Propulsion Community Research Articles

Selected Papers from 4th National Aerospace Propulsion Conference NAPC 2022

The National Aerospace Propulsion Conference (NAPC) is a national conference dedicated to aerospace propulsion technologies. Two erstwhile popular conference series, the National Propulsion Conference (NPC) and the National Conference on Air-Breathing Engines (NCABE) were merged to result in NAPC. This biennial conference is intended to bring together the entire propulsion community spanning the industry, academia and the research labs nationwide. The aerospace propulsion-centred conference is considered an ideal opportunity to showcase one’s research activities with peers and foster future collaborations through networking. The 4th edition of the conference, NAPC-2022, comprised a keynote lecture, several topical invited lectures and presentations of contributed papers in parallel sessions. This special issue contains eight manuscripts selected based on peer-reviews of selected papers presented during the conference. The manuscripts broadly cover several aspects pertaining to aerospace propulsion that includes airbreathing engine components such as intakes, compressors, combustors and nozzles. From the rocket propulsion (non-airbreathing), there are papers that investigate the rocket motor grain port alignment and another paper on the heat transfer characteristics of a supersonic rocket nozzle.

Evaluating New Liquid Storable Bipropellants: Safety and Performance Assessments

Conventional storable bipropellants make use of hydrazine and its derivatives as fuels and nitrogen tetroxide as an oxidizer. In recent years, the toxicity character of these chemicals pushed the propulsion community towards “green” alternatives. Several candidates have been proposed among existing and newly developed chemicals, highlighting the need for a common and robust selection methodology. This paper aims at reviewing the most important selection criteria in the field of toxicity and discusses how to objectively define a green propellant, considering both the health and environmental hazards caused by the chemicals. Additionally, consistent figures of merit in the field of safety and handling operations and performance are proposed. In particular, operating temperatures, flammability and stability issues are discussed in the framework of physical hazards and storage requirements, while vacuum impulses, adiabatic flame temperature and sooting occurrence of the investigated couples are compared to the UDMH/NTO benchmark case. Hydrogen peroxide and nitrous oxide, and light hydrocarbons, alcohols and kerosene are selected from the open literature as promising green oxidizers and fuels, respectively. The identified methodology highlights merits and limitations of each chemical, as well as the fact that the identification of a universally best suited green couple is quite impractical. On the contrary, the characteristics of each propellant lead to a scenario of several “sub-optimal” couples, each of them opportunely fitting into a specific mission class.

Overview of Al-based nanoenergetic ingredients for solid rocket propulsion

Abstract The introduction of nano-sized energetic ingredients first occurred in Russia about 60 years ago and arose great expectations in the rocket propulsion community, thanks to the higher energy densities and faster energy release rates exhibited with respect to conventional ingredients. But, despite intense worldwide research programs, still today mostly laboratory level applications are reported and often for scientific purposes only. A number of practical reasons prevent the applications at industrial level: inert native coating of the energetic particles, nonuniform dispersion, aging, excessive viscosity of the slurry propellant, possible limitations in mechanical properties, more demanding safety issues, cost, and so on. This paper describes the main features in terms of performance of solid rocket propellants loaded with nanometals and intends to emphasize the unique properties or operating conditions made possible by the addition of the nano-sized energetic ingredients. Steady and unsteady combustion regimes are examined.

A METHOD FOR SCREENING AND IDENTIFICATION OF GREEN HYPERGOLIC BIPROPELLANTS

As of now, the most widely used hypergolic fuels are hydrazine derivates with nitrogen tetroxide as the oxidizer. The toxicity and carcinogenic character of these substances make their handling challenging even if gelled. Therefore, there is great interest throughout the propulsion community in finding alternatives for these substances. However, despite a thorough literature review, clear selection criteria for 'green' hypergols could not be found. To tackle this issue, we developed a set of key requirements and criteria for propellant selection. The major drivers were hypergolic ignitability, propellant performance, reduced toxicity, and economic factors. Then we used these criteria to carry out a screening aimed at identifying potential oxidizers, fuels, and additives. Hypergolicity was assessed by capturing ignition events of a simple drop test experiment with a high-speed camera. This setup allowed us to find a set of promising novel hypergolic propellants and to verify our screening method. Not only do liquid and gelled diamine fuels with hydrogen peroxide as the oxidizer provide 96% of the specific impulse of the monomethylhydrazine/nitrogen tetroxide propellant system, but they also feature ignition delay times on the order of 15 ms.

