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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 235 records · Page 13

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NASA NTRS2005

Synthetic Vortex Generator Jets Used to Control Separation on Low-Pressure Turbine Airfoils

Low-pressure turbine (LPT) airfoils are subject to increasingly stronger pressure gradients as designers impose higher loading in an effort to improve efficiency and lower cost by reducing the number of airfoils in an engine. When the adverse pressure gradient on the suction side of these airfoils becomes strong enough, the boundary layer will separate. Separation bubbles, particularly those that fail to reattach, can result in a significant loss of lift and a subsequent degradation of engine efficiency. The problem is particularly relevant in aircraft engines. Airfoils optimized to produce maximum power under takeoff conditions may still experience boundary layer separation at cruise conditions because of the thinner air and lower Reynolds numbers at altitude. Component efficiency can drop significantly between takeoff and cruise conditions. The decrease is about 2 percent in large commercial transport engines, and it could be as large as 7 percent in smaller engines operating at higher altitudes. Therefore, it is very beneficial to eliminate, or at least reduce, the separation bubble.

Ashpis, David E.↗

NASA NTRSDate not supplied

Combustor-Turbine Interactions By Using the Open National Combustion Code (OpenNCC) and the Glenn-HT Code

We develop the methodology of coupling two different computational fluid dynamic (CFD) codes, OpenNCC and Glenn-HT, in order to investigate the unsteady flow fields inside the combustor and around the first-stage stator of a high-pressure turbine (HPT) from the energy efficient engine (E 3 ) program. In our coupling strategy, OpenNCC, a time accurate unstructured mesh multi-phase combustion CFD code, is applied to the combustor region and Glenn-HT, a structured mesh CFD code designed for turbine applications, is utilized inside the HPT region. As a proof of concept, we first simulate the three-dimensional airflow over the backward facing step and compared the predicted pressure coefficient against experimental data. We then considered a Jet-A/Air combustor with the NASA single learn direct injector (LDI) as a reacting flow test case. Here, Glenn-HT is used in the downstream of the combustor and assumes that the working fluid is a single gas (i.e., air) where the properties of the gas are taken from time- and area-averaged values at the combustor exit. A fully coupled combustor-turbine simulation using both OpenNCC and Glenn-HT was conducted on the E 3 combustor and a detailed analysis done pertaining to how a realistic combustor outflow affects migration of hot-streaks and its impact on the aerodynamic behavior inside the HPT. It is found that the combustion dynamics (i.e., the switching main and pilot flame strength) significantly alters the aerodynamic behavior inside the HPT including hot-streak migration. Variations of the temperature and the velocity magnitude in the passage could be up to ±75 and ±15. These large variations are an important part of the combustor-turbine interactions, which are overlooked by a single component simulation of the HPT while imposing the steady inflow boundary condition at the HPT inlet. Finally, it is observed the strong density gradient associated with the hot-streaks and the pressure gradient at the passage significantly enhances the baroclinic torque, affecting the vorticity field.

CFD hot-streaks↗

DOE OSTI2020

Assessment of turbulent combustion models for simulating pre-chamber ignition in a natural gas engine

Lean combustion in an internal combustion engine is a promising strategy to increase thermal efficiency by leveraging a more favorable specific heat ratio of the fresh mixture and simultaneously suppressing the heat losses to the cylinder wall. However, unstable ignition events and slow flame propagation at fuel-lean condition lead to high cycle-to-cycle variability and hence limit the high-efficiency engine operating range. Pre-chamber ignition is considered an effective concept to extend the lean operating limit, by providing spatially distributed ignition with multiple turbulent flame-jets and enabling faster combustion rate compared to the conventional spark ignition approach. From a numerical modeling perspective, to date, still the science base and available simulation tools are inadequate for understanding and predicting the combustion processes in pre-chamber ignited engines. In this paper, conceptually different RANS combustion models widely adopted in the engine modeling community were used to simulate the ignition and combustion processes in a medium-duty natural gas engine with a pre-chamber spark-ignition system. A flamelet-based turbulent combustion model, i.e., G-equation, and a multi-zone well-stirred reactor model were employed for the multi-dimensional study. Simulation results were compared with experimental data in terms of in-cylinder pressure and heat release rate. Finally, the analysis of the performance of the two models is carried out to highlight the strengths and limitations of the two formulations respectively.

