Relevant bibliographies by topics / Aerospike Nozzle

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  1. Journal articles
  2. Dissertations / Theses
  3. Books
  4. Book chapters
  5. Conference papers

Academic literature on the topic 'Aerospike Nozzle'

Author: Grafiati
Published: 4 June 2021
Last updated: 1 February 2022

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Journal articles on the topic "Aerospike Nozzle"

1

Lai, A., S. S. Wei, C. H. Lai, J. L. Chen, Y. H. Liao, J. S. Wu, and Y. S. Chen. "Comparison of the Propulsion Performance of Aerospike and Bell-Shaped Nozzle Using Hydrogen Peroxide Monopropellant Under Sea-Level Condition." Journal of Mechanics 35, no. 3 (July 2, 2018): 427–40. http://dx.doi.org/10.1017/jmech.2018.18.

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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.
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Menon, Pranav. "Investigation of Variation in the Performance of an Electro Thermal Thruster with Aerospike Nozzle." Advanced Engineering Forum 16 (April 2016): 91–103. http://dx.doi.org/10.4028/www.scientific.net/aef.16.91.

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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.
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Korte, J. J., A. O. Salas, H. J. Dunn, N. M. Alexandrov, W. W. Follett, G. E. Orient, and A. H. Hadid. "Multidisciplinary Approach to Linear Aerospike Nozzle Design." Journal of Propulsion and Power 17, no. 1 (January 2001): 93–98. http://dx.doi.org/10.2514/2.5712.

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Bogoi, Alina, Radu D. Rugescu, Valentin Ionut Misirliu, Florin Radu Bacaran, and Mihai Predoiu. "Inviscid Nozzle for Aerospike Rocket Engine Application." Applied Mechanics and Materials 811 (November 2015): 152–56. http://dx.doi.org/10.4028/www.scientific.net/amm.811.152.

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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.
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., Vinay Kumar Levaka. "DESIGN AND FLOW SIMULATION OF TRUNCATED AEROSPIKE NOZZLE." International Journal of Research in Engineering and Technology 03, no. 11 (November 25, 2014): 122–31. http://dx.doi.org/10.15623/ijret.2014.0311019.

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Naveen Kumar, K., M. Gopalsamy, Daniel Antony, R. Krishnaraj, and Chaparala B. V. Viswanadh. "Design and Optimization of Aerospike nozzle using CFD." IOP Conference Series: Materials Science and Engineering 247 (October 2017): 012008. http://dx.doi.org/10.1088/1757-899x/247/1/012008.

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Wang, Chang-Hui, Yu Liu, and Li-Zi Qin. "Aerospike nozzle contour design and its performance validation." Acta Astronautica 64, no. 11-12 (June 2009): 1264–75. http://dx.doi.org/10.1016/j.actaastro.2008.01.045.

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Chaudhari, Krunal C. "Analysis of Aerospike Nozzle Structural Contour Design Performance Optimization." International Journal for Research in Applied Science and Engineering Technology V, no. X (October 23, 2017): 1000–1004. http://dx.doi.org/10.22214/ijraset.2017.10144.

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Sankari Ashok Alshiya, K., M. Santhosh, V. K. Santhosh, and S. Sai Gopal. "Experimental analysis of jet flow in an aerospike nozzle." Materials Today: Proceedings 46 (2021): 3444–50. http://dx.doi.org/10.1016/j.matpr.2020.11.783.

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Ferlauto, Michele, Andrea Ferrero, Matteo Marsicovetere, and Roberto Marsilio. "Differential Throttling and Fluidic Thrust Vectoring in a Linear Aerospike." International Journal of Turbomachinery, Propulsion and Power 6, no. 2 (April 21, 2021): 8. http://dx.doi.org/10.3390/ijtpp6020008.

