Showing papers on "Rocket published in 2021"

MiniRocket: A Very Fast (Almost) Deterministic Transform for Time Series Classification

Abstract: Rocket achieves state-of-the-art accuracy for time series classification with a fraction of the computational expense of most existing methods by transforming input time series using random convolutional kernels, and using the transformed features to train a linear classifier. We reformulate Rocket into a new method, MiniRocket. MiniRocket is up to 75 times faster than Rocket on larger datasets, and almost deterministic (and optionally, fully deterministic), while maintaining essentially the same accuracy. Using this method, it is possible to train and test a classifier on all of 109 datasets from the UCR archive to state-of-the-art accuracy in under 10 minutes. MiniRocket is significantly faster than any other method of comparable accuracy (including Rocket), and significantly more accurate than any other method of remotely similar computational expense.

Numerical simulation of a methane-oxygen rotating detonation rocket engine

Abstract: The rotating detonation engine (RDE) is an important realization of pressure gain combustion for rocket applications. The RDE system is characterized by a highly unsteady flow field, with multiple reflected pressure waves following detonation and an entrainment of partially-burnt gases in the post-detonation region. While experimental efforts have provided macroscopic properties of RDE operation, limited accessibility for optical and flow-field diagnostic equipment constrain the understanding of mechanisms that lend to wave stability, controllability, and sustainability. To this end, high-fidelity numerical simulations of a methane-oxygen rotating detonation rocket engine (RDRE) with an impinging discrete injection scheme are performed to provide detailed insight into the detonation and mixing physics and anomalous behavior within the system. Two primary detonation waves reside at a standoff distance from the base of the channel, with peak detonation heat release at approximately 10 mm from the injection plane. The high plenum pressures and micro-nozzle injector geometry contribute to fairly stiff injectors that are minimally affected by the passing detonation wave. There is no large scale circulation observed in the reactant mixing region, and the fuel distribution is asymmetric with a rich mixture attached to the inner wall of the annulus. The detonation waves’ strengths spatially fluctuate, with large variations in local wave speed and flow compression. The flow field is characterized by parasitic combustion of the fresh reactant mixture as well as post-detonation deflagration of residual gases. By the exit plane of the RDRE, approximately 95.7% of the fuel has been consumed. In this work, a detailed statistical analysis of the interaction between mixing and detonation is presented. The results highlight the merit of high-fidelity numerical studies in investigating an RDRE system and the outcomes may be used to improve its performance.

Time-dependent inertia of self-propelled particles: The Langevin rocket.

Abstract: Many self-propelled objects are large enough to exhibit inertial effects but still suffer from environmental fluctuations. The corresponding basic equations of motion are governed by active Langevin dynamics, which involve inertia, friction, and stochastic noise for both the translational and orientational degrees of freedom coupled via the self-propulsion along the particle orientation. In this paper, we generalize the active Langevin model to time-dependent parameters and explicitly discuss the effect of time-dependent inertia for achiral and chiral particles. Realizations of this situation are manifold, ranging from minirockets (which are self-propelled by burning their own mass), to dust particles in plasma (which lose mass by evaporating material), to walkers with expiring activity. Here we present analytical solutions for several dynamical correlation functions, such as mean-square displacement and orientational and velocity autocorrelation functions. If the parameters exhibit a slow power law in time, we obtain anomalous superdiffusion with a nontrivial dynamical exponent. Finally, we constitute the ``Langevin rocket'' model by including orientational fluctuations in the traditional Tsiolkovsky rocket equation. We calculate the mean reach of the Langevin rocket and discuss different mass ejection strategies to maximize it. Our results can be tested in experiments on macroscopic robotic or living particles or in self-propelled mesoscopic objects moving in media of low viscosity, such as complex plasma.

Pseudo-boiling and heat transfer deterioration while heating supercritical liquid rocket engine propellants

Abstract: Heating of liquid propellants used as the coolant in rocket engines may lead to undesired phenomena such as pseudo-boiling or heat transfer deterioration under specific conditions. This can be an issue for propellants characterized at the same time by relatively low critical pressure and temperature. Light hydrocarbons, as for instance methane, belong to this family. In the present paper, a critical review is made of the main results obtained by Authors and their coworkers for the present application. Focus is on the correlations and trends inferred by their numerical simulations mainly carried out considering methane as the coolant, perhaps the most challenging one.

