerical results matched well all NPRs. The shock cell length, potential core length and supersonic core length at different NPRs were determined from experiments and compared to numerical predictions. These values are also compared to the predicted values based on the equations previously developed by researchers. It is observed that there is not much influence in nozzle profiles in the shock cell length, potential core length and supersonic core length.

High energy pulses of a CO2 laser are focused in a parabolic mirror yielding to a laser-supported detonation. The generated thrust acting on the reflector as a bell nozzle is studied in multiple pulse free flight experiments with respect to axial, lateral and angular momentum coupling. The employment of an ignition pin on the reflector's axis of symmetry lowering the ignition threshold by several orders of magnitude is found to provide for a reproducible detonation process. The axial momentum coupling of each pulse is analyzed with respect to initial lateral offset and tilt during the flight. High speed analyses of recorded flights indicate that lateral momentum components occur re-centering the thruster on the beam. Thrust vector steering can be realized by tilt of the ignition pin inside the thruster, thus shifting the detonation. A design model of a laser-driven rocket including a remotely accessible steering gear was developed and tested successfully.

In the present paper, the experimental and numerical investigations are described to clarify how the axial compression of the extension nozzle influences the unsteady flow fields and side load in startup and shutdown transients for a 2D dual bell nozzle. In the experiments, the flow fields were visualized and recorded by the schlieren method and the high-speed camera, respectively. Side loads were measured by a strain gauge. The flow visualizations show three flow patterns, i.e. an attached flow, an axial flow and a full flow, and the transients between the two flow patterns of those three. The strain gauge measurements indicate that side loads are produced by asymmetric transitions between the two flow patterns of the three The numerical results show the qualitative agreements with the experimental ones. They indicated that the jet oscillations are produced by vortices discharged from the extension nozzle exit alternatively and cause side loads. The axial compression of the extension nozzle changes the speed of the separation point movement.

The paper studies the definition of design-basis appearance of Electrothermal Microthrusters (ETMT) with Autonomous Heating Elements (AHE) for Correcting propulsion System (CPS) of manoeuvrable Small Space Vehicles (SSV). The findings of present study are of particular practical interest in connection with creation of CPS for manoeuvrable SSV of ultra-low mass (under 10 kg). In the framework of the taken approach, multipurpose method of structural design has defined design parameters and design appearance of ETMT. Different characteristics of ETMT are provided due to replaceable conical and bell nozzle linings’ usage without change of ETMT basic design. Design maps of ETMT with AHE have been developed, aimed to specific performances’ upgrade of CPS in conditions of low power consumption. The results of undertaken land experimental study are quoted, including conditions of vacuum-chamber using nitrogen and air as working medium of CPS. Basing on the undertaken experimental study the possibility in principle of the usage of ETMT with AHE as part of CPS for space vehicles of gross mass under 10 kg is shown. Moreover, design map, implementation approach and further upgrade of ETMT with AHE are defined, among which it is possible to emphasize the design map of ETMT integrated with evaporator, providing the increase of specific burst of power up to 30-35 % within allocated power consumption as part of SSV