Radiation shielding estimates for manned Mars space flight.

PubMed

Dudkin, V E; Kovalev, E E; Kolomensky, A V; Sakovich, V A; Semenov, V F; Demin, V P; Benton, E V

1992-01-01

In the analysis of the required radiation shielding protection of spacecraft during a Mars flight, specific effects of solar activity (SA) on the intensity of galactic and solar cosmic rays were taken into consideration. Three spaceflight periods were considered: (1) maximum SA; (2) minimum SA; and (3) intermediate SA, when intensities of both galactic and solar cosmic rays are moderately high. Scenarios of spaceflights utilizing liquid-propellant rocket engines, low- and intermediate-thrust nuclear electrojet engines, and nuclear rocket engines, all of which have been designed in the Soviet Union, are reviewed. Calculations were performed on the basis of a set of standards for radiation protection approved by the U.S.S.R. State Committee for Standards. It was found that the lowest estimated mass of a Mars spacecraft, including the radiation shielding mass, obtained using a combination of a liquid propellant engine with low and intermediate thrust nuclear electrojet engines, would be 500-550 metric tons.

DC-9 Flight Demonstration Program with Refanned JT8D Engines. Volume 3; Performance and Analysis

NASA Technical Reports Server (NTRS)

1975-01-01

The JT8D-109 engine has a sea level static, standard day bare engine takeoff thrust of 73,840 N. At sea level standard day conditions the additional thrust of the JT8D-109 results in 2,040 kg additional takeoff gross weight capability for a given field length. Range loss of the DC-9 Refan airplane for long range cruise was determined. The Refan airplane demonstrated stall, static longitudinal stability, longitudinal control, longitudinal trim, minimum control speeds, and directional control characteristics similar to the DC-9-30 production airplane and complied with airworthiness requirements. Cruise, climb, and thrust reverser performance were evaluated. Structural and dynamic ground test, flight test and analytical results substantiate Refan Program requirements that the nacelle, thrust reverser hardware, and the airplane structural modifications are flightworthy and certifiable and that the airplane meets flutter speed margins. Estimated unit cost of a DC-9 Refan retrofit program is 1.338 million in mid-1975 dollars with about an equal split in cost between airframe and engine.

Subsonic flight test evaluation of a propulsion system parameter estimation process for the F100 engine

NASA Technical Reports Server (NTRS)

Orme, John S.; Gilyard, Glenn B.

1992-01-01

Integrated engine-airframe optimal control technology may significantly improve aircraft performance. This technology requires a reliable and accurate parameter estimator to predict unmeasured variables. To develop this technology base, NASA Dryden Flight Research Facility (Edwards, CA), McDonnell Aircraft Company (St. Louis, MO), and Pratt & Whitney (West Palm Beach, FL) have developed and flight-tested an adaptive performance seeking control system which optimizes the quasi-steady-state performance of the F-15 propulsion system. This paper presents flight and ground test evaluations of the propulsion system parameter estimation process used by the performance seeking control system. The estimator consists of a compact propulsion system model and an extended Kalman filter. The extended Laman filter estimates five engine component deviation parameters from measured inputs. The compact model uses measurements and Kalman-filter estimates as inputs to predict unmeasured propulsion parameters such as net propulsive force and fan stall margin. The ability to track trends and estimate absolute values of propulsion system parameters was demonstrated. For example, thrust stand results show a good correlation, especially in trends, between the performance seeking control estimated and measured thrust.

Weight and cost estimating relationships for heavy lift airships

NASA Technical Reports Server (NTRS)

Gray, D. W.

1979-01-01

Weight and cost estimating relationships, including additional parameters that influence the cost and performance of heavy-lift airships (HLA), are discussed. Inputs to a closed loop computer program, consisting of useful load, forward speed, lift module positive or negative thrust, and rotors and propellers, are examined. Detail is given to the HLA cost and weight program (HLACW), which computes component weights, vehicle size, buoyancy lift, rotor and propellar thrust, and engine horse power. This program solves the problem of interrelating the different aerostat, rotors, engines and propeller sizes. Six sets of 'default parameters' are left for the operator to change during each computer run enabling slight data manipulation without altering the program.

Research on performance requirements of turbofan engine used on carrier-based UAV

NASA Astrophysics Data System (ADS)

Zhao, Shufan; Li, Benwei; Zhang, Wenlong; Wu, Heng; Feng, Tang

2017-05-01

According to the mission requirements of the carrier-based unmanned aerial vehicle (UAV), a mode level flight was established to calculate the thrust requirements from altitude 9 km to 13 km. Then, the estimation method of flight profile was used to calculate the weight of UAV in each stage to get the specific fuel consumption requirements of the UAV in standby stage. The turbofan engine of carrier-based UAV should meet the thrust and specific fuel consumption requirements. Finally, the GSP software was used to verify the simulation of a small high-bypass turbofan engine. The conclusion is useful for the turbofan engine selection of carrier-based UAV.

Adaptive Gas Turbine Engine Control for Deterioration Compensation Due to Aging

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Parker, Khary I.; Chatterjee, Santanu

2003-01-01

This paper presents an ad hoc adaptive, multivariable controller tuning rule that compensates for a thrust response variation in an engine whose performance has been degraded though use and wear. The upset appears when a large throttle transient is performed such that the engine controller switches from low-speed to high-speed mode. A relationship was observed between the level of engine degradation and the overshoot in engine temperature ratio, which was determined to cause the thrust response variation. This relationship was used to adapt the controller. The method is shown to work very well up to the operability limits of the engine. Additionally, since the level of degradation can be estimated from sensor data, it would be feasible to implement the adaptive control algorithm on-line.

Metaheuristic and Machine Learning Models for TFE-731-2, PW4056, and JT8D-9 Cruise Thrust

NASA Astrophysics Data System (ADS)

Baklacioglu, Tolga

2017-08-01

The requirement for an accurate engine thrust model has a major antecedence in airline fuel saving programs, assessment of environmental effects of fuel consumption, emissions reduction studies, and air traffic management applications. In this study, utilizing engine manufacturers' real data, a metaheuristic model based on genetic algorithms (GAs) and a machine learning model based on neural networks (NNs) trained with Levenberg-Marquardt (LM), delta-bar-delta (DBD), and conjugate gradient (CG) algorithms were accomplished to incorporate the effect of both flight altitude and Mach number in the estimation of thrust. For the GA model, the analysis of population size impact on the model's accuracy and effect of number of data on model coefficients were also performed. For the NN model, design of optimum topology was searched for one- and two-hidden-layer networks. Predicted thrust values presented a close agreement with real thrust data for both models, among which LM trained NNs gave the best accuracies.

IR signature study of aircraft engine for variation in nozzle exit area

NASA Astrophysics Data System (ADS)

Baranwal, Nidhi; Mahulikar, Shripad P.

2016-01-01

In general, jet engines operate with choked nozzle during take-off, climb and cruise, whereas unchoking occurs while landing and taxiing (when engine is not running at full power). Appropriate thrust in an aircraft in all stages of the flight, i.e., take-off, climb, cruise, descent and landing is achieved through variation in the nozzle exit area. This paper describes the effect on thrust and IR radiance of a turbojet engine due to variation in the exit area of a just choked converging nozzle (Me = 1). The variations in the nozzle exit area result in either choking or unchoking of a just choked converging nozzle. Results for the change in nozzle exit area are analyzed in terms of thrust, mass flow rate and specific fuel consumption. The solid angle subtended (Ω) by the exhaust system is estimated analytically, for the variation in nozzle exit area (Ane), as it affects the visibility of the hot engine parts from the rear aspect. For constant design point thrust, IR radiance is studied from the boresight (ϕ = 0°, directly from the rear side) for various percentage changes in nozzle exit area (%ΔAne), in the 1.9-2.9 μm and 3-5 μm bands.

A simplified gross thrust computing technique for an afterburning turbofan engine

NASA Technical Reports Server (NTRS)

Hamer, M. J.; Kurtenbach, F. J.

1978-01-01

A simplified gross thrust computing technique extended to the F100-PW-100 afterburning turbofan engine is described. The technique uses measured total and static pressures in the engine tailpipe and ambient static pressure to compute gross thrust. Empirically evaluated calibration factors account for three-dimensional effects, the effects of friction and mass transfer, and the effects of simplifying assumptions for solving the equations. Instrumentation requirements and the sensitivity of computed thrust to transducer errors are presented. NASA altitude facility tests on F100 engines (computed thrust versus measured thrust) are presented, and calibration factors obtained on one engine are shown to be applicable to the second engine by comparing the computed gross thrust. It is concluded that this thrust method is potentially suitable for flight test application and engine maintenance on production engines with a minimum amount of instrumentation.

Energy efficient engine: Propulsion system-aircraft integration evaluation

NASA Technical Reports Server (NTRS)

Owens, R. E.

1979-01-01

Flight performance and operating economics of future commercial transports utilizing the energy efficient engine were assessed as well as the probability of meeting NASA's goals for TSFC, DOC, noise, and emissions. Results of the initial propulsion systems aircraft integration evaluation presented include estimates of engine performance, predictions of fuel burns, operating costs of the flight propulsion system installed in seven selected advanced study commercial transports, estimates of noise and emissions, considerations of thrust growth, and the achievement-probability analysis.

Evaluation of various thrust calculation techniques on an F404 engine

NASA Technical Reports Server (NTRS)

Ray, Ronald J.

1990-01-01

In support of performance testing of the X-29A aircraft at the NASA-Ames, various thrust calculation techniques were developed and evaluated for use on the F404-GE-400 engine. The engine was thrust calibrated at NASA-Lewis. Results from these tests were used to correct the manufacturer's in-flight thrust program to more accurately calculate thrust for the specific test engine. Data from these tests were also used to develop an independent, simplified thrust calculation technique for real-time thrust calculation. Comparisons were also made to thrust values predicted by the engine specification model. Results indicate uninstalled gross thrust accuracies on the order of 1 to 4 percent for the various in-flight thrust methods. The various thrust calculations are described and their usage, uncertainty, and measured accuracies are explained. In addition, the advantages of a real-time thrust algorithm for flight test use and the importance of an accurate thrust calculation to the aircraft performance analysis are described. Finally, actual data obtained from flight test are presented.

Deimos Methane-Oxygen Rocket Engine Test Results

NASA Astrophysics Data System (ADS)

Engelen, S.; Souverein, L. J.; Twigt, D. J.

This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

Architecture-Led Safety Process

DTIC Science & Technology

2016-12-01

Action Hazard Guide 42 Table 18: Comparative Table of Safety and Reliability Terms 47 CMU/SEI-2016-TR-012 | SOFTWARE ENGINEERING INSTITUTE...provides too much thrust Engine is slow to pro- vide commanded thrust (increase or de- crease) Engine will not shut- down when com - manded...Thrust level must be provided at the com - manded level H4: Engine is slow to provide commanded thrust SC3: Engine must provide commanded thrust in

Characterization of Gas-Phase Organics Using Proton Transfer Reaction Time-of-Flight Mass Spectrometry: Aircraft Turbine Engines.

PubMed

Kilic, Dogushan; Brem, Benjamin T; Klein, Felix; El-Haddad, Imad; Durdina, Lukas; Rindlisbacher, Theo; Setyan, Ari; Huang, Rujin; Wang, Jing; Slowik, Jay G; Baltensperger, Urs; Prevot, Andre S H

2017-04-04

Nonmethane organic gas emissions (NMOGs) from in-service aircraft turbine engines were investigated using a proton transfer reaction time-of-flight mass spectrometer (PTR-ToF-MS) at an engine test facility at Zurich Airport, Switzerland. Experiments consisted of 60 exhaust samples for seven engine types (used in commercial aviation) from two manufacturers at thrust levels ranging from idle to takeoff. Emission indices (EIs) for more than 200 NMOGs were quantified, and the functional group fractions (including acids, carbonyls, aromatics, and aliphatics) were calculated to characterize the exhaust chemical composition at different engine operation modes. Total NMOG emissions were highest at idling with an average EI of 7.8 g/kg fuel and were a factor of ∼40 lower at takeoff thrust. The relative contribution of pure hydrocarbons (particularly aromatics and aliphatics) of the engine exhaust decreased with increasing thrust while the fraction of oxidized compounds, for example, acids and carbonyls increased. Exhaust chemical composition at idle was also affected by engine technology. Older engines emitted a higher fraction of nonoxidized NMOGs compared to newer ones. Idling conditions dominated ground level organic gas emissions. Based on the EI determined here, we estimate that reducing idle emissions could substantially improve air quality near airports.

14 CFR 25.934 - Turbojet engine thrust reverser system tests.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbojet engine thrust reverser system... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant General § 25.934 Turbojet engine thrust reverser system tests. Thrust reversers installed on turbojet engines must meet the...

Studies of Operating Frequency Effects On Ejector-based Thrust Augmentation in a Pulse Detonation Engine

NASA Technical Reports Server (NTRS)

Landry, K.

2005-01-01

Studies were performed in order to characterize the thrust augmentation potential of an ejector in a Pulse Detonation Engine application. A 49-mm diameter tube of 0.914-m length was constructed with one open end and one closed end. Ethylene, oxygen, and nitrogen were introduced into the tube at the closed end through the implementation of a fast mixing injector. The tube was completely filled with a stoichiometric mixture containing a one to one molar ratio of nitrogen to oxygen. Ethylene was selected as the fuel due to its detonation sensitivity and the molar ratio of the oxidizer was chosen for heat transfer purposes. Detonations were initiated in the tube through the use of a spark ignition system. The PDE was operated in a multi-cycle mode at frequencies ranging from 20-Hz to 50-Hz. Baseline thrust measurements with no ejector present were performed while operating the engine at various frequencies and compared to theoretical estimates. The baseline values were observed to agree with the theoretical model at low operating frequencies and proved to be increasingly lower than the predicted values as the operating frequency was increased. The baseline thrust measurements were observed to agree within 15 percent of the model for all operating frequencies. A straight 152-mm diameter ejector was installed and thrust augmentation percentages were measured. The length of the ejector was varied while the overlap percentage (percent of the ejector length which overlapped the tube) was maintained at 25 percent for all tests. In addition, the effect of ejector inlet geometry was investigated by comparing results with a straight inlet to those of a 38-mm inlet diameter. The thrust augmentation of the straight inlet ejector proved to be independent of engine operating frequency, augmenting thrust by 40 percent for the 0.914-m length ejector. In contrast, the rounded lip ejector of the same length seemed to be highly dependent on the engine operating frequency. An optimum operating frequency observed with the rounded inlet occurred at an operating frequency of 30-Hz, resulting in thrust augmentation percentages greater than 100 percent. The effect that the engine operating frequency had on thrust augmentation levels attained with an ejector was characterized and optimum performance parameters were established. Insight into the frequency dependent nature of the ejector performance was pursued. Suggestions for future experiments which are needed to fully understand the means in which thrust augmentation is achieved in a PDE-ejector configuration were noted.

Predicted Performance of a Thrust-Enhanced SR-71 Aircraft with an External Payload

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1997-01-01

NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime. To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement. This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.

14 CFR 23.934 - Turbojet and turbofan engine thrust reverser systems tests.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbojet and turbofan engine thrust... CATEGORY AIRPLANES Powerplant General § 23.934 Turbojet and turbofan engine thrust reverser systems tests. Thrust reverser systems of turbojet or turbofan engines must meet the requirements of § 33.97 of this...

Oxygen/Alcohol Dual Thrust RCS Engines

NASA Technical Reports Server (NTRS)

Angstadt, Tara; Hurlbert, Eric

1999-01-01

A non-toxic dual thrust RCS engine offers significant operational, safety, and performance advantages to the space shuttle and the next generation RLVs. In this concept, a single engine produces two thrust levels of 25 and 870 lbf. The low thrust level is provided by the spark torch igniter, which, with the addition of 2 extra valves, can also be made to function as a vernier. A dual thrust RCS engine allows 38 verniers to be packaged more efficiently on a vehicle. These 38 vemiers improve translation and reduce cross coupling, thereby providing more pure roll, pitch, and yaw maneuvers of the vehicle. Compared to the 6 vemiers currently on the shuttle, the 38 dual thrust engines would be 25 to 40% more efficient for the same maneuvers and attitude control. The vernier thrust level also reduces plume impingement and contamination concerns. Redundancy is also improved, thereby improving mission success reliability. Oxygen and ethanol are benign propellants which do not create explosive reaction products or contamination, as compared to hypergolic propellants. These characteristics make dual-thrust engines simpler to implement on a non-toxic reaction control system. Tests at WSTF in August 1999 demonstrated a dual-thrust concept that is successful with oxygen and ethanol. Over a variety of inlet pressures and mixture ratios at 22:1 area ratio, the engine produced between 230 and 297 sec Isp, and thrust levels from 8 lbf. to 50 lbf. This paper describes the benefits of dual-thrust engines and the recent results from tests at WSTF.

Developing stochastic model of thrust and flight dynamics for small UAVs

NASA Astrophysics Data System (ADS)

Tjhai, Chandra

This thesis presents a stochastic thrust model and aerodynamic model for small propeller driven UAVs whose power plant is a small electric motor. First a model which relates thrust generated by a small propeller driven electric motor as a function of throttle setting and commanded engine RPM is developed. A perturbation of this model is then used to relate the uncertainty in throttle and engine RPM commanded to the error in the predicted thrust. Such a stochastic model is indispensable in the design of state estimation and control systems for UAVs where the performance requirements of the systems are specied in stochastic terms. It is shown that thrust prediction models for small UAVs are not a simple, explicit functions relating throttle input and RPM command to thrust generated. Rather they are non-linear, iterative procedures which depend on a geometric description of the propeller and mathematical model of the motor. A detailed derivation of the iterative procedure is presented and the impact of errors which arise from inaccurate propeller and motor descriptions are discussed. Validation results from a series of wind tunnel tests are presented. The results show a favorable statistical agreement between the thrust uncertainty predicted by the model and the errors measured in the wind tunnel. The uncertainty model of aircraft aerodynamic coefficients developed based on wind tunnel experiment will be discussed at the end of this thesis.

Investigation of Thrust Augmentation of a 1600-pound Thrust Centrifugal-flow-type Turbojet Engine by Injection of Refrigerants at Compressor Inlets

NASA Technical Reports Server (NTRS)

Jones, William L.; Dowman, Harry W.

1947-01-01

Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.

Advanced space engine preliminary design

NASA Technical Reports Server (NTRS)

Cuffe, J. P. B.; Bradie, R. E.

1973-01-01

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

14 CFR 25.904 - Automatic takeoff thrust control system (ATTCS).

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Automatic takeoff thrust control system... Automatic takeoff thrust control system (ATTCS). Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when...

14 CFR 25.904 - Automatic takeoff thrust control system (ATTCS).

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Automatic takeoff thrust control system... Automatic takeoff thrust control system (ATTCS). Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when...

A Retro-Fit Control Architecture to Maintain Engine Performance With Usage

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Sowers, T. Shane; Garg, Sanjay

2007-01-01

An outer loop retrofit engine control architecture is presented which modifies fan speed command to obtain a desired thrust based on throttle position. This maintains the throttle-to-thrust relationship in the presence of engine degradation, which has the effect of changing the engine s thrust output for a given fan speed. Such an approach can minimize thrust asymmetry in multi-engine aircraft, and reduce pilot workload. The outer loop control is demonstrated under various levels of engine deterioration using a standard deterioration profile as well as an atypical profile. It is evaluated across various transients covering a wide operating range. The modified fan speed command still utilizes the standard engine control logic so all original life and operability limits remain in place. In all cases it is shown that with the outer loop thrust control in place, the deteriorated engine is able to match the thrust performance of a new engine up to the limits the controller will allow.

An algorithm for targeting finite burn maneuvers

NASA Technical Reports Server (NTRS)

Barbieri, R. W.; Wyatt, G. H.

1972-01-01

An algorithm was developed to solve the following problem: given the characteristics of the engine to be used to make a finite burn maneuver and given the desired orbit, when must the engine be ignited and what must be the orientation of the thrust vector so as to obtain the desired orbit? The desired orbit is characterized by classical elements and functions of these elements whereas the control parameters are characterized by the time to initiate the maneuver and three direction cosines which locate the thrust vector. The algorithm was built with a Monte Carlo capability whereby samples are taken from the distribution of errors associated with the estimate of the state and from the distribution of errors associated with the engine to be used to make the maneuver.

Thrust and Propulsive Efficiency from an Instructive Viewpoint

ERIC Educational Resources Information Center

Kaufman, Richard D.

