Aerospike Nozzle Research Articles

Aerospike nozzle contour design using Angelino’s method and CFD validation

The goal of the present study is to implement an algorithm capable of designing linear aerospike contours and develop a numerical investigation of the altitude adaptive behavior and performance of high enthalpy flow expanded with such contoured nozzle. The contour was designed using Angelino,s rapid method, targeting an exit Mach number of 4.16. An inviscid model and viscous model (employing the realizable k-ϵ\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\epsilon$$\\end{document} for turbulence) were considered and the results compared, showing that turbulent effects must be taken into account to properly simulate important flowfield aspects such as static pressure, static temperature, entropy and viscous losses. The specific impulse is used as the performance criterion and it was observed that it increased with altitude, with only a 2%\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\%$$\\end{document} decrease from an ideal bell nozzle for an underexpanding operation, situation where a nearly axial exhaust plume is obtained. Conversely, when turbulent effects are taken into account, overexpanded conditions have the lowest specific impulse, being lower than a quasi-1D analysis and the inviscid case. The findings summit the advantages of the aerospike nozzle over conventional nozzles. Conventional nozzles not only typically struggle in underexpansion scenarios, but also have high side-loads in overexpansion, due to flow detaching from the nozzle’s wall. A further conclusion is that the realizable k-ϵ\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\epsilon$$\\end{document} turbulent model proved a reliable and fast tool to evaluate turbulent effects that characterize the aerospike flowfield.

Performance Characterization of a Radial Rotating Detonation Combustor with Axial Exhaust

Abstract This experimental study characterizes the performance of a radial rotating detonation combustor (RDC). An aerospike nozzle for rocket propulsion has been integrated into the center of the combustor, although the same combustor could also be coupled with turbomachinery. The radial RDC utilized a rapid to gradual (RTG) area change in the flow direction to effectively confine the detonation region close to the inlet plane and to improve the uniformity of the flow exiting the RDC. Three test cases were analyzed, (a) a baseline case at a total reactant mass flow rate, m˙ = 0.136 kg/s and equivalence ratio, ϕ = 0.6, (b) a higher reactant flow rate, m˙ = 0.318 kg/s and ϕ = 0.6, and (c) a higher ϕ = 0.8 at m˙ = 0.318 kg/s. All tests were conducted using methane and a 67% oxygen and 33% nitrogen (by mole) oxidizer mixture. Measurements were acquired using CTAP probes inside the combustion channel and along the aerospike to characterize the performance, PCB and ion probes near the detonation region to identify wave modes and their variations during the test, and thrust measurements using a six-axis force sensor. Results show highly complex wave modes with multiple co-rotating and/or counter-rotating waves depending upon the reactant flow rate. The pressure and thrust measurements are consistent with the wave mode analysis. In general, a positive (combustor only) pressure gain was inferred when losses associated with the injection system were excluded.

Surface Pressure Measurement of Truncated, Linear Aerospike Nozzles Utilising Secondary Injection for Aerodynamic Thrust Vectoring

A cold-gas test campaign has been conducted at the DLR’s P6.2 test bench in Lampoldshausen, with the objective of investigating the linear aerospike nozzle flow in interaction with secondary injection thrust vector control (SITVC). In this campaign, the influence of nozzle truncation, injection position and injection pressure on the nozzle surface and base pressure is analysed using pressure probes and Schlieren flow-visualisation techniques. The effects of injection position and truncation on the nozzle surface pressure development are comparable for all geometric variations, resulting in a locally increased static pressure upstream and a locally decreased static pressure downstream of the injection. The magnitude and dimension of these high- and low-pressure regions are correlated with the injection pressure. However, the influence of injection position and truncation on the base pressure is not entirely predictable by the named parameters, indicating an interdependence between both geometric parameters. Finally, the required pressure ratio of injection to the primary flow to ensure sonic injection has been analysed on TUD’s cold-gas test bench. This allows the respective injection position-dependent threshold to be identified. The analysis reveals that these experiments have been conducted under transsonic injection conditions.

