Space Shuttle Main Engine Research Articles

A parametric study of squealer tip geometries applied in a hydraulic axial turbine used in a rocket engine turbopump

Axial turbines are machines widely used in different engineering applications. Due to their constructive characteristics, they must have a space between the rotor blades and the turbine casing, called tip clearance. Unfortunately, this gap allows a part of the fluid to leak from the pressure side to the suction side of the rotor blades. This leakage is undesirable and represents an energy loss. A way to avoid part of this loss is through the use of desensitization techniques. Although the use of these techniques is widely known, no studies in the open literature have evaluated these techniques in hydraulic turbines. This work presents a numerical analysis of squealer desensitization techniques applied in a hydraulic axial turbine. The turbomachine under study is the first stage of the hydraulic axial turbine used in the Low Pressure Oxidizer Turbopump (LPOTP) of the Space Shuttle Main Engine (SSME). Numerical simulations were performed using CFX v.19.2 software, and computational meshes were generated in ICEM v.19.2 software. Initially, the computational model was validated, using the experimental results published by the National Aeronautics and Space Administration (NASA). A parametric analysis was performed considering the variation in squealer cavity depth and rim thickness. The study found that the squealer cavity depth has a greater influence on the stage performance than its rim thickness. The tendency is that the greater the cavity depth, the greater the stage efficiency. One of the squealer geometries analyzed allowed an average increased efficiency of 1.43%, over the entire turbine operational range. The results obtained also show that the application of the proposed geometries would enable the reduction in cavitation close to the trailing edge of the rotor blades. This result is extremely valuable, as it can impact the life cycle of the turbine.

Effect of tip clearance on cavitating flow of a hydraulic axial turbine applied in turbopump

The requirements of Liquid Propellant Rocket Engine (LPRE) are high for thrust, specific impulse, and flow rate; thus, its components also have strict requirements. For the turbopumps (TPs), this means high flow rate, high rotational speed, and high pressure ratio, which makes their operations susceptible to the cavitation phenomenon, as observed in two previous works. In the first, cavitation regions were observed in the first stage of the Space Shuttle Main Engine (SSME) Liquid Oxygen (LOX) booster turbine, for 3.0, 5.5, and 8.0% tip clearances (relative to rotor blade height), using monophase flow (Lindquist Whitacker et al., 2017). In the second, the simulations were performed with multiphase flow, producing results more physically coherent for the 3.0% gap configuration (Whitacker et al., 2018). The characteristics of both types of simulations in space propulsion applications still require better understanding. Therefore, to compare monophase and multiphase results at various operating points and turbine configurations, steady-state turbulent 3-D Computational Fluid Dynamics (CFD) simulations were performed, based on Reynolds-Averaged Navier-Stokes (RANS) formulation. The same three tip configurations for the turbine first stage were simulated, and the calculations were validated using experimental results from the National Aeronautics and Space Administration (NASA) (Boynton and Rohlik, 1976). This made it possible to verify the effect of the tip clearance on the machine performance and internal flowfield. When the gap increased, the pressure loading decreased in a large region of the blade tip, the interaction was greater between the Tip Clearance Vortex (TCV) and a vortex generated around the shroud cavitation region (Cavitation Vortex - CV), and this interaction moved towards the middle of the blade-to-blade passage. Thus, the losses increased and the efficiency decreased. Various comparative aspects between the simulations using both mono and multiphase numerical schemes are also discussed.

Modeling the effect of reusability on the performance of an existing LPRE

The objective of this research is to study the impact of main engine parameters on the performance of a reusable liquid propellant rocket engine (LPRE). For this purpose, after a brief review over the available reusable liquid propellant rocket engines (RLPREs), the Space Shuttle Main Engine (SSME) is selected as a case study and a modular simulation software has been developed for analyzing the steady-state performance of LPRE in MATLAB Simulink. The evolution path of SSME is studied in detail and considering the reusability, two approaches have been selected. In the first approach, the performance of an expendable version of SSME is simulated and compared with the reusable version. It is found that the SSME working at 109% or higher rated thrust level (RTL) is the most possible option for the expendable version of this engine. The simulation results show that the components of the expendable SSME would work at 3–12% higher operating condition and the engine thrust would be 4.76% higher. In the second approach, in order to decrease the loads on the engine components and to enhance the reusability capability, the throat of the combustion chamber is enlarged. The simulation results show that by 11% increase in the combustion chamber throat area, the average load on the engine's different components would decrease 5–17% but for the same level of thrust, the specific impulse would be 2.43 s less. It is believed that the obtained results can help the designers to have a quantitative overview of the effect of reusability on the LPRE and launch vehicle performance during the conceptual design phase.