2010-01-01

In a typical engineering or physics curriculum, the momentum equation is used for the determination of jet engine thrust. Even a simple thrust analysis requires a heavy emphasis on mathematics that can cause students and engineers to lose a physical perspective on thrust. This article provides for this physical understanding using only static…

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1981-01-01

An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

Engine Cycle Analysis of Air Breathing Microwave Rocket with Reed Valves

DOE Office of Scientific and Technical Information (OSTI.GOV)

Fukunari, Masafumi; Komatsu, Reiji; Yamaguchi, Toshikazu

The Microwave Rocket is a candidate for a low cost launcher system. Pulsed plasma generated by a high power millimeter wave beam drives a blast wave, and a vehicle acquires impulsive thrust by exhausting the blast wave. The thrust generation process of the Microwave Rocket is similar to a pulse detonation engine. In order to enhance the performance of its air refreshment, the air-breathing mechanism using reed valves is under development. Ambient air is taken to the thruster through reed valves. Reed valves are closed while the inside pressure is high enough. After the time when the shock wave exhaustsmore » at the open end, an expansion wave is driven and propagates to the thrust-wall. The reed valve is opened by the negative gauge pressure induced by the expansion wave and its reflection wave. In these processes, the pressure oscillation is important parameter. In this paper, the pressure oscillation in the thruster was calculated by CFD combined with the flux through from reed valves, which is estimated analytically. As a result, the air-breathing performance is evaluated using Partial Filling Rate (PFR), the ratio of thruster length to diameter L/D, and ratio of opening area of reed valves to superficial area {alpha}. An engine cycle and predicted thrust was explained.« less

A study of variable thrust, variable specific impulse trajectories for solar system exploration

NASA Astrophysics Data System (ADS)

Sakai, Tadashi

A study has been performed to determine the advantages and disadvantages of variable thrust and variable Isp (specific impulse) trajectories for solar system exploration. There have been several numerical research efforts for variable thrust, variable Isp, power-limited trajectory optimization problems. All of these results conclude that variable thrust, variable Isp (variable specific impulse, or VSI) engines are superior to constant thrust, constant Isp (constant specific impulse; or CSI) engines. However, most of these research efforts assume a mission from Earth to Mars, and some of them further assume that these planets are circular and coplanar. Hence they still lack the generality. This research has been conducted to answer the following questions: (1) Is a VSI engine always better than a CSI engine or a high thrust engine for any mission to any planet with any time of flight considering lower propellant mass as the sole criterion? (2) If a planetary swing-by is used for a VSI trajectory, is the fuel savings of a VSI swing-by trajectory better than that of a CSI swing-by or high thrust swing-by trajectory? To support this research, an unique, new computer-based interplanetary trajectory calculation program has been created. This program utilizes a calculus of variations algorithm to perform overall optimization of thrust, Isp, and thrust vector direction along a trajectory that minimizes fuel consumption for interplanetary travel. It is assumed that the propulsion system is power-limited, and thus the compromise between thrust and Isp is a variable to be optimized along the flight path. This program is capable of optimizing not only variable thrust trajectories but also constant thrust trajectories in 3-D space using a planetary ephemeris database. It is also capable of conducting planetary swing-bys. Using this program, various Earth-originating trajectories have been investigated and the optimized results have been compared to traditional CSI and high thrust trajectory solutions. Results show that VSI rocket engines reduce fuel requirements for any mission compared to CSI rocket engines. Fuel can be saved by applying swing-by maneuvers for VSI engines; but the effects of swing-bys due to VSI engines are smaller than that of CSI or high thrust engines.

Saturn Apollo Program

NASA Image and Video Library

1968-01-09

A cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), initially had a thrust of 188,000 pounds each for a combined thrust of over 1,500,000 pounds. Later, the H-1 engine was upgraded to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program. This photo depicts a single modified H-1 engine. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Design and analysis report for the RL10-2B breadboard low thrust engine

NASA Technical Reports Server (NTRS)

Brown, J. R.; Foust, R. R.; Galler, D. E.; Kanic, P. G.; Kmiec, T. D.; Limerick, C. D.; Peckham, R. J.; Swartwout, T.

1984-01-01

The breadboard low thrust RL10-2B engine is described. A summary of the analysis and design effort to define the multimode thrust concept applicable to the requirements for the upper stage vehicles is provided. Baseline requirements were established for operation of the RL10-2B engine under the following conditions: (1) tank head idle at low propellant tank pressures without vehicle propellant conditioning or settling thrust; (2) pumped idle at a ten percent thrust level for low G deployment and/or vehicle tank pressurization; and (3) full thrust (15,000 lb.). Several variations of the engine configuration were investigated and results of the analyses are included.

Reverse thrust performance of the QCSEE variable pitch turbofan engine

NASA Technical Reports Server (NTRS)

Samanich, N. E.; Reemsnyder, D. C.; Blodmer, H. E.

1980-01-01

Results of steady state reverse and forward to reverse thrust transient performance tests are presented. The original quiet, clean, short haul, experimental engine four segment variable fan nozzle was retested in reverse and compared with a continuous, 30 deg half angle conical exlet. Data indicated that the significantly more stable, higher pressure recovery flow with the fixed 30 deg exlet resulted in lower engine vibrations, lower fan blade stress, and approximately a 20 percent improvement in reverse thrust. Objective reverse thrust of 35 percent of takeoff thrust was reached. Thrust response of less than 1.5 sec was achieved for the approach and the takeoff to reverse thrust transients.

Calculation of Turbine Axial Thrust by Coupled CFD Simulations of the Main Flow Path and Secondary Cavity Flow in an SLI LOX Turbine

NASA Technical Reports Server (NTRS)

Dorney, D. J.; Marci, Bogdan; Tran, Ken; Sargent, Scott

2003-01-01

Each single reusable Space Launch Initiative (SLI) booster rocket is an engine operating at a record vacuum thrust level of over 730,000 Ibf using LOX and LH2. This thrust is more than 10% greater than that of the Delta IV rocket, resulting in relatively large LOX and LH2 turbopumps. Since the SLI rocket employs a staged combustion cycle the level of pressure is very high (thousands of psia). This high pressure creates many engineering challenges, including the balancing of axial-forces on the turbopumps. One of the main parameters in the calculation of the axial force is the cavity pressure upstream of the turbine disk. The flow in this cavity is very complex. The lack of understanding of this flow environment hinders the accurate prediction of axial thrust. In order to narrow down the uncertainty band around the actual turbine axial force, a coupled, unsteady computational methodology has been developed to simulate the interaction between the turbine main flow path and the cavity flow. The CORSAIR solver, an unsteady three- dimensional Navier-Stokes code for turbomachinery applications, was used to solve for both the main and the secondary flow fields. Turbine axial thrust values are presented in conjunction with the CFD simulation, together with several considerations regarding the turbine instrumentation for axial thrust estimations during test.

Real-Time Aircraft Engine-Life Monitoring

NASA Technical Reports Server (NTRS)

Klein, Richard

2014-01-01

This project developed an inservice life-monitoring system capable of predicting the remaining component and system life of aircraft engines. The embedded system provides real-time, inflight monitoring of the engine's thrust, exhaust gas temperature, efficiency, and the speed and time of operation. Based upon this data, the life-estimation algorithm calculates the remaining life of the engine components and uses this data to predict the remaining life of the engine. The calculations are based on the statistical life distribution of the engine components and their relationship to load, speed, temperature, and time.

Thrust modeling for hypersonic engines

NASA Technical Reports Server (NTRS)

Riggins, D. W.; Mcclinton, C. R.

1995-01-01

Expressions for the thrust losses of a scramjet engine are developed in terms of irreversible entropy increases and the degree of incomplete combustion. A method is developed which allows the calculation of the lost vehicle thrust due to different loss mechanisms within a given flow-field. This analysis demonstrates clearly the trade-off between mixing enhancement and resultant increased flow losses in scramjet combustors. An engine effectiveness parameter is defined in terms of thrust loss. Exergy and the thrust-potential method are related and compared.

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

14 CFR 23.1155 - Turbine engine reverse thrust and propeller pitch settings below the flight regime.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Turbine engine reverse thrust and propeller... COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime. For turbine engine installations, each...

Review of Nuclear Thermal Propulsion Ground Test Options

NASA Technical Reports Server (NTRS)

Coote, David J.; Power, Kevin P.; Gerrish, Harold P.; Doughty, Glen

2015-01-01

High efficiency rocket propulsion systems are essential for humanity to venture beyond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rockets with relatively high thrust and twice the efficiency of highest performing chemical propellant engines. NTP utilizes the coolant of a nuclear reactor to produce propulsive thrust. An NTP engine produces thrust by flowing hydrogen through a nuclear reactor to cool the reactor, heating the hydrogen and expelling it through a rocket nozzle. The hot gaseous hydrogen is nominally expected to be free of radioactive byproducts from the nuclear reactor; however, it has the potential to be contaminated due to off-nominal engine reactor performance. NTP ground testing is more difficult than chemical engine testing since current environmental regulations do not allow/permit open air testing of NTP as was done in the 1960's and 1970's for the Rover/NERVA program. A new and innovative approach to rocket engine ground test is required to mitigate the unique health and safety risks associated with the potential entrainment of radioactive waste from the NTP engine reactor core into the engine exhaust. Several studies have been conducted since the ROVER/NERVA program in the 1970's investigating NTP engine ground test options to understand the technical feasibility, identify technical challenges and associated risks and provide rough order of magnitude cost estimates for facility development and test operations. The options can be divided into two distinct schemes; (1) real-time filtering of the engine exhaust and its release to the environment or (2) capture and storage of engine exhaust for subsequent processing.

Thrust Performance Evaluation of a Turbofan Engine Based on Exergetic Approach and Thrust Management in Aircraft

NASA Astrophysics Data System (ADS)

Yalcin, Enver

2017-05-01

The environmental parameters such as temperature and air pressure which are changing depending on altitudes are effective on thrust and fuel consumption of aircraft engines. In flights with long routes, thrust management function in airplane information system has a structure that ensures altitude and performance management. This study focused on thrust changes throughout all flight were examined by taking into consideration their energy and exergy performances for fuel consumption of an aircraft engine used in flight with long route were taken as reference. The energetic and exergetic performance evaluations were made under the various altitude conditions. The thrust changes for different altitude conditions were obtained to be at 86.53 % in descending direction and at 142.58 % in ascending direction while the energy and exergy efficiency changes for the referenced engine were found to be at 80.77 % and 84.45 %, respectively. The results revealed here can be helpful to manage thrust and reduce fuel consumption, but engine performance will be in accordance with operation requirements.

Wind-Tunnel Investigation of Effects of Unsymmetrical Horizontal-Tail Arrangements on Power-on Static Longitudinal Stability of a Single-Engine Airplane Model

NASA Technical Reports Server (NTRS)

Purser, Paul E.; Spear, Margaret F.

1947-01-01

A wind-tunnel investigation has been made to determine the effects of unsymmetrical horizontal-tail arrangements on the power-on static longitudinal stability of a single-engine single-rotation airplane model. Although the tests and analyses showed that extreme asymmetry in the horizontal tail indicated a reduction in power effects on longitudinal stability for single-engine single-rotation airplanes, the particular "practical" arrangement tested did not show marked improvement. Differences in average downwash between the normal tail arrangement and various other tail arrangements estimated from computed values of propeller-slipstream rotation agreed with values estimated from pitching-moment test data for the flaps-up condition (low thrust and torque) and disagreed for the flaps-down condition (high thrust and torque). This disagreement indicated the necessity for continued research to determine the characteristics of the slip-stream behind various propeller-fuselage-wing combinations. Out-of-trim lateral forces and moments of the unsymmetrical tail arrangements that were best from consideration of longitudinal stability were no greater than those of the normal tail arrangement.

Quiet Clean Short-haul Experimental Engine (QCSEE) over-the-wing engine and control simulation results

NASA Technical Reports Server (NTRS)

1978-01-01

A hybrid-computer simulation of the over the wing turbofan engine was constructed to develop the dynamic design of the control. This engine and control system includes a full authority digital electronic control using compressor stator reset to achieve fast thrust response and a modified Kalman filter to correct for sensor failures. Fast thrust response for powered-lift operations and accurate, fast responding, steady state control of the engine is provided. Simulation results for throttle bursts from 62 to 100 percent takeoff thrust predict that the engine will accelerate from 62 to 95 percent takeoff thrust in one second.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Parametric Model of an Aerospike Rocket Engine

NASA Technical Reports Server (NTRS)

Korte, J. J.

2000-01-01

A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

NASA Astrophysics Data System (ADS)

Forde, Scott; Bulman, Mel; Neill, Todd

2006-07-01

Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

A Study on Aircraft Engine Control Systems for Integrated Flight and Propulsion Control

NASA Astrophysics Data System (ADS)

Yamane, Hideaki; Matsunaga, Yasushi; Kusakawa, Takeshi; Yasui, Hisako

The Integrated Flight and Propulsion Control (IFPC) for a highly maneuverable aircraft and a fighter-class engine with pitch/yaw thrust vectoring is described. Of the two IFPC functions the aircraft maneuver control utilizes the thrust vectoring based on aerodynamic control surfaces/thrust vectoring control allocation specified by the Integrated Control Unit (ICU) of a FADEC (Full Authority Digital Electronic Control) system. On the other hand in the Performance Seeking Control (PSC) the ICU identifies engine's various characteristic changes, optimizes manipulated variables and finally adjusts engine control parameters in cooperation with the Engine Control Unit (ECU). It is shown by hardware-in-the-loop simulation that the thrust vectoring can enhance aircraft maneuverability/agility and that the PSC can improve engine performance parameters such as SFC (specific fuel consumption), thrust and gas temperature.

Lateral dampers for thrust bearings

NASA Technical Reports Server (NTRS)

Hibner, D. H.; Szafir, D. R.

1985-01-01

The development of lateral damping schemes for thrust bearings was examined, ranking their applicability to various engine classes, selecting the best concept for each engine class and performing an in-depth evaluation. Five major engine classes were considered: large transport, military, small general aviation, turboshaft, and non-manrated. Damper concepts developed for evaluation were: curved beam, constrained and unconstrained elastomer, hybrid boost bearing, hydraulic thrust piston, conical squeeze film, and rolling element thrust face.

Low heat transfer oxidizer heat exchanger design and analysis

NASA Technical Reports Server (NTRS)

Kanic, P. G.; Kmiec, T. D.; Peckham, R. J.

1987-01-01

The RL10-IIB engine, a derivative of the RLIO, is capable of multi-mode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2 percent of full thrust, and pumped idle (PI), which is 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for log-g deployment. Stable combustion of the RL10-IIB engine at THI and PI thrust levels can be accomplished by providing gaseous oxygen at the propellant injector. Using gaseous hydrogen from the thrust chamber jacket as an energy source, a heat exchanger can be used to vaporize liquid oxygen without creating flow instability. This report summarizes the design and analysis of a United Aircraft Products (UAP) low-rate heat transfer heat exchanger concept for the RL10-IIB rocket engine. The design represents a second iteration of the RL10-IIB heat exchanger investigation program. The design and analysis of the first heat exchanger effort is presented in more detail in NASA CR-174857. Testing of the previous design is detailed in NASA CR-179487.

Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft

NASA Technical Reports Server (NTRS)

Hess, J. R.; Bear, R. L.

1982-01-01

A viable, single engine, supersonic V/STOL fighter/attack aircraft concept was defined. This vectored thrust, canard wing configuration utilizes an advanced technology separated flow engine with fan stream burning. The aerodynamic characteristics of this configuration were estimated and performance evaluated. Significant aerodynamic and aerodynamic propulsion interaction uncertainties requiring additional investigation were identified. A wind tunnel model concept and test program to resolve these uncertainties and validate the aerodynamic prediction methods were defined.

Electric propulsion options for 10 kW class earth space missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

A simple dynamic engine model for use in a real-time aircraft simulation with thrust vectoring

NASA Technical Reports Server (NTRS)

Johnson, Steven A.

1990-01-01

A simple dynamic engine model was developed at the NASA Ames Research Center, Dryden Flight Research Facility, for use in thrust vectoring control law development and real-time aircraft simulation. The simple dynamic engine model of the F404-GE-400 engine (General Electric, Lynn, Massachusetts) operates within the aircraft simulator. It was developed using tabular data generated from a complete nonlinear dynamic engine model supplied by the manufacturer. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. Included is a description of a method to account for axial thrust loss resulting from thrust vectoring. In addition, the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000 ft altitude and at Mach 0.7, 35,000 ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.

DC-9 flight demonstration program with refanned JT8D engines. Volume 1: Summary

NASA Technical Reports Server (NTRS)

1975-01-01

The design, analysis, fabrication, and ground and flight testing of DC-9 airframe/nacelle hardware with prototype JT8D-109 engines are discussed. The installation of the JT8D-109 engine on the DC-9 Refan airplane required new or modified hardware for the pylon, nacelle, and fuselage. The acoustic material used in the nose cowl was bonded aluminum honeycomb sandwich and the exhaust duct acoustic material was Inconel 625 Stresskin. The sea level static, standard day bare engine takeoff thrust, the cruise TSFC and the maximum available cruise thrust for the JT8D-109 engine were compared with those of the JT8D-9 engine. The range capabilities of the DC-9 Refan and the production DC-9 airplane were also compared. The Refan airplane demonstrated flight characteristics similar to the production DC-9-30 and satisfied airworthiness requirements. Flyover noise levels were determined for the DC-9 Refan and the DC-9 C-9A airplane for takeoff and landing conditions. Cost estimates were also made.

Breadboard RL10-11B low thrust operating mode

NASA Technical Reports Server (NTRS)

Kmiec, Thomas D.; Galler, Donald E.

1987-01-01

Cryogenic space engines require a cooling process to condition engine hardware to operating temperature before start. This can be accomplished most efficiently by burning propellants that would otherwise be dumped overboard after cooling the engine. The resultant low thrust operating modes are called Tank Head Idle and Pumped Idle. During February 1984, Pratt & Whitney conducted a series of tests demonstrating operation of the RL10 rocket engines at low thrust levels using a previously untried hydrogen/oxygen heat exchanger. The initial testing of the RL10-11B Breadboard Low Thrust Engine is described. The testing demonstrated operation at both tank head idle and pumped idle modes.

Evaluation of a simplified gross thrust calculation method for a J85-21 afterburning turbojet engine in an altitude facility

NASA Technical Reports Server (NTRS)

Baer-Riedhart, J. L.

1982-01-01

A simplified gross thrust calculation method was evaluated on its ability to predict the gross thrust of a modified J85-21 engine. The method used tailpipe pressure data and ambient pressure data to predict the gross thrust. The method's algorithm is based on a one-dimensional analysis of the flow in the afterburner and nozzle. The test results showed that the method was notably accurate over the engine operating envelope using the altitude facility measured thrust for comparison. A summary of these results, the simplified gross thrust method and requirements, and the test techniques used are discussed in this paper.

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

NASA Technical Reports Server (NTRS)

Henneberry, Hugh M.; Snyder, Christopher A.

1993-01-01

An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

High-speed engine/component performance assessment using exergy and thrust-based methods

NASA Technical Reports Server (NTRS)

Riggins, D. W.

1996-01-01

This investigation summarizes a comparative study of two high-speed engine performance assessment techniques based on energy (available work) and thrust-potential (thrust availability). Simple flow-fields utilizing Rayleigh heat addition and one-dimensional flow with friction are used to demonstrate the fundamental inability of conventional energy techniques to predict engine component performance, aid in component design, or accurately assess flow losses. The use of the thrust-based method on these same examples demonstrates its ability to yield useful information in all these categories. Energy and thrust are related and discussed from the stand-point of their fundamental thermodynamic and fluid dynamic definitions in order to explain the differences in information obtained using the two methods. The conventional definition of energy is shown to include work which is inherently unavailable to an aerospace Brayton engine. An engine-based energy is then developed which accurately accounts for this inherently unavailable work; performance parameters based on this quantity are then shown to yield design and loss information equivalent to the thrust-based method.

Application of a Constant Gain Extended Kalman Filter for In-Flight Estimation of Aircraft Engine Performance Parameters

NASA Technical Reports Server (NTRS)

Kobayashi, Takahisa; Simon, Donald L.; Litt, Jonathan S.

2005-01-01

An approach based on the Constant Gain Extended Kalman Filter (CGEKF) technique is investigated for the in-flight estimation of non-measurable performance parameters of aircraft engines. Performance parameters, such as thrust and stall margins, provide crucial information for operating an aircraft engine in a safe and efficient manner, but they cannot be directly measured during flight. A technique to accurately estimate these parameters is, therefore, essential for further enhancement of engine operation. In this paper, a CGEKF is developed by combining an on-board engine model and a single Kalman gain matrix. In order to make the on-board engine model adaptive to the real engine s performance variations due to degradation or anomalies, the CGEKF is designed with the ability to adjust its performance through the adjustment of artificial parameters called tuning parameters. With this design approach, the CGEKF can maintain accurate estimation performance when it is applied to aircraft engines at offnominal conditions. The performance of the CGEKF is evaluated in a simulation environment using numerous component degradation and fault scenarios at multiple operating conditions.

Real-time in-flight thrust calculation on a digital electronic engine control-equipped F100 engine in an F-15 airplane

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1984-01-01

Computer algorithms which calculate in-flight engine and aircraft performance real-time are discussed. The first step was completed with the implementation of a real-time thrust calculation program on a digital electronic engine control (DEEC) equiped F100 engine in an F-15 aircraft. The in-flight thrust modifications that allow calculations to be performed in real-time, to compare results to predictions, are presented.

Thrust Area Report, Engineering Research, Development and Technology

DOE Office of Scientific and Technical Information (OSTI.GOV)

Langland, R. T.

1997-02-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the knowledge base, process technologies, specialized equipment, tools and facilities to support current and future LLNL programs. Engineering`s efforts are guided by a strategy that results in dual benefit: first, in support of Department of Energy missions, such as national security through nuclear deterrence; and second, in enhancing the nation`s economic competitiveness through our collaboration with U.S. industry in pursuit of the most cost- effective engineering solutions to LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Programmore » has two important goals: (1) identify key technologies relevant to LLNL programs where we can establish unique competencies, and (2) conduct high-quality research and development to enhance our capabilities and establish ourselves as the world leaders in these technologies. To focus Engineering`s efforts technology {ital thrust areas} are identified and technical leaders are selected for each area. The thrust areas are comprised of integrated engineering activities, staffed by personnel from the nine electronics and mechanical engineering divisions, and from other LLNL organizations. This annual report, organized by thrust area, describes Engineering`s activities for fiscal year 1996. The report provides timely summaries of objectives, methods, and key results from eight thrust areas: Computational Electronics and Electromagnetics; Computational Mechanics; Microtechnology; Manufacturing Technology; Materials Science and Engineering; Power Conversion Technologies; Nondestructive Evaluation; and Information Engineering. Readers desiring more information are encouraged to contact the individual thrust area leaders or authors. 198 refs., 206 figs., 16 tabs.« less

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

H-1 engine characteristics: The H-1 engine was developed under the management of the Marshall Space Flight Center (MSFC). The cluster of eight H-1 engines was used to power the first stage of the Saturn I (S-I stage) and Saturn IB (S-IVB stage) launch vehicles, and produced 188,00 pounds of thrust, a combined thrust of 1,500,000 pounds, later uprated to 205,000 pounds of thrust and a combined total thrust of 1,650,000 pounds for the Saturn IB program.