Onboard Guidance for Reusable Rockets: Aerodynamic Descent and Powered Landing

Abstract: This paper describes a novel general on-board guidance strategy which can be applied toboth the aerodynamically-controlled descent and the powered landing phase of reusable rockets.The proposed guidance method is based on sequential convex optimization applied to a Cartesianrepresentation of the equations of motion. The contributions are an exploitation of convexand non-convex contributions, which are processed separately to maximize the computationalefficiency of the approach, the inclusion of highly nonlinear terms represented by aerodynamicaccelerations, a complete reformulation of the problem based on the use of Euler angle rates ascontrol means, an improved transcription based on the use of a generalized hp pseudospectralmethod, and a dedicated formulation of the aerodynamic guidance problem for reusable rockets.The problem is solved for a 40 kN-class reusable rocket. Results show that the proposedtechnique is a very effective methodology able to satisfy all the constraints acting on the system,and can be potentially employed online to solve the entire descent phase of reusable rockets inreal-time.

Computational and Experimental Study of Nozzle Performance for Rotating Detonation Rocket Engines

Abstract: A computational and experimental study was conducted on nozzle geometries for rocket application rotating detonation engines (RDEs). Three geometries, including a nozzleless blunt body typically em...

Design of Low-Cost Solid Propellant Engine Test Bench and Electronic Embedded System used for Small Rockets

Abstract: The development of experimental propulsion projects by undergraduate students introduces new practical knowledge to improve competitiveness in the rocket industry. This work addresses a methodology for testing solid engines based on an embedded and mechanical system at low cost from the modeling of the solid potassium nitrate propellant with candy-rocket sucrose propellant (KNSB). The purpose of this study was modeling of the rocket engine, the design and manufacture of the test bench and the design of the Integrated system to measure the thrust of the rocket engine. Complementarily, the result of an experimental measurement is detailed and compared with the analytical study of the thrust. Using small rockets to study atmospheric precipitation requires several engine tests. For this reason, it is important to study different engines for the academic missions that are proposed in the future. The Universidad Nacional de Ingenieria promotes this research that also participates in the “CSPACE” launch campaigns organized by the French Space Agency in Tarbes, France and won the Technological Innovation Award from the Peruvian Oil and Energy Mining Society in 2018.

Development of Green Storable Hybrid Rocket Propulsion Technology Using 98% Hydrogen Peroxide as Oxidizer

Abstract: This paper presents the development of indigenous hybrid rocket technology, using 98% hydrogen peroxide as an oxidizer. Consecutive steps are presented, which started with interest in hydrogen peroxide and the development of technology to obtain High Test Peroxide, finally allowing concentrations of up to 99.99% to be obtained in-house. Hydrogen peroxide of 98% concentration (mass-wise) was selected as the workhorse for further space propulsion and space transportation developments. Over the course nearly 10 years of the technology’s evolution, the Lukasiewicz Research Network—Institute of Aviation completed hundreds of subscale hybrid rocket motor and component tests. In 2017, the Institute presented the first vehicle in the world to have demonstrated in-flight utilization for 98% hydrogen peroxide. This was achieved by the ILR-33 AMBER suborbital rocket, which utilizes a hybrid rocket propulsion as the main stage. Since then, three successful consecutive flights of the vehicle have been performed, and flights to the Von Karman Line are planned. The hybrid rocket technology developments are described. Advances in hybrid fuel technology are shown, including the testing of fuel grains. Theoretical studies and sizing of hybrid propulsion systems for spacecraft, sounding rockets and small launch vehicles have been performed, and planned further developments are discussed.

Formation mechanism of high-frequency combustion oscillations in a model rocket engine combustor

Abstract: We study the formation mechanism of high-frequency combustion oscillations in a model rocket combustor from the viewpoints of symbolic dynamics and complex networks. The flow velocity fluctuations in the fuel injector generated by the pressure fluctuations in the combustor give rise to the periodic ignition of the unburnt fuel/oxidizer mixture, resulting in a significant change in the heat release rate fluctuations in the combustor. The heat release rate fluctuations drive the pressure fluctuations in the combustor before a transition state, while the pressure fluctuations in the combustor gradually begin to significantly affect the heat release rate fluctuations during the transition to combustion oscillations. The directional feedback process during the transition and subsequent combustion oscillations is identified by the directionality index of the symbolic transfer entropy. The thermoacoustic power network enables us to understand the physical mechanism behind the transition and subsequent combustion oscillations.

Thrust Validation of Rotating Detonation Engine System by Moving Rocket Sled Test

Abstract: This Paper presents the results of a rotating detonation engine rocket sled test demonstration on a 127 mm gauge rail at 100 m long high-speed running track test equipment to evaluate the thrust pe...

