Aerospike Nozzle Research Articles

Comparison of the Propulsion Performance of Aerospike and Bell-Shaped Nozzle Using Hydrogen Peroxide Monopropellant Under Sea-Level Condition

ABSTRACTThis study investigates numerically the performance of applying aerospike nozzle in a hydrogen peroxide mono-propellant propulsion system. A set of governing equations, including continuity, momentum, energy and species conservation equations with extended k-ε turbulence equations, are solved using the finite-volume method. The hydrogen peroxide mono-propellant is assumed to be fully decomposed into water vapor and oxygen after flowing through a catalyst bed before entering the nozzle. The aerospike nozzle is expected to have high performance even in deep throttling cases due to its self-compensating characteristics in a wide range of ambient pressure environments. The results show that the thrust coefficient efficiency (Cf,η) of this work exceeds 90% of the theoretical value with a nozzle pressure ratio (PR) in the range of 20 ~ 45. Many complex gas dynamics phenomena in the aerospike nozzle are found and explained in the paper. In addition, performance of the aerospike nozzle is compared with that of the bell-shape nozzle.

Design of axisymmetric aerospike nozzle based on modified MOC

In this study a direct method based on the method of characteristics with the assumptions of non-viscous flow and thermal perfect gas is provided to design of axisymmetric plug nozzle. Inputs for numeric code are including output Mach number, specific heat ratio, global gas constant and the number of discrete steps of Prandtl–Meyer expansion fan. In the design process, there has been no simplification and only to create unique spike geometry. Based on the criterion of locating the tip of the spike on the axis of symmetry, for each output Mach number, a correction factor has been extracted that it was the sole source of theoretical errors. Studied parameters include spike geometry, exhaust nozzle flow parallelism and uniformity. In the results by numerical analysis, consistent results with desired output Mach number and parallel of output flow with the axis of symmetry are obtained. The simple and direct formulation is used and so it performs very fast. Due to recognizing of the error resulted in the surface slope reduction factor, some movement in spike geometry relative to ideal conditions was raised so the effect of reduction factor will tend to zero by increasing of the design Mach number.

Computational study of performance characteristics for truncated conical aerospike nozzles

Aerospike nozzles are advanced rocket nozzles that can maintain its aerodynamic efficiency over a wide range of altitudes. It belongs to class of altitude compensating nozzles. A vehicle with an aerospike nozzle uses less fuel at low altitudes due to its altitude adaptability, where most missions have the greatest need for thrust. Aerospike nozzles are better suited to Single Stage to Orbit (SSTO) missions compared to conventional nozzles. In the current study, the flow through 20% and 40% aerospike nozzle is analyzed in detail using computational fluid dynamics technique. Steady state analysis with implicit formulation is carried out. Reynolds averaged Navier-Stokes equations are solved with the Spalart-Allmaras turbulence model. The results are compared with experimental results from previous work. The transition from open wake to closed wake happens in lower Nozzle Pressure Ratio for 20% as compared to 40% aerospike nozzle.

Analysis of Aerospike Nozzle Structural Contour Design Performance Optimization

Analysis of Aerospike Nozzle Structural Contour Design Performance Optimization

Design and Optimization of Aerospike nozzle using CFD

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 behaviour is analysed by the Computational Fluid Dynamics (CFD) software called FLUENT. Comparison also takes place between the full length and the optimized models.

Three Dimensional Computational Flow Simulation of Truncated Aerospike Nozzle Considering Different Plug Lengths

Propulsive systems for planetary landing and ascent needs mass reduction since the cost of delivering the required propellant and structural mass is very high. Aerospike nozzle has demonstrated a significant potential advantage for purely in-space applications. This paper mainly concentrates on the CFD analysis of aerospike nozzle for 100%, 75%, 50% and 25% plug length with the CFD analysis at cases considering multiple altitudes representing ideal, under-expansion and over-expansion conditions of the flow had been carried out. ANSYS CFX was used to simulate the flow at these different conditions. The visualisation obtained through the CFD simulation was used to compare the performance of different spike lengths at different above mentioned operation conditions. A deduction had been made that at overexpansion condition the spike nozzle is efficient of producing five percent more thrust and the loss of the thrust is proportional to the truncation.