Design and Evaluation of Aerospike Nozzles for an Upper-Stage Application

Aerospike nozzles are often studied because of their intrinsic self-adaptation behavior to the atmospheric pressure. Another unique feature of aerospike nozzles is their mass- and volume-efficient integration with the underlying vehicle. This latter aspect, if fully exploited, could be of interest for both lower stages operating in the atmosphere and for upper stages operating in vacuum. In this paper, a numerical analysis is carried out of the possible advantages of using an aerospike engine under specific volume constraints of a reference upper-stage configuration, discussing basic aspects of aerospike and internal expansion design, clustering, and plug truncation length. The results of the present preliminary analysis show the extent of possible performance advantages of aerospikes in terms of the specific impulse delivered for a given available volume for the nozzle with respect to a conventional bell configuration.

Low mass-flow operation of small-scale rotating detonation engine

The primary goal of this research was to utilize a range of oxygen/ethylene mixtures in a small-scale Rotating Detonation Engine (RDE) at low mass flow rates to achieve operational detonation frequencies above the 20 kHz threshold of human hearing. A Micro-RDE was designed with an outer diameter of 28 mm and channel gap of 2 mm that yielded detonations at a wide range of testing conditions. Flow was induced via a Jets-in-Crossflow (JIC) scheme to achieve appropriate mixing of reactants before ignition with a pre-detonator device. Two aerospike nozzles of different diameter were tested to change the backpressure on the device and detonation wave speed and frequency were determined using high-speed imagery. Most successful tests yielded two detonation heads. Wave speeds were between 1100 and 1400 m/s and test frequencies were found to be above 31 kHz, well above the audible frequency. A secondary goal of this investigation was to investigate the mixing process within the RDE. Optical access was achieved by implementing a quartz outerbody. Optical imagery of the detonation channel was captured with a Shimadzu camera for detonations ranging from equivalence ratios (Φ) of 0.2–1.7. The fuel and air mixing processes were imaged at MHz rate revealing the local variations in the equivalence ratio that the bulk fluid flow was predominantly axial.

Development and Numerical validation of an Aerospike nozzle Contour Design

Aerospike nozzle simplified design and analysis, as well as test results, are presented. New techniques in nozzle design have been used to improve the performance of current rocket engines. The use of an aerospike nozzle is one such technique. It outperforms the existing proven bell nozzle in terms of overall performance. This study analyses the development and flow analysis of an aerospike nozzle at sea level, validating their results by taking into account currently used bell nozzles. The flow over the aerospike pump is investigated in higher depth using computational fluid dynamics under an autonomous fire testing condition. This test showed that the approximate design method used to determine an aerospike nozzle shape can lead to extremely effective nozzles. A model of the current 7.45 kN bell nozzle was selected as a specimen to develop an aerospike Nozzle. Gambit and Fluent software were used to model and examine the flow characteristics of building the aerospike nozzle. The developed aerospike nozzle shows a nearly 2% increase in performance over the existing nozzle.

Computational Aeroacoustics for a Cold, Non-Ideally Expanded Aerospike Nozzle

Abstract In supersonic aerospace applications, aerospike nozzles have been subject of growing interest. This study sheds light on the noise components of a cold jet exhausting an aerospike nozzle. Implicit large eddy simulations (ILES) are deployed to simulate the jet at a nozzle pressure ratio (NPR)=3. For far-field acoustic computation, the Ffowcs Williams–Hawkings (FWH) equation is applied. A mesh sensitivity study is performed and the jet instantaneous and time-averaged flow characteristics are analyzed. The annular shock structure displays short non-attached shock-cells and longer attached shock-cells. Downstream of the aerospike, a circular shock-cell structure is formed with long shock-cells. Two-point cross-correlations of data acquired at monitoring points located along the shear layers allow to identify upstream propagating waves associated to screech. Power spectral density at monitoring points in the annular shock-cell structure allows to identify its radial oscillation modes. Furthermore, a vortex sheet model is adapted to predict the annular shock-cells length and the BBSAN central frequency. High sound pressure levels (SPL) are detected at the determined BBSAN central frequencies. Finally, high SPL are obtained at the radial oscillation frequencies for the annular shock-cell structure.

Experimental research of the performance and pressure gain in continuous detonation engines with aerospike nozzles

Several continuous detonation engines (CDEs) with aerospike nozzles were designed, and nearly 100 effective experimental tests with methane/GOX were carried out. All tests exhausted to sea level backpressure and were conducted using a calibrated thrust stand and Coriolis mass flow meters. Throughout testing, three key parameters (area ratio, equivalence ratio, and mass flux) affecting CDE operability are discussed in detail. Three normalized parameters were proposed and analyzed to evaluate the performance of CDEs, called the normalized characteristic velocity, the normalized thrust coefficient, and the normalized specific impulse. A new evaluation method for total pressure gain is proposed and implemented. The experimental results showed that lower area ratio (less than 20%) and lower mass flux contribute to form a stable detonation and enjoys a wider range of equivalence ratio. The studied CDEs presented higher performance in the range of area ratio 12.8% - 18.8% and equivalence ratio 1.0 - 1.2. The pressure gain of detonation mode ranged between 15% and 25%, and decreased slightly with the increase of equivalence ratio (0.7 - 1.2). A suitable area ratio is necessary to improve the pressure gain.