Two dimensional rotor blade numerical analysis using ANSYS

Turbo-pump is an essential component in any rocket engine. The airfoil (NGTE 10C4 series) section of a rotor blade in the SSME (Space Shuttle Main Engine) was modelled along with a far-field of suitable shape and dimensions. The model was suitably discretized with proper amount of refinement, etc. A numerical simulation was then carried-out using ANSYS software with several iterations of varying inlet boundary conditions to check for any improvement in the performance of the rotor blade. Inlet velocity for optimum functionality of the blade was identified.

Static performance modeling and simulation of the staged combustion cycle LPREs

The objective of the current research is to analyze the steady state performance of staged combustion cycle liquid propellant rocket engines (SCCLPREs) and to provide a comprehensive case study with all the required technical input data. For this purpose, using the modular concept, engine components are classified into 14 modules and a modeling and simulation platform for SCCLPRE integrated system simulation has been developed in MATLAB Simulink. Space shuttle main engine (SSME) has been selected as a case study, and its static performance has been studied. A total of 34 elements has been taken into account, and using 110 linear/non-linear equations the engine's system modular model has been established. Most of the equations are solved by direct substitution of the initial guesses and iteration through eleven nested iterative loops and a few nonlinear complex equations are solved by Newton–Raphson method. The SSME's throttled performance at four different throttling levels has been simulated, and the effect of throttling has been studied on engine major elements. It is found that by throttling, the pressure of preburners, combustion chamber and high-pressure turbopumps are influenced more than other elements. The simulation mean error at 109% rated thrust level (RTL), 100% RTL, 60% RTL and 50% RTL is 1.63%, 5.76%, 12.91% and 15.28% respectively. Comparing to the previous studies, the simulation accuracy has been enhanced 3% to 7%. The developed model is a basis for further research over SSME and other SCCLPRE.

Characterization of Cavitation Instabilities in a Four-Bladed Turbopump Inducer

Rocket engine turbopump inducers can suffer from cavitation instabilities, dramatically limiting performance and life of the turbomachinery components. An unrestricted inducer geometry, representative of the Space Shuttle main engine low-pressure oxidizer pump inducer, is defined and serves as a platform for characterization of the cavitation dynamics. Experiments demonstrate rotating cavitation, alternate blade cavitation, and cavitation surge. Borrowing ideas from the modeling and assessment of rotating stall in aeroengine compressors, a previously established traveling wave energy analysis is applied. At flow coefficients higher than design, a single cell cavity occurs rotating at supersynchronous frequencies between 1.2 and 1.5 times rotor frequency. Optical measurements corroborate the findings and underlying mechanisms; the onset of rotating cavitation is governed by the interaction of the blade cavity with the leading edge of the adjacent blade. This leads to a change of incidence, which causes sheet cavity breakoff with periodic growth and collapse of cavities, yielding the apparent supersynchronous rotation of the cavities around the annulus. The paper breaks new ground in diagnostics of inducer dynamic behavior during rotating cavitation, alternate blade cavitation, and cavitation surge. The details of the unrestricted inducer geometry are provided for advancement of international research on rocket engine cavitation dynamics.

An evaluation of the tip clearance effects on turbine efficiency for space propulsion applications considering liquid rocket engine using turbopumps

Large launch vehicles use their propulsion systems based on Liquid Rocket Engines (LRE) equipped with turbopumps. Turbopumps are complex rotary machines that supply high power, mass flow, and pressures in the engine system to reach the thrust requirements as determined in the rocket engine thermodynamic cycle. Strong engines need a secondary turbopump system called a booster. These boosters have pumps and turbines smaller than those of the main engine turbopumps, and their important function is to increase the fluid pressure at the inlet of the main turbopumps, mainly to avoid cavitation. In the present work, the influence of the tip clearance issues in an axial turbine installed to operate as oxidizer booster in the Space Shuttle Main Engine (SSME) were evaluated numerically. The results are compared with experimental data from National Aeronautics and Space Administration (NASA). The flow characteristics and the variation in the turbine efficiency for different jet velocities were determined for three different tip clearance values associated with the percentage of turbine blade height: 3.0%, 5.5%, and 8.0%. The turbine design, numerical issues, mesh generation and results are described and discussed. The methodology and numerical simulations used in the present work was consistent with the experimental data and can be extended for other correlated numerical simulations related to axial hydraulic turbines.

