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Showing papers on "Rocket published in 2017"


Journal ArticleDOI
01 Jan 2017
TL;DR: In this paper, the authors present recent progress in the field of thermoacoustic combustion instabilities in propulsion engines such as rockets or gas turbines, and show that LES is not sufficient and that theory, even in these complex systems, plays a major role to understand both experimental and LES results and to identify mitigation techniques.
Abstract: This paper presents recent progress in the field of thermoacoustic combustion instabilities in propulsion engines such as rockets or gas turbines. Combustion instabilities have been studied for more than a century in simple laminar configurations as well as in laboratory-scale turbulent flames. These instabilities are also encountered in real engines but new mechanisms appear in these systems because of obvious differences with academic burners: larger Reynolds numbers, higher pressures and power densities, multiple inlet systems, complex fuels. Other differences are more subtle: real engines often feature specific unstable modes such as azimuthal instabilities in gas turbines or transverse modes in rocket chambers. Hydrodynamic instability modes can also differ as well as the combustion regimes, which can require very different simulation models. The integration of chambers in real engines implies that compressor and turbine impedances control instabilities directly so that the determination of the impedances of turbomachinery elements becomes a key issue. Gathering experimental data on combustion instabilities is difficult in real engines and Large Eddy Simulation (LES) has become a major tool in this field. Recent examples, however, show that LES is not sufficient and that theory, even in these complex systems, plays a major role to understand both experimental and LES results and to identify mitigation techniques.

445 citations


Book
22 Jun 2017
TL;DR: Aircraft Propulsion and Gas Turbine Engines, Second Edition as discussed by the authors is the most widely used book on turbine engines, fuels, and combustion, with the addition of three major topic areas: Piston engines with integrated propeller coverage; Pump Technologies; and Rocket Propulsion.
Abstract: Aircraft Propulsion and Gas Turbine Engines, Second Edition builds upon the success of the book’s first edition, with the addition of three major topic areas: Piston Engines with integrated propeller coverage; Pump Technologies; and Rocket Propulsion. The rocket propulsion section extends the text’s coverage so that both Aerospace and Aeronautical topics can be studied and compared. Numerous updates have been made to reflect the latest advances in turbine engines, fuels, and combustion. The text is now divided into three parts, the first two devoted to air breathing engines, and the third covering non-air breathing or rocket engines.

194 citations


Journal ArticleDOI
TL;DR: In this paper, a UHTC-based thermal protection system for re-entry vehicles or components for space propulsion has been proposed for high-enthalpy environments, where a hypersonic arc-jet facility allows performing tests in simulated atmospheric reentry conditions.
Abstract: Ultra-High-Temperature Ceramic (UHTC) materials, because of their high temperature resistance, are suitable as thermal protection systems for re-entry vehicles or components for space propulsion. Massive UHTC materials are characterized by poor thermal shock resistance, which may be overcome using C or SiC fibers in a UHTC matrix (UHTCMC). The University of Naples “Federico II” has a proven experience in the field of material characterization in high-enthalpy environments. A hypersonic arc-jet facility allows performing tests in simulated atmospheric re-entry conditions. The Aerospace Propulsion Laboratory is employed for testing rocket components in a representative combustion environment. Ad-hoc computational models are developed to characterize the flow field in both facilities and perform thermal analysis of solid samples. Current research programs are related to a new-class of UHTCMC materials, for rocket nozzles and thermal protection systems. The activities include design of the prototypes for the test campaign, numerical simulations and materials characterizations.

100 citations


Journal ArticleDOI
TL;DR: Hniličková et al. as discussed by the authors monitored the effect of salt stress induced by the NaCl solution (0, deionized water; 50, 100, 200, 300 mmol/L) in rocket (Eruca sativa (L.) mill.) cv. Astro over the course of 50 days.
Abstract: Hniličková H., Hnilička F., Martinková J., Kraus K. (2017): Effects of salt stress on water status, photosynthesis and chlorophyll fluorescence of rocket. Plant Soil Environ., 63: 362–367. Salinity is a significant environmental factor affecting physiological processes in plants. This study monitors the effect of salt stress induced by the NaCl solution (0 – deionized water; 50, 100, 200, 300 mmol/L) in rocket (Eruca sativa (L.) Mill.) cv. Astro over the course of 50 days. Salt stress significantly affected the monitored parameters. The osmotic potential decreased with increasing NaCl concentrations, while relative water content decrease did not take place until 200 mmol/L NaCl. Compared to the control group, transpiration (E) decreased at the concentration of 50 mmol/L NaCl and stomatal conductance (gs) and net photosynthetic rate (Pn) decreased at 100 mmol/L NaCl. Further increase of salt concentrations did not affect Pn and no significant differences gs, E and substomatal concentration CO2 were measured between the concentrations of 200 and 300 mmol/L NaCl. A decrease of Fv/Fm took place from the concentration of 100 mmol/L NaCl, while differences between 200 and 300 mmol/L NaCl were also not significant. The obtained results therefore prove the tolerance of the E. sativa cv. Astro to salt stress.