PERFORMANCE OF SMALL SIZE AEROSPIKE. ENGINE

In this paper we are showing how aerospike nozzle can play an important role in rocket INDUSTRY as compare to other nozzles. We are using New rocket designs and increasing the performance of the current rockets But, the aerospike nozzle concept that has been under development since the 1950s is yet to be utilized on a launching. Due to its ability to adjust the environment by altering the outer jet boundary, the aerospike nozzle has better performance compared to present day others nozzle. An aerospike nozzle is designed for 20 bar pressure ratio to improve the performance of the aerospike nozzle for various conditions, some important design parameters and their performances were studied for low temperature flow conditions.

System Design and Simulation of Air-spacecraft Aerospike Nozzle by Utilizing the Computational Fluid Dynamics (CFD) Method

System Design and Simulation of Air-spacecraft Aerospike Nozzle by Utilizing the Computational Fluid Dynamics (CFD) Method

Profiling cross-sectional area of a radial rotating detonation combustor to increase pressure gain

The detonation wave(s) in a rotating detonation combustor (RDC) produce highly unsteady flow with varying supersonic and subsonic segments. The non-homogeneous RDC flow must be conditioned to improve its uniformity, which can increase the performance and operation of the downstream hardware for propulsion and power generation applications. In this study, the RDC channel and converging nozzle are combined into a single flow conditioning device, and Gradual to rapid (GTR), Linear, and Rapid to Gradual (RTG) strategies to profile the cross-sectional area of the combined channel are introduced. An inward flowing disk-RDC (D-RDC) operated on gaseous methane-oxygen is numerically investigated using a quasi-two-dimensional computational fluid dynamic (CFD) simulation with a curvilinear coordinate system. The study focuses on the combined channel to isolate the effects of area profiling and neglects radial to axial turning of the flow necessary to integrate it, for example, with an aerospike nozzle located downstream. Simulations show that the RTG area profile yields a higher performance compared to GTR and Linear cases. The RTG profile is also shown to improve pressure and temperature uniformity of the RDC flow at the throat. Results are analyzed to show that a rapid decrease in the cross-sectional area closer to the detonation for the RTG profile improves the stability of the detonation process and constrains the expansion of product gases behind detonation to orient the flow radially inwards. Therefore, profiling the cross-sectional area of an RDC can be vital in improving the operability and effective integration of the downstream hardware.

Experimental research on the performance of hollow and annular rotating detonation engines with nozzles

• The performance of hollow and annular RDEs with nozzles are compared. • The hollow RDE can reduce the injection pressure loss. • The hollow chamber brings a higher performance loss. Rotating detonation engines (RDEs), which are hoped to achieve significant performance improvements in propulsion systems, have received much attention in recent years. The annular RDE and hollow RDE are two of the major configurations of which the structure is very similar to each other. However, there has been a lack of experimental research discussing the differences in their performance. In the present study, experiments on hollow and annular RDEs are conducted. Four configurations in total are tested at ambient conditions: hollow and annular combustion chambers with two exit throat areas. The gaseous methane and gaseous oxygen are used as propellants. The results reveal that the c*-efficiency of the two hollow combustion chamber configurations is on average 0.7 and 0.79, which is lower than the 0.75 and 0.84 of annular chambers. However, it is exhibited that the hollow RDE can ensure high detonation efficiency at a lower injection pressure ratio. From this perspective, the hollow RDE can be considered a more promising configuration for practical application because it provides a new approach to reducing injection pressure loss. In addition, the analysis of nozzle indicates that the Laval nozzles overperform the aerospike nozzles, which goes against previous studies. The mechanism that accounts for this phenomenon needs further investigation.