combustion model evaluation↗

NASA NTRS1982

NASA ECI programs - Benefits to Pratt and Whitney engines

Proprietary research conducted under the NASA-sponsored Engine Component Improvement program is shown to have advanced the state of such component technologies as thermal barrier coatings, ceramic seal systems, advanced turbine clearance control, and JT9D engine fan design. In addition, the tests conducted have demonstrated the advantages of updated cooling, sealing and aerodynamic designs in the high pressure turbine and compressor of the JT8D engine. These technological advances are being transferred to such novel engine configurations as the PW2037 and NASA Energy Efficient Engine. Attention is given to the results of the thermal coating refinement tests of convectively cooled vanes, and to the development of a single shroud fan design applicable to the JT9D and next-generation engines.

Gaffin, W. O.↗

NASA NTRS1982

Cruise Missile Engines

Williams International's F107 fanjet engine is used in two types of cruise missiles, Navy-sponsored Tomahawk and the Air Force AGM-86B Air Launched Cruise Missile (ALCM). Engine produces about 600 pounds thrust, is one foot in diameter and weighs only 141 pounds. Design was aided by use of a COSMIC program in calculating airflows in engine's internal ducting, resulting in a more efficient engine with increased thrust and reduced fuel consumption.

Source record↗

NASA NTRS2013

Implementation of Sub-Cooling of Cryogenic Propellants by Injection of Non-condensing Gas to the Generalized Fluid Systems Simulation Program (GFSSP)

Cryogenic propellants are readily heated when used. This poses a problem for rocket engine efficiency and effective boot-strapping of the engine, as seen in the "hot" LOX (Liquid Oxygen) problem on the S-1 stage of the Saturn vehicle. In order to remedy this issue, cryogenic fluids were found to be sub-cooled by injection of a warm non-condensing gas. Experimental results show that the mechanism behind the sub-cooling is evaporative cooling. It has been shown that a sub-cooled temperature difference of approximately 13 deg F below saturation temperature [1]. The phenomenon of sub-cooling of cryogenic propellants by a non-condensing gas is not readily available with the General Fluid System Simulation Program (GFSSP) [2]. GFSSP is a thermal-fluid program used to analyze a wide variety of systems that are directly impacted by thermodynamics and fluid mechanics. In order to model this phenomenon, additional capabilities had to be added to GFSSP in the form of a FORTRAN coded sub-routine to calculate the temperature of the sub-cooled fluid. Once this was accomplished, the sub-routine was implemented to a GFSSP model that was created to replicate an experiment that was conducted to validate the GFSSP results.

Huggett, Daniel J.↗

DOE OSTI2022

Optimizing the fuel efficiency of an opposed piston engine for electric power generation

This paper investigates the optimal crankshaft motion for an opposed piston (OP) engine in a novel hybrid architecture to maximize fuel efficiency. The OP engine was selected for this work due to its inherent thermodynamic benefits and the balanced nature of the engine which can achieve downsizing through reducing the number of cylinders rather than the individual cylinder volume. The typical geartrain required on an OP engine was exchanged for two electric motors, reducing friction loss and decoupling the crankshafts. Using the motors to control the crankshaft motion profiles, this architecture introduces capabilities to dynamically vary compression ratio, combustion volume, and scavenging dynamics. To leverage these opportunities, an optimization scheme was developed utilizing nonlinear optimization of a 0-D model to compute the crankshaft motion profile that maximizes the work generated by the system. This optimization was then iteratively coupled with a high fidelity model which supplies the cylinder flow boundary conditions. This iterative approach reduces the model complexity used in the optimal control problem (OCP) while capturing the gas exchange dynamics critical to the 2-stroke cycle of the OP engine. By using the rate of change of motor torque as the input to the OCP, the torque fluctuation in a single cycle can be limited to ensure tracking feasibility. The results show crankshaft velocity slows during the compression stroke and conversely accelerates during the expansion stroke, reducing the peak motor torque required for control and thus reducing the motor losses. The extended residence time at top dead center, however, leads to an increase in heat transfer, illustrating the trade-off between the work extraction efficiency and the indicated engine efficiency.