Recommended Practice for Pressure Measurement and Calculation of Effective Pumping Speed in Electric Propulsion Testing

The electric propulsion community has been implored to establish and implement a set of universally applicable test standards during the research, development, and qualification of electric propulsion systems. Variability between facility-to-facility and more importantly ground-to-flight performance can result in large margins in application or aversion to mission infusion. Performance measurements and life testing under appropriate conditions can be costly and lengthy. Measurement practices must be consistent, accurate, and repeatable. Additionally, the measurements must be universally transportable across facilities throughout the development, qualification, spacecraft integration, and on-orbit performance. A recommended practice for making pressure measurements, pressure diagnostics, and calculating effective pumping speeds with justification is presented.

Recommended Practice for Flow Control and Measurement in Electric Propulsion Testing

Accurate control and measurement of propellant flow to a thruster is one of the most basic and fundamental requirements for operation of electric propulsion systems, whether they be in the laboratory or on flight spacecraft. Hence, it is important for the electric propulsion community to have a common understanding of typical methods for flow control and measurement. This paper addresses the topic of propellant flow primarily for the gaseous propellant systems that have dominated laboratory research and flight application over the last few decades, although other types of systems are also briefly discussed. Although most flight systems have employed a type of pressure-fed flow restrictor for flow control, both thermal-based and pressure-based mass flow controllers are routinely used in laboratories. Fundamentals and theory of operation of these types of controllers are presented, along with sources of uncertainty associated with their use. Methods of calibration and recommendations for calibration processes are presented. Finally, details of uncertainty calculations are presented for some common calibration methods and for the linear fits to calibration data that are commonly used.

A radial non-uniform helicon equilibrium discharge model

Helicon discharges have attracted great attention in the electric propulsion community in recent years. To acquire the equilibrium properties, a self-consistent model is developed, which combines the helicon/Trivelpiece—Gould (TG) waves-plasma interaction mechanism and the plasma flow theory under the confinement of the magnetic field. The calculations reproduce the central peak density phenomenon observed in the experiments. The results show that when operating in the wave coupling mode, high magnetic field strength B0 results in the deviation of the central density versus B0 from the linear relationship, while the density rise becomes flatter as the radiofrequency (rf) input power Prf grows, and the electron temperature Te radial profile is mainly determined by the characteristic of the rf energy deposition. The model could provide suggestions in choosing the B0 and Prf for medium power helicon thrusters.

Status of turbulence modeling for hypersonic propulsion flowpaths

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier–Stokes (RANS) methods, but some discussion of newer methods such as large eddy simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath, including laminar-to-turbulent boundary layer transition, shock wave/turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers), and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

Techniques for Solving Stiff Chemical Kinetics on Graphical Processing Units

The cost of integrating detailed finite rate chemical kinetics mechanisms can be prohibitive in turbulent combustion simulations. Techniques that can significantly reduce these simulation times are of great interest to model developers and the broader propulsion and power community. Two new integration methods using graphical processing units are presented that can rapidly integrate the nonlinear ordinary differential equations at each grid point. The explicit graphical processing-unit-enabled fourth-order-accurate Runge–Kutta–Fehlberg ordinary differential equation solver achieved a maximum speed up of 20.2x over the baseline implicit fifth-order-accurate DVODE CPU run time for large numbers of independent ordinary differential equation systems with comparable accuracy. The graphical processing unit implementation of DVODE achieved a maximum speed up of 7.7x over the baseline CPU run time. The performance impact of mapping one graphical processing unit thread to each ordinary differential equation system was compared with mapping an entire graphical processing unit thread block per ordinary differential equation (i.e., multiple threads per ordinary differential equation). The one-thread-per-ordinary-differential-equation approach achieved greater overall speed up but only when the number of independent ordinary differential equations was large. The one-block-per-ordinary-differential-equation implementation of Runge–Kutta–Fehlberg and DVODE both achieved lower peak speed ups, but outperformed the serial CPU performance with as few as 100 ordinary differential equations. The new graphical processing-unit-enabled ordinary differential equation solvers demonstrate a method to significantly reduce the computational cost of detailed finite rate combustion simulations.