70 citations


Journal ArticleDOI
TL;DR: In this article, the authors showed that the maximum velocity of a photon propulsion rocket is a function of the reduced Compton wavelength of the heavy subatomic particles in the rocket, which is the same as the maximum velocities of a Planck mass particle.

42 citations


Journal ArticleDOI
TL;DR: In this paper, a pintle injector was used for developing throttleable rocket engines, where variable-area injectors are suitable choices for developing a rocket engine because it is difficult to efficiently control thrust when fixed-area injectionors are used.
Abstract: Variable-area injectors are suitable choices for developing throttleable rocket engines because it is difficult to efficiently control thrust when fixed-area injectors are used. A pintle injector i...

41 citations


Journal ArticleDOI
TL;DR: In this paper, an axisymmetric, multispecies compressible flow solver was used to solve the longitudinal combustion instability in liquid-propellant rocket engines.
Abstract: Longitudinal combustion instability in liquid-propellant rocket engines is investigated using an in-house axisymmetric, multispecies compressible flow solver. Turbulence is treated using a hybrid R...

40 citations


Journal ArticleDOI
TL;DR: HEROS 3 was launched from the European Space and Sounding Rocket Range (ESRrange) Space Center to an apogee altitude of 32 km at 1030hrs on 8 November 2016 as discussed by the authors.
Abstract: On 8 November 2016 at 1030 hrs, the Hybrid Experimental Rocket Stuttgart (HEROS) 3 was launched from the European Space and Sounding Rocket Range (ESRANGE) Space Center to an apogee altitude of 32,...

38 citations


Journal ArticleDOI
TL;DR: In this paper, a low-order analytical performance model for the rocket-mode rotating detonation engine was developed, starting from the general form of the continuity, momentum, and energy equations and integrati...
Abstract: A low-order analytical performance model for the rocket-mode rotating detonation engine was developed. Starting from the general form of the continuity, momentum, and energy equations and integrati...

37 citations


Journal ArticleDOI
TL;DR: In this paper, a combustion model of boron particles for detailed Computational Fluid Dynamics (CFD) based simulations of ducted rocket combustion chambers is studied, which includes all main physical processes required to define an accurate particle combustion simulation.

36 citations


Journal ArticleDOI
TL;DR: In this article, a 3D fluid-structural coupling computational methodology is developed to predict the thermal and structural responses of the thrust chamber wall under cyclic work, and the results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber.

Journal ArticleDOI
TL;DR: In this article, the authors reported the first observation of concentric traveling ionospheric disturbances (CTIDs) triggered by the launch of a SpaceX Falcon 9 rocket on 17 January 2016, which showed a shock acoustic wave signature in the time rate change (time derivative) of total electron content (TEC), followed by CTIDs in the 8-15 minutes bandpass filtering of TEC.
Abstract: We report the first observation of concentric traveling ionospheric disturbances (CTIDs) triggered by the launch of a SpaceX Falcon 9 rocket on 17 January 2016. The rocket triggered ionospheric disturbances show shock acoustic wave signature in the time rate change (time derivative) of total electron content (TEC), followed by CTIDs in the 8–15 minutes bandpass filtering of TEC. The CTIDs propagated northward with phase velocity of 241–617 m/s and reached distances more than 1000 km away from the source on the rocket trajectory. The wave characteristics of CTIDs with periods of 10.5–12.7 minutes and wavelength ~200–400 km agree well with the gravity wave dispersion relation. The optimal wave source searching and gravity wave ray-tracing technique suggested that the CTIDs have multiple sources which are originated from ~38–120 km altitude before and after the ignition of the 2nd stage rocket, ~200 seconds after the rocket was launched.

Journal ArticleDOI
TL;DR: In this paper, the acceleration autopilot with a rate loop is the most commonly implemented autopilot, which has been extensively applied to high-performance missiles, and the design of the guidance and control modules is a challenging task because the rapid spinning of the body creates a heavy coupling between the normal and lateral rocket dynamics.