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Aerospike nozzles represent an interesting solution for Single-Stage-To-Orbit or clustered launchers owing to their self-adapting capability, which can lead to better performance compared to classical nozzles. Furthermore, they can provide thrust vectoring in several ways. A simple solution consists of applying differential throttling when multiple combustion chambers are used. An alternative solution is represented by fluidic thrust vectoring, which requires the injection of a secondary flow from a slot. In this work, the flow field in a linear aerospike nozzle was investigated numerically and both differential throttling and fluidic thrust vectoring were studied. The flow field was predicted by solving the Reynolds-averaged Navier–Stokes equations. The thrust vectoring performance was evaluated in terms of side force generation and axial force reduction. The effectiveness of fluidic thrust vectoring was investigated by changing the mass flow rate of secondary flow and injection location. The results show that the response of the system can be non-monotone with respect to the mass flow rate of the secondary injection. In contrast, differential throttling provides a linear behaviour but it can only be applied to configurations with multiple combustion chambers. Finally, the effects of different plug truncation levels are discussed.
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Dissertations / Theses on the topic "Aerospike Nozzle"

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Case, William Scott. "AEROSPIKE THRUST VECTORING SLOT-TYPE COMPOUND NOZZLE." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/316.

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A study of thrust vectoring techniques of annular aerospike nozzles was conducted. Cold-flow blow-down testing along with solid modeling and rapid prototyping technology were used to investigate the effects of slot size, placement, geometry and orientation. The use of slot-type compound nozzles proved to be a feasible approach to thrust vectoring. Previous methods of thrust vectoring have proved to be difficult to implement in a cost effective manner or have had limited effectiveness or durability.
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Beebe, Stanley Ikuo. "HOLE-TYPE AEROSPIKE COMPOUND NOZZLE THRUST VECTORING." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/164.

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Compound aerospike nozzles were designed and tested as part of an ongoing experimental study to determine the feasibility of thrust vectoring an aerospike nozzle with the addition of a secondary port. Earlier phases of the study have indicated that a compound aerospike nozzle could provide sufficient thrust vectoring. The addition of a hole-type secondary port was found to provide effective thrust vectoring. Experiments were carried out to determine the effects of secondary port size, secondary port inlet geometry and compound aerospike nozzle chamber pressure. Results show good predictability, axisymmetric flow, and emphasize the importance of a radius on secondary port inlet geometry.
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Gould, Cedric Oldfield. "DESIGN OF AN AEROSPIKE NOZZLE FOR A HYBRID ROCKET." MSSTATE, 2008. http://sun.library.msstate.edu/ETD-db/theses/available/etd-04292008-102320/.

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This document describes the design of an axisymmetric aerospike nozzle to replace the conical converging-diverging nozzle of a commercially available hybrid rocket motor. The planar method of characteristics is used with isentropic flow assumptions to design the nozzle wall. Axisymmetric adjustments are made with quasi-one-dimensional flow approximations. Computational Fluid Dynamics (CFD) simulations verify these assumptions, and illustrate viscous effects within the flow. Nozzle truncations are also investigated. Development of a hybrid-rocket-specific data acquisition system is also detailed.
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Imbaratto, David Michael. "The Interaction between Throttling and Thrust Vectoring of an Annular Aerospike Nozzle." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/159.

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Applied research and testing has been conducted at the Cal Poly San Luis Obispo High-pressure Blow-Down facility to study the affects of throttling in a thrust-vectored aerospike nozzle. This study supports the ongoing research at Cal Poly to effectively thrust vector a hybrid rocket motor. Such thrust vectoring is achieved by small secondary ports in the nozzle body that are perpendicular to the main nozzle. The testing conducted included characterizing and comparing the performance of a straight aerospike nozzle to that of a thrust-vectored aerospike nozzle. Throttling effects on the aerospike nozzle in an unvectored and in a vectored configuration were also investigated. The interaction between throttling and thrust vectoring of an aerospike nozzle is the focus of this thesis research. This research shows that large-throat/high-thrust operation of an aerospike nozzle provides little thrust vector generation. Conversely, small-throat/low-thrust operation provides ample thrust vector generation. These results have implications in the effectiveness of thrust vectoring an aerospike nozzle with secondary ports. Rockets having an aerospike nozzle with throttling capabilities will be subject to the minimum and maximum turn angles for a given throttle position. As such, certain vehicle maneuvers might not be obtainable at certain throttle operations. Conversely, at lower throttling conditions, higher turn angles will be achievable.
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Zahn, Alexander R. "Characterization and Examination of Performance Parameters of a Back-pressurized RDC." University of Cincinnati / OhioLINK, 2019. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1554119639742205.