Measuring Model Rocket Engine Thrust Curves

ERIC Educational Resources Information Center

Penn, Kim; Slaton, William V.

2010-01-01

This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

An engine trade study for a supersonic STOVL fighter-attack aircraft, volume 1

NASA Technical Reports Server (NTRS)

Beard, B. B.; Foley, W. H.

1982-01-01

The best main engine for an advanced STOVL aircraft flight demonstrator was studied. The STOVL aircraft uses ejectors powered by engine bypass flow together with vectored core exhaust to achieve vertical thrust capability. Bypass flow and core flow are exhausted through separate nozzles during wingborne flight. Six near term turbofan engines were examined for suitability for this aircraft concept. Fan pressure ratio, thrust split between bypass and core flow, and total thrust level were used to compare engines. One of the six candidate engines was selected for the flight demonstrator configuration. Propulsion related to this aircraft concept was studied. A preliminary candidate for the aircraft reaction control system for hover attitude control was selected. A mathematical model of transfer of bypass thrust from ejectors to aft directed nozzle during the transition to wingborne flight was developed. An equation to predict ejector secondary air flow rate and ram drag is derived. Additional topics discussed include: nozzle area control, ejector to engine inlet reingestion, bypass/core thrust split variation, and gyroscopic behavior during hover.

Hamilton Standard Q-fan demonstrator dynamic pitch change test program, volume 1

NASA Technical Reports Server (NTRS)

Demers, W. J.; Nelson, D. J.; Wainauski, H. S.

1975-01-01

Tests of a full scale variable pitch fan engine to obtain data on the structural characteristics, response times, and fan/core engine compatibility during transient changes in blade angle, fan rpm, and engine power is reported. Steady state reverse thrust tests with a take off nozzle configuration were also conducted. The 1.4 meter diameter, 13 bladed controllable pitch fan was driven by a T55 L 11A engine with power and blade angle coordinated by a digital computer. The tests demonstrated an ability to change from full forward thrust to reverse thrust in less than one (1) second. Reverse thrust was effected through feather and through flat pitch; structural characteristics and engine/fan compatibility were within satisfactory limits.

Electric Propulsion Options for 10 kW Class Earth-Space Missions

NASA Technical Reports Server (NTRS)

Patterson, M. J.; Curran, Francis M.

1989-01-01

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

Test Stand at the Rocket Engine Test Facility

NASA Image and Video Library

1973-02-21

The thrust stand in the Rocket Engine Test Facility at the National Aeronautics and Space Administration (NASA) Lewis Research Center in Cleveland, Ohio. The Rocket Engine Test Facility was constructed in the mid-1950s to expand upon the smaller test cells built a decade before at the Rocket Laboratory. The $2.5-million Rocket Engine Test Facility could test larger hydrogen-fluorine and hydrogen-oxygen rocket thrust chambers with thrust levels up to 20,000 pounds. Test Stand A, seen in this photograph, was designed to fire vertically mounted rocket engines downward. The exhaust passed through an exhaust gas scrubber and muffler before being vented into the atmosphere. Lewis researchers in the early 1970s used the Rocket Engine Test Facility to perform basic research that could be utilized by designers of the Space Shuttle Main Engines. A new electronic ignition system and timer were installed at the facility for these tests. Lewis researchers demonstrated the benefits of ceramic thermal coatings for the engine’s thrust chamber and determined the optimal composite material for the coatings. They compared the thermal-coated thrust chamber to traditional unlined high-temperature thrust chambers. There were more than 17,000 different configurations tested on this stand between 1973 and 1976. The Rocket Engine Test Facility was later designated a National Historic Landmark for its role in the development of liquid hydrogen as a propellant.

Advanced engine study for mixed-mode orbit-transfer vehicles

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1978-01-01

Engine design, performance, weight and envelope data were established for three mixed-mode orbit-transfer vehicle engine candidates. Engine concepts evaluated are the tripropellant, dual-expander and plug cluster. Oxygen, RP-1 and hydrogen are the propellants considered for use in these engines. Theoretical performance and propellant properties were established for bipropellant and tripropellant mixes of these propellants. RP-1, hydrogen and oxygen were evaluated as coolants and the maximum attainable chamber pressures were determined for each engine concept within the constraints of the propellant properties and the low cycle thermal fatigue (300 cycles) requirement. The baseline engine design and component operating characteristics are determined at a thrust level of 88,964N (20,000 lbs) and a thrust split of 0.5. The parametric data is generated over ranges of thrust and thrust split of 66.7 to 400kN (15 to 90 klb) and 0.4 to 0.8, respectively.

Tests of a D vented thrust deflecting nozzle behind a simulated turbofan engine

NASA Technical Reports Server (NTRS)

Watson, T. L.

1982-01-01

A D vented thrust deflecting nozzle applicable to subsonic V/STOL aircraft was tested behind a simulated turbofan engine in the verticle thrust stand. Nozzle thrust, fan operating characteristics, nozzle entrance conditions, and static pressures were measured. Nozzle performance was measured for variations in exit area and thrust deflection angle. Six core nozzle configurations, the effect of core exit axial location, mismatched core and fan stream nozzle pressure ratios, and yaw vane presence were evaluated. Core nozzle configuration affected performance at normal and engine out operating conditions. Highest vectored nozzle performance resulted for a given exit area when core and fan stream pressure were equal. Its is concluded that high nozzle performance can be maintained at both normal and engine out conditions through control of the nozzle entrance Mach number with a variable exit area.

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Thrust reversers. 33.97 Section 33.97 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the...

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

Reusable thrust chamber and injector concepts were evaluated for the space shuttle orbit maneuvering engine (OME). Parametric engine calculations were carried out by computer program for N2O4/amine, LOX/amine and LOX/hydrocarbon propellant combinations for engines incorporating regenerative cooled and insulated columbium thrust chambers. The calculation methods are described including the fuel vortex film cooling method of combustion gas temperature control, and performance prediction. A method of acceptance of a regeneratively cooled heat rejection reduction using a silicone oil additive was also demonstrated by heated tube heat transfer testing. Regeneratively cooled thrust chamber operation was also demonstrated where the injector was characterized for the OME application with a channel wall regenerative thrust chamber. Bomb stability testing of the demonstration chambers/injectors demonstrated recovery for the nominal design of acoustic cavities. Cavity geometry changes were also evaluated to assess their damping margin. Performance and combustion stability was demonstrated of the originally developed 10 inch diameter combustion pattern operating in an 8 inch diameter thrust chamber.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. The H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

A Cluster of eight H-1 engines were used to thrust the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern. Four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. The H-1 engine, fueled with liquid oxygen (LOX) and kerosene (RP-1), had a thrust of 188,000 pound each for a combined thrust of over 1,500,000 pounds. Each H-1 engine was developed under the direction of Marshall Space Flight Center (MSFC).

Definition, technology readiness, and development cost of the orbit transfer vehicle engine integrated control and health monitoring system elements

NASA Technical Reports Server (NTRS)

Cannon, I.; Balcer, S.; Cochran, M.; Klop, J.; Peterson, S.

1991-01-01

An Integrated Control and Health Monitoring (ICHM) system was conceived for use on a 20 Klb thrust baseline Orbit Transfer Vehicle (OTV) engine. Considered for space used, the ICHM was defined for reusability requirements for an OTV engine service free life of 20 missions, with 100 starts and a total engine operational time of 4 hours. Functions were derived by flowing down requirements from NASA guidelines, previous OTV engine or ICHM documents, and related contracts. The elements of an ICHM were identified and listed, and these elements were described in sufficient detail to allow estimation of their technology readiness levels. These elements were assessed in terms of technology readiness level, and supporting rationale for these assessments presented. The remaining cost for development of a minimal ICHM system to technology readiness level 6 was estimated. The estimates are within an accuracy range of minus/plus 20 percent. The cost estimates cover what is needed to prepare an ICHM system for use on a focussed testbed for an expander cycle engine, excluding support to the actual test firings.

High heat transfer oxidizer heat exchanger design and analysis. [RL10-2B engine

NASA Technical Reports Server (NTRS)

Kmiec, Thomas D.; Kanic, Paul G.; Peckham, Richard J.

1987-01-01

The RL10-2B engine, a derivative of the RL10, is capable of multimode thrust operation. This engine operates at two low thrust levels: tank head idle (THI), which is approximately 1 to 2% of full thrust, and pumped idle (PI), which is 10% of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient engine thermal conditioning; PI operation provides vehicle tank pre-pressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-2B engine during the low thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidizer heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. The design, concept verification testing and analysis for such a heat exchanger is discussed. The design presented uses a high efficiency compact core to vaporize the oxygen, and in the self-contained unit, attenuates any pressure and flow oscillations which result from unstable boiling in the core. This approach is referred to as the high heat transfer design. An alternative approach which prevents unstable boiling of the oxygen by limiting the heat transfer is referred to as the low heat transfer design and is reported in Pratt & Whitney report FR-19135-2.

Full Flight Envelope Direct Thrust Measurement on a Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Conners, Timothy R.; Sims, Robert L.

1998-01-01

Direct thrust measurement using strain gages offers advantages over analytically-based thrust calculation methods. For flight test applications, the direct measurement method typically uses a simpler sensor arrangement and minimal data processing compared to analytical techniques, which normally require costly engine modeling and multisensor arrangements throughout the engine. Conversely, direct thrust measurement has historically produced less than desirable accuracy because of difficulty in mounting and calibrating the strain gages and the inability to account for secondary forces that influence the thrust reading at the engine mounts. Consequently, the strain-gage technique has normally been used for simple engine arrangements and primarily in the subsonic speed range. This paper presents the results of a strain gage-based direct thrust-measurement technique developed by the NASA Dryden Flight Research Center and successfully applied to the full flight envelope of an F-15 aircraft powered by two F100-PW-229 turbofan engines. Measurements have been obtained at quasi-steady-state operating conditions at maximum non-augmented and maximum augmented power throughout the altitude range of the vehicle and to a maximum speed of Mach 2.0 and are compared against results from two analytically-based thrust calculation methods. The strain-gage installation and calibration processes are also described.

Statistical error model for a solar electric propulsion thrust subsystem

NASA Technical Reports Server (NTRS)

Bantell, M. H.

1973-01-01

The solar electric propulsion thrust subsystem statistical error model was developed as a tool for investigating the effects of thrust subsystem parameter uncertainties on navigation accuracy. The model is currently being used to evaluate the impact of electric engine parameter uncertainties on navigation system performance for a baseline mission to Encke's Comet in the 1980s. The data given represent the next generation in statistical error modeling for low-thrust applications. Principal improvements include the representation of thrust uncertainties and random process modeling in terms of random parametric variations in the thrust vector process for a multi-engine configuration.

Computation of optimal Mars trajectories via combined chemical/electrical propulsion, Part 3: Compromise solutions

NASA Astrophysics Data System (ADS)

Miele, A.; Wang, T.; Williams, P. N.

2005-12-01

The success of the solar-electric ion engine powering the DS1 spacecraft has paved the way toward the use of low-thrust electrical engines in future planetary/interplanetary missions. Vis-à-vis a chemical engine, an electrical engine has a higher specific impulse, implying a possible decrease in propellant mass; however, the low-thrust aspect discourages the use of an electrical engine in the near-planet phases of a trip, since this might result in an increase in flight time. Therefore, a fundamental design problem is to find the best combination of chemical propulsion and electrical propulsion for a given mission, for example, a mission from Earth to Mars. With this in mind, this paper is the third of a series dealing with the optimization of Earth Mars missions via the use of hybrid engines, namely the combination of high-thrust chemical engines for planetary flight and low-thrust electrical engines for interplanetary flight. We look at the deep-space interplanetary portion of the trajectory under rather idealized conditions. The two major performance indexes, the propellant mass and the flight time, are in conflict with one another for the following reason: any attempt at reducing the former causes an increase in the latter and vice versa. Therefore, it is natural to consider a compromise performance index involving the scaled values of the propellant mass and flight time weighted respectively by the compromise factor C and its complement 1-C. We use the compromise factor as the parameter of the one-parameter family of compromise trajectories. Analyses carried out with the sequential gradient-restoration algorithm for optimal control problems lead to results which can be highlighted as follows. Thrust profile. Generally speaking, the thrust profile of the compromise trajectory includes three subarcs: the first subarc is characterized by maximum thrust in conjunction with positive (upward) thrust direction; the second subarc is characterized by zero thrust (coasting flight); the third subarc is characterized by maximum thrust in conjunction with negative (downward) thrust direction. Effect of the compromise factor. As the compromise factor increases, the propellant mass decreases and the flight time increases; correspondingly, the following changes in the thrust profile take place: (a) the time lengths of the first and third subarcs (powered phases) decrease slightly, meaning that thrust application occurs for shorter duration; also, the average value of the thrust direction in the first and third subarcs decreases, implying higher efficiency of thrust application wrt the spacecraft energy level; as a result, the total propellant mass decreases; (b) the time length of the second subarc (coasting) increases considerably, resulting in total time increase. Minimum time trajectory. If C=0, the resulting minimum time trajectory has the following characteristics: (a) the time length of the coasting subarc reduces to zero and the three-subarc trajectory degenerates into a two-subarc trajectory; (b) maximum thrust is applied at all times and the thrust direction switches from upward to downward at midcourse. Minimum propellant mass trajectory. If C=1, the resulting minimum propellant mass trajectory has the following characteristics: (a) the thrust magnitude has a bang-zero-bang profile; (b) for the powered subarcs, the thrust direction is tangent to the flight path at all times.

Thrust Augmentation of a Turbojet Engine at Simulated Flight Conditions by Introduction of a Water-Alcohol Mixture into the Compressor

NASA Technical Reports Server (NTRS)

Useller, James W.; Auble, Carmon M.; Harvey, Ray W., Sr.

1952-01-01

An investigation was conducted at simulated high-altitude flight conditions to evaluate the use of compressor evaporative cooling as a means of turbojet-engine thrust augmentation. Comparison of the performance of the engine with water-alcohol injection at the compressor inlet, at the sixth stage of the compressor, and at the sixth and ninth stages was made. From consideration of the thrust increases achieved, the interstage injection of the coolant was considered more desirable preferred over the combined sixth- and ninth-stage injection because of its relative simplicity. A maximum augmented net-thrust ratio of 1.106 and a maximum augmented jet-thrust ratio of 1.062 were obtained at an augmented liquid ratio of 2.98 and an engine-inlet temperature of 80 F. At lower inlet temperatures (-40 to 40 F), the maximum augmented net-thrust ratios ranged from 1.040 to 1.076 and the maximum augmented jet-thrust ratios ranged from 1.027 to 1.048, depending upon the inlet temperature. The relatively small increase in performance at the lower inlet-air temperatures can be partially attributed to the inadequate evaporation of the water-alcohol mixture, but the more significant limitation was believed to be caused by the negative influence of the liquid coolant on engine- component performance. In general, it is concluded that the effectiveness of the injection of a coolant into the compressor as a means of thrust augmentation is considerably influenced by the design characteristics of the components of the engine being used.

Jet noise suppressor nozzle development for augmentor wing jet STOL research aircraft (C-8A Buffalo)

NASA Technical Reports Server (NTRS)

Harkonen, D. L.; Marks, C. C.; Okeefe, J. V.

1974-01-01

Noise and performance test results are presented for a full-scale advanced design rectangular array lobe jet suppressor nozzle (plain wall and corrugated). Flight design and installation considerations are also discussed. Noise data are presented in terms of peak PNLT (perceived noise level, tone corrected) suppression relative to the existing airplane and one-third octave-band spectra. Nozzle performance is presented in terms of velocity coefficient. Estimates of the hot thrust available during emergency (engine out) with the suppressor nozzle installed are compared with the current thrust levels produced by the round convergent nozzles.

Sensory Perception in the Human Research and Engineering Directorate: Thrust Areas and Recent Research 2011-2014

DTIC Science & Technology

2014-09-01

estimation in an open space. In: Glotin, editor. Soundscape semiotics: localisation and categorisation. Rijeka (Croatia): InTech; 2014. Available at...Glotin H, editor. Soundscape semiotics – localisation and categorisation. Rijeka (Croatia): InTech; 2014. Available at: http://www.intechopen.com...books/ soundscape -semiotics-localisation -and-categorisation/auditory-distance-estimation-in-an-open-space. The purpose of the study was to expand our

Experimental Determination of Exhaust Gas Thrust, Special Report

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Voss, Fred

1940-01-01

This investigation presents the results of tests made on a radial engine to determine the thrust that can be obtained from the exhaust gas when discharged from separate stacks and when discharged from the collector ring with various discharge nozzles. The engine was provided with a propeller to absorb the power and was mounted on a test stand equipped with scales for measuring the thrust and engine torque. The results indicate that at full open throttle at sea level, for the engine tested, a gain in thrust horsepower of 18 percent using separate stacks, and 9.5 percent using a collector ring and discharge nozzle, can be expected at an air speed of 550 miles per hour.

Influence of Thrust Level on the Architecture and Optimal Working Process Parameters of a Small-scale Turbojet for UAV

NASA Astrophysics Data System (ADS)

Kuz`michev, V. S.; Filinov, E. P.; Ostapyuk, Ya A.

2018-01-01

This article describes how the thrust level influences the turbojet architecture (types of turbomachines that provide the maximum efficiency) and its working process parameters (turbine inlet temperature (TIT) and overall pressure ratio (OPR)). Functional gasdynamic and strength constraints were included, total mass of fuel and the engine required for mission and the specific fuel consumption (SFC) were considered optimization criteria. Radial and axial turbines and compressors were considered. The results show that as the engine thrust decreases, optimal values of working process parameters decrease too, and the regions of compromise shrink. Optimal engine architecture and values of working process parameters are suggested for turbojets with thrust varying from 100N to 100kN. The results show that for the thrust below 25kN the engine scale factor should be taken into the account, as the low flow rates begin to influence the efficiency of engine elements substantially.

Large-Scale Wind-Tunnel Tests of Exhaust Ingestion Due to Thrust Reversal on a Four-Engine Jet Transport during Ground Roll

NASA Technical Reports Server (NTRS)

Tolhurst, William H., Jr.; Hickey, David H.; Aoyagi, Kiyoshi

1961-01-01

Wind-tunnel tests have been conducted on a large-scale model of a swept-wing jet transport type airplane to study the factors affecting exhaust gas ingestion into the engine inlets when thrust reversal is used during ground roll. The model was equipped with four small jet engines mounted in nacelles beneath the wing. The tests included studies of both cascade and target type reversers. The data obtained included the free-stream velocity at the occurrence of exhaust gas ingestion in the outboard engine and the increment of drag due to thrust reversal for various modifications of thrust reverser configuration. Motion picture films of smoke flow studies were also obtained to supplement the data. The results show that the free-stream velocity at which ingestion occurred in the outboard engines could be reduced considerably, by simple modifications to the reversers, without reducing the effective drag due to reversed thrust.

Exhaust-stack nozzle area and shape for individual cylinder exhaust-gas jet-propulsion system

NASA Technical Reports Server (NTRS)

Pinkel, Benjamin; Turner, Richard; Voss, Fred; Humble, Leroy V

1943-01-01

This report presents the results of an investigation conducted on the effect of exhaust-stack nozzle area, shape, and length on engine power, jet thrust, and gain in net thrust (engine propeller plus jet). Single-cylinder engine data were obtained using three straight stacks 25, 44, and 108 inches in length; an S-shaped stack, a 90 degree bend, a 180 degree bend, and a short straight stack having a closed branch faired into it. Each stack was fitted with nozzles varying in exit area from 0.91 square inch to the unrestricted area of the stack of 4.20 square inches. The engine was generally operated over a range of engine speeds from 1300 to 2100 r.p.m, inlet-manifold pressures from 22 to 30 inches of mercury absolute, and a fuel-air ratio of 0.08. The loss in engine power, the jet thrust, and the gain in net thrust are correlated in terms of several simple parameters. An example is given for determining the optimum nozzle area and the overall net thrust.

Do trout swim better than eels? Challenges for estimating performance based on the wake of self-propelled bodies

NASA Astrophysics Data System (ADS)

Tytell, Eric D.

2007-11-01

Engineers and biologists have long desired to compare propulsive performance for fishes and underwater vehicles of different sizes, shapes, and modes of propulsion. Ideally, such a comparison would be made on the basis of either propulsive efficiency, total power output or both. However, estimating the efficiency and power output of self-propelled bodies, and particularly fishes, is methodologically challenging because it requires an estimate of thrust. For such systems traveling at a constant velocity, thrust and drag are equal, and can rarely be separated on the basis of flow measured in the wake. This problem is demonstrated using flow fields from swimming American eels, Anguilla rostrata, measured using particle image velocimetry (PIV) and high-speed video. Eels balance thrust and drag quite evenly, resulting in virtually no wake momentum in the swimming (axial) direction. On average, their wakes resemble those of self-propelled jet propulsors, which have been studied extensively. Theoretical studies of such wakes may provide methods for the estimation of thrust separately from drag. These flow fields are compared with those measured in the wakes of rainbow trout, Oncorhynchus mykiss, and bluegill sunfish, Lepomis macrochirus. In contrast to eels, these fishes produce wakes with axial momentum. Although the net momentum flux must be zero on average, it is neither spatially nor temporally homogeneous; the heterogeneity may provide an alternative route for estimating thrust. This review shows examples of wakes and velocity profiles from the three fishes, indicating challenges in estimating efficiency and power output and suggesting several routes for further experiments. Because these estimates will be complicated, a much simpler method for comparing performance is outlined, using as a point of comparison the power lost producing the wake. This wake power, a component of the efficiency and total power, can be estimated in a straightforward way from the flow fields. Although it does not provide complete information about the performance, it can be used to place constraints on the relative efficiency and cost of transport for the fishes.