Impact of plasma noise on a direct thrust measurement system

In order to evaluate the accuracy and sensitivity of a pendulum-type thrust measurement system, a linear variable differential transformer (LVDT) and a laser optical displacement sensor have been used simultaneously to determine the displacement resulting from an applied thrust. The LVDT sensor uses an analog interface, whereas the laser sensor uses a digital interface to communicate the displacement readings to the data acquisition equipment. The data collected by both sensors show good agreement for static mass calibrations and validation with a cold gas thruster. However, the data obtained using the LVDT deviate significantly from that of the laser sensor when operating two varieties of plasma thrusters: a radio frequency (RF) driven plasma thruster, and a DC powered plasma thruster. Results establish that even with appropriate shielding and signal filtering the LVDT sensor is subject to plasma noise and radio frequency interactions which result in anomalous thrust readings. Experimental data show that the thrust determined using the LVDT system in a direct current plasma environment and a RF discharge is approximately a factor of three higher than the thrust values obtained using a laser sensor system for the operating conditions investigated. These findings are of significance to the electric propulsion community as LVDT sensors are often utilized in thrust measurement systems and accurate thrust measurement and the reproducibility of thrust data is key to analyzing thruster performance. Methods are proposed to evaluate system susceptibility to plasma noise and an effective filtering scheme presented for DC discharges.

Empirical correlations for the solubility of pressurant gases in cryogenic propellants

We have analyzed data published by others reporting the solubility of helium in liquid hydrogen, oxygen, and methane, and of nitrogen in liquid oxygen, to develop empirical correlations for the mole fraction of these pressurant gases in the liquid phase as a function of temperature and pressure. The data, compiled and provided by NIST, are from a variety of sources and covers a large range of liquid temperatures and pressures. The correlations were developed to yield accurate estimates of the mole fraction of the pressurant gas in the cryogenic liquid at temperature and pressures of interest to the propulsion community, yet the correlations developed are applicable over a much wider range. The mole fraction solubility of helium in all these liquids is less than 0.3% at the temperatures and pressures used in propulsion systems. When nitrogen is used as a pressurant for liquid oxygen, substantial contamination can result, though the diffusion into the liquid is slow.

Analysis and design of a friction brake for momentum exchange propulsion tethers

The purpose of this paper is to present an engineering assessment of a proposal for a deployment brake for a momentum exchange tether. Tethers have attracted considerable interest amongst the propulsion community for many years for orbit raising, energy conversion, de-orbiting, and injection into interplanetary transfers, and the motivation for the work described in the paper is directly mission-oriented. The YES2 SpaceMail mission has been planned as a low-cost international effort to provide a facility for the return of small samples from the ISS back to Earth, based on an inflatable re-entry capsule and a mechanical tether system. The system will be lofted into orbit as part of a payload on a Foton/Bion rocket launcher. The brake design concept is based on the use of friction between the deploying tether and a short pole onto which several turns are wound, and the paper summarises the modelling work that has been done, and then discusses an experiment in which the premise was tested in the laboratory. The paper concludes with certain practical proposals for implementation.

Mo-Si-B Alloys: Developing a Revolutionary Turbine-Engine Material

AbstractThis article discusses three main points. First, ultrahigh-temperature materials offer a large return to the propulsion community and to the economy, but replacing Ni-based alloys is a very difficult problem for both technical and financial reasons. Second, the oxidation resistance of selected Mo-Si-B alloys and the prospects for combining this with useful structural properties are remarkable, given the nature of conventional Mo alloys; however, there are nagging issues associated with their low-temperature (∼700°C) behavior. Third, further advances in the processing of such materials, together with assessments of their affordability and reliability, are vital for achieving a successful engineering technology.