Designing and Fabricating a Vortex Aerospike Rocket Engine

On average, 51% of a rocket engine’s weight is in the cooling system, so finding anengine design that would drastically reduce the engine's weight, complexity, and production costwould be extremely beneficial. Small one-stage rockets will be the main method for smallerorganizations to reach orbit shortly. The engine must be efficient at sea level and in space toachieve this. Since the start of spaceflight in the 1960s, many innovations on the typical rocketengine have been theorized, but few have been adopted. This project takes two innovationspreviously theorized and tested, vortex cooling and aerospike nozzle, and combines them in aproof of concept prototype. First, the prototype was designed in Computer-Aided-Design andthen validated in a Compusonal-Fluid-Dynamic simulation. Next, the design was fabricated outof mostly hardware store materials. The custom nozzle design was made throughlost-PLA-casting. Many tests were conducted with this prototype, and 4 data points werecollected: nozzle temperature, combustion chamber temperature, chamber pressure, and thrust.The data collected showed the success of the vortex cooling method, as the chamber temperaturewas much cooler than the nozzle temperature. The engine's success as a whole was validated bythe thrust data. This project shows that a vortex aerospike design is a viable and useful idea. Afunctioning rocket engine can be created relatively inexpensively and with widely available partsand proves that departing from typical rocket engine design can lead to beneficial findings.

Computational investigation of the effect of variable cowl length on the flowfield of conical aerospike nozzles

Computational investigation of the effect of variable cowl length on the flowfield of conical aerospike nozzles

Study on Control of Unsteadiness and Flow Asymmetry in Linear Aerospike Nozzles

The unsteadiness in flow over a linear aerospike nozzle is analyzed by a two-dimensional numerical simulation. Variations in the separation locations on two sides of the full-length plug nozzle, with the increase in nozzle pressure ratios (NPRs), are compared to study the asymmetry. For the linear aerospike nozzle considered, the flow becomes unsteady after . The amplitude of the pressure fluctuation over time is compared at different locations on the surface of the nozzle, starting from the separation point to a few locations downstream. The amplitude of the pressure fluctuation increases from the separation location, reaches a peak value, and then starts decreasing. Interaction between the trapped recirculation bubble and the overexpansion shock causes the asymmetry. An optimum truncation at is able to arrest the separation shock foot at the tip of the truncated base, thereby eliminating the pressure fluctuation over the plug surface. Interaction of the recirculation zones behind the truncated base causes a minor fluctuation from until . Nevertheless, the fluctuations are negligible when compared to the fluctuation over the full-length plug surface. The fluctuations die out at when the truncated plug is nearing closed-wake operation.

Experimental study on wave propagations in a rotating detonation chamber with different outlet configurations

To improve the propulsive performance, exhaust nozzles are desired for rotating detonation engines in practical applications. Therefore, their effects on the upstream behaviors inside the combustion chamber are required to be clarified. This study focuses on the characteristics of rotating detonation waves influenced by different outlet configurations. Three types of aerospike nozzle configurations are developed with different throat areas. Ethylene and oxygen-enriched air with an oxygen volume fraction of 50% are utilized in this study. When nozzles are employed, five wave propagation modes (i.e., the single-wave mode, the counter-rotating wave mode, the self-adjustive mode, the periodic oscillation mode, and the deflagration mode) are observed with the equivalence ratio ranging from 0.33 to 1.39, while only single-wave mode appears in the nozzle-less case under the same operating conditions. When increasing the equivalence ratio or reducing the nozzle throat area, those unstable propagation modes are more likely to happen and three boundaries of the equivalence ratio for the wave propagation transformation are clarified. On the contrary, the standard deviations of propagation velocities in the stable propagation process are smaller when employing nozzle configurations compared to the nozzle-less cases, which indicates a more stable state of the wave propagation although the stable propagation range is narrowed. Besides, the pressure rise and specific thrust are analyzed under different outlet configurations.

Study on CerAMfacturing of Novel Alumina Aerospike Nozzles by Lithography-Based Ceramic Vat Photopolymerization (CerAM VPP).