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Armstrong, Isaac W. "Development and Testing of Additively Manufactured Aerospike Nozzles for Small Satellite Propulsion." DigitalCommons@USU, 2019. https://digitalcommons.usu.edu/etd/7428.

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Automatic altitude compensation has been a holy grail of rocket propulsion for decades. Current state-of-the-art bell nozzles see large performance decreases at low altitudes, limiting rocket designs, shrinking payloads, and overall increasing costs. Aerospike nozzles are an old idea from the 1960’s that provide superior altitude-compensating performance and enhanced performance in vacuum, but have survivability issues that have stopped their application in satellite propulsion systems. A growing need for CubeSat propulsion systems provides the impetus to study aerospike nozzles in this application. This study built two aerospike nozzles using modern 3D metal printing techniques to test aerospikes at a size small enough to be potentially used on a CubeSat. Results indicated promising in-space performance, but further testing to determine thermal limits is deemed necessary.
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Grieb, Daniel Joseph. "Design and Analysis of a Reusable N2O-Cooled Aerospike Nozzle for Labscale Hybrid Rocket Motor Testing." DigitalCommons@CalPoly, 2012. https://digitalcommons.calpoly.edu/theses/692.

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A reusable oxidizer-cooled annular aerospike nozzle was designed for testing on a labscale PMMA-N20[1] hybrid rocket motor at Cal Poly-SLO.[2] The detailed design was based on the results of previous research involving cold-flow testing of annular aerospike nozzles and hot-flow testing of oxidizer-cooled converging-diverging nozzles. In the design, nitrous oxide is routed to the aerospike through a tube that runs up the middle of the combustion chamber. The solid fuel is arranged in an annular configuration, with a solid cylinder of fuel in the center of the combustion chamber and a hollow cylinder of fuel lining the circumference of the combustion chamber. The center fuel grain insulates the coolant from the heat of the combustion chamber. The two-phase mixture of nitrous oxide then is routed through channels that cool the copper surface of the aerospike. The outer copper shell is brazed to a stainless steel core that provides structural rigidity. The gaseous N2O flows from the end of aerospike to provide base bleed, compensating for the necessary truncation of the spike. Sequential and fully-coupled thermal-mechanical finite element models developed in Abaqus CAE were used to analyze the design of the cooled aerospike. The stress and temperature distributions in the aerospike were predicted for a 10-sec burn time of the hybrid rocket motor. [1] PMMA stands for polymethyl methacrylate, a thermoplastic commonly known by the brand name Plexiglas®. N2O is the molecular formula for nitrous oxide. [2] California Polytechnic State University, San Luis Obispo
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Pearl, Jason M. "Two-Dimensional Numerical Study of Micronozzle Geometry." ScholarWorks @ UVM, 2016. http://scholarworks.uvm.edu/graddis/579.

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Supersonic micronozzles operate in the unique viscosupersonic flow regime, characterized by large Mach numbers (M>1) and low Reynolds numbers (Re<1000). Past research has primarily focused on the design and analysis of converging-diverging de Laval nozzles; however, plug (i.e. centerbody) designs also have some promising characteristics that might make them amenable to microscale operation. In this study, the effects of plug geometry on plug micronozzle performance are examined for the Reynolds number range Re = 80-640 using 2D Navier-Stokes-based simulations. Nozzle plugs are shortened to reduce viscous losses via three techniques: one - truncation, two - the use of parabolic contours, and three - a geometric process involving scaling. Shortened nozzle are derived from a full length geometry designed for optimal isentropic performance. Expansion ratio (ε = 3.19 and 6.22) and shortened plug length (%L = 10-100%) are varied for the full Reynolds number range. The performance of plug nozzles is then compared to that of linear-walled nozzles for equal pressure ratios, Reynolds numbers, and expansion ratios. Linear-walled nozzle half-angle is optimized to to ensure plug nozzles are compared against the best-case linear-walled design. Results indicate that the full length plug nozzle delivers poor performance on the microscale, incurring excessive viscous losses. Plug performance is increased by shortening the nozzle plug, with the scaling technique providing the best performance. The benefit derived from reducing plug length depends upon the Reynolds number, with a 1-2% increase for high Reynolds numbers an up to 14% increase at the lowest Reynolds number examined. In comparison to Linear-walled nozzle, plug nozzles deliver superior performance when under-expanded, however, this trend reverses at low pressure ratios when the nozzles become over-expanded.
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Brennen, Peter Alexander. "SIMULATION OF AN OXIDIZER-COOLED HYBRID ROCKET THROAT: METHODOLOGY VALIDATION FOR DESIGN OF A COOLED AEROSPIKE NOZZLE." DigitalCommons@CalPoly, 2009. https://digitalcommons.calpoly.edu/theses/166.