Engineering↗

NASA NTRS1982

Improved Stirling engine performance using jet impingement

Of the many factors influencing the performance of a Stirling engine, that of transferring the combustion gas heat into the working fluid is crucial. By utilizing the high heat transfer rates obtainable with a jet impingement heat transfer system, it is possible to reduce the flame temperature required for engine operation. Also, the required amount of heater tube surface area may be reduced, resulting in a decrease in the engine nonswept volume and a related increase in engine efficiency. A jet impingement heat transfer system was designed by Rasor Associates, Inc., and tested in the GPU-3 Stirling engine at the NASA Lewis Research Center. For a small penalty in pumping power (less than 0.5% of engine output) the jet impingement heat transfer system provided a higher combustion-gas-side heat transfer coefficient and a smoothing of heater temperature profiles resulting in lower combustion system temperatures and a 5 to 8% increase in engine power output and efficiency.

Johnson, D. C.↗

NASA NTRS2009

Analysis of Turbofan Design Options for an Advanced Single-Aisle Transport Aircraft

The desire for higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets, to low bypass ratio, first generation turbofans, to today's high bypass ratio turbofans. It is possible that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines. Although increased bypass ratio has clear benefits in terms of propulsion system metrics such as specific fuel consumption, these benefits may not translate into aircraft system level benefits due to integration penalties. In this study, the design trade space for advanced turbofan engines applied to a single-aisle transport (737/A320 class aircraft) is explored. The benefits of increased bypass ratio and associated enabling technologies such as geared fan drive are found to depend on the primary metrics of interest. For example, bypass ratios at which fuel consumption is minimized may not require geared fan technology. However, geared fan drive does enable higher bypass ratio designs which result in lower noise. Regardless of the engine architecture chosen, the results of this study indicate the potential for the advanced aircraft to realize substantial improvements in fuel efficiency, emissions, and noise compared to the current vehicles in this size class.

Guynn, Mark D.↗

DOE OSTI2022

Achieving complete electrooxidation of ethanol by single atomic Rh decoration of Pt nanocubes

Significance Direct ethanol fuel cells are attracting growing attention as portable power sources due to their advantages such as higher mass-energy density than hydrogen and less toxicity than methanol. However, it is challenging to achieve the complete electrooxidation to generate 12 electrons per ethanol, resulting in a low fuel utilization efficiency. This manuscript reports the complete ethanol electrooxidation by engineering efficient catalysts via single-atom modification. The combined electrochemical measurements, in situ characterization, and density functional theory calculations unravel synergistic effects of single Rh atoms and Pt nanocubes and identify reaction pathways leading to the selective C–C bond cleavage to oxidize ethanol to CO 2 . This study provides a unique single-atom approach to tune the activity and selectivity toward complicated electrocatalytic reactions.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗

NASA NTRS2006

Solid Oxide Fuel Cell APU Feasibility Study for a Long Range Commercial Aircraft Using UTC ITAPS Approach: Aircraft Propulsion and Subsystems Integration Evaluation - Volume 1

The objective of this contract effort was to define the functionality and evaluate the propulsion and power system benefits derived from a Solid Oxide Fuel Cell (SOFC) based Auxiliary Power Unit (APU) for a future long range commercial aircraft, and to define the technology gaps to enable such a system. The study employed technologies commensurate with Entry into Service (EIS) in 2015. United Technologies Corporation (UTC) Integrated Total Aircraft Power System (ITAPS) methodologies were used to evaluate system concepts to a conceptual level of fidelity. The technology benefits were captured as reductions of the mission fuel burn and emissions. The baseline aircraft considered was the Boeing 777-200ER airframe with more electric subsystems, Ultra Efficient Engine Technology (UEET) engines, and an advanced APU with ceramics for increased efficiency. In addition to the baseline architecture, four architectures using an SOFC system to replace the conventional APU were investigated. The mission fuel burn savings for Architecture-A, which has minimal system integration, is 0.16 percent. Architecture-B and Architecture-C employ greater system integration and obtain fuel burn benefits of 0.44 and 0.70 percent, respectively. Architecture-D represents the highest level of integration and obtains a benefit of 0.77 percent.