Advanced ceramics are recognized as key enabling materials possessing combinations of properties not achievable in other material classes. They provide very high thermal, chemical and mechanical resistance and typically exhibit lower densities than metals. These properties predestine ceramics for many different applications, especially those in space. Aerospike nozzles promise an increased performance compared to classic bell nozzles but are also inherently more complex to manufacture due to their shape. Additive manufacturing (AM) drastically simplifies or even enables the fabrication of very complex structures while minimizing the number of individual parts. The applicability of ceramic AM (“CerAMfacturing”) on rocket engines and especially nozzles is consequently investigated in the frame of the “MACARONIS” project, a cooperation of the Institute of Aerospace Engineering at Technische Universität Dresden and the Fraunhofer Institute for Ceramic Technologies and Systems (IKTS) in Dresden. The goal is to develop novel filigree aerospike nozzles with 2.5 N and 10 N thrust. For this purpose, CerAM VPP (ceramic AM via Vat Photopolymerization) using photoreactive and highly particle-filled suspensions was utilized. This contribution gives an overview of the component development starting from CAD modeling, suspension development based on alumina AES-11C, heat treatment and investigation of the microstructure of the sintered components. It could be shown that modifying the suspension composition significantly reduced the formation of cracks during processing, resulting in defect-free filigree aerospike nozzles for application in space.

Design and optimization of aerospike nozzle for rotating detonation engine

Rotating detonation engine (RDE) is increasingly viewed as a viable technology to improve the performance of propulsion devices. At now, the design of RDE faces challenges when integrated with aerospike nozzle due to the large pressure fluctuations of rotating detonation waves. To optimize the nozzle configuration for thrust generation, the current work investigates the performance of RDE combined with various types of aerospike nozzles. The method of designing optimal aerospike nozzle for RDE with annulus combustion chamber is proposed. A series of three-dimensional numerical simulations are conducted to obtain the thrust performance of nozzles based on open-source software OpenFOAM. The flow-field structure, pressure gain in combustor and thrust performance of RDE are investigated. The results suggest that although the instantaneous flow-field in RDE is time-varying, the averaged plume field can be viewed as axisymmetric under-expanded supersonic jets. It's reasonable to use the time-averaged value of physical quantities for nozzle design. For nozzle with inward-bend cowl lip and isentropic ramp surface, the supersonic flow in nozzle undergoes approximate isentropic expansion with minimum loss. The flow at nozzle exit is basically parallel to chamber axis, generating the optimal thrust. This indicates that the aerospike nozzle is a good match for efficient expansion of rotating detonation products for propulsion. Moreover, the stagnation pressure gain in combustor has been greatly improved with the constriction nozzle throat. This suggests that in order to elevate the chamber pressure, the convergence section with nozzle throat should be adopted in RDE just as in traditional rocket engines.

RDE research and development in Poland

A very short survey of research conducted in Poland on the development of the rotating detonation engine (RDE) is presented. Initial studies conducted in cooperation with Japanese partners lead to development of a joint patent on RDE. Then, an intensive basic and applied research was started at the Institute of Heat Engineering of the Warsaw University of Technology. One of the first achievements was the demonstration of performance of the rocket engine with an aerospike nozzle utilizing continuously rotating detonation (CRD), and research was directed into development of a small turbofan engine utilizing such a combustion regime. These activities promoted international cooperation and stimulated RDE development not only in Poland but also in other countries. A research directed to measure and calculate flow parameters as well as to analyze the use of liquid fuels was conducted. In the Institute of Aviation in Warsaw, research on the application of the CRD to turbine engines as well as rocket, ramjet, and combined cycle engines was carried out. In the paper, a special emphasis is given to international cooperation in this area with partners from many countries engaged in the development of the pressure gain combustion to propulsion systems.

Numerical Analysis of Flowfield in Linear Plug Nozzle with Base Bleed

A linear aerospike nozzle is a kind of altitude adaptive nozzle that has optimum performance over the entire operational range. In this work, flow characteristics of a linear plug nozzle and a truncated plug with 40% plug length are studied numerically. Different flow types associated with an increase in nozzle pressure ratio, corresponding wall pressure variation and shock patterns are studied. The effect of base pressure improvement with a secondary flow from the base (base bleed) of the truncated nozzle is also examined. Four base bleed points are considered, and the optimum bleed point is selected based on performance and base pressure improvement. The two-dimensional numerical model is solved using Reynolds-averaged Navier–Stokes equations with the Transition SST turbulence model. Validation shows that the present model is good enough at predicting the flow features associated with a linear plug nozzle. From the three different flow types associated with this nozzle, the transition from one type to other happens at lower nozzle pressure ratios with truncation due to the reduction in length of the plug. Thrust coefficient increased by 3.54% in the case of the bleed 3 configuration with a secondary flow of 2% of inlet mass flow rate when compared with 40% plug length without base bleed.