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A study was undertaken to create a finite element model of a cooled throat converging/diverging rocket nozzle to be used as a tool in designing a cooled aerospike nozzle. Using ABAQUS, a simplified 2D axisymmetric model was created featuring only the copper throat and stainless steel support ring, which were brazed together for the experimental test firings. This analysis was a sequentially coupled thermal/mechanical model. The steady state thermal data matched closely to experimental data. The subsequent mechanical model predicted a life of over 300 cycles using the Manson-Halford fatigue life criteria. A mesh convergence study was performed to establish solution mesh independence. This model was expanded by adding the remainder of the parts of the nozzle aft of the rocket motor so as to attempt to match the transient nature of the experimental data. This model included variable hot gas side coefficients in the nozzle calculated using the Bartz coefficients and mapped onto the surface of the model using a FORTRAN subroutine. Additionally, contact resistances were accounted for between the additional parts. The results from the preliminary run suggested the need for a parameter re-evaluation for cold side gas conditions. Parametric studies were performed on contact resistance and cold side film coefficient. This data led to the final thermal contact conductance of k=0.005 BTU/s•in.•°R for contact between metals, k=0.001 BTU/s•in.•°R for contact between graphite and metal, and h=0.03235 BTU/s2•in.•°R for the cold side film coefficient. The transient curves matched closely and the results were judged acceptable. Finally, a 3D sector model was created using identical parameters as the 2D model except that a variable cold side film condition was added. Instead of modeling a symmetric one or two inlet/one or two outlet cooling channel, this modeled a one inlet/one outlet nozzle in which the coolant traveled almost the full 360° around the cooling annulus. To simplify the initial simulation, the model was cut at the barrier between inlet and outlet to form one large sector, rather than account for thermal gradients across this barrier. This simplified nozzle produced expected data, and a 3D full nozzle model was created. The cold side film coefficients were calculated from previous experimental data using a simplified 2D finite difference approach. The full nozzle model was created in the same manner as the 2D full nozzle model. A mesh convergence study was performed to establish solution mesh independence. The 3D model results matched well to experimental data, and the model was considered a useful tool for the design of an oxidizer cooled aerospike nozzle.
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Papp, John Laszlo. "SIMULATION OF TURBULENT SUPERSONIC SEPARATED BASE FLOWS USING ENHANCED TURBULENCE MODELING TECHNIQUES WITH APPLICATION TO AN X-33 AEROSPIKE ROCKET NOZZLE SYSTEM." University of Cincinnati / OhioLINK, 2000. http://rave.ohiolink.edu/etdc/view?acc_num=ucin962118912.

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Books on the topic "Aerospike Nozzle"

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Ruffin, Stephen M. Computational design aspects of a NASP nozzle/afterbody experiment. Washington, D. C: American Institute of Aeronautics and Astronautics, 1989.

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Wardell, T. C. Final report for the evaluation of a metered mixer for RTV silicone for RSRM nozzle backfill operations. Marshall Space Flight Center, Ala: National Aeronautics and Space Administration, George C. Marshall Space Flight Center, 1989.

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Adamovsky, Grigory. Optical techniques for determination of normal shock position in supersonic flows for aerospace applications. [Washington, D.C.]: NASA, 1990.

