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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.

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.

The application of dual-bell nozzles to existing space launchers offers the possibility to significantly increase their payload capability. This benefit is related to the ability of this kind of nozzles to switch between a low-altitude and a high-altitude working mode. However, the transition between the working modes introduces several problems like for example the possible development of high side-loads and a significant jump in the thrust magnitude. Furthermore, the natural transition takes place earlier with respect to the ideal case and this limits the performance gain. Among the different strategies proposed to control the transition, the injection of a secondary flow close to the inflection point represents a promising solution. In this work, the use of fluidic control is investigated by means of numerical simulations. After a preliminary calibration of a modified version of the Spalart-Allmaras turbulence model, the effects of the injection on the transitional nozzle pressure ratio are investigated.

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.

Space exploration and space tourism have now become a raging competition among the developed nations. For this reason, different types of advanced rocket nozzles with prospective privileges are introduced. Altitude adaptive dual bell nozzle will soon replace the conventional nozzles for the first stage rocket launcher. Indeed, this nozzle has auto adaption capability based on altitude. The major feature of a dual bell nozzle is the two bell-shaped contours separated by an inflection point. This nozzle has left rooms for researchers to test different flight conditions and transition characteristics. In this paper, a dual bell nozzle contour has been developed in MATLAB and analysed for different thermodynamic parameters. ANSYS Fluent is used in analysing flow through the nozzle. Shadowgraph imaging technique is used for measuring density gradient and compared it with fluent results. The simulations were performed by using the k-epsilon turbulence model.

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...

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 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.

International audience; The main objective of a dual bell nozzle is the enhancement of performances based on the principle of auto-adaptation in accordance with the altitude. Indeed, this system has as advantage the auto-adaptation of the flow for two operating modes (at low and high altitude) without mechanical activation. The principle is theoretically simple but structural forces involved can be significant. In this study, a numerical method for the design of this type of nozzle is developed. On the one hand, it is based on a transonic flow approaches to define the starting line on which the supersonic calculations will be initiated. On the other hand, the method of characteristics is used to draw the base nozzle profile. Knowing that the latter is assimilated as a polynomial of the second degree, its constants are calculated from initial conditions. In order to minimize the weight of this nozzle, its truncation proves necessary; this is performed at a point where the best compromise (weight / performances) was respected. The profile of the second curve is calculated to give a constant wall pressure. This is achieved by using the direct method of characteristics applied for a centered expansion wave that the intensity is P2/P1 at the junction. Once the profile is generated, an analysis of the thermodynamic-parameters evolution (such as: pressure, Mach number) and aerodynamic performances is conducted. For more consistency, our results are compared with numerical databases of ONERA nozzle. Simulations of flow in the nozzle with Ansys 13.0 environment for different types of meshes are presented. Also, to offset the effects of the boundary layer, the simulations were performed by using the k-ω SST turbulence model. The obtained results by the method of characteristics and numerical simulation are compared to the computed results of the literature and it was found good agreement and similarity.

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.

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.

New rocket designs are being adopted to increase the performance of the current satellite launch vehicles (SLVs). But, the aerospike (or plug) nozzle concept that has been under development since the 1950s is yet to be utilized on a launch platform. Due to its ability to adjust the environment by altering the outer jet boundary, the aerospike nozzle delivers better performance compared to present day bell nozzle. An aerospike nozzle is designed for 20 bar pressure ratio. In order to improve the performance of the aerospike nozzle for various conditions, optimization of the nozzle was carried out for some important design parameters and their performances were studied for cold flow conditions. Initially a model of an aerospike nozzle is created for certain parameters, then the optimization process is carried out for the nozzle (Truncated model & Base bleed model). Optimized model is designed by the software GAMBIT and the flow behavior is analyzed by the Computational Fluid Dynamics (CFD) software called FLUENT. Comparison also takes place between the full length and the optimized models.