Srinivasan, Hari↗

NASA NTRS2016

Dynamic Systems Analysis for Turbine Based Aero Propulsion Systems

The aircraft engine design process seeks to optimize the overall system-level performance, weight, and cost for a given concept. Steady-state simulations and data are used to identify trade-offs that should be balanced to optimize the system in a process known as systems analysis. These systems analysis simulations and data may not adequately capture the true performance trade-offs that exist during transient operation. Dynamic systems analysis provides the capability for assessing the dynamic tradeoffs at an earlier stage of the engine design process. The dynamic systems analysis concept, developed tools, and potential benefit are presented in this paper. To provide this capability, the Tool for Turbine Engine Closed-loop Transient Analysis (TTECTrA) was developed to provide the user with an estimate of the closed-loop performance (response time) and operability (high pressure compressor surge margin) for a given engine design and set of control design requirements. TTECTrA along with engine deterioration information, can be used to develop a more generic relationship between performance and operability that can impact the engine design constraints and potentially lead to a more efficient engine.

dynamic analysis↗

NASA NTRSDate not supplied

Upgrades to the Single Stage Axial Compressor and Fan Facility For Low Pressure Ratio and Boundary Layer Ingesting Fan Research

Next generation aircraft engine efficiency goals will be met through high bypass ratio engines. To achieve these goals, a combination of small core and low pressure ratio fan technologies will be necessary. Additionally, advanced propulsion air-frame integration concepts, such as boundary layer propulsion, and distributed electrically driven fans with short inlets or wing integration strategies, are beginning to be investigated for their potential fuel burn savings. In order to realize the potential fuel savings from these technologies, it is necessary to have testing capabilities for the components. The NASA Glenn Research Center (GRC) Single StageAxial Fan and Compressor Facility (also called the GRC W-8Facility) has recently undergone several significant upgrades in order to effectively test low pressure ratio fans and various inlet distortions. This paper will outline various challenges to testing low pressure ratio fans and the methods implemented to ensure that the W-8 facility is capable of performing these tests and acquiring high quality data. The upgrades include significant pip-ing upgrades to reduce pressure loss and turbulence, capability for inlet total pressure and swirl distortion, increased test section length including a longer shaft, a two component balance, an up-graded rotating data system, and a new temperature characterization to reduce uncertainties in thermocouple measurements

W-8 Facility↗

NASA NTRSDate not supplied

Upgrades to the Single Stage Axial Compressor and Fan Facility For Low Pressure Ratio and Boundary Layer Ingesting Fan Research

Next generation aircraft engine efficiency goals will be met through high bypass ratio engines. To achieve these goals, a combination of small core and low pressure ratio fan technologies will be necessary. Additionally, advanced propulsion airframe integration concepts, such as boundary layer propulsion, and distributed electrically driven fans with short inlets or wing integration strategies, are beginning to be investigated for their potential fuel burn savings. In order to realize the potential fuel savings from these technologies, it is necessary to have testing capabilities for the components. The NASA Glenn Research Center (GRC) Single Stage Axial Fan and Compressor Facility (also called the GRC W-8 Facility) has recently undergone several significant upgrades in order to effectively test low pressure ratio fans and various inlet distortions. This paper will outline various challenges to testing low pressure ratio fans and the methods implemented to ensure that the W-8 facility is capable of performing these tests and acquiring high quality data. The upgrades include significant piping upgrades to reduce pressure loss and turbulence, capability for inlet total pressure and swirl distortion, increased test section length including a longer shaft, a two component balance, an upgraded rotating data system, and a new temperature characterization to reduce uncertainties in thermocouple measurements.