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Adamovsky, Grigory. Optical techniques for determination of normal shock position in supersonic flows for aerospace applications. [Washington, D.C.]: NASA, 1990.

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J, Korte J., and Langley Research Center, eds. Multidisciplinary optimization of an aerospike nozzle. Hampton, Va: National Aeronautics and Space Administration, Langley Research Center, 1997.

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Multidisciplinary approach to aerospike nozzle design. Hampton, Va: National Aeronautics and Space Administration, Langley Research Center, 1997.

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Comparison of response surface and kriging models in the multidisciplinary design of an aerospike nozzle. Hampton, VA: Institute for Computer Applications in Science and Engineering, NASA Langley Research Center, 1998.

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A survey of challenges in aerodynamic exhaust nozzle technology for aerospace propulsion applications. Cleveland, Ohio: National Aeronautics and Space Administration, Glenn Research Center, 2002.

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Leonard, Schoenman, and United States. National Aeronautics and Space Administration., eds. Advanced small rocket chambers option 3: 110 1bf Ir-Re rocket. [Washington, DC]: National Aeronautics and Space Administration, 1995.

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Book chapters on the topic "Aerospike Nozzle"

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Shenoy, Aswith R., T. S. Sreekumar, Pranav Menon, and Gerogi Alex. "Computational Analysis of Dual Expander Aerospike Nozzle." In Advances in Mechanical Engineering, 151–58. Singapore: Springer Singapore, 2020. http://dx.doi.org/10.1007/978-981-15-3639-7_18.

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Sequeira, Clavin Wilton, and M. V. Sanjay. "Efficiency Analysis of Aerospike Nozzle by Comparison with a de-Lavel Nozzle Using Computational Fluid Dynamics." In Lecture Notes in Mechanical Engineering, 467–76. Singapore: Springer Singapore, 2021. http://dx.doi.org/10.1007/978-981-16-0159-0_41.

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Aso, S., and K. Sugimoto. "Improvement of aerospike nozzle efficiency due to change of plug base configuration." In Shock Waves, 125–30. Berlin, Heidelberg: Springer Berlin Heidelberg, 2005. http://dx.doi.org/10.1007/978-3-540-27009-6_15.

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Ito, Takashi, and Kozo Fujii. "Numerical Analysis of the Clustered Type Aerospike Nozzle Flow: Flow Structures and Thrust Performance." In Computational Fluid Dynamics 2002, 285–90. Berlin, Heidelberg: Springer Berlin Heidelberg, 2003. http://dx.doi.org/10.1007/978-3-642-59334-5_41.

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Sforza, Pasquale M. "Nozzles." In Theory of Aerospace Propulsion, 161–98. Elsevier, 2012. http://dx.doi.org/10.1016/b978-1-85617-912-6.00005-0.

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Sforza, Pasquale M. "Nozzles for Airbreathing Engines." In Theory of Aerospace Propulsion, 216–68. Elsevier, 2017. http://dx.doi.org/10.1016/b978-0-12-809326-9.00005-1.

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Silvestroni, Laura, and Diletta Sciti. "Effect of Transition Metal Silicides on Microstructure and Mechanical Properties of Ultra-High Temperature Ceramics." In MAX Phases and Ultra-High Temperature Ceramics for Extreme Environments, 125–79. IGI Global, 2013. http://dx.doi.org/10.4018/978-1-4666-4066-5.ch005.