Modular simulation software development for liquid propellant rocket engines based on MATLAB Simulink

Focusing on Liquid Propellant Rocket Engine (LPRE) major components, a steady state modular simulation software has been established in MATLAB Simulink. For integrated system analysis, a new algorithm dependant on engine inlet mass flow rate and pressure is considered. Using the suggested algorithm, it is possible to evaluate engine component operation, similar to the known initial parameters during hot fire test of the engine on stand. As a case study, the reusable Space Shuttle Main Engine (SSME) has been selected and the simulation has been performed to predict engine’s throttled operation at 109 percent of the nominal thrust value. For this purpose the engine actual flow diagram has been converted to 34 numerical modules and the engine has been modelled by solving a total of 101 steady state mathematic equations. The mean error of the simulation results is found to be less than 5% compared with the published SSME data. Using the presented idea and developed modules, it is possible to build up the numerical model and simulate other LPREs.

Dynamic Simulation and Analysis of the Space Shuttle Main Engine with Artificially Injected Faults

Securing the safety and the reliability of liquid-propellant rocket engines (LREs) for space vehicles is indispensable as engines consist of many complex components and operate under extremely high energy-dense conditions. Thus, health monitoring has become a mandatory requirement, especially for the reusable LREs that are currently being developed. In this context, a dynamic simulation program based on MATLAB/Simulink was developed in the current research on the Space Shuttle Main Engine (SSME), a partly reusable engine. Then, a series of fault simulations using this program was conducted: at a steady state operating condition (104% Rated Propulsion Level), various simulated fault conditions were artificially injected into the simulation models for the five major valves, the pumps, and the turbines of the SSME. The consequent effects due to each fault were analyzed based on the time responses of the major parameters of the engine. It is believed that this research topic is an essential pre-step for the development of fault detection and diagnosis algorithms for reusable engines in the future.

Application of fault factor method to fault detection and diagnosis for space shuttle main engine

This paper deals with an application of the multiple linear regression algorithm to fault detection and diagnosis for the space shuttle main engine (SSME) during a steady state. In order to develop the algorithm, the energy balance equations, which balances the relation among pressure, mass flow rate and power at various locations within the SSME, are obtained. Then using the measurement data of some important parameters of the engine, fault factors which reflects the deviation of each equation from the normal state are estimated. The probable location of each fault and the levels of severity can be obtained from the estimated fault factors. This process is numerically demonstrated for the SSME at 104% Rated Propulsion Level (RPL) by using the simulated measurement data from the mathematical models of the engine. The result of the current study is particularly important considering that the recently developed reusable Liquid Rocket Engines (LREs) have staged-combustion cycles similarly to the SSME.

Application of Constraint Force Equation Methodology for Launch Vehicle Stage Separation

A = axial location of the solid rocket booster (SRB) reference point in Space Shuttle main engine (SSME) plume coordinate system, ft ax, ay, az = acceleration components along body axes (excluding components due to gravity), ft=s CA = isolated (freestream) axial force coefficient Cm = isolated (freestream) pitching moment coefficient CN = isolated (freestream) normal force coefficient Cn = isolated (freestream) yawing moment coefficient CY = isolated (freestream) side force coefficient Cl = isolated (freestream) rolling moment coefficient eA, eB = unit vectors in body A and body B, respectively Fp = plume impingement force, lb Fx, Fy, Fz = aerodynamic forces in axial, lateral, and normal directions, lb F CON A , F CON B = joint constraint force vector for body A and body B

An optical sensor for multi-species impurity monitoring in hydrogen fuel

Abstract Hydrogen is the fuel used by the National Aeronautics and Space Administration (NASA) for the Space Shuttle Main Engine (SSME), among others. The purity of the hydrogen fuel is very important for the performance of hydrogen-fueled engines. In the present work, we have designed, developed, and tested a state-of-the-art laser-induced breakdown spectroscopy (LIBS) based optical sensor to detect four important trace level impurities (i.e. nitrogen, argon, oxygen, and helium) in hydrogen as a rocket fuel. The sensor can simultaneously measure the concentrations of nitrogen, argon, oxygen, and helium in hydrogen fuel storage tanks and supply lines. LABVIEW ® based software is also developed to control and operate the optical sensor. The software performs the data analysis and provides elemental concentrations (in ppm) of these impurities. The sensor has estimated limits of detection of 80 ppm for nitrogen, 97 ppm for argon, 10 ppm for oxygen, and 25 ppm for helium. This prototype sensor is easy to deploy in the field.

Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME) or Vulcain (Ariane 5), the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL), spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

Space shuttle main propulsion system anomaly detection: A case study

The space shuttle main engine (SSME) is part of the Main Propulsion System (MPS) which is an extremely complex system containing several sub-systems and components, each of which must work precisely in order to achieve a successful mission. A critical component under study is the flow control valve (FCY) which controls the pressure of the gaseous hydrogen between the SSME and the external fuel tank. The FCV has received added attention since a Space Shuttle Mission in November 2008, where it was discovered during the mission that an anomaly had occurred in one of the three FCV's. Subsequent inspection revealed that one FCV cracked during ascent. This type of fault is of high criticality because it can lead to potentially catastrophic gaseous hydrogen leakage. A supervised learning method known as Virtual Sensors (VS), and an unsupervised learning method known as the Inductive Monitoring System (IMS) were used to detect anomalies related to the FCV in the MPS. Both algorithms identify the time of the anomaly in a multi-dimensional time series of temperatures, pressures, and control signals related to the FCV. This discovery corroborates the results of the inspection and also reveals the time at which the anomaly likely occurred. The methods were applied to data obtained from the March 2009 launch of Space Shuttle Discovery to determine whether an anomaly occurred in the same sub-system. According to our models, the FCV sub-system showed nominal behavior during ascent.

발사체 개발의 연차별 비용 추정을 위한 비용모델 개발

성공적인 발사체 개발을 위해서는 개발비용을 정확하게 추정하는 것도 중요하지만 연차별 예산 배분을 적절히 계획하는 것도 중요하다. 본 논문에서는 통계적 기법을 이용하여 개발기간 중 연차별 예산 배분에 대하여 분석하였다. 연차별 비용 모델링을 위해서 통계적 모델의 분석을 통해 적합성을 판단하여 베타분포모델을 선택하였고, 프로젝트의 설계와 기술의 성숙도로 추정이 가능한 모델을 개발하였다. 연차별 비용 모델의 타당성을 검증하기 위하여 우주왕복선 주 엔진(Space Shuttle Main Engine)의 실제 개발비용 분포와 본 모델을 통한 개발비용 분포의 비교 연구를 통해 적합성을 확인하였다. 또한, 본 논문의 개발 비용 모델을 한국형발사체 개발에 적용하여 연차별 소요 비용을 추정하였다. 본 논문에서 제시하는 연차별 분포 모델은 발사체뿐만 아니라 다른 대형복합시스템의 개발에도 활용이 가능할 것으로 예상된다. In order to develop a launch vehicle successfully, it is important to estimate development costs accurately but it is also important to plan the annual budget. In this paper, the statistical method was utilized for cost spreading. For cost spread modeling, the suitability of the model by analyzing several statistical models was evaluated and consequently, the beta-distribution model has been selected. In this study, the validity of the annual estimation cost model was verified through the comparison of the actual development cost distribution and the estimating cost distribution of Space Shuttle Main Engine. In addition, this paper estimated the annual budget required for the development of the KSLV-II using currently allocated cost for successful development. It is anticipated that the present cost spread model can be applied to not only launch vehicle development but also other large complex system development.

Numerical study of liquid propellants combustion for space applications

The present study focuses on the combustion of liquid propellants used in engines for space propulsion applications (satellites and launch vehicles). The combustion of the following propellant combinations was studied: hydrogen/oxygen, hydrocarbon fuel/oxygen, hydrocarbon fuel/hydrogen peroxide, hydrazine/nitrogen tetroxide, Mono-methyl hydrazine/nitrogen tetroxide and Unsymmetrical dimethyl hydrazine/nitrogen tetroxide. The purpose of this paper is to determine the combustion flame temperature and the other thermochemical parameters (combustion products politropic parameter and molecular mass) as function of mixture ratio by mass of oxidizer to fuel (O/F). Furthermore, the vacuum specific impulse was also calculated at assumed pressure conditions of propulsion engines to show the effect of mixture ratio of the propellants on the performances. For the determination of the equilibrium composition at assumed temperature, nonlinear systems of equations were solved numerically using Lieberstein’s method. Some results were presented and compared with previous research in this area, the comparison shows good agreement between the results and the difference is less than 5%. As example, three cases of bi-propellant engines were studied using the programs of combustion and frozen fluid flow approximation: Space Shuttle Main Engine (SSME), Zenit second stage engine (RD-120) and Ariane-5 upper stage engine (Aestus).