Three-dimensional reactive full Navier-Stokes simulations of hydrogen/oxygen flow in a nozzle with area ratio of 140 are evaluated using a detailed chemical kinetics reaction model. The flows over each overlapped grid are computed using a unified zonal method technique. The flow field computed by the present CFD code shows good agreement with the flow field computed by the TDK code except the position of the weak compression wave focusing on the nozzle axis. Isp profiles of the subscale nozzle computed by the CFD code assuming a laminar condition agree well with the experiment and those calculated by the TDK code. The measured heat flux distributions at higher nozzle expansion ratios also agree well with the laminar CFD results. The inconsistency of the prediction of turbulent boundary layer loss between TDK and CFD leads to some difference in Isp. The effect of grid resolution on Isp turned out to be small in the present conditions.

A sub-scale test campaign has been conducted in the framework of cooperation between the German Aerospace Center (DLR) and the Japan Aerospace Exploration Agency (JAXA). Both partners had great experience in cold flow dual bell nozzle investigation. The present campaign took place at P6.1 test facility at DLR Lampoldshausen with a common combustion chamber. Two test specimens were tested with similar conditions and measurement systems. The conditions provided at the bench were the combustion of gaseous methane with liquid oxygen at pressure up to 60 bar. The combustion chamber pressure was varied up and down with constant ramps while keeping the mixture ratio constant. Objective of the study was, a part from demonstrating the precision of the design tool, to investigate the sensitivity of the transition to flow property variations (like the one due to the mixture ratio). The transition and retransition conditions, the range of hysteresis and the transition velocity were of particular interest for DLR side. The work of JAXA concentrates on heat flux and combustion instabilities and is presented in the second part of this work.

Single-stage-to-orbit (SSTO) rocket technology offers the potential to substantially reduce launch costs, but is not considered practical for conventional launch vehicles. However, new research in composite propellant tank technology opens the field for reevaluation. One technology that increases the efficiency and feasibility of SSTO flight is an altitude compensating rocket engine nozzle, as opposed to a conventional constant-area, bell nozzle. By implementing an altitude compensation nozzle, such as an aerospike nozzle for inatmosphere flight, the propellant mass fraction (PMF) may be reduced by as much as two percent compared to a conventional rocket engine. In this paper, Optimal Trajectories by Implicit Simulation (OTIS) is used to model SSTO flight trajectories by comparing an aerospike nozzle to a conventional bell nozzle; this includes vacuum thrust, vacuum specific impulse, and nozzle combinations to show that nozzle variability increases the efficiency of SSTO flight through a reduction in PMF. Results suggest that limiting the altitude compensation to lower altitudes further decreases PMF by 0.1-0.2%. Uncertainty estimates suggest that this added benefit is within the range of uncertainty of OTIS, and performs at least as well as the theoretical, infinite nozzle expansion. It is apparent that altitude compensating nozzles outperform conventional bell nozzle designs.

The proliferation of earth orbital missions enabled by nanosatellites, of which CubeSats are the most common, has focused increased attention on low-cost, mobile launch systems. Advanced nozzle design can address the demand for higher booster performance, and has resulted in the investigation of alternate nozzle geometry. A dual-bell nozzle represents a novel approach to this problem by utilizing two theoretically ideal design points over the ascent trajectory instead of just one. The primary focus for this project was to investigate the performance of a representative dual-bell nozzle both computationally and experimentally. First, a dual-bell nozzle was designed which could be used on a representative, nanosatellite launch system with design altitudes of 3 km and 17 km. The nozzle contours were modified to facilitate machining, and incorporated in a supersonic wind tunnel. The dual-bell nozzle fabricated for testing in the wind tunnel had a “first contour” expansion ratio of 1:9.85 corresponding to an ideal design pressure ratio of 134. The downstream, “second contour” had an expansion ratio of 1:19.7 corresponding to an ideal design pressure ratio of 374. The conventional nozzle fabricated for testing in the wind tunnel had an expansion ratio of 1:11.3 corresponding to an ideal design pressure ratio of 102. Imaging of flow structures was achieved using a schlieren optical system. In addition, the computational fluid dynamics program ANSYS Fluent was used to model both a hypothetical, full-scale nozzle as well as the scaled version tested in the wind tunnel. Simulations were performed for full-scale, dual-bell nozzles operating at pressure ratios (chamber to ambient) of 20, 50, 100, 150 and 1000. These results were then compared to a similar conventional nozzle for each case. Preliminary schlieren results indicate that the scale wind tunnel dual-bell contours show flow structures consistent with the Fluent results.