Reply to Robert L. Glick's Comment: Physicality of Core Flow Models in Rocket Motors

T HE followingexpositionis intended to clarify the fundamental issues raised in Glick’s commentary. Despite the commentary’s focus on one individual study,2 the multiple questions that were raised apply just as easily to most theoretical, numerical, and experimental simulations of idealized motors based on simpliŽ ed geometric conŽ gurations and smooth porous walls. In fact, all reported core  ow studies have so far relied, to some extent, on idealistic conditions. The main concern by Glick1 is that mathematicalmodels may be inadequate as vehicles for physical understanding, especially that burning surfaces in production rocket motors are rough, heterogeneous, and pliable. From this perspective, a core  ow that does not account for all conceivable features, including surface vibrations and complex boundary conditions, may be deemed impractical or uncertain. Unfortunately,Glick’s commentary does not explore the physical beneŽ ts of mathematical models nor does it recognize the role, scope, objectives, and recent successes of core  ow idealizations. These have been motivated by important technologicalapplications that will now be overviewed. As a problem of real concern in rocket motors and large gas turbines, aeroacoustic instabilities have received much scrutiny in the propulsion community. Models of these instabilities have invariably pointed to the importance of providing judicious assessments of correspondingcore  ow details.3i8 The desire for explicit  ow models has also been motivated by the need to understand physically the intricate coupling between unsteady pressure waves and gas motions.9i12 As ultimately suggestedby repeated tests, not only does this inevitable pairing provoke unsteady burning, but it also generates intense sound-pressure levels and boundary-driven vortices. This delicateinterplayof underpinningcore  owmechanismshas inspiredover the years severalcapable theoreticiansto seekphysical idealizations.The goal has been generallyset to isolatecarefully the intricate mechanisms by parametric linearization or vector decomposition. In the midst of this unusually complex problem, the quest for basic answers has often become a central focus. Pioneeredby Culick13i19 andFlandro,3i8 theoreticalstudiessimilar to the one under consideration2 have elucidated a number of physical features in rocket motors. Among them were the multidimensional spatial and temporal velocity, vorticity, and stress distributions along the length of simulated motors. Points exhibiting maximum and minimum stress disturbances were identiŽ ed, and the acoustic character in the chamber was being disclosed.21i27 These analyses have been accomplished by Ž rst unraveling the problem’s principal convection–diffusion equations and their un-

Modeling of supersonic combustor flows using parallel computing

Computational Fluid Dynamics (CFD) has matured rapidly in the past 20 years and is now an important tool for analyzing and understanding complex fluid flows. Since 1985, CFD has played a vital role in the study of hypersonic flight. It has provided the capability for scientists and engineers to model both internal and external hypersonic flow-fields. Such flows are often impractical or impossible to analyze in laboratory conditions. In particular, the recent application of CFD to the modeling of internal reacting supersonic combustor flows has significantly advanced the understanding of such flows and has increased confidence in the predictive ability of codes. The purpose of these efforts has been to provide the hypersonic propulsion community with realistic large-scale applications of CFD and to use these solutions in direct support of engineering analysis and design of hypersonic vehicles. Although these applications have been successful to date, expectations and requirements are increasing dramatically for both faster turn-around of solutions and for more detailed and accurate solutions (hence requiring greater computational mesh refinement, more complete chemistry and turbulence models, etc.). In order to begin to meet these requirements, a ten-fold or greater increase in computational efficiency is required, relative to current supercomputing capabilities. This increase can be achieved easily by suitably programming existing CFD technology on existing distributed memory parallel computing machines or multicomputers. This paper presents and analyzes the results obtained to date in an investigation aimed at the application of parallel computing to the simulation of scramjet combustor flow-fields.