Do trout swim better than eels? Challenges for estimating performance based on the wake of self-propelled bodies

NASA Astrophysics Data System (ADS)

Tytell, Eric D.

Engineers and biologists have long desired to compare propulsive performance for fishes and underwater vehicles of different sizes, shapes, and modes of propulsion. Ideally, such a comparison would be made on the basis of either propulsive efficiency, total power output or both. However, estimating the efficiency and power output of self-propelled bodies, and particularly fishes, is methodologically challenging because it requires an estimate of thrust. For such systems traveling at a constant velocity, thrust and drag are equal, and can rarely be separated on the basis of flow measured in the wake. This problem is demonstrated using flow fields from swimming American eels, Anguilla rostrata, measured using particle image velocimetry (PIV) and high-speed video. Eels balance thrust and drag quite evenly, resulting in virtually no wake momentum in the swimming (axial) direction. On average, their wakes resemble those of self-propelled jet propulsors, which have been studied extensively. Theoretical studies of such wakes may provide methods for the estimation of thrust separately from drag. These flow fields are compared with those measured in the wakes of rainbow trout, Oncorhynchus mykiss, and bluegill sunfish, Lepomis macrochirus. In contrast to eels, these fishes produce wakes with axial momentum. Although the net momentum flux must be zero on average, it is neither spatially nor temporally homogeneous; the heterogeneity may provide an alternative route for estimating thrust. This review shows examples of wakes and velocity profiles from the three fishes, indicating challenges in estimating efficiency and power output and suggesting several routes for further experiments. Because these estimates will be complicated, a much simpler method for comparing performance is outlined, using as a point of comparison the power lost producing the wake. This wake power, a component of the efficiency and total power, can be estimated in a straightforward way from the flow fields. Although it does not provide complete information about the performance, it can be used to place constraints on the relative efficiency and cost of transport for the fishes.

The use of laterally vectored thrust to counter thrust asymmetry in a tactical jet aircraft

NASA Technical Reports Server (NTRS)

1983-01-01

A nonlinear, six degree-of-freedom flight simulator for a twin engine tactical jet was built on a hybrid computer to investigate lateral vectoring of the remaining thrust component for the case of a single engine failure at low dynamic pressures. Aircraft control was provided by an automatic controller rather than a pilot, and thrust vector control was provided by an open-loop controller that deflected a vane (located on the periphery of each exhaust jet and normally streamlined for noninterference with the flow). Lateral thrust vectoring decreased peak values of lateral control deflections, eliminated the requirement for steady-state lateral aerodynamic control deflections, and decreased the amount of altitude lost for a single engine failure.

Engineering Research and Development and Technology thrust area report FY92

DOE Office of Scientific and Technical Information (OSTI.GOV)

Langland, R.T.; Minichino, C.

1993-03-01

The mission of the Engineering Research, Development, and Technology Program at Lawrence Livermore National Laboratory (LLNL) is to develop the technical staff and the technology needed to support current and future LLNL programs. To accomplish this mission, the Engineering Research, Development, and Technology Program has two important goals: (1) to identify key technologies and (2) to conduct high-quality work to enhance our capabilities in these key technologies. To help focus our efforts, we identify technology thrust areas and select technical leaders for each area. The thrust areas are integrated engineering activities and, rather than being based on individual disciplines, theymore » are staffed by personnel from Electronics Engineering, Mechanical Engineering, and other LLNL organizations, as appropriate. The thrust area leaders are expected to establish strong links to LLNL program leaders and to industry; to use outside and inside experts to review the quality and direction of the work; to use university contacts to supplement and complement their efforts; and to be certain that we are not duplicating the work of others. This annual report, organized by thrust area, describes activities conducted within the Program for the fiscal year 1992. Its intent is to provide timely summaries of objectives, theories, methods, and results. The nine thrust areas for this fiscal year are: Computational Electronics and Electromagnetics; Computational Mechanics; Diagnostics and Microelectronics; Emerging Technologies; Fabrication Technology; Materials Science and Engineering; Microwave and Pulsed Power; Nondestructive Evaluation; and Remote Sensing and Imaging, and Signal Engineering.« less

NASA Orbit Transfer Rocket Engine Technology Program

NASA Technical Reports Server (NTRS)

1984-01-01

The advanced expander cycle engine with a 15,000 lb thrust level and a 6:1 mixture ratio and optimized performance was used as the baseline for a design study of the hydrogen/oxgyen propulsion system for the orbit transfer vehicle. The critical components of this engine are the thrust chamber, the turbomachinery, the extendible nozzle system, and the engine throttling system. Turbomachinery technology is examined for gears, bearing, seals, and rapid solidification rate turbopump shafts. Continuous throttling concepts are discussed. Components of the OTV engine described include the thrust chamber/nozzle assembly design, nozzles, the hydrogen regenerator, the gaseous oxygen heat exchanger, turbopumps, and the engine control valves.

Characterization of Low-Frequency Combustion Stability of the Fastrac Engine

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Jones, Preston (Technical Monitor)

2002-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.

Regeneratively Cooled Liquid Oxygen/Methane Technology Development

NASA Technical Reports Server (NTRS)

Robinson, Joel W.; Greene, Christopher B.; Stout, Jeffrey

2012-01-01

The National Aeronautics & Space Administration (NASA) has identified Liquid Oxygen (LOX)/Liquid Methane (LCH4) as a potential propellant combination for future space vehicles based upon exploration studies. The technology is estimated to have higher performance and lower overall systems mass compared to existing hypergolic propulsion systems. NASA-Marshall Space Flight Center (MSFC) in concert with industry partner Pratt & Whitney Rocketdyne (PWR) utilized a Space Act Agreement to test an oxygen/methane engine system in the Summer of 2010. PWR provided a 5,500 lbf (24,465 N) LOX/LCH4 regenerative cycle engine to demonstrate advanced thrust chamber assembly hardware and to evaluate the performance characteristics of the system. The chamber designs offered alternatives to traditional regenerative engine designs with improvements in cost and/or performance. MSFC provided the test stand, consumables and test personnel. The hot fire testing explored the effective cooling of one of the thrust chamber designs along with determining the combustion efficiency with variations of pressure and mixture ratio. The paper will summarize the status of these efforts.

Space shuttle orbit maneuvering engine reusable thrust chamber program

NASA Technical Reports Server (NTRS)

Senneff, J. M.

1975-01-01

The feasibility of potential reusable thrust chamber concepts is studied. Propellant condidates were examined and analytically combined with potential cooling schemes. A data base of engine data which would assist in a configuration selection was produced. The data base verification was performed by the demonstration of a thrust chamber of a selected coolant scheme design. A full scale insulated columbium thrust chamber was used for propellant coolant configurations. Combustion stability of the injectors and a reduced size thrust chamber were experimentally verified as proof of concept demonstrations of the design and study results.

Operationalizing Special Operations Aviation in Indonesia

DTIC Science & Technology

2006-12-15

special operations forces Builder: Lockheed Power Plant: Four Allison T56 -A-15 turboprop engines Thrust: 4,910 shaft horsepower each engine...Builder: Lockheed Power Plant: Four Allison T56 -A-15 turboprop engines Thrust: 4,910 shaft horsepower each engine Length: 98 feet, 9 inches (30.09

Task 12 data dump (phase 2) OME integrated thrust chamber test report

NASA Technical Reports Server (NTRS)

Tobin, R. D.; Pauckert, R. P.

1974-01-01

The characteristics and performance of the orbit maneuvering engine for the space shuttle are discussed. Emphasis is placed on the regeneratively cooled thrust chamber of the engine. Tests were conducted to determine engine operating parameters during the start, shutdown, and restart. Characteristics of the integrated thrust chamber and the performance and thermal conditions for blowdown operation without supplementary boundary layer cooling were investigated. The results of the test program are presented.

14 CFR 33.8 - Selection of engine power and thrust ratings.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Selection of engine power and thrust ratings. 33.8 Section 33.8 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES General § 33.8 Selection of engine power and...

14 CFR 33.8 - Selection of engine power and thrust ratings.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Selection of engine power and thrust ratings. 33.8 Section 33.8 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES General § 33.8 Selection of engine power and...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

14 CFR 25.1143 - Engine controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Engine controls. 25.1143 Section 25.1143... STANDARDS: TRANSPORT CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust...

14 CFR Appendix I to Part 25 - Installation of an Automatic Takeoff Thrust Control System (ATTCS)

Code of Federal Regulations, 2010 CFR

2010-01-01

...) This appendix specifies additional requirements for installation of an engine power control system that... crew to increase thrust or power. I25.2Definitions. (a) Automatic Takeoff Thrust Control System (ATTCS... mechanical and electrical, that sense engine failure, transmit signals, actuate fuel controls or power levers...

Intelligent, Robust Control of Deteriorated Turbofan Engines via Linear Parameter Varying Quadratic Lyapunov Function Design

NASA Technical Reports Server (NTRS)

Turso, James A.; Litt, Jonathan S.

2004-01-01

A method for accommodating engine deterioration via a scheduled Linear Parameter Varying Quadratic Lyapunov Function (LPVQLF)-Based controller is presented. The LPVQLF design methodology provides a means for developing unconditionally stable, robust control of Linear Parameter Varying (LPV) systems. The controller is scheduled on the Engine Deterioration Index, a function of estimated parameters that relate to engine health, and is computed using a multilayer feedforward neural network. Acceptable thrust response and tight control of exhaust gas temperature (EGT) is accomplished by adjusting the performance weights on these parameters for different levels of engine degradation. Nonlinear simulations demonstrate that the controller achieves specified performance objectives while being robust to engine deterioration as well as engine-to-engine variations.

The Control System for the X-33 Linear Aerospike Engine

NASA Technical Reports Server (NTRS)

Jackson, Jerry E.; Espenschied, Erich; Klop, Jeffrey

1998-01-01

The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

Orbit transfer vehicle engine study, phase A extension. Volume 2A: Study results

NASA Technical Reports Server (NTRS)

1980-01-01

Engine trade studies and systems analyses leading to a baseline engine selection for advanced expander cycle engine are discussed with emphasis on: (1) performance optimization of advanced expander cycle engines in the 10 to 20K pound thrust range; (2) selection of a recommended advanced expander engine configuration based on maximized performance and minimized mission risk, and definition of the components for this configuration; (3) characterization of the low thrust adaptation requirements and performance for the staged combustion engine; (4) generation of a suggested safety and reliability approach for OTV engines independent of engine cycle; (5) definition of program risk relationships between expander and staged combustion cycle engines; and (6) development of schedules and costs for the DDT&E, production, and operation phases of the 10K pound thrust expander engine program.

Optimization of transfer trajectories to the Apophis asteroid for spacecraft with high and low thrust

NASA Astrophysics Data System (ADS)

Ivashkin, V. V.; Krylov, I. V.

2014-03-01

The problem of optimization of a spacecraft transfer to the Apophis asteroid is investigated. The scheme of transfer under analysis includes a geocentric stage of boosting the spacecraft with high thrust, a heliocentric stage of control by a low thrust engine, and a stage of deceleration with injection to an orbit of the asteroid's satellite. In doing this, the problem of optimal control is solved for cases of ideal and piecewise-constant low thrust, and the optimal magnitude and direction of spacecraft's hyperbolic velocity "at infinity" during departure from the Earth are determined. The spacecraft trajectories are found based on a specially developed comprehensive method of optimization. This method combines the method of dynamic programming at the first stage of analysis and the Pontryagin maximum principle at the concluding stage, together with the parameter continuation method. The estimates are obtained for the spacecraft's final mass and for the payload mass that can be delivered to the asteroid using the Soyuz-Fregat carrier launcher.

Schlieren Imaging of a Single-Ejector, Multi-Tube Pulsed Detonation Engine (Postprint)

DTIC Science & Technology

2009-01-01

studies have shown the potential of an ejector to almost double the thrust of a pulsed detonation engine ( PDE ) tube [1-3]. Axial misalignment of the... Detonation Research Facility in the Air Force Research Laboratory were used for this study. The PDE utilizes automotive valving to feed up to four... detonation tubes. The damped thrust stand was setup to measure PDE thrust alone for baseline tests or total thrust from ejector and PDE . This

Comparison of Performance and Component Frontal Areas of Hypothetical Two-spool and One-spool Turbojet Engines

NASA Technical Reports Server (NTRS)

Dugan, James F , Jr

1956-01-01

For constant-mechanical-speed operation, the two-spool thrust values are as great as or greater than the one-spool thrust values over the entire flight range considered, while the specific fuel consumption for the two engines agrees within 1 percent. The maximum difference in thrust occurs at Mach 2.8 in the stratosphere, where the two-spool thrust advantage is about 9 percent for operation with the after burning.

Static Performance of a Wing-Mounted Thrust Reverser Concept

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

1998-01-01

An experimental investigation was conducted in the Jet-Exit Test Facility at NASA Langley Research Center to study the static aerodynamic performance of a wing-mounted thrust reverser concept applicable to subsonic transport aircraft. This innovative engine powered thrust reverser system is designed to utilize wing-mounted flow deflectors to produce aircraft deceleration forces. Testing was conducted using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0, a supercritical left-hand wing section attached via a pylon, and wing-mounted flow deflectors attached to the wing section. Geometric variations of key design parameters investigated for the wing-mounted thrust reverser concept included flow deflector angle and chord length, deflector edge fences, and the yaw mount angle of the deflector system (normal to the engine centerline or parallel to the wing trailing edge). All tests were conducted with no external flow and high pressure air was used to simulate core and fan engine exhaust flows. Test results indicate that the wing-mounted thrust reverser concept can achieve overall thrust reverser effectiveness levels competitive with (parallel mount), or better than (normal mount) a conventional cascade thrust reverser system. By removing the thrust reverser system from the nacelle, the wing-mounted concept offers the nacelle designer more options for improving nacelle aero dynamics and propulsion-airframe integration, simplifying nacelle structural designs, reducing nacelle weight, and improving engine maintenance access.

Method and apparatus for rapid thrust increases in a turbofan engine

NASA Technical Reports Server (NTRS)

Cornett, J. E.; Corley, R. C.; Fraley, T. O.; Saunders, A. A., Jr. (Inventor)

1980-01-01

Upon a landing approach, the normal compressor stator schedule of a fan speed controlled turbofan engine is temporarily varied to substantially close the stators to thereby increase the fuel flow and compressor speed in order to maintain fan speed and thrust. This running of the compressor at an off-design speed substantially reduces the time required to subsequently advance the engine speed to the takeoff thrust level by advancing the throttle and opening the compressor stators.

CAMELOT: Computational-Analytical Multi-fidElity Low-thrust Optimisation Toolbox

NASA Astrophysics Data System (ADS)

Di Carlo, Marilena; Romero Martin, Juan Manuel; Vasile, Massimiliano

2018-03-01

Computational-Analytical Multi-fidElity Low-thrust Optimisation Toolbox (CAMELOT) is a toolbox for the fast preliminary design and optimisation of low-thrust trajectories. It solves highly complex combinatorial problems to plan multi-target missions characterised by long spirals including different perturbations. To do so, CAMELOT implements a novel multi-fidelity approach combining analytical surrogate modelling and accurate computational estimations of the mission cost. Decisions are then made using two optimisation engines included in the toolbox, a single-objective global optimiser, and a combinatorial optimisation algorithm. CAMELOT has been applied to a variety of case studies: from the design of interplanetary trajectories to the optimal de-orbiting of space debris and from the deployment of constellations to on-orbit servicing. In this paper, the main elements of CAMELOT are described and two examples, solved using the toolbox, are presented.

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

NASA Technical Reports Server (NTRS)

Tomsik, Thomas M.

1994-01-01

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

NASA Technical Reports Server (NTRS)

Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

1977-01-01

This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 transport aircraft approaches its first landing under engine power only on Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 approaches the first landing ever of a transport aircraft under engine power only on Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when it normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

Predicted performance of an integrated modular engine system

NASA Technical Reports Server (NTRS)

Binder, Michael; Felder, James L.

1993-01-01

Space vehicle propulsion systems are traditionally comprised of a cluster of discrete engines, each with its own set of turbopumps, valves, and a thrust chamber. The Integrated Modular Engine (IME) concept proposes a vehicle propulsion system comprised of multiple turbopumps, valves, and thrust chambers which are all interconnected. The IME concept has potential advantages in fault-tolerance, weight, and operational efficiency compared with the traditional clustered engine configuration. The purpose of this study is to examine the steady-state performance of an IME system with various components removed to simulate fault conditions. An IME configuration for a hydrogen/oxygen expander cycle propulsion system with four sets of turbopumps and eight thrust chambers has been modeled using the Rocket Engine Transient Simulator (ROCETS) program. The nominal steady-state performance is simulated, as well as turbopump thrust chamber and duct failures. The impact of component failures on system performance is discussed in the context of the system's fault tolerant capabilities.

Replacement of chemical rocket launchers by beamed energy propulsion.

PubMed

Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

2014-11-01

Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

Hypersonic trajectory control of aerospace plane with integrated SCRAMJET engine

NASA Astrophysics Data System (ADS)

Yonemoto, Koichi

The aerospace plane is an airbreathing 'propulsion configured' vehicle having proper forebody contour for inflow pre-compression to the inlet and afterbody that operates as an external expansion nozzle. Since the whole lower side of the body acts as important compression and expansion elements for the airbreathing engine, the flight attitude influences its performance such as specific impulse and thrust coefficient considerably. The stability of ascent trajectory controlling dynamic pressure or heat-input rate is analyzed considering the performance change due to attitude fluctuation. The performance of scramjet engine, a typical hypersonic airbreathing engine, is estimated by a rapid prediction methodology of the combustor proposed by Ikawa.

Background and principles of throttles-only flight control

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.

1995-01-01

There have been many cases in which the crew of a multi-engine airplane had to use engine thrust for emergency flight control. Such a procedure is very difficult, because the propulsive control forces are small, the engine response is slow, and airplane dynamics such as the phugoid and dutch roll are difficult to damp with thrust. In general, thrust increases are used to climb, thrust decreases to descend, and differential thrust is used to turn. Average speed is not significantly affected by changes in throttle setting. Pitch control is achieved because of pitching moments due to speed changes, from thrust offset, and from the vertical component of thrust. Roll control is achieved by using differential thrust to develop yaw, which, through the normal dihedral effect, causes a roll. Control power in pitch and roll tends to increase as speed decreases. Although speed is not controlled by the throttles, configuration changes are often available (lowering gear, flaps, moving center-of-gravity) to change the speed. The airplane basic stability is also a significant factor. Fuel slosh and gyroscopic moments are small influences on throttles-only control. The background and principles of throttles-only flight control are described.

Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60K-lb Thrust Fastrac Rocket Engine

NASA Technical Reports Server (NTRS)

Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

2000-01-01

A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location approximately equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal 0.7 microgram/cc, and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal 2,200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60k-lb Thrust Fastrac Rocket Engine

NASA Technical Reports Server (NTRS)

Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

2000-01-01

A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location about equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal to 0.7 micrograms/cubic cm and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal to 2.200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

NASA Technical Reports Server (NTRS)

Betts, Erin M.; Frederick, Robert A., Jr.

2010-01-01

This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

Maximum thrust mode evaluation

NASA Technical Reports Server (NTRS)

Orme, John S.; Nobbs, Steven G.

1995-01-01

Measured reductions in acceleration times which resulted from the application of the F-15 performance seeking control (PSC) maximum thrust mode during the dual-engine test phase is presented as a function of power setting and flight condition. Data were collected at altitudes of 30,000 and 45,000 feet at military and maximum afterburning power settings. The time savings for the supersonic acceleration is less than at subsonic Mach numbers because of the increased modeling and control complexity. In addition, the propulsion system was designed to be optimized at the mid supersonic Mach number range. Recall that even though the engine is at maximum afterburner, PSC does not trim the afterburner for the maximum thrust mode. Subsonically at military power, time to accelerate from Mach 0.6 to 0.95 was cut by between 6 and 8 percent with a single engine application of PSC, and over 14 percent when both engines were optimized. At maximum afterburner, the level of thrust increases were similar in magnitude to the military power results, but because of higher thrust levels at maximum afterburner and higher aircraft drag at supersonic Mach numbers the percentage thrust increase and time to accelerate was less than for the supersonic accelerations. Savings in time to accelerate supersonically at maximum afterburner ranged from 4 to 7 percent. In general, the maximum thrust mode has performed well, demonstrating significant thrust increases at military and maximum afterburner power. Increases of up to 15 percent at typical combat-type flight conditions were identified. Thrust increases of this magnitude could be useful in a combat situation.

Evaluation of a ducted-fan power plant designed for high output and good cruise fuel economy

NASA Technical Reports Server (NTRS)

Behun, M; Rom, F E; Hensley, R V

1950-01-01

Theoretical analysis of performance of a ducted-fan power plant designed both for high-output, high-altitude operation at low supersonic Mach numbers and for good fuel economy at lower fight speeds is presented. Performance of ducted fan is compared with performance (with and without tail-pipe burner) of two hypothetical turbojet engines. At maximum power, the ducted fan has propulsive thrust per unit of frontal area between thrusts obtained by turbojet engines with and without tail-pipe burners. At cruise, the ducted fan obtains lowest thrust specific fuel consumption. For equal maximum thrusts, the ducted fan obtains cruising flight duration and range appreciably greater than turbojet engines.