Feasibility Studies on the Cooling of Hybrid Rocket Nozzles Using Supercritical Nitrous Oxide

Abstract The fundamental requirement in the development of Single-Stage-To-Orbit (SSTO) system is to implement a nozzle contour with altitude compensation characteristics which leads to increase in aerodynamic efficiencies. Also, the reusability of the SSTO is another challenge due to the thermal fatigue loads experienced by the launch vehicle. To overcome such challenges, cooling of the throat of the hybrid aerospike nozzle becomes essential as the nozzles experiencing greater heat fluxes at the outlet. In 2009, Patrick suggested the use of nitrous oxide (N2O) in sub-critical regime as an active coolant for the nozzle component cooling. However, using sub-critical N2O causes more complications by introducing multi phase flow and heat transfer characteristics. The science of supercritical fluids is the only possible solution which consists of homogeneous single phase chemistry. So, in the present study, Supercritical N2O (TC=309.58K, PC=7.254 MPa) is proposed as an active coolant for the cooling of aerospike nozzles. Thermophysical properties of N2O are compared in both sub-critical and supercritical state and also an attempt has been made in order to correlate some of them as a continuous function of temperature and pressure. Statistical parameters are used to ensure the accuracy of developed correlations.

Design and Analysis of Aerospike Nozzle to Improve Thrust in Hybrid Rocket Engine

Design and Analysis of Aerospike Nozzle to Improve Thrust in Hybrid Rocket Engine

Investigation of Variation in the Performance of an Electro Thermal Thruster with Aerospike Nozzle

One of the most recently developed modes of propulsion is electric propulsion. The commonly used chemical propulsion systems have the advantage of a high Specific Impulse as compared to that of ion propulsion systems. However, owing to the efficacy of ion propulsion systems, it is considered the future of space exploration.Electro thermal thrusters produce thrust by using electrical fields to force hot plasma out of the nozzle with certain exit velocity. The plasma’s exit velocity and the system’s thrust capacity, as of now, are insufficient for space travel to be conducted within a reasonable time. I intend to study the possibility of improving the thruster’s performance by using an aerospike nozzle as an exit nozzle which meets the conditions required for the thruster to function appropriately. I shall be studying the plasma plume exit velocity variation with respect to the nozzles used. Also, a thermal analysis will be conducted in order to find the correct material for the nozzle.

Performance Evaluation of Airframe-Integrated Aerospike Propulsion Systems in Off-Design Flight Conditions

The thrust performance of airframe-integrated linear aerospike propulsion systems for high-speed aircraft operated under off-design in-flight conditions was evaluated, based on performance prediction models established for linear aerospike nozzles by accounting for the freestream effect. The airframe configuration, equipped with a boattail appearing to the right upstream of the primary cell nozzle exit, enabled the freestream effect to be used for increasing the spike surface pressure. Performances of each thrust component were predicted by using the physics-based models established in separately performed studies. First, a parametric study was conducted on a rocket primary thruster with various design parameters. Then, an advanced concept with an airbreathing primary propulsion system was evaluated for its feasibility by comparing its thrust performance with that of a conventional rocket primary thruster. The benefit of the aerospike nozzle integrated into the airframe was also evaluated by comparing its thrust to a configuration without the aerospike nozzle. Results showed that the employment of the aerospike nozzle offered improved performance from takeoff to cruise phases. The advanced concept using the bypass airbreathing system for the primary thruster followed by the aerospike configuration exhibited further improvement in thrust by compensating for the boattail drag.

Inviscid Nozzle for Aerospike Rocket Engine Application

A computational method for the steady 2-D flow in axially symmetrical rocket nozzles with a given profile is developed, in order to determine the Maximum thrust contour of rocket engine nozzles with large expansion ratio. The optimized nozzles proved a more than 10% increase in the integral specific impulse recorded during the variable altitude atmospheric flight of rocket vehicles. The method is well suited for application in the design of the optimum contour for axially-symmetric nozzles for atmospheric rocket ascent, specifically for aerospike type nozzles, as for other similar industrial applications in gas and steam turbine technology.