Can molecular diffusion explain Space Shuttle plume spreading?

The satellite‐borne Global Ultraviolet Imager (GUVI) has produced more than 20 images of NASA Space Shuttle main engine plumes in the lower thermosphere. These reveal atomic hydrogen and, by inference, water vapor transport over hemispherical‐scale distances with speeds much faster than expected from models of thermospheric wind motions. Furthermore, the hydrogen plumes expand rapidly. We find rates that exceed the horizontal diffusion speed at nominal plume altitudes of 104‐112 km. Kelley et al. (2009) have proposed a 2‐D turbulence mechanism to explain the observed spreading rates (and rapid advection) of the plumes. But upon further investigation, we conclude that H atom diffusion can indeed account for the observed expansion rates by recognizing that vertical diffusion quickly conveys atoms to higher altitudes where horizontal diffusion is much more rapid. We also find evidence for H atom production directly during the Shuttle's main engine burn.

Unsupervised Anomaly Detection for Liquid-Fueled Rocket Propulsion Health Monitoring

This article describes the results of applying four unsupervised anomaly detection algorithms to data from two rocket propulsion testbeds.The first testbed uses historical data from the Space Shuttle Main Engine. The second testbed uses data from an experimental rocket engine test stand located at NASA Stennis Space Center. The article describes nine anomalies detected by the four algorithms.The four algorithms use four different definitions of anomalousness. Orca uses a nearest-neighbor approach, defining a point to be an anomaly if its nearest neighbors in the data space are far away from it. The Inductive Monitoring System clusters the training data, and then uses the distance to the nearest cluster as its measure of anomalousness. GritBot learns rules from the training data, and then classifies points as anomalous if they violate these rules. One-class support vector machines map the data into a high-dimensional space in which most of the normal points are on one side of a hyperplane, and then classify points on the other side of the hyperplane as anomalous. Because of these different definitions of anomalousness, different algorithms detect different anomalies. We therefore conclude that it is useful to use multiple algorithms.

Replica-based crack inspection

Surface replication has been proposed as a method for crack detection in space-shuttle main engine flowliner slots. The results of a feasibility study show that examination of surface replicas with a scanning electron microscope can result in the detection of cracks as small as 0.005 inch, and surface flaws as small as 0.001 inch, for the flowliner material.

Low friction and severe wear of alumina in cryogenic environment: A first report

Structural ceramics are considered as potential candidate materials for use in hybrid bearings in rocket turbopumps, operated under high stress in cryogenic environment. The friction and wear-related surface failure is considered as one of the critical factors in selecting the materials for cryo-turbopumps of Space Shuttle Main Engine (SSME). To obtain fundamental understanding of the tribological properties of ceramics in cryogenic environment, a very first set of sliding wear tests were carried out on self-mated Al2O3, a model brittle ceramic material, in liquid nitrogen (LN2) under varying load (2–10 N) and high rotational speed of 2550 rpm, using a newly designed cryogenic tribometer. The present research attempts to answer some important questions: (i) What would be the influence of LN2 on frictional and fracture behavior at sliding contacts? (ii) How does the material removal process occur in LN2 environment? Our experimental results reveal that self-mated alumina exhibits low steady-state coefficient of friction ∼0.13–0.18 and suffers from high wear rate (10−5 mm3/Nm) under the selected testing conditions. The novelty of the present work also lies in presenting some interesting results, for the first time, concerning the deformation and fracture of alumina at cryogenic temperature under high speed sliding conditions. Detailed scanning electronic microscope observation of the worn surfaces indicates that severe damage of both ball and flat occurs in cryogenic environment by transgranular and intergranular fracture. The observed wear behavior is explained in terms of thermal heat dissipation and brittle fracture of alumina in LN2.