Flow fluctuations in or around rocket nozzles can lead to a deformation of their structure. In case of separated nozzle flow, an inward bended nozzle wall shifts the flow separation position downstream, increasing the bending force lever arm. This flow-structure interaction is a self-reinforcing effect, which can potentially lead to heavy structural damages. These disturbances can be produced either by combustion instabilities or ambient pressure fluctuations. Under certain conditions, they may excite the bell nozzle eigenfrequency. As a result, the structure responds and the nozzle ovalizes. For this reason, the investigation of flow-structure interaction in ovalized nozzles under separated flow condition is difficult to realize [1–4].

Three-dimensional reactive full Navier-Stokes simulations of hydrogen/oxygen flow in a nozzle having area ratio of 140 is evaluated using a detailed chemical kinetics reaction model. The flow over each overlapped grids is computed using a unified zonal method technique. The flow field computed by the present CFD code shows good agreement with the flow field computed by the TDK code except the position of the weak compression wave focusing on the nozzle axis. Isp profiles of the subscale nozzle computed by the CFD code assuming a laminar condition agree well with the experiment and those calculated by the TDK code. The measured heat flux distributes at higher nozzle expansion ratios also agree well with the laminar CFD results. The inconsistency of the prediction of turbulent boundary layer loss between TDK and CFD leads to some difference on Isp. The effect of grid resolution turned out to be small on Isp in the present conditions.

Dual-bell nozzles represent a possible solution to improve the performance of large liquid rocket engines for launcher first stages. The present paper studies the role of the second bell shape on the side loads that can occur during the transition between the two main operating modes of the nozzle. In particular, the design of the second bell profile is critically discussed on the basis of results obtained from suitable test cases. The analysis of performance and behavior during the transition is carried out by a validated turbulent Navier-Stokes solver. Geometries were generated by the method of characteristics. Results show that slightly different geometries designed by the method of characteristics yield modest performance changes but severe differences in behavior during the transition phase.

近い将来の再使用型ロケットエンジンの性能向上の一案として，伸展ノズルやデュアルベルノズルが有望であると考えられている．そこで，これらのノズルの基本特性を把握するために，４種のノズル（標準ノズル，伸展途中の過渡状態を模擬したステップノズル，デュアルベルノズルおよび可動伸展ノズル）を用いて，高空性能燃焼試験を実施した．試験で計測したノズル性能，圧力分布および熱伝達率等のデータをＣＦＤ解析結果と比較検討した．その結果，デュアルベルノズルの性能は標準ノズルやステップノズルよりも低く，現状のノズルコンターはさらに改良の余地があることが判明した．また，可動伸展ノズルの伸展時には，固定ノズルと伸展ノズル隙間からの燃焼ガスの逆流は認められず，ノズル壁面の熱伝達率は過渡的に約20％増加することが判明した．これらの現象はＣＦＤ解析結果とも一致し，ＣＦＤ解析によってノズルのステップ流れや逆流が予測できる目処を得た．

The dual bell nozzle belongs to the family of Altitude Compensating Nozzles (ACN), which constitutes an option of great interest to improve launcher first stage performance. The internal flow of this nozzle adapts to the external pressure by separating at the wall inflection at low altitude (first operating mode) and by full flowing at high altitude (second operating mode). Therefore the dual bell nozzle operates with a separated flow at take off and in the initial portion of the flight trajectory. This characteristic is of important concern since the external flow is neither steady nor axisymmetric, and its coupling with the internal flow separation can cause dangerous side loads. This work presents the results of a time accurate numerical analysis of the effect of unsteady ambient pressure on the separation point and on the shock system inside the dual bell nozzle, operating in the first mode. The external computational domain is characterized by an acoustically open downstream boundary. The numerical simulations show that the oscillation of the separation point is very sensitive to frequencies near the acoustic wave characteristic frequencies.