Solar-Thermal Engine Testing

NASA Technical Reports Server (NTRS)

Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)

2001-01-01

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

Thrust Augmentation Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Pal, S.; Santoro, Robert J.; Shehadeh, R.; Saretto, S.; Lee, S.-Y.

2005-01-01

Thrust augmentation results of an ongoing study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) setup with various ejector configurations. The PDE used in these experiments utilizes ethylene (C2H4) as the fuel, and an equi-molar mixture of oxygen and nitrogen as the oxidizer at an equivalence ratio of one. High fidelity thrust measurements were made using an integrated spring damper system. The baseline thrust of the PDE engine was first measured and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The parameter space for the study included ejector length, PDE tube exit to ejector tube inlet overlap distance, and straight versus rounded ejector inlets. The relationship between the thrust augmentation results and various physical phenomena is described. To further understand the flow dynamics, shadow graph images of the exiting shock wave front from the PDE were also made. For the studied parameter space, the results showed a maximum augmentation of 40%. Further increase in augmentation is possible if the geometry of the ejector is tailored, a topic currently studied by numerous groups in the field.

Quiet Clean Short-haul Experimental Engine (QCSEE) under-the-wing engine simulation report

NASA Technical Reports Server (NTRS)

1977-01-01

Hybrid computer simulations of the under-the-wing engine were constructed to develop the dynamic design of the controls. The engine and control system includes a variable pitch fan and a digital electronic control. Simulation results for throttle bursts from 62 to 100 percent net thrust predict that the engine will accelerate 62 to 95 percent net thrust in one second.

Mariner Venus/Mercury 1973 rocket engine assembly

NASA Technical Reports Server (NTRS)

Snoke, D. R.; Williams, R. S.

1972-01-01

The fabrication and test of rocket engine assemblies (REA) for Mariner Venus/Mercury 1973 are reported. The fabrication, assembly and flight acceptance test of seven REA's including the type approval test of one engine and fabrication of one additional kit consisting of detail parts for an engine ready for catalyst loading are presented. The MV/M '73 REA which is a nominal 51 lbs thrust monopropellant engine is described. Under steady state operation the specific impulse is not less than 228 lb-sec at 55 lb and 218.5 lb-sec at 10 lb thrust varying linearly between these limits. The characteristic velocity is not less than 4100 ft/sec at any thrust level.

On spacecraft maneuvers control subject to propellant engine modes.

PubMed

Mazinan, A H

2015-09-01

The paper attempts to address a new control approach to spacecraft maneuvers based upon the modes of propellant engine. A realization of control strategy is now presented in engine on mode (high thrusts as well as further low thrusts), which is related to small angle maneuvers and engine off mode (specified low thrusts), which is also related to large angle maneuvers. There is currently a coarse-fine tuning in engine on mode. It is shown that the process of handling the angular velocities are finalized via rate feedback system in engine modes, where the angular rotations are controlled through quaternion based control (QBCL)strategy in engine off mode and these ones are also controlled through an optimum PID (OPIDH) strategy in engine on mode. Copyright © 2015 ISA. Published by Elsevier Ltd. All rights reserved.

JT8D-100 turbofan engine, phase 1. [noise reduction

NASA Technical Reports Server (NTRS)

1974-01-01

The JT8D turbofan engine, widely used in short and medium range transport aircraft, contributes substantially to airport community noise. The jet noise is predominant in the JT8D engine and may be reduced in a modified engine, without loss of thrust, by increasing the airflow to reduce jet velocity. A configuration study evaluated the effects of fan airflow, fan pressure ratio, and bypass ratio on noise, thrust, and fuel comsumption. The cycle selected for the modified engine was based upon an increased diameter, single-stage fan and two additional core engine compressor stages, which replace the existing two-stage fan. Modifications were also made to the low pressure turbine to provide the increased torque required by the larger diameter fan. The resultant JT8D-100 engine models have the following characteristics at take-off thrust, compared to the current JT8D engine: Airflow and bypass ratio are increased, and fan pressure ratio and engine speed are reduced. The resultant engine is also longer, larger in diameter, and heavier than the JT8D base model, but these latter changes are compensated by the increased thrust and decreased fuel comsumption of the modified engine, thus providing the capability for maintaining the performance of the current JT8D-powered aircraft.

Plasma Engines,

DTIC Science & Technology

1982-09-08

low thrust, long duration power device, the plasma engine 6 has certain distinct advantages. For a chemical fuel rocket engine , a thrust of M.’)1...PLASMA ENGINES.CU) UNCLASSZICD FTO-ZIftS)T-0636-98 NL * UUUUU UUMile ~ FTD-ID(RS)T-0636-82 FOREIGN TECHNOLOGY DIVISION q 14 PLASMA ENGINES bv Sung...8 September 1982 MICROFICHE NR: FTD-82-C-001198 PLASMA ENGINES By: Sung Yuyang English pages: 7 Source: Hangkong Zhishi, March 1982, pp. 12-13 Country

Flight evaluation of an extended engine life mode on an F-15 airplane

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Conners, Timothy R.

1992-01-01

An integrated flight and propulsion control system designed to reduce the rate of engine deterioration was developed and evaluated in flight on the NASA Dryden F-15 research aircraft. The extended engine life mode increases engine pressure ratio while reducing engine airflow to lower the turbine temperature at constant thrust. The engine pressure ratio uptrim is modulated in real time based on airplane maneuver requirements, flight conditions, and engine information. The extended engine life mode logic performed well, significantly reducing turbine operating temperature. Reductions in fan turbine inlet temperature of up to 80 F were obtained at intermediate power and up to 170 F at maximum augmented power with no appreciable loss in thrust. A secondary benefit was the considerable reduction in thrust-specific fuel consumption. The success of the extended engine life mode is one example of the advantages gained from integrating aircraft flight and propulsion control systems.

A guidance and navigation system for continuous low thrust vehicles. M.S. Thesis

NASA Technical Reports Server (NTRS)

Tse, C. J. C.

1973-01-01

A midcourse guidance and navigation system for continuous low thrust vehicles is described. A set of orbit elements, known as the equinoctial elements, are selected as the state variables. The uncertainties are modelled statistically by random vector and stochastic processes. The motion of the vehicle and the measurements are described by nonlinear stochastic differential and difference equations respectively. A minimum time nominal trajectory is defined and the equation of motion and the measurement equation are linearized about this nominal trajectory. An exponential cost criterion is constructed and a linear feedback guidance law is derived to control the thrusting direction of the engine. Using this guidance law, the vehicle will fly in a trajectory neighboring the nominal trajectory. The extended Kalman filter is used for state estimation. Finally a short mission using this system is simulated. The results indicate that this system is very efficient for short missions.

Saturn Apollo Program

NASA Image and Video Library

1965-03-01

This photograph shows F-1 engines being stored in the F-1 Engine Preparation Shop, building 4666, at the Marshall Space Flight Center. Each F-1 engine produced a thrust of 1,500,000 pounds. A cluster of five engines was mounted on the thrust structure of the S-IC stage of a 364-foot long Saturn V launch vehicle that ultimately took astronauts to the Moon.

n/a

NASA Image and Video Library

1961-01-01

The static firing of a Saturn F-1 engine at the Marshall Space Flight Center's Static Test Stand. The F-1 engine is a single-start, 1,5000,000 Lb fixed-thrust, bipropellant rocket system. The engine uses liquid oxygen as the oxidizer and RP-1 (kerosene) as fuel. The five-engine cluster used on the first stage of the Saturn V produces 7,500,000 lbs of thrust.

Dynamics of high-bypass-engine thrust reversal using a variable-pitch fan

NASA Technical Reports Server (NTRS)

Schaefer, J. W.; Sagerser, D. R.; Stakolich, E. G.

1977-01-01

The test program demonstrated that successful and rapid forward-to reverse-thrust transients can be performed without any significant engine operational limitations for fan blade pitch changes through either feather pitch or flat pitch. For through-feather-pitch operation with a flight inlet, fan stall problems were encountered, and a fan blade overshoot technique was used to establish reverse thrust.

Model-Based Fault Tolerant Control

NASA Technical Reports Server (NTRS)

Kumar, Aditya; Viassolo, Daniel

2008-01-01

The Model Based Fault Tolerant Control (MBFTC) task was conducted under the NASA Aviation Safety and Security Program. The goal of MBFTC is to develop and demonstrate real-time strategies to diagnose and accommodate anomalous aircraft engine events such as sensor faults, actuator faults, or turbine gas-path component damage that can lead to in-flight shutdowns, aborted take offs, asymmetric thrust/loss of thrust control, or engine surge/stall events. A suite of model-based fault detection algorithms were developed and evaluated. Based on the performance and maturity of the developed algorithms two approaches were selected for further analysis: (i) multiple-hypothesis testing, and (ii) neural networks; both used residuals from an Extended Kalman Filter to detect the occurrence of the selected faults. A simple fusion algorithm was implemented to combine the results from each algorithm to obtain an overall estimate of the identified fault type and magnitude. The identification of the fault type and magnitude enabled the use of an online fault accommodation strategy to correct for the adverse impact of these faults on engine operability thereby enabling continued engine operation in the presence of these faults. The performance of the fault detection and accommodation algorithm was extensively tested in a simulation environment.

Development of a CFRP Engine Thrust Frame for the Next Generation Launchers

NASA Astrophysics Data System (ADS)

Fatemi, Javad; van der Bas, Finn; Cruijssen, Henk

2012-07-01

This paper addresses the activities related to the development of technologies for a composite Engine Thrust Frame (ETF) for the next generation launchers. In particular, the design and analyses of a full Carbon Fibre Reinforced Plastic (CFRP) engine thrust frame are presented in more detail. The ETF concept is composed of three main parts, i.e. an aluminium top-ring which connects the ETF to the upper-stage tank, a CFRP cone, and a CFRP cone-cap which connects the Vinci engine to the ETF. The main challenging requirements for development of a CFRP ETF are recalled. The ETF concept and its mechanical performances are assessed.

Cruise Missile Engines

NASA Technical Reports Server (NTRS)

1982-01-01

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.

A Layman's Guide to Thrust Engine Development for Super/Hyper Sonic Flight.

ERIC Educational Resources Information Center

Thiesse, James L.

The intention of this paper is to discuss the advances in thrust engines from the initial development of the J58/SR-71 (JT11D-20) of the U.S. Air Force's SR-71 Blackbird to the development of the RAM and SCRAM engines necessary to propel the new generations of high-flying super-speed aircraft. Engineering complexities suggest that the engines and…

A Computational Study to Investigate the Effect of Altitude on Deteriorated Engine Performance

NASA Astrophysics Data System (ADS)

Koh, W. C.; Mazlan, N. M.; Rajendran, P.; Ismail, M. A.

2018-05-01

This study presents an investigation on the effect of operational altitudes on the performance of the deteriorated engine. A two-spool high bypass ratio turbofan engine is used as the test subject for this study. The engine is modelled in Gas Turbine Simulation Program (GSP) based on an existing engine model from literature. Real flight data were used for the validation. Deterioration rate of 0.1% per day is applied for all turbofan components engine. The simulation is performed by varying the altitude from sea level until 9000m. Results obtained show reduction in air mass flow rate and engine thrust as altitude increases. The reduction in air mass flow rate is due to the lower air density at higher altitude hence reduces amount of engine thrust. At 1000m to 4000m, thrust specific fuel consumption (TSFC) of the engine is improved compared to sea level. However depleted in TSFC is shown when the aircraft flies at altitude higher than 4000m. At this altitude, the effect of air density is dominant. As a result, the engine is required to burn more fuel to provide a higher thrust to sustain the aircraft speed. More fuel is consumed hence depletion in TSFC is obtained.

NACA Conference on Turbojet-Engine Thrust Augmentation Research: A Compilation of the Papers Presented by NACA Staff Members

NASA Technical Reports Server (NTRS)

1948-01-01

The conference on Turbojet-Engine Thrust-Augmentation Research was organized by the NACA to present in summarized form the results of the latest experimental and analytical investigations conducted at the Lewis Flight Propulsion Laboratory on methods of augmenting the thrust of turbojet engines. The technical discussions are reproduced herewith in the same form in which they were presented. The original presentation in this record are considered as complementary to, rather than substitutes for, the committee's system of complete and formal reports.

Advances in Thrust-Based Emergency Control of an Airplane

NASA Technical Reports Server (NTRS)

Creech, Gray; Burken, John J.; Burcham, Bill

2003-01-01

Engineers at NASA's Dryden Flight Research Center have received a patent on an emergency flight-control method implemented by a propulsion-controlled aircraft (PCA) system. Utilizing the preexisting auto-throttle and engine-pressure-ratio trim controls of the airplane, the PCA system provides pitch and roll control for landing an airplane safely without using aerodynamic control surfaces that have ceased to function because of a primary-flight-control-system failure. The installation of the PCA does not entail any changes in pre-existing engine hardware or software. [Aspects of the method and system at previous stages of development were reported in Thrust-Control System for Emergency Control of an Airplane (DRC-96-07), NASA Tech Briefs, Vol. 25, No. 3 (March 2001), page 68 and Emergency Landing Using Thrust Control and Shift of Weight (DRC-96-55), NASA Tech Briefs, Vol. 26, No. 5 (May 2002), page 58.]. Aircraft flight-control systems are designed with extensive redundancy to ensure low probabilities of failure. During recent years, however, several airplanes have exhibited major flight-control-system failures, leaving engine thrust as the last mode of flight control. In some of these emergency situations, engine thrusts were successfully modulated by the pilots to maintain flight paths or pitch angles, but in other situations, lateral control was also needed. In the majority of such control-system failures, crashes resulted and over 1,200 people died. The challenge lay in creating a means of sufficient degree of thrust-modulation control to safely fly and land a stricken airplane. A thrust-modulation control system designed for this purpose was flight-tested in a PCA an MD-11 airplane. The results of the flight test showed that without any operational control surfaces, a pilot can land a crippled airplane (U.S. Patent 5,330,131). The installation of the original PCA system entailed modifications not only of the flight-control computer (FCC) of the airplane but also of each engine-control computer. Inasmuch as engine-manufacturer warranties do not apply to modified engines, the challenge became one of creating a PCA system that does not entail modifications of the engine computers.

Full scale technology demonstration of a modern counterrotating unducted fan engine concept. Engine test

NASA Technical Reports Server (NTRS)

1987-01-01

The Unducted Fan (UDF) engine is an innovative aircraft engine concept based on an ungeared, counterrotating, unducted, ultra-high-bypass turbofan configuration. This engine is being developed to provide a high thrust-to-weight ratio power plant with exceptional fuel efficiency for subsonic aircraft application. This report covers the successful ground testing of this engine. A test program exceeding 100-hr duration was completed, in which all the major goals were achieved. The following accomplishments were demonstrated: (1) full thrust (25,000 lb); (2) full counterrotating rotor speeds (1393+ rpm); (3) low specific fuel consumption (less than 0.24 lb/hr/lb); (4) new composite fan design; (5) counterrotation of structures, turbines, and fan blades; (6) control system; (7) actuation system; and (8) reverse thrust.

The 727 airplane target thrust reverser static performance model test for refanned JT8D engines

NASA Technical Reports Server (NTRS)

Chow, C. T. P.; Atkey, E. N.

1974-01-01

The results of a scale model static performance test of target thrust reverser configurations for the Pratt and Whitney Aircraft JT8D-100 series engine are presented. The objective of the test was to select a series of suitable candidate reverser configurations for the subsequent airplane model wind tunnel ingestion and flight controls tests. Test results indicate that adequate reverse thrust performance with compatible engine airflow match is achievable for the selected configurations. Tapering of the lips results in loss of performance and only minimal flow directivity. Door pressure surveys were conducted on a selected number of lip and fence configurations to obtain data to support the design of the thrust reverser system.

Low-thrust chemical rocket engine study

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

MD-11 PCA - View of aircraft on ramp

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 is taxiing to a position on the flightline at NASA's Dryden Flight Research Center, Edwards, California, following its completion of the first and second landings ever performed by a transport aircraft under engine power only (on Aug. 29, 1995). The milestone flight, with NASA research pilot and former astronaut Gordon Fullerton at the controls, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple. For pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

A transport aircraft lands for the first time under engine power only, as this McDonnell Douglas MD-11 touches down at 11:38 a.m., Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - Closeup view of aircraft on ramp

NASA Technical Reports Server (NTRS)

1995-01-01

This McDonnell Douglas MD-11 has taxied to a position on the flightline at NASA's Dryden Flight Research Center, Edwards, California, following its completion of the first and second landings ever performed by a transport aircraft under engine power only (on Aug. 29, 1995). The milestone flight, with NASA research pilot and former astronaut Gordon Fullerton at the controls, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple. For pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

MD-11 PCA - First Landing at Edwards

NASA Technical Reports Server (NTRS)

1995-01-01

A transport aircraft lands for the first time under engine power only, as this McDonnell Douglas MD-11 touches down at 11:38 a.m., Aug. 29, 1995, at NASA's Dryden Flight Research Center, Edwards, California. The milestone flight, flown by NASA research pilot and former astronaut Gordon Fullerton, was part of a NASA project to develop a computer-assisted engine control system that enables a pilot to land a plane safely when its normal control surfaces are disabled. The Propulsion-Controlled Aircraft (PCA) system uses standard autopilot controls already present in the cockpit, together with the new programming in the aircraft's flight control computers. The PCA concept is simple--for pitch control, the program increases thrust to climb and reduces thrust to descend. To turn right, the autopilot increases the left engine thrust while decreasing the right engine thrust. The initial Propulsion-Controlled Aircraft studies by NASA were carried out at Dryden with a modified twin-engine F-15 research aircraft.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Jr., Lee H. (Inventor)

1991-01-01

The invention is a method and system for monitoring and directly displaying the actual thrust produced by a jet aircraft engine under determined operating conditions and the available thrust and predicted (commanded) thrust of a functional model of an ideal engine under the same determined operating conditions. A first set of actual value output signals representative of a plurality of actual performance parameters of the engine under the determined operating conditions is generated and compared with a second set of predicted value output signals representative of the predicted value of corresponding performance parameters of a functional model of the engine under the determined operating conditions to produce a third set of difference value output signals within a range of normal, caution, or warning limit values. A thrust indicator displays when any one of the actual value output signals is in the warning range while shaping function means shape each of the respective difference output signals as each approaches the limit of the respective normal, caution, and warning range limits.

Regeneratively Cooled Liquid Oxygen/Methane Technology Development Between NASA MSFC and PWR

NASA Technical Reports Server (NTRS)

Robinson, Joel W.; Greene, Christopher B.; Stout, Jeffrey B.

2012-01-01

The National Aeronautics & Space Administration (NASA) has identified Liquid Oxygen (LOX)/Liquid Methane (LCH4) as a potential propellant combination for future space vehicles based upon exploration studies. The technology is estimated to have higher performance and lower overall systems mass compared to existing hypergolic propulsion systems. NASA-Marshall Space Flight Center (MSFC) in concert with industry partner Pratt & Whitney Rocketdyne (PWR) utilized a Space Act Agreement to test an oxygen/methane engine system in the Summer of 2010. PWR provided a 5,500 lbf (24,465 N) LOX/LCH4 regenerative cycle engine to demonstrate advanced thrust chamber assembly hardware and to evaluate the performance characteristics of the system. The chamber designs offered alternatives to traditional regenerative engine designs with improvements in cost and/or performance. MSFC provided the test stand, consumables and test personnel. The hot fire testing explored the effective cooling of one of the thrust chamber designs along with determining the combustion efficiency with variations of pressure and mixture ratio. The paper will summarize the status of these efforts.

A supersonic fan equipped variable cycle engine for a Mach 2.7 supersonic transport

NASA Technical Reports Server (NTRS)

Tavares, T. S.

1985-01-01

The concept of a variable cycle turbofan engine with an axially supersonic fan stage as powerplant for a Mach 2.7 supersonic transport was evaluated. Quantitative cycle analysis was used to assess the effects of the fan inlet and blading efficiencies on engine performance. Thrust levels predicted by cycle analysis are shown to match the thrust requirements of a representative aircraft. Fan inlet geometry is discussed and it is shown that a fixed geometry conical spike will provide sufficient airflow throughout the operating regime. The supersonic fan considered consists of a single stage comprising a rotor and stator. The concept is similar in principle to a supersonic compressor, but differs by having a stator which removes swirl from the flow without producing a net rise in static pressure. Operating conditions peculiar to the axially supersonic fan are discussed. Geometry of rotor and stator cascades are presented which utilize a supersonic vortex flow distribution. Results of a 2-D CFD flow analysis of these cascades are presented. A simple estimate of passage losses was made using empirical methods.

Innovative Double Bypass Engine for Increased Performance

NASA Astrophysics Data System (ADS)

Manoharan, Sanjivan

Engines continue to grow in size to meet the current thrust requirements of the civil aerospace industry. Large engines pose significant transportation problems and require them to be split in order to be shipped. Thus, large amounts of time have been spent in researching methods to increase thrust capabilities while maintaining a reasonable engine size. Unfortunately, much of this research has been focused on increasing the performance and efficiencies of individual components while limited research has been done on innovative engine configurations. This thesis focuses on an innovative engine configuration, the High Double Bypass Engine, aimed at increasing fuel efficiency and thrust while maintaining a competitive fan diameter and engine length. The 1-D analysis was done in Excel and then compared to the results from Numerical Propulsion Simulation System (NPSS) software and were found to be within 4% error. Flow performance characteristics were also determined and validated against their criteria.

Oxidizer heat exchanger component test

NASA Technical Reports Server (NTRS)

Kanic, P. G.