Journal ArticleDOI
TL;DR: In this paper, the authors examined ablative material behavior in a laboratory-scale solid rocket motor with a planar, two-dimensional flow channel in which flat ablative materials samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography.
Abstract: Ablative insulators are used in the interior surfaces of solid rocket motors to prevent the mechanical structure of the rocket from failing due to intense heating by the high-temperature solid-propellant combustion products. The complexity of the ablation process underscores the need for ablative material response data procured from a realistic solid rocket motor environment, where all of the potential contributions to material degradation are present and in their appropriate proportions. For this purpose, the present study examines ablative material behavior in a laboratory-scale solid rocket motor. The test apparatus includes a planar, two-dimensional flow channel in which flat ablative material samples are installed downstream of an aluminized solid propellant grain and imaged via real-time X-ray radiography. In this way, the in-situ transient thermal response of an ablator to all of the thermal, chemical, and mechanical erosion mechanisms present in a solid rocket environment can be observed and recorded. The ablative material is instrumented with multiple micro-thermocouples, so that in-depth temperature histories are known. Both total heat flux and thermal radiation flux gauges have been designed, fabricated, and tested to characterize the thermal environment to which the ablative material samples are exposed. These tests not only allow different ablative materials to be compared in a realistic solid rocket motor environment but also improve the understanding of the mechanisms that influence the erosion behavior of a given ablative material.

Journal ArticleDOI
TL;DR: In this article, an infrared signature analysis tool (IRSAT) was developed to understand the spectral characteristics of exhaust plumes in detail through a finite volume technique, flow field properties were obtained through the solution of axisymmetric Navier-Stokes equations with the Reynolds-averaged approach.

Journal ArticleDOI
TL;DR: A promising technology to solve both problems is swirling oxidizer injection, which enhances the wall hea... as discussed by the authors, which is used in hybrid rockets to solve the problem of low regression rate and combustion inefficiencies.
Abstract: Hybrid rockets present some disadvantages, mainly low regression rate and combustion inefficiencies. A promising technology to solve both is swirling oxidizer injection, which enhances the wall hea...

Journal ArticleDOI
TL;DR: In this article, a comparison of the numerical predictions of several groups modeling the reacting flow inside a gaseous methane/gaseous oxygen single-element rocket combustion chamber is conducted.
Abstract: A comparison of the numerical predictions of several groups modeling the reacting flow inside a gaseous methane/gaseous oxygen single-element rocket combustion chamber is conducted. The focus is pl...

Journal ArticleDOI
TL;DR: In this article, a planar pintle injector was used to investigate the effects of total momentum ratio and O/F on the combustion characteristics of the planar injector.
Abstract: The pintle injector is a promising candidate for propellant-injection systems for various rocket engines. However, fundamental studies focusing on the pintle injector are rather limited, and the effects of key design parameters on combustion behaviors of the injector are still not identified. Therefore, combustion tests of an ethanol/liquid-oxygen rocket-engine combustor with a planar pintle injector are conducted to investigate the effects of total momentum ratio and O/F on the combustion characteristics of the pintle injector. The total momentum ratio and O/F vary from 1.0 to 2.2 and from 0.32 to 3.65, respectively, whereas the combustion pressure is kept constant at approximately 0.4 MPa. The spray structures are observed with high-speed cameras. The characteristic exhaust velocity (C*) efficiency increases with the increase in O/F in the fuel-centered configuration, and the opposite tendency is observed in the oxidizer-centered configuration. The C* efficiency decreases with the increase in the total ...

Journal ArticleDOI
Yue Li1, Chunbo Hu1, Zhe Deng1, Chao Li1, Haijun Sun1, Yupeng Cai1 
TL;DR: In this article, the performance characteristics of ammonium perchlorate/aluminum powder rocket motor were investigated experimentally based on a powder rocket testing system and the multiple-pulse performance characteristics were analyzed accordingly.

Journal ArticleDOI
TL;DR: In this paper, the authors modeled the emissions associated with a hydrogen fueled reusable rocket system based on the launch requirements of developing a space-based solar power system that generates present-day global electric energy demand.
Abstract: Modern reusable launch vehicle technology may allow high flight rate space transportation at low cost. Emissions associated with a hydrogen fueled reusable rocket system are modeled based on the launch requirements of developing a space-based solar power system that generates present-day global electric energy demand. Flight rates from 104 to 106 per year are simulated and sustained to a quasisteady state. For the assumed rocket engine, H2O and NOX are the primary emission products; this also includes NOX produced during reentry heating. For a base case of 105 flights per year, global stratospheric and mesospheric water vapor increase by approximately 10 and 100%, respectively. As a result, high-latitude cloudiness increases in the lower stratosphere and near the mesopause by as much as 20%. Increased water vapor also results in global effective radiative forcing of about 0.03 W/m2. NOX produced during reentry exceeds meteoritic production by more than an order of magnitude, and along with in situ stratospheric emissions, results in a 0.5% loss of the globally averaged ozone column, with column losses in the polar regions exceeding 2%.