Performance Analysis and Mass Estimation of a Small-Sized Liquid Rocket Engine with Electric-Pump Cycle

Abstract: The propellant supply system of a liquid rocket engine using an electric pump has high reliability because of the relatively small number of components. The system also has the merit of a quick response and ease of control owing to its simple configuration. Recently, the rocket lab developed the Rutherford engine, which has an electric pump cycle, because of the improved technology in the electric motor and battery. This paper examined the development of the electric-pump cycle and compared the performance with other cycles for a small-sized low-thrust rocket engine. Performance analysis and mass estimation were conducted using the developed analysis program, in which reliability in mass estimation was improved based on the designed configuration or real performance data from commercial products. In addition, the modeling method and analysis procedure were described in detail. The results showed that it is possible to develop a small-sized engine with an electric-pump cycle when the present technologies are applied. The electric-pump cycle had a smaller dry mass than the gas-generator cycle, even at a low thrust level of 500 N, and showed higher performance in specific impulse and speed increments.

Free variable mass nonlocal systems, jerks, and snaps, and their implications in rotating fluids in rockets

Abstract: Dynamical systems with position varying mass are archetypal examples of classical mechanical systems with rocket engine being a typical realistic model In the present study, we extend the classical Newtonian mechanics by replacing the kinetic energy by a nonlocal-in-time kinetic energy and the standard velocity by a fractional velocity These replacements lead to an extension of Newton’s second law of motion which has interesting implications in incompressible fluid dynamics As an application, we discuss the rotating fluid problem subject to a position varying fluid mass which occurs in rocket dynamics Several features were observed, mainly the transition from order to disorder in rotating fluids in rockets

Trajectory Design for the ESA LISA Mission

Abstract: The three Laser Interferometer Space Antenna (LISA) spacecraft are going to be placed in a triangular formation in an Earth-trailing or Earth-leading orbit. They will be launched together on a single rocket and transferred to that science orbit using Solar Electric Propulsion. Since the transfer Δv depends on the chosen science orbit, both transfer and science orbit have been optimised together. For a thrust level of 90 mN, an allocation of 1092 m/s per spacecraft is sufficient for an all-year launch in 2034. For every launch month a dedicated science orbit is designed with a corner angle variation of 60° ± 1.0° and an arm length rate of maximum 10 m/s. Moreover, a detailed navigation analysis of the science orbit insertion and the impact on insertion errors on the constellation stability has been conducted. The analysis shows that Range/Doppler measurements together with a series of correction manoeuvres at the beginning of the science orbit phase can reduce insertion dispersions to a level where corner angle variations remain at about 60° ± 1.1° at 99% C.L. However, the situation can become significantly worse if the self-gravity accelerations acting during the science orbit phase are not sufficiently characterised prior to science orbit insertion.

Active vibration suppression for large space structure assembly: A distributed adaptive model predictive control approach:

Abstract: Subject to the limited envelope size and carrying capacity of a rocket, it could be one of the most promising solutions to construct a large space structure through multilaunching and on-orbit asse...

Experimental Investigation of Rotating Detonation Rocket Engines for Space Propulsion

Abstract: The performance (specific impulse, Isp) of a modular, 150-lbf-class rotating detonation rocket engine (RDRE) was measured with three gaseous fuels (methane, ethane, and ethylene) and gaseous oxygen...

Rankine Vortex Suppression in Tanks with Conical Base: A Numerical Investigation

Abstract: During draining of liquid propellants from propellant tanks of spacecraft rocket engines, air core vortex can be formed and this can result in blockage of drain port. Supply of required amount of f...

Early detection of thermoacoustic instabilities in a cryogenic rocket thrust chamber using combustion noise features and machine learning

Abstract: We present a data-driven method for the early detection of thermoacoustic instabilities. Recurrence quantification analysis is used to calculate characteristic combustion features from short-length time series of dynamic pressure sensor data. Features like recurrence rate are used to train support vector machines to detect the onset of instability a few hundred milliseconds in advance. The performance of the proposed method is investigated on experimental data from a representative LOX/H 2 research thrust chamber. In most cases, the method is able to timely predict two types of thermoacoustic instabilities on test data not used for training. The results are compared with state-of-the-art early warning indicators.

Investigation of CO recombination in the boundary layer of CH4/O2 rocket engines

Abstract: In this work, the combustion and thermal recombination in the boundary layer of a single-element methane/oxygen rocket combustor is investigated using large-eddy simulations. The experimental configuration consists of a coaxial injector and an operating point with a nominal pressure of 20 bar and gaseous injection of both propellants are considered. A non-adiabatic flamelet model is utilized with the purpose of examining its capability to predict the wall heat transfer. Good agreement of the simulation results with measurements of heat flux and pressure profiles is obtained using the non-adiabatic model. By comparing results with a frozen flamelet model, the importance of the recombination reactions in the cold boundary layer is investigated. The species profiles of CO and CO2 are examined and the reaction pathways leading to the recombination reactions at the wall were analyzed. Results from this analysis show that the reaction of OH with CO forming CO2 is the main contributor to the additional heat release in the boundary layer.