1988-01-01

The RL10-IIB engine, is capable of multimode thrust operation. The engine operates at two low-thrust levels: tank head idle (THI), approximately 1 to 2 percent of full thrust; and pumped idle, 10 percent of full thrust. Operation at THI provides vehicle propellant settling thrust and efficient thermal conditioning; PI operation provides vehicle tank prepressurization and maneuver thrust for low-g deployment. Stable combustion of the RL10-IIB engine during the low-thrust operating modes can be accomplished by using a heat exchanger to supply gaseous oxygen to the propellant injector. The oxidized heat exchanger (OHE) vaporizes the liquid oxygen using hydrogen as the energy source. This report summarizes the test activity and post-test data analysis for two possible heat exchangers, each of which employs a completely different design philosophy. One design makes use of a low-heat transfer (PHT) approach in combination with a volume to attenuate pressure and flow oscillations. The test data showed that the LHT unit satisfied the oxygen exit quality of 0.95 or greater in both the THI and PI modes while maintaining stability. The HHT unit fulfilled all PI requirements; data for THI satisfactory operation is implied from experimental data that straddle the exact THI operating point.

Wind tunnel test of model target thrust reversers for the Pratt and Whitney aircraft JT8D-100 series engines installed on a 727-200 airplane

NASA Technical Reports Server (NTRS)

Hambly, D.

1974-01-01

The results of a low speed wind tunnel test of 0.046 scale model target thrust reversers installed on a 727-200 model airplane are presented. The full airplane model was mounted on a force balance, except for the nacelles and thrust reversers, which were independently mounted and isolated from it. The installation had the capability of simulating the inlet airflows and of supplying the correct proportions of primary and secondary air to the nozzles. The objectives of the test were to assess the compatibility of the thrust reversers target door design with the engine and airplane. The following measurements were made: hot gas ingestion at the nacelle inlets; model lift, drag, and pitching moment; hot gas impingement on the airplane structure; and qualitative assessment of the rudder effectiveness. The major parameters controlling hot gas ingestion were found to be thrust reverser orientation, engine power setting, and the lip height of the bottom thrust reverser doors on the side nacelles. The thrust reversers tended to increase the model lift, decrease the drag, and decrease the pitching moment.

A Regeneratively Cooled Thrust Chamber For The Fastrac Engine

NASA Technical Reports Server (NTRS)

Brown, Kendall K.; Sparks, Dave; Woodcock, Gordon

2000-01-01

Abstract This paper presents the development of a low-cost, regeneratively-cooled thrust chamber for the Fastrac engine. The chamber was fabricated using hydraformed copper tubing to form the coolant jacket and wrapped with a fiber reinforced polymer composite Material to form a structural jacket. The thrust chamber design and fabrication approach was based upon Space America. Inc.'s 12,000 lb regeneratively-cooled LOX/kerosene rocket engine. Fabrication of regeneratively cooled thrust chambers by tubewall construction dates back to the early US ballistic missile programs. The most significant innovations in this design was the development of a low-cost process for fabrication from copper tubing (nickel alloy was the usual practice) and use of graphite composite overwrap as the pressure containment, which yields an easily fabricated, lightweight pressure jacket around the copper tubes A regeneratively-cooled reusable thrust chamber can benefit the Fastrac engine program by allowing more efficient (cost and scheduler testing). A proof-of-concept test article has been fabricated and will he tested at Marshall Space Flight Center in the late Summer or Fall of 2000.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

NASA Astrophysics Data System (ADS)

Tian, Lu; Chen, Li-Hong; Chen, Qiang; Zhong, Feng-Quan; Chang, Xin-Yu

2016-02-01

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5-3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Characterization of advanced electric propulsion systems

NASA Technical Reports Server (NTRS)

Ray, P. K.

1982-01-01

Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

Emergency Flight Control Using Only Engine Thrust and Lateral Center-of-Gravity Offset: A First Look

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Burken, John; Maine, Trindel A.; Bull, John

1997-01-01

Normally, the damage that results in a total loss of the primary flight controls of a jet transport airplane, including all engines on one side, would be catastrophic. In response, NASA Dryden has conceived an emergency flight control system that uses only the thrust of a wing-mounted engine along with a lateral center-of-gravity (CGY) offset from fuel transfer. Initial analysis and simulation studies indicate that such a system works, and recent high-fidelity simulation tests on the MD-11 and B-747 suggest that the system provides enough control for a survivable landing. This paper discusses principles of flight control using only a wing engine thrust and CGY offset, along with the amount of CGY offset capability of some transport airplanes. The paper also presents simulation results of the throttle-only control capability and closed-loop control of ground track using computer-controlled thrust.

4. COMPLETE X15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

4. COMPLETE X-15 VEHICLE TEST STAND, DETAIL OF THRUST MOUNTING STRUCTURE AT ENGINE END OF PLANE. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

Low thrust vehicle concept study

NASA Technical Reports Server (NTRS)

1980-01-01

Low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low acceleration orbit transfer of large space systems were defined. Results indicate that it is cost effective and least risk to combine the OTV and stowed spacecraft in a single 65 K Shuttle. The study shows that the engine for an optimized low thrust stage (1) does not require very low thrust; (2) 1-3 K thrust range appears optimum; (3) thrust transient is not a concern; (4) throttling probably not worthwhile; and (5) multiple thrusters complicate OTV/LSS design and aggravate LSS loads. Regarding the optimum vehicle for low acceleration missions, the single shuttle launch (LSS and expendable OTV) is most cost effective and least risky. Multiple shuttles increase diameter 20%. The space based radar structure short OTV (which maximizes space available for packaged LSS) favors use of torus tank. Propellant tank pressures/vapor residuals are little affected by engine thrust level or number of burns.

Improving Space Project Cost Estimating with Engineering Management Variables

NASA Technical Reports Server (NTRS)

Hamaker, Joseph W.; Roth, Axel (Technical Monitor)

2001-01-01

Current space project cost models attempt to predict space flight project cost via regression equations, which relate the cost of projects to technical performance metrics (e.g. weight, thrust, power, pointing accuracy, etc.). This paper examines the introduction of engineering management parameters to the set of explanatory variables. A number of specific engineering management variables are considered and exploratory regression analysis is performed to determine if there is statistical evidence for cost effects apart from technical aspects of the projects. It is concluded that there are other non-technical effects at work and that further research is warranted to determine if it can be shown that these cost effects are definitely related to engineering management.

Study of aerodynamic technology for VSTOL fighter/attack aircraft, phase 1

NASA Technical Reports Server (NTRS)

Driggers, H. H.

1978-01-01

A conceptual design study was performed of a vertical attitude takeoff and landing (VATOL) fighter/attack aircraft. The configuration has a close-coupled canard-delta wing, side two-dimensional ramp inlets, and two augmented turbofan engines with thrust vectoring capability. Performance and sensitivities to objective requirements were calculated. Aerodynamic characteristics were estimated based on contractor and NASA wind tunnel data. Computer simulations of VATOL transitions were performed. Successful transitions can be made, even with series post-stall instabilities, if reaction controls are properly phased. Principal aerodynamic uncertainties identified were post-stall aerodynamics, transonic aerodynamics with thrust vectoring and inlet performance in VATOL transition. A wind tunnel research program was recommended to resolve the aerodynamic uncertainties.

Evaluation of Fuel Character Effects on J79 Engine Combustion System

DTIC Science & Technology

1979-06-01

A. Overall Engine Description The J79 engine is a lightweight, high-thrust, axial - flow turbojet engine with variable afterburner thrust. This engine...thimbles are arranged to provide flow patterns for flame stabilization in the primary zone and mixing and turbine inlet temperature profile control at...measured with stainard )S𔃾Z orifices- Fuel flow races uere measured with calibrated turbine flotaMcers corrected for the density aan viscosity of each

Multiphysics Computational Analysis of a Solid-Core Nuclear Thermal Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Canabal, Francisco; Cheng, Gary; Chen, Yen-Sen

2007-01-01

The objective of this effort is to develop an efficient and accurate computational heat transfer methodology to predict thermal, fluid, and hydrogen environments for a hypothetical solid-core, nuclear thermal engine - the Small Engine. In addition, the effects of power profile and hydrogen conversion on heat transfer efficiency and thrust performance were also investigated. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics platform, while formulations of conjugate heat transfer were implemented to describe the heat transfer from solid to hydrogen inside the solid-core reactor. The computational domain covers the entire thrust chamber so that the afore-mentioned heat transfer effects impact the thrust performance directly. The result shows that the computed core-exit gas temperature, specific impulse, and core pressure drop agree well with those of design data for the Small Engine. Finite-rate chemistry is very important in predicting the proper energy balance as naturally occurring hydrogen decomposition is endothermic. Locally strong hydrogen conversion associated with centralized power profile gives poor heat transfer efficiency and lower thrust performance. On the other hand, uniform hydrogen conversion associated with a more uniform radial power profile achieves higher heat transfer efficiency, and higher thrust performance.

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bruce G. Schnitzler; Stanley K. Borowski

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of both the Strategic Defense Initiative (SDI) and Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified asmore » the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.« less

Small Fast Spectrum Reactor Designs Suitable for Direct Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Schnitzler, Bruce G.; Borowski, Stanley K.

2012-01-01

Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. Past studies, in particular those in support of the Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. The recent NASA Design Reference Architecture (DRA) 5.0 Study re-examined mission, payload, and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal propulsion was again identified as the preferred in-space transportation system. A common nuclear thermal propulsion stage with three 25,000-lbf thrust engines was used for all primary mission maneuvers. Moderately lower thrust engines may also have important roles. In particular, lower thrust engine designs demonstrating the critical technologies that are directly extensible to other thrust levels are attractive from a ground testing perspective. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. Both graphite and refractory metal alloy fuel types were pursued. Reactors and engines employing graphite based fuels were designed, built and ground tested. A number of fast spectrum reactor and engine designs employing refractory metal alloy fuel types were proposed and designed, but none were built. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art graphite based fuel design incorporating lessons learned from the very successful technology development program. The SNRE was a nominal 16,000-lbf thrust engine originally intended for unmanned applications with relatively short engine operations and the engine and stage design were constrained to fit within the payload volume of the then planned space shuttle. The SNRE core design utilized hexagonal fuel elements and hexagonal structural support elements. The total number of elements can be varied to achieve engine designs of higher or lower thrust levels. Some variation in the ratio of fuel elements to structural elements is also possible. Options for SNRE-based engine designs in the 25,000-lbf thrust range were described in a recent (2010) Joint Propulsion Conference paper. The reported designs met or exceeded the performance characteristics baselined in the DRA 5.0 Study. Lower thrust SNRE-based designs were also described in a recent (2011) Joint Propulsion Conference paper. Recent activities have included parallel evaluation and design efforts on fast spectrum engines employing refractory metal alloy fuels. These efforts include evaluation of both heritage designs from the Argonne National Laboratory (ANL) and General Electric Company GE-710 Programs as well as more recent designs. Results are presented for a number of not-yet optimized fast spectrum engine options.

Orbit transfer vehicle engine study. Volume 2: Technical report

NASA Technical Reports Server (NTRS)

1980-01-01

The orbit transfer vehicle (OTV) engine study provided parametric performance, engine programmatic, and cost data on the complete propulsive spectrum that is available for a variety of high energy, space maneuvering missions. Candidate OTV engines from the near term RL 10 (and its derivatives) to advanced high performance expander and staged combustion cycle engines were examined. The RL 10/RL 10 derivative performance, cost and schedule data were updated and provisions defined which would be necessary to accommodate extended low thrust operation. Parametric performance, weight, envelope, and cost data were generated for advanced expander and staged combustion OTV engine concepts. A prepoint design study was conducted to optimize thrust chamber geometry and cooling, engine cycle variations, and controls for an advanced expander engine. Operation at low thrust was defined for the advanced expander engine and the feasibility and design impact of kitting was investigated. An analysis of crew safety and mission reliability was conducted for both the staged combustion and advanced expander OTV engine candidates.

75 FR 14377 - Airworthiness Directives; Pratt & Whitney (PW) Model PW2037, PW2037(M), and PW2040 Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-25

...The FAA proposes to adopt a new airworthiness directive (AD) for PW Model PW2037, PW2037(M), and PW2040 turbofan engines. This proposed AD would require removing erosion damage on fan blades with cutback leading edges and restoring the leading edge contour. This proposed AD results from reports from PW that fan blade leading edge erosion can result in a fan thrust deterioration mode (FTDM) condition, which reduces the engine's capability of producing full rated take-off thrust. We are proposing this AD to prevent loss of engine thrust from an FTDM condition, which could result in an inability to maintain safe flight.

Performance (Off-Design) Cycle Analysis for a Turbofan Engine With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This report presents the performance of a steady-state, dual-spool, separate-exhaust turbofan engine, with an interstage turbine burner (ITB) serving as a secondary combustor. The ITB, which is located in the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet-engine propulsion. A detailed off-design performance analysis of ITB engines is written in Microsoft(Registered Trademark) Excel (Redmond, Washington) macrocode with Visual Basic Application to calculate engine performances over the entire operating envelope. Several design-point engine cases are pre-selected using a parametric cycle-analysis code developed previously in Microsoft(Registered Trademark) Excel, for off-design analysis. The off-design code calculates engine performances (i.e. thrust and thrust-specific-fuel-consumption) at various flight conditions and throttle settings.

Attitude Control Flight Experience: Coping with Solar Radiation and Ion Engines Leak Thrust in Hayabusa (MUSES-C)

NASA Technical Reports Server (NTRS)

Kawaguchi, Jun'ichiro; Kominato, Takashi; Shirakawa, Ken'ichi

2007-01-01

The paper presents the attitude reorientation taking the advantage of solar radiation pressure without use of any fuel aboard. The strategy had been adopted to make Hayabusa spacecraft keep pointed toward the Sun for several months, while spinning. The paper adds the above mentioned results reported in Sedona this February showing another challenge of combining ion engines propulsion tactically balanced with the solar radiation torque with no spin motion. The operation has been performed since this March for a half year successfully. The flight results are presented with the estimated solar array panel diffusion coefficient and the ion engine's swirl torque.

Performance Cycle Analysis of a Two-Spool, Separate-Exhaust Turbofan With Interstage Turbine Burner

NASA Technical Reports Server (NTRS)

Liew, K. H.; Urip, E.; Yang, S. L.; Mattingly, J. D.; Marek, C. J.

2005-01-01

This paper presents the performance cycle analysis of a dual-spool, separate-exhaust turbofan engine, with an Interstage Turbine Burner serving as a secondary combustor. The ITB, which is located at the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet engine propulsion. A detailed performance analysis of this engine has been conducted for steady-state engine performance prediction. A code is written and is capable of predicting engine performances (i.e., thrust and thrust specific fuel consumption) at varying flight conditions and throttle settings. Two design-point engines were studied to reveal trends in performance at both full and partial throttle operations. A mission analysis is also presented to assure the advantage of saving fuel by adding ITB.

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2012 CFR

2012-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2014 CFR

2014-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2010 CFR

2010-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2013 CFR

2013-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

14 CFR 33.79 - Fuel burning thrust augmentor.

Code of Federal Regulations, 2011 CFR

2011-01-01

... thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of... range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to...

Preliminary flight evaluation of an engine performance optimization algorithm

NASA Technical Reports Server (NTRS)

Lambert, H. H.; Gilyard, G. B.; Chisholm, J. D.; Kerr, L. J.

1991-01-01

A performance seeking control (PSC) algorithm has undergone initial flight test evaluation in subsonic operation of a PW 1128 engined F-15. This algorithm is designed to optimize the quasi-steady performance of an engine for three primary modes: (1) minimum fuel consumption; (2) minimum fan turbine inlet temperature (FTIT); and (3) maximum thrust. The flight test results have verified a thrust specific fuel consumption reduction of 1 pct., up to 100 R decreases in FTIT, and increases of as much as 12 pct. in maximum thrust. PSC technology promises to be of value in next generation tactical and transport aircraft.

Solar thermal rocket engine (STRE) thrust characteristics at the change of engine operation mode and of the flight vehicle attitude in the solar system

NASA Astrophysics Data System (ADS)

Kudrin, O. I.

1993-10-01

Relationships are presented which describe changes in the thrust and specific impulse of a solar thermal rocket engine due to a change in the flow rate of the working fluid (hydrogen). Expressions are also presented which describe the variation of the STRE thrust and specific impulse with the distance between the flight vehicle and the sun. Results of calculations are presented for an STRE with afterburning of the working fluid (hydrogen + oxygen) using hydrogen heating by solar energy to a temperature of 2360 K.

Nozzle Side Load Testing and Analysis at Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Ruf, Joseph H.; McDaniels, David M.; Brown, Andrew M.

2009-01-01

Realistic estimates of nozzle side loads, the off-axis forces that develop during engine start and shutdown, are important in the design cycle of a rocket engine. The estimated magnitude of the nozzle side loads has a large impact on the design of the nozzle shell and the engine s thrust vector control system. In 2004 Marshall Space Flight Center (MSFC) began developing a capability to quantify the relative magnitude of side loads caused by different types of nozzle contours. The MSFC Nozzle Test Facility was modified to measure nozzle side loads during simulated nozzle start. Side load results from cold flow tests on two nozzle test articles, one with a truncated ideal contour and one with a parabolic contour are provided. The experimental approach, nozzle contour designs and wall static pressures are also discussed

Design issues for lunar in situ aluminum/oxygen propellant rocket engines

NASA Technical Reports Server (NTRS)

Meyer, Michael L.

1992-01-01

Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

Dynamic Imbalance Would Counter Offcenter Thrust

NASA Technical Reports Server (NTRS)

Mccanna, Jason

1994-01-01

Dynamic imbalance generated by offcenter thrust on rotating body eliminated by shifting some of mass of body to generate opposing dynamic imbalance. Technique proposed originally for spacecraft including massive crew module connected via long, lightweight intermediate structure to massive engine module, such that artificial gravitation in crew module generated by rotating spacecraft around axis parallel to thrust generated by engine. Also applicable to dynamic balancing of rotating terrestrial equipment to which offcenter forces applied.

Direct Fusion Drive for a Human Mars Orbital Mission

DOE Office of Scientific and Technical Information (OSTI.GOV)

Paluszek, Michael; Pajer, Gary; Razin, Yosef

2014-08-01

The Direct Fusion Drive (DFD) is a nuclear fusion engine that produces both thrust and electric power. It employs a field reversed configuration with an odd-parity rotating magnetic field heating system to heat the plasma to fusion temperatures. The engine uses deuterium and helium-3 as fuel and additional deuterium that is heated in the scrape-off layer for thrust augmentation. In this way variable exhaust velocity and thrust is obtained.

Using Engine Thrust for Emergency Flight Control: MD-11 and B-747 Results

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Maine, Trindel A.; Burken, John J.; Bull, John

1998-01-01

With modern digital control systems, using engine thrust for emergency flight control to supplement or replace failed aircraft normal flight controls has become a practical consideration. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control. An F-15 and an MD-11 airplane have been landed without using any flight control surfaces. Preliminary studies have also been conducted that show that engines on only one wing can provide some flight control capability if the lateral center of gravity can be shifted toward the side of the airplane that has the operating engine(s). Simulator tests of several airplanes with no flight control surfaces operating and all engines out on the left wing have all shown positive control capability within the available range of lateral center-of-gravity offset. Propulsion-controlled aircraft systems that can operate without modifications to engine control systems, thus allowing PCA technology to be installed on less capable airplanes or at low cost, are also desirable. Further studies have examined simplified 'PCA Lite' and 'PCA Ultralite' concepts in which thrust control is provided by existing systems such as auto-throttles or a combination of existing systems and manual pilot control.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 Engine major components with callouts (chart 1): The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

2004-04-15

H-1 engine major components with callouts (chart 1). The H-1 engine was used in a cluster of eight on the the first stage of Saturn I (S-I stage) and Saturn IB (S-IB stage). The engines were arranged in a double pattern: four engines, located inboard, were fixed in a square pattern around the stage axis, while the remaining four engines were located outboard in a larger square pattern and each outer engine was gimbaled. Each H-1 engine had a thrust of 188,000 pounds for a combined thrust of over 1,500,000 pounds.

Airflow and thrust calibration of an F100 engine, S/N P680059, at selected flight conditions

NASA Technical Reports Server (NTRS)

Biesiadny, T. J.; Lee, D.; Rodriguez, J. R.

1978-01-01

An airflow and thrust calibration of an F100 engine, S/N P680059, was conducted to study airframe propulsion system integration losses in turbofan-powered high-performance aircraft. The tests were conducted with and without thrust augmentation for a variety of simulated flight conditions with emphasis on the transonic regime. The resulting corrected airflow data generalized into one curve with corrected fan speed while corrected gross thrust increased as simulated flight conditions increased. Overall agreement between measured data and computed results was 1 percent for corrected airflow and -1 1/2 percent for gross thrust. The results of an uncertainty analysis are presented for both parameters at each simulated flight condition.

Model-Based Control of a Nonlinear Aircraft Engine Simulation using an Optimal Tuner Kalman Filter Approach

NASA Technical Reports Server (NTRS)

Connolly, Joseph W.; Csank, Jeffrey Thomas; Chicatelli, Amy; Kilver, Jacob

2013-01-01

This paper covers the development of a model-based engine control (MBEC) methodology featuring a self tuning on-board model applied to an aircraft turbofan engine simulation. Here, the Commercial Modular Aero-Propulsion System Simulation 40,000 (CMAPSS40k) serves as the MBEC application engine. CMAPSS40k is capable of modeling realistic engine performance, allowing for a verification of the MBEC over a wide range of operating points. The on-board model is a piece-wise linear model derived from CMAPSS40k and updated using an optimal tuner Kalman Filter (OTKF) estimation routine, which enables the on-board model to self-tune to account for engine performance variations. The focus here is on developing a methodology for MBEC with direct control of estimated parameters of interest such as thrust and stall margins. Investigations using the MBEC to provide a stall margin limit for the controller protection logic are presented that could provide benefits over a simple acceleration schedule that is currently used in traditional engine control architectures.