Projectile launch experiments were conducted using a new laser-driven, in-tube accelerator (LITA) in which the propellant was supplied not as pre-filled gas but from laser ablation. The launch tube had an inner diameter of 25 mm and was 1 m in length, made of acrylic. The 7.1-gram projectile had a parabolic nozzle and an ablator rod made of polyacetal that was located at the focus of the parabola. Laser pulses of 3.1 J to 3.8 were repetitively sent at a frequency of 50 Hz from a Transversely-Excited Atmospheric (TEA) CO2 laser (wavelength; 10.6 μm). Time-modulating the laser-ablative pressure, a large impulse performance was obtained with a moderate acceleration level. A momentum coupling coefficient of 1.5 mNs/J and a gross acceleration of 2.5g were obtained.

The paper presents the design of a 30kN multi-thruster annular aerospike engine as candidate for the first stage of a two-stage pressure-fed Nanosat Launch Vehicle (NLV). Detailed description of the injector and layout of single thruster are provided. A major enhancement compared to a classical multi-chamber design with straight thrusters is achieved by introducing nozzle curvature in the circumferential and in the longitudinal direction. This approach allows increasing the available exit area, or pressure ratio, thus improving performance. The design work is supported by numerical analyses of the flow inside the curved thrusters and along the engine including slipstream effects using Reynolds-Averaged Navier-Stokes simulations with realizable k-e turbulence modeling. The computations indicate an improved performance of the dual curved thrusters engine, which is about 2% above the linear design and above an ideal bell nozzle of equal area ratio. Simultaneously, the use of a larger expansion ratio reduces base drag in the transonic flight regime.

The dual bell nozzle features two operation modes permitting an altitude adaption. The geometry of the nozzle extension defines the transitional behavior from one operating mode to the other. In particular, the wall pressure distribution along the extension contour is critical for a fast transition. A contour extension featuring a constant wall pressure (CP) or with a positive pressure gradient (PP) ensures, in theory, an optimal transition. Experimental and numerical researchs have been conducted at the German Aerospace Center to optimize the transitional behavior through extension geometry variation. Three nozzles have been investigated: a CP-reference contour and two PP-extensions. The positive wall pressure gradient in the extension leads to a better stability of the operation modes (wider range of hysteresis). The wall pressure gradient has also a critical influence on the sneak transition length and duration and should be therefore chosen accordingly.

The present research is conducted to design the optimum nozzles for RCX250 engine, that is designed to produce maximum thrust of 250 kgf with combination of LOX and Kerosene as its propellant. The new nozzles were determined to be parabolic nozzle, with conical nozzle as its comparison. The parabolic nozzle was designed using Thrust Optimized Parabolic (TOP) method invented by G.V.R.Rao. TOP nozzle design method is performed by approximating a Thrust Optimized Contoured (TOC) Nozzle using parabolic equation. The method would result more efficient nozzle than conical or ideal bell nozzle. Further, the parabolic nozzle were modified in its initial and exit angle to create uniform velocities distribution at nozzle exit. A Computational Fluid Dynamics Method (CFD) is used to simulate the nozzle designs. The simulation was carried out in axis-symmetric condition using commercial CFD software. The simulation results show that MOD 1 nozzle, with initial angle (θN) 26 deg and exit angle (θe) 12 deg, gives maximum thrust, which is 4.67 % higher than reference conical nozzle. Key words:Liquid rocket, Parabolic nozzle, Thrust, CFD