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The IV and V group transition metals borides, carbides, and nitrides are widely known as ultra-high temperature ceramics (UHTCs), owing to their high melting point above 2500°C. These ceramics possess outstanding physical and engineering properties, such as high hardness and strength, low electrical resistivity and good chemical inertness which make them suitable structural materials for applications under high heat fluxes. Potential applications include aerospace manufacturing; for example sharp leading edge parts on hypersonic atmospheric re-entry vehicles, rocket nozzles, and scramjet components, where operating temperatures can exceed 3000°C. The extremely high melting point and the low self-diffusion coefficient make these ceramics very difficult to sinter to full density: temperatures above 2000°C and the application of pressure are necessary conditions. However these processing parameters lead to coarse microstructures, with mean grain size of the order of 20 µm and trapped porosity, all features which prevent the achievement of the full potential of the thermo-mechanical properties of UHTCs. Several activities have been performed in order to decrease the severity of the processing conditions of UHTCs introducing sintering additives, such as metals, nitrides, carbides or silicides. In general the addition of such secondary phases does decrease the sintering temperature, but some additives have some drawbacks, especially during use at high temperature, owing to their softening and the following loss of integrity of the material. In this chapter, composites based on borides and carbides of Zr, Hf and Ta were produced with addition of MoSi2 or TaSi2. These silicides were selected as sintering aids owing to their high melting point (>2100°C), their ductility above 1000°C and their capability to increase the oxidation resistance. The microstructure of fully dense hot pressed UHTCs containing 15 vol% of MoSi2 or TaSi2, was characterized by x-ray diffraction, scanning, and transmission electron microscopy. Based on microstructural features detected by TEM, thermodynamical calculations, and the available phase diagrams, a densification mechanism for these composites is proposed. The mechanical properties, namely hardness, fracture toughness, Young’s modulus and flexural strength at room and high temperature, were measured and compared to the properties of other ultra-high temperature ceramics produced with other sintering additives. Further, the microstructural findings were used to furnish possible explanations for the excellent high temperature performances of these composites.
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Conference papers on the topic "Aerospike Nozzle"

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Kumakawa, A., T. Onodera, M. Yoshida, M. Atsumi, and I. Igarashi. "A study of aerospike-nozzle engines." In 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1998. http://dx.doi.org/10.2514/6.1998-3526.

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Zhao, Qingwei, J. Mo, Alan Chow, and K. He. "Numerical modeling of aerospike nozzle characteristics." In 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2000. http://dx.doi.org/10.2514/6.2000-3849.

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Korte, J., A. Salas, H. Dunn, N. Alexandrov, W. Follett, G. Orient, A. Hadid, et al. "Multidisciplinary approach to linear aerospike nozzle optimization." In 33rd Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1997. http://dx.doi.org/10.2514/6.1997-3374.

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Ferlauto, Michele, Andrea Ferrero, and Roberto Marsilio. "Fluidic Thrust Vectoring for Annular Aerospike Nozzle." In AIAA Propulsion and Energy 2020 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2020. http://dx.doi.org/10.2514/6.2020-3777.

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Wuye, Dai, Liu Yu, Cheng Xianchen, and Tang Haibin. "Aerospike nozzle performance study and its contour optimization." In 37th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2001. http://dx.doi.org/10.2514/6.2001-3237.

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Kumar, K. Naveen, M. Gopalsamy, Daniel Antony, R. Krishnaraj, Chaparala B. V. Viswanadh, and J. Livil Lyle. "Design and optimization of aerospike nozzle using CFD." In 2017 First International Conference on Recent Advances in Aerospace Engineering (ICRAAE). IEEE, 2017. http://dx.doi.org/10.1109/icraae.2017.8297246.

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Ito, Takashi, and Kozo Fujii. "The Effect of Module Clustering on Aerospike Nozzle Performance." In 32nd AIAA Fluid Dynamics Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-3119.

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Hall, Joshua, Carl Hartsfield, Joseph Simmons, and Richard Branam. "Optimized Dual-Expander Aerospike Nozzle Upper Stage Rocket Engine." In 49th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2011. http://dx.doi.org/10.2514/6.2011-419.

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Stewart, Kyle J., Periklis Papadopoulos, and Jordan Pollard. "Nuclear Thermal Rocket Engine with a Toroidal Aerospike Nozzle." In AIAA Propulsion and Energy 2020 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2020. http://dx.doi.org/10.2514/6.2020-3841.

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Reza, Md Saquib, and Konark Arora. "Contour design of aerospike nozzle and comparison of performance." In 2017 International Conference on Infocom Technologies and Unmanned Systems (Trends and Future Directions) (ICTUS). IEEE, 2017. http://dx.doi.org/10.1109/ictus.2017.8286122.

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