Advanced Aeroservoelastic Testing and Data Analysis (Les Essais Aeroservoelastiques et l’Analyse des Donnees).

DTIC Science & Technology

1995-11-01

network - based AFS concepts. Neural networks can addition of vanes in each engine exhaust for thrust provide...parameter estimation programs 19-11 8.6 Neural Network Based Methods unknown parameters of the postulated state space model Artificial neural network ...Forward Neural Network the network that the applicability of the recurrent neural and ii) Recurrent Neural Network [117-119]. network to

Preliminary supersonic flight test evaluation of performance seeking control

NASA Technical Reports Server (NTRS)

Orme, John S.; Gilyard, Glenn B.

1993-01-01

Digital flight and engine control, powerful onboard computers, and sophisticated controls techniques may improve aircraft performance by maximizing fuel efficiency, maximizing thrust, and extending engine life. An adaptive performance seeking control system for optimizing the quasi-steady state performance of an F-15 aircraft was developed and flight tested. This system has three optimization modes: minimum fuel, maximum thrust, and minimum fan turbine inlet temperature. Tests of the minimum fuel and fan turbine inlet temperature modes were performed at a constant thrust. Supersonic single-engine flight tests of the three modes were conducted using varied after burning power settings. At supersonic conditions, the performance seeking control law optimizes the integrated airframe, inlet, and engine. At subsonic conditions, only the engine is optimized. Supersonic flight tests showed improvements in thrust of 9 percent, increases in fuel savings of 8 percent, and reductions of up to 85 deg R in turbine temperatures for all three modes. The supersonic performance seeking control structure is described and preliminary results of supersonic performance seeking control tests are given. These findings have implications for improving performance of civilian and military aircraft.

Speciation and chemical evolution of nitrogen oxides in aircraft exhaust near airports.

PubMed

Wood, Ezra C; Herndon, Scott C; Timko, Michael T; Yelvington, Paul E; Miake-Lye, Richard C

2008-03-15

Measurements of nitrogen oxides from a variety of commercial aircraft engines as part of the JETS-APEX2 and APEX3 campaigns show that NOx (NOx [triple bond] NO + NO2) is emitted primarily in the form of NO2 at idle thrust and NO at high thrust. A chemical kinetics combustion model reproduces the observed NO2 and NOx trends with engine power and sheds light on the relevant chemical mechanisms. Experimental evidence is presented of rapid conversion of NO to NO2 in the exhaust plume from engines at low thrust. The rapid conversion and the high NO2/NOx emission ratios observed are unrelated to ozone chemistry. NO2 emissions from a CFM56-3B1 engine account for approximately 25% of the NOx emitted below 3000 feet (916 m) and 50% of NOx emitted below 500 feet (153 m) during a standard ICAO (International Civil Aviation Organization) landing-takeoff cycle. Nitrous acid (HONO) accounts for 0.5% to 7% of NOy emissions from aircraft exhaust depending on thrust and engine type. Implications for photochemistry near airports resulting from aircraft emissions are discussed.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Nuclear Thermal Rocket Simulation in NPSS

NASA Technical Reports Server (NTRS)

Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas L.

2013-01-01

Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic- metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

Feasibility Study of a Pressure-fed Engine for a Water Recoverable Space Shuttle Booster

NASA Technical Reports Server (NTRS)

Gerstl, E.

1972-01-01

Detailed mass properties are presented for a gimbaled, fixed thrust, regeneratively cooled engine having a coaxial pintle injector. The baseline design parameters for this engine are tabulated. Mass properties are also summarized for several other engine configurations i.e., a hinge nozzle using a Techroll seal, a gimbaled duct cooled engine and a regeneratively cooled engine using liquid injection thrust vector control (LITVC). Detailed engine analysis and design trade studies leading to the selection of a regeneratively cooled gimbaled engine and pertaining to the selection of the baseline design configuration are also given.

75 FR 8471 - Airworthiness Directives; Airbus Model A340-200 and A340-300 Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-02-25

... condition as: A review of A340 missions has demonstrated that CFM56-5C forward engine mount thrust links... thrust link would reduce the forward engine mounts' structural integrity and could eventually lead to... products. The MCAI states: A review of A340 missions has demonstrated that CFM56-5C forward engine mount...

Stage Separation Failure: Model Based Diagnostics and Prognostics

NASA Technical Reports Server (NTRS)

Luchinsky, Dmitry; Hafiychuk, Vasyl; Kulikov, Igor; Smelyanskiy, Vadim; Patterson-Hine, Ann; Hanson, John; Hill, Ashley

2010-01-01

Safety of the next-generation space flight vehicles requires development of an in-flight Failure Detection and Prognostic (FD&P) system. Development of such system is challenging task that involves analysis of many hard hitting engineering problems across the board. In this paper we report progress in the development of FD&P for the re-contact fault between upper stage nozzle and the inter-stage caused by the first stage and upper stage separation failure. A high-fidelity models and analytical estimations are applied to analyze the following sequence of events: (i) structural dynamics of the nozzle extension during the impact; (ii) structural stability of the deformed nozzle in the presence of the pressure and temperature loads induced by the hot gas flow during engine start up; and (iii) the fault induced thrust changes in the steady burning regime. The diagnostic is based on the measurements of the impact torque. The prognostic is based on the analysis of the correlation between the actuator signal and fault-induced changes in the nozzle structural stability and thrust.

J85 Rejuvenation Through Technology Insertion

DTIC Science & Technology

2000-10-01

and Sabre 75 business addition to military production, the J85 was jets . Number Model Produced Aircraft Type(s) Engine Type Thrust (lbs) J85-GE-4 740...REJUVENATION THROUGH TECHNOLOGY INSERTION T.A. Brisken, P.N. Howell, A.C. Ewing Military Engines Operation GE Aircraft Engines 1 Neumann Way Cincinnati...OH 45215, USA Summary thrust to weight ratio turbojet engines with potential application to early cruise missiles and drones. The history of the

Orbital Transfer Rocket Engine Technology. Advanced Engine Study, Task D.6 Final Report

DTIC Science & Technology

1992-06-01

PROPERTIES _- -,mr m" , MANUAL a PAQ *E,- 7.3.2.1.2. IA .A.2 ,C -- 70-t’ i Rl I _ N -’.±v-j-. .......-441I 0.2% YS Design Allowable • -’Moo 0 2W0" 6W...Storage External Radiation Environment ( Buried Engine) The engine thrust chamber would be cold to the touch even at full thrust operation from the

A new generation of high performance engines for spacecraft propulsion

NASA Technical Reports Server (NTRS)

Rosenberg, Sanders D.; Schoenman, Leonard

1991-01-01

Experimental data validating advanced engine designs at three thrust levels (5, 15, and 100 lbF) is presented. All of the three engine designs considered employ a Moog bipropellant torque motor valve, platelet injector design, and iridium-lined rhenium combustion chamber. Attention is focused on the performance, robustness, duration, and flexibility characteristics of the engines. It is noted that the 5- and 15-lbF thrust engines can deliver a steady state specific impulse in excess of 310 lbF-sec/lbm at an area ratio of 150:1, while the 150-lbF thrust engines deliver a steady state specific impulse of 320 lbF-sec/lbm at an area ratio of 250:1. The hot-fire test results reveal specific impulse improvements of 15 to 25 sec over conventional fuel film cooled columbium chamber designs while operating at maximum chamber temperatures.

The Benefits of Nuclear Thermal Propulsion (NTP) in an Evolvable Mars Campaign

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; Mccurdy, David R.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2propellant which is then exhausted to produce thrust. Conventional chemical engine LH2tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used --NTR is next evolutionary step in high performance liquid rocket engines During the Rover program, a common fuel element tie tube design was developed and used in the design of the 50 klbf Kiwi-B4E (1964), 75 klbf Phoebus-1B (1967), 250 klbf Phoebus-2A (June 1968), then back down to the 25 klbf Pewee engine (Nov-Dec 1968) NASA and DOE are using this same approach: design, build, ground then flight test a small engine using a common fuel element that is scalable to a larger 25 klbf thrust engine needed for human missions

State variable modeling of the integrated engine and aircraft dynamics

NASA Astrophysics Data System (ADS)

Rotaru, Constantin; Sprinţu, Iuliana

2014-12-01

This study explores the dynamic characteristics of the combined aircraft-engine system, based on the general theory of the state variables for linear and nonlinear systems, with details leading first to the separate formulation of the longitudinal and the lateral directional state variable models, followed by the merging of the aircraft and engine models into a single state variable model. The linearized equations were expressed in a matrix form and the engine dynamics was included in terms of variation of thrust following a deflection of the throttle. The linear model of the shaft dynamics for a two-spool jet engine was derived by extending the one-spool model. The results include the discussion of the thrust effect upon the aircraft response when the thrust force associated with the engine has a sizable moment arm with respect to the aircraft center of gravity for creating a compensating moment.

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2012 CFR

2012-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the... this subpart must be run with the reverser installed. In complying with this section, the power control... regimes of control operations are incorporated necessitating scheduling of the power-control lever motion...

14 CFR 33.97 - Thrust reversers.

Code of Federal Regulations, 2010 CFR

2010-01-01

... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.97 Thrust reversers. (a) If the... this subpart must be run with the reverser installed. In complying with this section, the power control... regimes of control operations are incorporated necessitating scheduling of the power-control lever motion...

Engineering research, development and technology report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Langland, R T

1999-02-01

Nineteen ninety-eight has been a transition year for Engineering, as we have moved from our traditional focus on thrust areas to a more focused approach with research centers. These five new centers of excellence collectively comprise Engineering's Science and Technology program. This publication summarizes our formative year under this new structure. Let me start by talking about the differences between a thrust area and a research center. The thrust area is more informal, combining an important technology with programmatic priorities. In contrast, a research center is directly linked to an Engineering core technology. It is the purer model, for itmore » is more enduring yet has the scope to be able to adapt quickly to evolving programmatic priorities. To put it another way, the mission of a thrust area was often to grow the programs in conjunction with a technology, whereas the task of a research center is to vigorously grow our core technologies. By cultivating each core technology, we in turn enable long-term growth of new programs.« less

Energy Efficient Engine program advanced turbofan nacelle definition study

NASA Technical Reports Server (NTRS)

Howe, David C.; Wynosky, T. A.

1985-01-01

Advanced, low drag, nacelle configurations were defined for some of the more promising propulsion systems identified in the earlier Benefit/Cost Study, to assess the benefits associated with these advanced technology nacelles and formulate programs for developing these nacelles and low volume thrust reversers/spoilers to a state of technology readiness in the early 1990's. The study results established the design feasibility of advanced technology, slim line nacelles applicable to advanced technology, high bypass ratio turbofan engines. Design feasibility was also established for two low volume thrust reverse/spoiler concepts that meet or exceed the required effectiveness for these engines. These nacelle and thrust reverse/spoiler designs were shown to be applicable in engines with takeoff thrust sizes ranging from 24,000 to 60,000 pounds. The reduced weight, drag, and cost of the advanced technology nacelle installations relative to current technology nacelles offer a mission fuel burn savings ranging from 3.0 to 4.5 percent and direct operating cost plus interest improvements from 1.6 to 2.2 percent.

Selected Performance Measurements of the F-15 Active Axisymmetric Thrust-vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1998-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Selected Performance Measurements of the F-15 ACTIVE Axisymmetric Thrust-Vectoring Nozzle

NASA Technical Reports Server (NTRS)

Orme, John S.; Sims, Robert L.

1999-01-01

Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.

Orbital transfer rocket engine technology 7.5K-LB thrust rocket engine preliminary design

NASA Technical Reports Server (NTRS)

Harmon, T. J.; Roschak, E.

1993-01-01

A preliminary design of an advanced LOX/LH2 expander cycle rocket engine producing 7,500 lbf thrust for Orbital Transfer vehicle missions was completed. Engine system, component and turbomachinery analysis at both on design and off design conditions were completed. The preliminary design analysis results showed engine requirements and performance goals were met. Computer models are described and model outputs are presented. Engine system assembly layouts, component layouts and valve and control system analysis are presented. Major design technologies were identified and remaining issues and concerns were listed.

Nuclear-Rocket Propulsion

NASA Technical Reports Server (NTRS)

Rom, Frank E.

1968-01-01

The three basic types of nuclear power-plants (solid, liquid, and gas core) are compared on the bases of performance potential and the status of current technology. The solid-core systems are expected to have impulses in the range of 850 seconds, any thrust level (as long as it is greater than 10,000 pounds (44,480 newtons)), and thrust-to-engine-weight ratios of 2 to 20 pounds per pound (19.7 to 197 newtons per kilogram). There is negligible or no fuel loss from the solid-core system. The solid-core system, of course, has had the most work done on it. Large-scale tests have been performed on a breadboard engine that has produced specific impulses greater than 700 seconds at thrust levels of about 50,000 pounds (222,000 newtons). The liquid-core reactor would be interesting in the specific impulse range of 1200 to 1500 seconds. Again, any thrust level can be obtained depending on how big or small the reactor is made. The thrust-to-engine weight ratio for these systems would be in the range of 1 to 10. The discouraging feature of the liquid-core system is the high fuel-loss ratio anticipated. Values of 0.01 to 0.1 pound (0.00454 to 0.0454 kilograms) or uranium loss per pound (0.454 kilograms) of hydrogen are expected, if impulses in the range of 1200 to 1500 seconds are desired. The gas-core reactor shows specific impulses in the range of 1500 to 2500 seconds. The thrust levels should be at least as high as the weight so that the thrust-to-weight ratio does not go below 1. Because the engine weight is not expected to be under 100,000 pounds (444,800 newtons), thrust levels higher than 100,000 pounds (448,000 newtons) are of interest. The thrust-to-engine weights, in that case, would run from 1 to 20 pounds per pound (9.8 to 19.7 kilograms). Gas-core reactors tend to be very large, and can have high thrust-to-weight ratios. As in the case of the liquid-core system, the fuel loss that will be attendant with gas cores as envisioned today will be rather high. The loss rates will be 0.01 to 0.1 pound of uranium (0.00454 to 0.0454 kilograms) for each pound (0.454 kilograms) of hydrogen.

Rapid deceleration mode evaluation

NASA Technical Reports Server (NTRS)

Conners, Timothy R.; Nobbs, Steven G.; Orme, John S.

1995-01-01

Aircraft with flight capability above 1.4 normally have an RPM lockup or similar feature to prevent inlet buzz that would occur at low engine airflows. This RPM lockup has the effect of holding the engine thrust level at the intermediate power (maximum non-afterburning). For aircraft such as military fighters or supersonic transports, the need exists to be able to rapidly slow from supersonic to subsonic speeds. For example, a supersonic transport that experiences a cabin decompression needs to be able to slow/descend rapidly, and this requirement may size the cabin environmental control system. For a fighter, there may be a desire to slow/descend rapidly, and while doing so to minimize fuel usage and engine exhaust temperature. Both of these needs can be aided by achieving the minimum possible overall net propulsive force. As the intermediate power thrust levels of engines increase, it becomes even more difficult to slow rapidly from supersonic speeds. Therefore, a mode of the performance seeking control (PSC) system to minimize overall propulsion system thrust has been developed and tested. The rapid deceleration mode reduces the engine airflow consistent with avoiding inlet buzz. The engine controls are trimmed to minimize the thrust produced by this reduced airflow, and moves the inlet geometry to degrade the inlet performance. As in the case of the other PSC modes, the best overall performance (in this case the least net propulsive force) requires an integrated optimization of inlet, engine, and nozzle variables. This paper presents the predicted and measured results for the supersonic minimum thrust mode, including the overall effects on aircraft deceleration.

Kadenancy effect, acoustical resonance effect valveless pulse jet engine

NASA Astrophysics Data System (ADS)

Ismail, Rafis Suizwan; Jailani, Azrol; Haron, Muhammad Adli

2017-09-01

A pulse jet engine is a tremendously simple device, as far as moving parts are concerned, that is capable of using a range of fuels, an ignition device, and the ambient air to run an open combustion cycle at rates commonly exceeding 100 Hz. The pulse jet engine was first recognized as a worthy device for aeronautics applications with the introduction of the German V-1 Rocket, also known as the "Buzz Bomb." Although pulse jets are somewhat inefficient compared to other jet engines in terms of fuel usage, they have an exceptional thrust to weight ratio if the proper materials are chosen for its construction. For this reason, many hobbyists have adopted pulse jet engines for a propulsive device in RC planes, go-karts, and other recreational applications. The concept behind the design and function of propulsion devices are greatly inspired by the Newton's second and third laws. These laws quantitatively described thrust as a reaction force. Basically, whenever a mass is accelerated or expelled from one direction by a system, such a mass will exert the same force which will be equal in magnitude, however that will be opposite in direction over the same system. Thrust is that force utilized over a facade in a direction normal and perpendicular to the facade which is known as the thrust. This is the simplest explanation of the concept, on which propulsion devices functions. In mechanical engineering, any force that is orthogonal to the main load is generally referred to as thrust [1].

Application of Solar Electric Propulsion to a Comet Surface Sample Return Mission

NASA Technical Reports Server (NTRS)

Cupples, Mike; Coverstone, Victoria; Woo, Byoungsam

2004-01-01

Current NSTAR (planned for the Discovery Mission: Dawn) and NASA's Evolutionary Xenon Thruster based propulsion systems were compared for a comet surface sample return mission to Tempe1 1. Mission and systems analyses were conducted over a range of array power for each propulsion system with an array of 12 kW EOL at 1 AU chosen for a baseline. Engine configurations investigated for NSTAR included 4 operational engines with 1 spare and 5 operational engines with 1 spare. The NEXT configuration investigated included 2 operational engines plus 1 spare, with performance estimated for high thrust and high Isp throttling modes. Figures of merit for this comparison include Solar Electric Propulsion dry mass, average engine throughput, and net non-propulsion payload returned to Earth flyby.

Performance of 4600-pound-thrust centrifugal-flow-type turbojet engine with water-alcohol injection at inlet

NASA Technical Reports Server (NTRS)

Glasser, Philip W

1950-01-01

An experimental investigation of the effects of injecting a water-alcohol mixture of 2:1 at the compressor inlet of a centrifugal-flow type turbojet engine was conducted in an altitude test chamber at static sea-level conditions and at an altitude of 20,000 feet with a flight Mach number of 0.78 with an engine operating at rated speed. The net thrust was augmented by 0.16 for both flight conditions with a ratio of injected liquid to air flow of 0.05. Further increases in the liquid-air ratio did not give comparable increases in thrust.

Fuzzy/Neural Software Estimates Costs of Rocket-Engine Tests

NASA Technical Reports Server (NTRS)

Douglas, Freddie; Bourgeois, Edit Kaminsky

2005-01-01

The Highly Accurate Cost Estimating Model (HACEM) is a software system for estimating the costs of testing rocket engines and components at Stennis Space Center. HACEM is built on a foundation of adaptive-network-based fuzzy inference systems (ANFIS) a hybrid software concept that combines the adaptive capabilities of neural networks with the ease of development and additional benefits of fuzzy-logic-based systems. In ANFIS, fuzzy inference systems are trained by use of neural networks. HACEM includes selectable subsystems that utilize various numbers and types of inputs, various numbers of fuzzy membership functions, and various input-preprocessing techniques. The inputs to HACEM are parameters of specific tests or series of tests. These parameters include test type (component or engine test), number and duration of tests, and thrust level(s) (in the case of engine tests). The ANFIS in HACEM are trained by use of sets of these parameters, along with costs of past tests. Thereafter, the user feeds HACEM a simple input text file that contains the parameters of a planned test or series of tests, the user selects the desired HACEM subsystem, and the subsystem processes the parameters into an estimate of cost(s).

Reverse Core Engine with Thrust Reverser

NASA Technical Reports Server (NTRS)

Chandler, Jesse M. (Inventor); Suciu, Gabriel L. (Inventor)

2017-01-01

An engine system has a gas generator, a bi-fi wall surrounding at least a portion of the gas generator, a casing surrounding a fan, and the casing having first and second thrust reverser doors which in a deployed position abut each other and the bi-fi wall.

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Power or thrust response. 33.73 Section 33.73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or...

Orbital Transfer Vehicle (OTV) engine study. Phase A: Extension

NASA Technical Reports Server (NTRS)

Sobin, A. J.

1980-01-01

The current Phase A-Extension of the OTV engine study program aims to provide additional expander and staged combustion cycle data that will lead to design definition of the OTV engine. The proposed program effort seeks to optimize the expander cycle engine concept (consistent with identified OTV engine requirements), investigate the feasibility of kitting the staged combustion cycle engine to provide extended thrust operation, and conduct in-depth analysis of development risk, crew safety, and reliability for both cycles. Additional tasks address the costing of a 10/K thrust expander cycle engine and support of OTV systems study contractors.

Saturn Apollo Program

NASA Image and Video Library

1962-06-07

This photograph depicts the Rocketdyne static firing of the F-1 engine at the towering 76-meter Test Stand 1-C in Area 1-125 of the Edwards Air Force Base in California. The Saturn V S-IC (first) stage utilized five F-1 engines for its thrust. Each engine provided 1,500,000 pounds, for a combined thrust of 7,500,000 pounds with liquid oxygen and kerosene as its propellants.