Journal ArticleDOI
TL;DR: In this paper, a set of simplified chemical kinetics mechanisms for hybrid rocket applications using gaseous oxygen (GOX) and hydroxyl-terminated polybutadiene (HTPB) is proposed.

Journal ArticleDOI
TL;DR: In this paper, the formation of high-velocity compact elements of shaped charges with a liner of a combined hemisphere-cylinder shape has been analyzed by numerical simulations of a two-dimensional axisymmetric problem of continuum mechanics.


Journal ArticleDOI
TL;DR: In this paper, a fluid-solid coupling numerically method is established based on the conserved form of the three-dimensional unsteady Navier-Stokes (N-S) equations, considering gas fluid with chemical reactions and heat transfer between the fluid and solid region.

Journal ArticleDOI
TL;DR: In this article, a review of specific features of experimental development of liquid rocket engines (LRE) is presented, including the 11D56, 11D57, RD0120, KVD1, and a number of propulsion units and power plants.

Journal ArticleDOI
TL;DR: P pitch angle control parameter is optimized using improved differential evolution algorithm (IDEA) and the performance of pitch angle of rocket system is enhanced by using PID controller based on IDEA.
Abstract: Pitch angle of rocket system is the important parts of the rocket. This part corresponds to the movement of the rocket system. Rocket system is fell into multi-input and multi-output (MIMO) system. The most challenge factor in MIMO system is designing the controller, if the design is not appropriate, it may lead to the unstable condition. Hence, appropriate and robust control design is inevitable. This paper introduces PID controller as pitch angle control of rocket system. Furthermore, PID controller parameter is optimized using improved differential evolution algorithm (IDEA). To analyze the performance of rocket system, time domain simulation is implemented. From the simulation result, it is found by using PID controller based on IDEA, the performance of pitch angle of rocket system is enhanced.


Journal ArticleDOI
TL;DR: In this article, a solid-fuel rocket scramjet with metalized solid fuel as a propellant was tested experimentally, and preliminary evaluation results showed that the combustion efficiency of the propellant is about 90% and the total pressure recovery coefficient in the supersonic combustor is about 0.6.
Abstract: Liquid or gaseous fuel scramjet technology has made great progress, and there has been some research attention to solid-fuel scramjet. A new scramjet configuration using solid fuel as propellant, namely solid-fuel rocket scramjet, is tested experimentally. It consists of two combustors. One is a rocket combustor used as gas generator, and the other is a supersonic combustor used for secondary combustion. The experiment simulates a flight Mach number of 4 at high altitude (stagnation temperature and pressure are 1170 K and 1.16 MPa, respectively), and metalized solid fuel is used as propellant. The results reveal that fuel-rich gas from the gas generator can burn with air in the supersonic combustor. Preliminary evaluation results show that the combustion efficiency of the propellant is about 90%, and the total pressure recovery coefficient in the supersonic combustor is about 0.6. These results indicate that the configuration of solid-fuel rocket scramjet is feasible.

Journal ArticleDOI
TL;DR: In this paper, a two-dimensional ablation analysis code (MOPAR-MD) capable of modeling pyrolyzing thermal protection system materials is presented, which can be coupled to the LeMANS reacting flow solver.
Abstract: A new two-dimensional ablation analysis code (MOPAR-MD) capable of modeling pyrolyzing thermal protection system materials is presented. Favorable agreement with analytical solutions and results from other (one-dimensional) ablation solvers for a wide range of test cases indicates a correct implementation consistent with other codes. This new material response code can be coupled to the LeMANS reacting flow solver. New capabilities required for modeling nozzle flowfields are added to LeMANS, including the Menter baseline and shear stress transport turbulence models and a “two-gas” method for capturing the thermodynamics of gas–particle flow found in many rocket nozzles. These updated codes are used to perform uncoupled simulations, predicting the thermal and ablation response of the HIPPO nozzle test case. Radiation is found to have minimal impact on the response of the throat and downstream portions of the rocket nozzle, but remains significant for the motor chamber and upstream portions of the nozzle. E...

Journal ArticleDOI
01 Jun 2017-Fuel
TL;DR: In this article, the behavior of ram-scram combustion mode transition was examined using direct-connect experiments of a model RBCC combustor along with pressure measurements, and four combustion modes were classified through wall pressure distributions, combustor inlet Mach number, and flame plume visualization.