Publisher Summary This chapter explores liquid rocket motors, rocket propellants, propellant tank, and feed system design. The rocket is the simplest jet propulsion device in that it carries all its propellant and liberates the chemical energy in that propellant under virtually quiescent conditions in the combustion chamber. The rocket motor is basically a nozzle used to transform the heat energy in the combustion chamber gas into kinetic energy by guiding the nearly isentropic expansion of that gas through the increasing cross-sectional area of the nozzle. Liquid fuel and oxidizer combinations have high energy density on a volumetric basis, and therefore the packaging of such propellants is efficient in terms of vehicle frontal area. The nozzle on the other hand produces a steady flow of high-velocity gas parallel to its centerline in order to produce thrust along that line of action. Because all aerospace vehicle operations are constrained by weight considerations, it is desired that the nozzle be as light as possible. Capacities of bell nozzle, plug nozzle, and extendable nozzles are also calculated. Combustion chamber, propellant injectors, cryogenic and hypergolic propellants are also discussed.

Flow instabilities under low-altitude operation mode in dual-bell nozzles are examined. Influences of base nozzle contour are investigated with two types of base nozzle contours: a truncated perfect (TP) nozzle and a compressed truncated perfect (CTP) nozzle. Experimental and numerical results show that the Mach disk oscillation in the axial direction occurs in the dual-bell nozzle with the CTP base nozzle. Such Mach disk oscillation causes higher level wall pressure oscillations than those in the dual bell nozzle with the TP base nozzle. Influences of geometric area ratio of the dual bell nozzle are also examined. As a result, the geometric area ratio does not affect the characteristics of the flow oscillation modes except for the amplitude of the wall pressure fluctuations.

The altitude adaptive dual bell nozzles can increase the overall performance of existing main stage rocket engines like Europe’s Vulcain 2. The Institute of Space propulsion of the German Aerospace Center in Lampoldshausen carried out in the past various cold and warm flow subscale tests to study the flow physics of dual bell nozzles fundamentally. Following this line and in order to increase the technology readiness level hot flow subscale tests were conducted to study the effects of a cooling film on the characteristics of a dual bell nozzle. The experimental results of a first campaign, performed with a representative Vulcain 2 like subscale engine, are presented and discussed.

The dual bell is a nozzle concept for altitude adaption. The flow separates at the contour inflection in sea level mode in a mainly controlled and symmetrical way, reducing the side load generation and increasing the thrust. The transition to altitude mode is reached when the flow suddenly attaches to the extension for an improved altitude thrust. The conditions of this transition and its evolution are the key for the study of dual bell nozzles. For a better understanding of the flow behavior, a 2D subscale dual bell model has been designed and tested at the German Aerospace Center (DLR). The tests were divided into two campaigns and performed under cold and warm flow conditions. The evolution of the shock system at the inflection during the transition was observed using schlieren optics. The planar nozzle was tested under various conditions in pressure and temperature. Both test campaigns have been recalculated in cooperation with Astrium. Numerical and experimental results are presented in this paper.

The dual bell nozzle offers a simple and efficient altitude adaption through its contour inflection, which insures symmetrical and controlled separation at sea level and a large area ratio at high altitude. The understanding of the transition from one operating mode to the other is the key to its prediction. Intensive cold flow investigation has led to an empirical criterion for a precise transition prediction. However, the testing conditions may have a great influence on the actual transition. The temperature of the flow and the nozzle wall, especially when driving test series, can shift the nozzle pressure ratio leading to the transition. A test campaign in a high altitude chamber has also shown the dependence of the transition condition with the chamber pressure, i.e. with the density of the separated backflow. For a reliable prediction of the dual-bell nozzle flow behavior, both on test facility and in real flight conditions, all these influences have to be identified and taken into account.

The dual bell nozzle offers a simple and efficient altitude adaption through its contour inflection, which insures a symmetrical and controlled separation at sea level and a large area ratio at high altitude. The conditions for the flow transition from one operating mode to the other have been intensively investigated in cold flow conditions. A test campaign has been conducted on two dual bell nozzle geometries to verify the theoretical prediction criterion and to investigate the influence of the temperature variations. Therefore, a thin walled axisymmetric and a planar nozzle were designed and tested under hot flow conditions. Wall temperature measurements have been made using both thermocouples and thermal imaging. The region of the contour inflection was of particular interest for the evaluation of thermal behavior in the nozzle wall.