W-8 facitlity↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E^3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P(41) = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E^3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P(41) = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRSDate not supplied

Numerical Study of Combustor-Turbine Interactions Considering a Two-Stage High-Pressure Turbine

In this study, preliminary results are presented from the Open National Combustion Code (OpenNCC) applied to the two-stage high-pressure turbine (HTP) from the energy efficient engine (E3) program. Although the combustor is designed for efficient fuel–air mixing and a relatively uniform exit temperature profile, temporal and spatial variation in combustor exit conditions inevitably lead to hot-streaks in the HPT. These hot streaks can cause local hot spots on the HPT blade surfaces that reduce blade life. To improve understanding of these effects and the overall design of combustor and HPT, we are developing a fully-coupled combustor-turbine computational fluid dynamics (CFD) capability. Previous work has simulated the E3 combustor at the simulated Sea-Level Take-off (SLTO) condition (P41 = 2.4 [atm]) with and without the first-stage HPT stator included. This work addresses HPT simulations using inflow conditions based on previous simulation results of the E3 combustor, as well as investigating the effects of spatial and temporal temperature non-uniformity of the combustor-exit flow on the HPT performance for a more realistic SLTO (P41 = 27.4 [atm])). First, we analyze existing results from the combustor and the HPT 1st-stage stator simulation, where the unsteady hot gas coming out of the combustor exit travels between the stationary stators and then directly exits the computational domain (i.e., no effect from the rotors). Second, we simulate the HPT including the 1st and 2nd stage stators and rotors with the time-averaged, spatially-nonuniform inflow to the HPT taken from the combustor (with HPT 1st-stage stators) solution. Differences in simulations results for the flow field in the 1st-stage stator region are presented. The results of these two simulations (combustor and HPT simulated in a sequential manner) will be compared to a fully-coupled EEE combustor/HPT simulation to understand the impacts of fully-coupled simulations on combustor and HPT performance for two different operating conditions.

LES↗

NASA NTRSDate not supplied

Computational Study of Modeling Fully-Coupled Combustor-Turbine Interactions by the Open National Combustion Code (OpenNCC)

LES↗

NASA NTRS2004

Strategy Developed for Selecting Optimal Sensors for Monitoring Engine Health

Sensor indications during rocket engine operation are the primary means of assessing engine performance and health. Effective selection and location of sensors in the operating engine environment enables accurate real-time condition monitoring and rapid engine controller response to mitigate critical fault conditions. These capabilities are crucial to ensure crew safety and mission success. Effective sensor selection also facilitates postflight condition assessment, which contributes to efficient engine maintenance and reduced operating costs. Under the Next Generation Launch Technology program, the NASA Glenn Research Center, in partnership with Rocketdyne Propulsion and Power, has developed a model-based procedure for systematically selecting an optimal sensor suite for assessing rocket engine system health. This optimization process is termed the systematic sensor selection strategy. Engine health management (EHM) systems generally employ multiple diagnostic procedures including data validation, anomaly detection, fault-isolation, and information fusion. The effectiveness of each diagnostic component is affected by the quality, availability, and compatibility of sensor data. Therefore systematic sensor selection is an enabling technology for EHM. Information in three categories is required by the systematic sensor selection strategy. The first category consists of targeted engine fault information; including the description and estimated risk-reduction factor for each identified fault. Risk-reduction factors are used to define and rank the potential merit of timely fault diagnoses. The second category is composed of candidate sensor information; including type, location, and estimated variance in normal operation. The final category includes the definition of fault scenarios characteristic of each targeted engine fault. These scenarios are defined in terms of engine model hardware parameters. Values of these parameters define engine simulations that generate expected sensor values for targeted fault scenarios. Taken together, this information provides an efficient condensation of the engineering experience and engine flow physics needed for sensor selection. The systematic sensor selection strategy is composed of three primary algorithms. The core of the selection process is a genetic algorithm that iteratively improves a defined quality measure of selected sensor suites. A merit algorithm is employed to compute the quality measure for each test sensor suite presented by the selection process. The quality measure is based on the fidelity of fault detection and the level of fault source discrimination provided by the test sensor suite. An inverse engine model, whose function is to derive hardware performance parameters from sensor data, is an integral part of the merit algorithm. The final component is a statistical evaluation algorithm that characterizes the impact of interference effects, such as control-induced sensor variation and sensor noise, on the probability of fault detection and isolation for optimal and near-optimal sensor suites.

Source record↗