Numerical investigation of flow separation behavior in an over-expanded annular conical aerospike nozzle

A three-part numerical investigation has been conducted in order to identify the flow separation behavior—the progression of the shock structure, the flow separation pattern with an increase in the nozzle pressure ratio (NPR), the prediction of the separation data on the nozzle wall, and the influence of the gas density effect on the flow separation behavior are included. The computational results reveal that the annular conical aerospike nozzle is dominated by shock/shock and shock/boundary layer interactions at all calculated NPRs, and the shock physics and associated flow separation behavior are quite complex. An abnormal flow separation behavior as well as a transition process from no flow separation at highly over-expanded conditions to a restricted shock separation and finally to a free shock separation even at the deign condition can be observed. The complex shock physics has further influence on the separation data on both the spike and cowl walls, and separation criteria suggested by literatures developed from separation data in conical or bell-type rocket nozzles fail at the prediction of flow separation behavior in the present asymmetric supersonic nozzle. Correlation of flow separation with the gas density is distinct for highly over-expanded conditions. Decreasing the gas density or reducing mass flow results in a smaller adverse pressure gradient across the separation shock or a weaker shock system, and this is strongly coupled with the flow separation behavior. The computational results agree well with the experimental data in both shock physics and static wall pressure distribution at the specific NPRs, indicating that the computational methodology here is advisable to accurately predict the flow physics.

小型高輝度短波長光源に向けたパルス発生器の設計と超音速Heガス流による液体金属流の制御法の検討

To demonstrate repetitive-intense X-ray source by using pulsed-power discharge, we proposed an innovative X-pinch method using compact pulse generator and liquid metal injection. The compact pulse generator using small pulse-forming-network modules (PFNMs) has been developed. The current rising rate of the PFNMs with X-pinch system is over 0.1 kA/ns at low operation voltage. To suppress the expansion of liquid metal in a vacuum chamber, we have demonstrated liquid metal injection with coaxial supersonic flow generated by an aerospike nozzle. The results indicate that the expansion of liquid metal is suppressed with the coaxial supersonic flow.

DESIGN AND FLOW SIMULATION OF TRUNCATED AEROSPIKE NOZZLE

Aerospike nozzles are being considered in the development of the Single Stage to Orbit launching vehicles because of their prominent features and altitude compensating characteristics. This paper presents the design of aerospike nozzles using characteristic method in conjunction with streamline function, and performance study through numerical simulation using commercial Computational Fluid Dynamics (CFD) code ANSYS FLUENT. For this purpose nozzles with truncation lengths of 25%, 40%, 50% are choosen, because of the thermal and structural complications in the ideal aerospike nozzle. Simulation of the flow is carried out at three different altitude conditions representing Under-expansion, Ideal, and over-expansion conditions of the flow. FLUENT predictions were used to verify the isentropic flow assumption and that the working fluid reached the design exit Mach number. The flow-fields obtained through the numerical simulation are analysed to know the effect of truncation on the performance of aerospike nozzle. Optimum percentage of the truncation is selected by the comparison of nozzles with different lengths of truncation under various altitude parameters. The results show that the flow pattern of the nozzles under different altitude conditions are almost similar. The 40 % truncated nozzle is found to give optimum performance and it has achieved the desired exit Mach number in all the three altitude conditions.

Aerodynamic Characterization of Linear Aerospike Nozzles in Off-Design Flight Conditions

Spike surface flowfields and base wake behaviors of clustered linear aerospike nozzles under the presence of external flow were characterized using surface pressure measurements, flow visualizations using the background-oriented schlieren method, and theoretical modeling. The test model consisted of three clustered cell nozzle modules or a nonclustered two-dimensional nozzle with an exit Mach number of 3.5, followed by a straight section and a contoured spike. The model was exposed to an external flow of Mach 2.0 to simulate off-design transonic flight conditions. The measured surface pressure distributions indicated that the entire flowfield on the spike surface became more two-dimensional under the presence of external flow. Periodic compressions and expansions of the cell nozzle jet were eliminated, but the pressure continued to increase downstream. The results of the flow visualizations found that this was mainly governed by the external flow that controls the jet expansion by forming a slip surface. This mechanism also made the spike base pressure higher than environmental pressure. The physics-based model predicted the measured surface pressure distribution well and explained the mechanism behind the influence of external flow. The effects of the external flow showed improvements in thrust performance in off-design flight conditions.