A dual-bell nozzle has been tested in the ONERA-R2Ch wind tunnel within the CNES PERSEUS program. The wall pressure distributions and the thrust for the two §ow regimes have been characterized in the nozzle pressure ratio (NPR) range from 51 up to 597. A hysteresis on the transition NPR between the two §ow regimes has been observed according to the evolution of NPR. The duration for the switch between the two §ow regimes is less than 10 ms. The hysteresis of about 20%

An aerospike nozzle has the ability to perfectly expand the combustion gas to ambient conditions at all altitudes maximizing the efficiency of the nozzle and increasing the performance of the overall system. However, one of the inherent problems with this type of nozzle is the large amount of heat transferred to vital components due to its complex geometry, which could cause performance degradation or even nozzle failure. An aerospike nozzle design was developed that incorporates graphite inserts at the nozzle throat which reduces substantial material erosion at vital areas of the nozzle. This design was integrated with a custom hybrid rocket motor utilizing nitrous oxide and hydroxyl-terminated polybutadiene (HTPB) providing an excellent reusable test bed producing increased performance over that of a solid rocket motor. The unique integration between the hybrid and the aerospike provides a high performance rocket propulsion system capable of consistent experimental testing. A series of hot fire static tests were conducted in order to demonstrate that the aerospike nozzle can not only withstand the intense combustion gas temperatures, but furthermore yield a design that has higher efficiency over a range of design altitudes in comparison to a standard bell nozzle with fixed expansion ratio geometry. The aerospike nozzle was able to outperform an over-expanded bell nozzle by 6.4%, and perform within 0.9% of a perfectly expanded nozzle based on each nozzle’s CF. These tests prove that utilization of an aerospike nozzle has the potential to substantially increase the range or payload capabilities of modern launch vehicles or rockets.

The contour inflection of the dual bell nozzle forces the flow to a symmetrical and controlled separation under sea level conditions. At a certain altitude, the transition to high altitude mode takes place: the flow attaches rapidly the nozzle extension wall down to the exit plane. During this transition, the separation point moves in the extension generating potential high side load peaks due to its asymmetrical position. A campaign of cold flow tests has been conducted to evaluate the generation of side load. The phenomenology during each phase of the flow has been presented in the first part of this work. The influence of the various geometrical parameters on flow behavior and side load generation is investigated in this study. The extension length is shown to be the critical parameter for flow stability, transition duration and side load generation. A trade o has to be found for the optimization of the dual bell concept to rocket applications.

An experimental investigation has been carried out to study the flow characteristics of a dual-bell nozzle during its transition modes. Cold gas tests using gaseous nitrogen were carried out in (i) a horizontal test rig with nozzle exhausting into atmospheric conditions and, (ii) also inside a high altitude simulation chamber with nozzle operation under self-evacuation mode. Transient tests indicate that increasing 0P0 /0t reduces the transition time as well as the amplitude of pressure fluctuations of separation shock at wall inflection point. Sea-level tests show 15-17% decrease in transition NPR during subsequent tests in a single run primarily due to frost formation in the nozzle extension up to the wall inflection location. This reduces the wall inflection angle and hence, the transition NPR. However, tests inside the altitude chamber show nearly constant NPR value during subsequent runs primarily due to decrease in back temperature with decrease in back pressure that prevents any frost formation.

*† ‡ For a dual-bell nozzle to be employed in a future reusable launch vehicle, active control of working mode transition is essential. To establish a technique for such control, the effectiveness of secondary injection was investigated by conducting cold-flow tests. The secondary injection was introduced from the inflection position of the nozzle contour perpendicular to the thrust axis. As a result, the secondary injection successfully changed the transition condition. Based on the results, a system and a sequence to realize the active control of the transition between two working modes of a dual-bell nozzle are herein discussed.