Influence of External Flow on Plume Physics of Clustered Linear Aerospike Nozzles

The influence of external flow on the plume physics of cell nozzle flows and the cell base wake behavior of simplified clustered linear aerospike nozzles were characterized using surface pressure measurement, flow visualization using the background oriented schlieren method, and theoretical modeling. The test model comprised three or four clustered cell nozzles, or a nonclustered two-dimensional nozzle with a designed Mach number of 3.5, and was connected to a straight section followed by a straight ramp inclined at 12 deg. In the cases with or without external flow, the measured pressure distributions on the surface of the test model were found to be different in the presence of external flow: The ramp surface pressures were maintained at a certain level above the environmental pressure according to the cell nozzle flow conditions, whereas the pressure distributions in the cases without external flow demonstrated highly periodical compression/expansion features. The results of flow visualization indicated that the pressure maintained on the ramp surface depends mainly on how the external flow controls the jet expansion by forming a slip surface. The cell base wake was found to behave as a closed wake in the presence of external flow. Those pressure behaviors were successfully predicted using the physics-based theoretical models established in this study.

808 数値解析によるエアロスパイクノズルの性能評価 : 内部ノズルの性能評価

808 数値解析によるエアロスパイクノズルの性能評価 : 内部ノズルの性能評価

Regeneratively Cooled Multiple-Use Plug Hybrid for Nanosatellites

The multiple-use plug hybrid for nanosatellites prototype thruster is being developed to fill a niche application for nanosat scale spacecraft propulsion. The prototype thruster uses safe-handling and inexpensive nitrous oxide and acrylonitrile-butadiene-styrene as propellants. The thruster system provides attitude control using secondary-injection thrust vectoring and allows large-impulse and small-impulse attitude control and proximity burns to be performed with the same system. To ensure survivability during extended duration burns, the multiple-use plug hybrid for nanosatellites incorporates a novel regenerative cooling design in which the nitrous oxide oxidizer flows through a cooling path embedded in the aerospike nozzle, before being injected into the combustion chamber near the nozzle base. Digital manufacturing was used to fabricate the nozzle components. Fused deposition modeling was used to fabricate the solid fuel grain from the acrylonitrile-butadiene-styrene thermoplastic. When fully developed, the multiple-use plug hybrid for nanosatellites thruster could provide an enhanced propulsive capability that would enable multiple nanosatellites to be independently repositioned after deployment from the parent launch vehicle. Because the environmentally benign propellants are mixed only within the combustion chamber after ignition, the system is inherently safe and can be piggybacked on a secondary payload with very little overall mission risk increase to the primary payload.

Effect of Spike Truncation on the Acoustic Behavior of Annular Aerospike Nozzles

The work toward reusable single-stage-to-orbit launch-vehicle concepts has initiated several research activities toward a suitable nozzle design, which (unlike bell nozzles) is expected to perform optimally over the entire flight regime. Of these concepts, the annular aerospike nozzle and its variants, linear and truncated, are among the most promising. Given the concern over the intensity of acoustic loading during takeoff, it is vital to carry out a detailed aeroacoustic characterization of aerospike nozzles with special attention to overexpanded conditions, which are generally encountered during liftoff. In this work, an experimental investigation into aeroacoustics of an annular aerospike nozzle of design Mach number () of 2.0 has been carried out, and the results are presented. The effect of truncation of the aerospike to 40 and 60% of the spike length on the acoustic field is also studied and compared with the full-length aerospike. Results show that, in the overexpansion regime, truncated conical aerospike nozzles have reduced acoustic () levels until wake closure compared to the full-length aerospike and conventional convergent–divergent nozzle, all operating at identical conditions. This noise reduction is predominantly in the region upstream of the nozzle exit. At the peak radiation angle ( from the jet upstream direction), the acoustic levels are almost same for all of the nozzles investigated in the present study.

THRUST FORCE ANALYSIS OF SPIKE BELL NOZZLE

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.