U.S. Air Force Bomber Sustainment and Modernization: Background and Issues for Congress

DTIC Science & Technology

2014-06-04

turbofan Thrust: Each engine up to 17,000 pounds Wingspan: 185 feet (56.4 meters) Length: 159 feet, 4 inches (48.5 meters) Height: 40 feet, 8...precision and non-precision weapons. Features The B-1B’s blended wing and body configuration, variable-geometry wings, and turbofan afterburning engines... turbofan engine with afterburner Thrust: 30,000-plus pounds with afterburner, per engine Wingspan: 137 feet (41.8 meters) extended forward, 79 feet

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2012 CFR

2012-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2011 CFR

2011-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

14 CFR 33.201 - Design and test requirements for Early ETOPS eligibility.

Code of Federal Regulations, 2013 CFR

2013-01-01

... maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other...-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if...

Engine-propeller power plant aircraft community noise reduction key methods

NASA Astrophysics Data System (ADS)

Moshkov P., A.; Samokhin V., F.; Yakovlev A., A.

2018-04-01

Basic methods of aircraft-type flying vehicle engine-propeller power plant noise reduction were considered including single different-structure-and-arrangement propellers and piston engines. On the basis of a semiempirical model the expressions for blade diameter and number effect evaluation upon propeller noise tone components under thrust constancy condition were proposed. Acoustic tests performed at Moscow Aviation institute airfield on the whole qualitatively proved the obtained ratios. As an example of noise and detectability reduction provision a design-and-experimental estimation of propeller diameter effect upon unmanned aircraft audibility boundaries was performed. Future investigation ways were stated to solve a low-noise power plant design problem for light aircraft and unmanned aerial vehicles.

Application of Chaboche Model in Rocket Thrust Chamber Analysis

NASA Astrophysics Data System (ADS)

Asraff, Ahmedul Kabir; Suresh Babu, Sheela; Babu, Aneena; Eapen, Reeba

2017-06-01

Liquid Propellant Rocket Engines are commonly used in space technology. Thrust chamber is one of the most important subsystems of a rocket engine. The thrust chamber generates propulsive thrust force for flight of the rocket by ejection of combustion products at supersonic speeds. Often double walled construction is employed for these chambers. The thrust chamber investigated here has its hot inner wall fabricated out of a high thermal conductive material like copper alloy and outer wall made of stainless steel. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Main reasons for the failure of such chambers are fatigue in the plastic range (called as low cycle fatigue since the number of cycles to failure will be low in plastic range), creep and thermal ratcheting. Elasto plastic material models are required to simulate the above effects through a cyclic stress analysis. This paper gives the details of cyclic stress analysis carried out for the thrust chamber using different plasticity model combinations available in ANSYS (Version 15) FE code. The best model among the above is applied in the cyclic stress analysis of two dimensional (plane strain and axisymmetric) and three dimensional finite element models of thrust chamber. Cyclic life of the chamber is calculated from stress-strain graph obtained from above analyses.

Method and system for monitoring and displaying engine performance parameters

NASA Technical Reports Server (NTRS)

Abbott, Terence S. (Inventor); Person, Lee H., Jr. (Inventor)

1988-01-01

The invention is believed a major improvement that will have a broad application in governmental and commercial aviation. It provides a dynamic method and system for monitoring and simultaneously displaying in easily scanned form the available, predicted, and actual thrust of a jet aircraft engine under actual operating conditions. The available and predicted thrusts are based on the performance of a functional model of the aircraft engine under the same operating conditions. Other critical performance parameters of the aircraft engine and functional model are generated and compared, the differences in value being simultaneously displayed in conjunction with the displayed thrust values. Thus, the displayed information permits the pilot to make power adjustments directly while keeping him aware of total performance at a glance of a single display panel.

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Alternate low-thrust capability task report

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1980-01-01

The feasibility and design impact of a requirement for the advanced expander cycle engine to be adaptable to extended low thrust operation of approximately 1K to 2K lb is assessed. It is determined that the orbit transfer vehicle point design engine can be reduced in thrust with minor injector modifications from 15K to 1K without significantly affecting combustion performance efficiency or injector face/chamber wall thermal compatibility. Likewise, high frequency transverse mode combustion instability is not expected to be detrimentally affected. Primarily, the operational limitations consist of feed system chugging instabilities and potential coupling of the injector response with the chamber longitudinal mode resonances under certain operating conditions. The recommended injector modification for low thrust operation is a change in the oxidizer injector element orifice size. Analyses also indicate that chamber coolant flow stability may be a concern below 2K 1bF operation and oxidizer pump stability could be a problem below a 2K thrust level although a recirculation flow could alleviate the problem.

Acoustics and Trust of Separate-Flow Exhaust Nozzles With Mixing Devices for High-Bypass-Ratio Engines

NASA Technical Reports Server (NTRS)

Saiyed, Naseem H.; Mikkelsen, Kevin L.; Bridges, James E.

2000-01-01

The NASA Glenn Research Center recently completed an experimental study to reduce the jet noise from modern turbofan engines. The study concentrated on exhaust nozzle designs for high-bypass-ratio engines. These designs modified the core and fan nozzles individually and simultaneously. Several designs provided an ideal jet noise reduction of over 2.5 EPNdB for the effective perceived noise level (EPNL) metric. Noise data, after correcting for takeoff thrust losses, indicated over a 2.0-EPNdB reduction for nine designs. Individually modifying the fan nozzle did not provide attractive EPNL reductions. Designs in which only the core nozzle was modified provided greater EPNL reductions. Designs in which core and fan nozzles were modified simultaneously provided the greatest EPNL reduction. The best nozzle design had a 2.7-EPNdB reduction (corrected for takeoff thrust loss) with a 0.06-point cruise thrust loss. This design simultaneously employed chevrons on the core and fan nozzles. In comparison with chevrons, tabs appeared to be an inefficient method for reducing jet noise. Data trends indicate that the sum of the thrust losses from individually modifying core and fan nozzles did not generally equal the thrust loss from modifying them simultaneously. Flow blockage from tabs did not scale directly with cruise thrust loss and the interaction between fan flow and the core nozzle seemed to strongly affect noise and cruise performance. Finally, the nozzle configuration candidates for full-scale engine demonstrations are identified.

Effects of Altitude on Turbojet Engine Performance

NASA Technical Reports Server (NTRS)

Fleming, William A

1951-01-01

Component and over-all performance characteristics of several turbojet engines investigated in the altitude facilities of the NACA Lewis Laboratory during the last several years are summarized to indicate the effects of altitude on turbojet engine performance. Data presented show that failure of turbojet engine performance to generalize for all altitudes can be traced to reductions in compressor efficiency, corrected air flow, and combustion efficiency at altitude. In addition, it is shown that although engines of different design may have equal thrusts at sea level, the thrusts at altitude may vary widely because of differences in compressor performance characteristics from one engine to another.

Full Scale Technology Demonstration of a Modern Counterrotating Unducted Fan Engine Concept. Design Report

NASA Technical Reports Server (NTRS)

1987-01-01

The Unducted Fan engine (UDF trademark) concept is based on an ungeared, counterrotating, unducted, ultra-high-bypass turbofan configuration. This engine is being developed to provide a high thrust-to-weight ratio power plant with exceptional fuel efficiency for subsonic aircraft application. This report covers the design methodology and details for the major components of this engine. The design intent of the engine is to efficiently produce 25,000 pounds of static thrust while meeting life and stress requirements. The engine is required to operate at Mach numbers of 0.8 or above.

Flight Measurements of the Effect of a Controllable Thrust Reverser on the Flight Characteristics of a Single-Engine Jet Airplane

NASA Technical Reports Server (NTRS)

Anderson, Seth B.; Cooper, George E.; Faye, Alan E., Jr.

1959-01-01

A flight investigation was undertaken to determine the effect of a fully controllable thrust reverser on the flight characteristics of a single-engine jet airplane. Tests were made using a cylindrical target-type reverser actuated by a hydraulic cylinder through a "beep-type" cockpit control mounted at the base of the throttle. The thrust reverser was evaluated as an in-flight decelerating device, as a flight path control and airspeed control in landing approach, and as a braking device during the ground roll. Full deflection of the reverser for one reverser configuration resulted in a reverse thrust ratio of as much as 85 percent, which at maximum engine power corresponded to a reversed thrust of 5100 pounds. Use of the reverser in landing approach made possible a wide selection of approach angles, a large reduction in approach speed at steep approach angles, improved control of flight path angle, and more accuracy in hitting a given touchdown point. The use of the reverser as a speed brake at lower airspeeds was compromised by a longitudinal trim change. At the lower airspeeds and higher engine powers there was insufficient elevator power to overcome the nose-down trim change at full reverser deflection.

The Design and Testing of a Miniature Turbofan Engine

NASA Technical Reports Server (NTRS)

Cosentino, Gary B.; Murray, James E.

2009-01-01

Off-the-shelf jet propulsion in the 50 - 500 lb thrust class sparse. A true twin-spool turbofan in this range does not exist. Adapting an off-the-shelf turboshaft engine is feasible. However the approx.10 Hp SPT5 can t quite make 50 lbs. of thrust. Packaging and integration is challenging, especially the exhaust. Building on our engine using a 25 Hp turboshaft seems promising if the engine becomes available. Test techniques used, though low cost, adequate for the purpose.

Development of HIDEC adaptive engine control systems

NASA Technical Reports Server (NTRS)

Landy, R. J.; Yonke, W. A.; Stewart, J. F.

1986-01-01

The purpose of NASA's Highly Integrated Digital Electronic Control (HIDEC) flight research program is the development of integrated flight propulsion control modes, and the evaluation of their benefits aboard an F-15 test aircraft. HIDEC program phases are discussed, with attention to the Adaptive Engine Control System (ADECS I); this involves the upgrading of PW1128 engines for operation at higher engine pressure ratios and the production of greater thrust. ADECS II will involve the development of a constant thrust mode which will significantly reduce turbine operating temperatures.

Air Taxi at Your Service

NASA Technical Reports Server (NTRS)

2002-01-01

Under an exclusive agreement with Eclipse Aviation Corporation, Williams International is manufacturing the EJ22 engine, a commercial version of the NASA/Williams FJX-2, for the Eclipse 500 aircraft. The new engine, which weighs approximately 85 pounds and delivers over 770 pounds of thrust, provides a higher thrust-to-weight ratio than any commercial turbofan ever produced. Being the smallest, quietest, and lightest commercial aircraft engine currently available, the EJ22 engine makes a whole new class of twinjet light aircraft feasible.

Multiphysics Analysis of a Solid-Core Nuclear Thermal Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Canabal, Francisco; Cheng, Gary; Chen, Yen-Sen

2006-01-01

The objective of this effort is to develop an efficient and accurate thermo-fluid computational methodology to predict environments for a hypothetical solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics methodology. Formulations for heat transfer in solids and porous media were implemented and anchored. A two-pronged approach was employed in this effort: A detailed thermo-fluid analysis on a multi-channel flow element for mid-section corrosion investigation; and a global modeling of the thrust chamber to understand the effect of hydrogen dissociation and recombination on heat transfer and thrust performance. The formulations and preliminary results on both aspects are presented.

Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

2014-01-01

NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

Camera Layout Design for the Upper Stage Thrust Cone

NASA Technical Reports Server (NTRS)

Wooten, Tevin; Fowler, Bart

2010-01-01

Engineers in the Integrated Design and Analysis Division (EV30) use a variety of different tools to aid in the design and analysis of the Ares I vehicle. One primary tool in use is Pro-Engineer. Pro-Engineer is a computer-aided design (CAD) software that allows designers to create computer generated structural models of vehicle structures. For the Upper State thrust cone, Pro-Engineer was used to assist in the design of a layout for two camera housings. These cameras observe the separation between the first and second stage of the Ares I vehicle. For the Ares I-X, one standard speed camera was used. The Ares I design calls for two separate housings, three cameras, and a lighting system. With previous design concepts and verification strategies in mind, a new layout for the two camera design concept was developed with members of the EV32 team. With the new design, Pro-Engineer was used to draw the layout to observe how the two camera housings fit with the thrust cone assembly. Future analysis of the camera housing design will verify the stability and clearance of the camera with other hardware present on the thrust cone.

Simulator Evaluation of Simplified Propulsion-Only Emergency Flight Control Systems on Transport Aircraft

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Kaneshige, John; Bull, John; Maine, Trindel A.

1999-01-01

With the advent of digital engine control systems, considering the use of engine thrust for emergency flight control has become feasible. Many incidents have occurred in which engine thrust supplemented or replaced normal aircraft flight controls. In most of these cases, a crash has resulted, and more than 1100 lives have been lost. The NASA Dryden Flight Research Center has developed a propulsion-controlled aircraft (PCA) system in which computer-controlled engine thrust provides emergency flight control capability. Using this PCA system, an F-15 and an MD-11 airplane have been landed without using any flight controls. In simulations, C-17, B-757, and B-747 PCA systems have also been evaluated successfully. These tests used full-authority digital electronic control systems on the engines. Developing simpler PCA systems that can operate without full-authority engine control, thus allowing PCA technology to be installed on less capable airplanes or at lower cost, is also a desire. Studies have examined simplified ?PCA Ultralite? concepts in which thrust control is provided using an autothrottle system supplemented by manual differential throttle control. Some of these concepts have worked well. The PCA Ultralite study results are presented for simulation tests of MD-11, B-757, C-17, and B-747 aircraft.

Thrust stand evaluation of engine performance improvement algorithms in an F-15 airplane

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1992-01-01

Results are presented from the evaluation of the performance seeking control (PSC) optimization algorithm developed by Smith et al. (1990) for F-15 aircraft, which optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. Comparisons are presented between the load cell measurements, PSC onboard model thrust calculations, and posttest state variable model computations. Actual performance improvements using the PSC algorithm are presented for its various modes. The results of using PSC algorithm are compared with similar test case results using the HIDEC algorithm.

The Rolls Royce Allison RB580 turbofan - Matching the market requirement for regional transport

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sadler, J.H.R.; Peacock, N.J.; Snyder, L.

1989-01-01

The RB580 high bypass turbofan engine has a thrust growth capability to 10,000 lb and has been optimized for efficient operation in regional markets involving 50-70 seat airliners with higher-than-turboprop cruise speeds. The two-spool engine configuration achieves an overall pressure ratio of 24 and features a single-stage wide-chord fan for high efficiency/low noise operation. The highly modular design of the configuration facilitates maintenance and repair; a dual-redundant full-authority digital electronic control system is incorporated. An SFC reduction of the order of 10 percent at cruise thrust is achieved, relative to current engines of comparable thrust class.

Performance Evaluation of the T6 Ion Engine

NASA Technical Reports Server (NTRS)

Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

2010-01-01

The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

Subsonic Performance of Ejector Systems

NASA Astrophysics Data System (ADS)

Weil, Samuel

Combined cycle engines combining scramjets with turbo jets or rockets can provide efficient hypersonic flight. Ejectors have the potential to increase the thrust and efficiency of combined cycle engines near static conditions. A computer code was developed to support the design of a small-scale, turbine-based combined cycle demonstrator with an ejector, built around a commercially available turbojet engine. This code was used to analyze the performance of an ejector system built around a micro-turbojet. With the use of a simple ejector, net thrust increases as large as 20% over the base engine were predicted. Additionally the specific fuel consumption was lowered by 10%. Increasing the secondary to primary area ratio of the ejector lead to significant improvements in static thrust, specific fuel consumption (SFC), and propulsive efficiency. Further ejector performance improvements can be achieved by using a diffuser. Ejector performance drops off rapidly with increasing Mach number. The ejector has lower thrust and higher SFC than the turbojet core at Mach numbers above 0.2. When the nozzle chokes a significant drop in ejector performance is seen. When a diffuser is used, higher Mach numbers lead to choking in the mixer and a shock in the nozzle causing a significant decrease in ejector performance. Evaluation of different turbo jets shows that ejector performance depends significantly on the properties of the turbojet. Static thrust and SFC improvements can be achieved with increasing ejector area for all engines, but size of increase and change in performance at higher Mach numbers depend heavily on the turbojet. The use of an ejector in a turbine based combined cycle configuration also increases performance at static conditions with a thrust increase of 5% and SFC decrease of 5% for the tested configuration.

Easy method of matching fighter engine to airframe for use in aircraft engine design courses

DOE Office of Scientific and Technical Information (OSTI.GOV)

Mattingly, J.D.

1989-01-01

The proper match of the engine(s) to the airframe affects both aircraft size and life cycle cost. A fast and straightforward method is developed and used for the matching of fighter engine(s) to airframes during conceptual design. A thrust-lapse equation is developed for the dual-spool, mixed-flow, afterburning turbofan type of engine based on the installation losses of 'Aircraft Engine Design' and the performance predictions of the cycle analysis programs ONX and OFFX. Using system performance requirements, the effects of aircraft thrust-to-weight, wing loading, and engine cycle on takeoff weight are analyzed and example design course results presented. 5 refs.

Thrust Augmentation with Mixer/Ejector Systems

NASA Technical Reports Server (NTRS)

Presz, Walter M., Jr.; Reynolds, Gary; Hunter, Craig

2002-01-01

Older commercial aircraft often exceed FAA (Federal Aviation Administration) sideline noise regulations. The major problem is the jet noise associated with the high exhaust velocities of the low bypass ratio engines on such aircraft. Mixer/ejector exhaust systems can provide a simple means of reducing the jet noise on these aircraft by mixing cool ambient air with the high velocity engine gases before they are exhausted to ambient. This paper presents new information on thrust performance predictions, and thrust augmentation capabilities of mixer/ejectors. Results are presented from the recent development program of the patented Alternating Lobe Mixer Ejector Concept (ALMEC) suppressor system for the Gulfstream GII, GIIB and GIII aircraft. Mixer/ejector performance procedures are presented which include classical control volume analyses, compound compressible flow theory, lobed nozzle loss correlations and state of the art computational fluid dynamic predictions. The mixer/ejector thrust predictions are compared to subscale wind tunnel test model data and actual aircraft flight test measurements. The results demonstrate that a properly designed mixer/ejector noise suppressor can increase effective engine bypass ratio and generate large thrust gains at takeoff conditions with little or no thrust loss at cruise conditions. The cruise performance obtained for such noise suppressor systems is shown to be a strong function of installation effects on the aircraft.

Results of solar electric thrust vector control system design, development and tests

NASA Technical Reports Server (NTRS)

Fleischer, G. E.

1973-01-01

Efforts to develop and test a thrust vector control system TVCS for a solar-energy-powered ion engine array are described. The results of solar electric propulsion system technology (SEPST) III real-time tests of present versions of TVCS hardware in combination with computer-simulated attitude dynamics of a solar electric multi-mission spacecraft (SEMMS) Phase A-type spacecraft configuration are summarized. Work on an improved solar electric TVCS, based on the use of a state estimator, is described. SEPST III tests of TVCS hardware have generally proved successful and dynamic response of the system is close to predictions. It appears that, if TVCS electronic hardware can be effectively replaced by control computer software, a significant advantage in control capability and flexibility can be gained in future developmental testing, with practical implications for flight systems as well. Finally, it is concluded from computer simulations that TVCS stabilization using rate estimation promises a substantial performance improvement over the present design.

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 deg, respectively.

Performance improvements of an F-15 airplane with an integrated engine-flight control system

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1988-01-01

An integrated flight and propulsion control system has been developed and flight demonstrated on the NASA Ames-Dryden F-15 research aircraft. The highly integrated digital control (HIDEC) system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power. The amount of EPR uptrim is modulated based on airplane maneuver requirements, flight conditions, and engine information. Engine thrust was increased as much as 10.5 percent at subsonic flight conditions by uptrimming EPR. The additional thrust significantly improved aircraft performance. Rate of climb was increased 14 percent at 40,000 ft and the time to climb from 10,000 to 40,000 ft was reduced 13 percent. A 14 and 24 percent increase in acceleration was obtained at intermediate and maximum power, respectively. The HIDEC logic performed fault free. No engine anomalies were encountered for EPR increases up to 12 percent and for angles of attack and sideslip of 32 and 11 degrees, respectively.

Heat pipe technology for advanced rocket thrust chambers

NASA Technical Reports Server (NTRS)

Rousar, D. C.

1971-01-01

The application of heat pipe technology to the design of rocket engine thrust chambers is discussed. Subjects presented are: (1) evaporator wick development, (2) specific heat pipe designs and test results, (3) injector design, fabrication, and cold flow testing, and (4) preliminary thrust chamber design.

Tubular copper thrust chamber design study

NASA Technical Reports Server (NTRS)

Masters, A. I.; Galler, D. E.

1992-01-01

The use of copper tubular thrust chambers is particularly important in high performance expander cycle space engines. Tubular chambers have more surface area than flat wall chambers, and this extra surface area provides enhanced heat transfer for additional energy to power the cycle. This paper was divided into two sections: (1) a thermal analysis and sensitivity study; and (2) a preliminary design of a selected thrust chamber configuration. The thermal analysis consisted of a statistical optimization to determine the optimum tube geometry, tube booking, thrust chamber geometry, and cooling routing to achieve the maximum upper limit chamber pressure for a 25,000 pound thrust engine. The preliminary design effort produced a layout drawing of a tubular thrust chamber that is three inches shorter than the Advanced Expander Test Bed (AETB) milled channel chamber but is predicted to provide a five percent increase in heat transfer. Testing this chamber in the AETB would confirm the inherent advantages of tubular chamber construction and heat transfer.

Thrust law effects on the long-period modes of aerospace craft

NASA Technical Reports Server (NTRS)

Markopoulos, Nikos; Mease, Kenneth D.; Vinh, Nguyen X.

1989-01-01

An analytical study is presented of the longitudinal long-period dynamics of an aerospace craft in a nearly circular orbit, with a thrust law depending arbitrarily on the speed and altitude. A plane of engine possibilities is first defined, with points corresponding to propulsion systems having prescribed thrust slopes with respect to speed and altitude. Approximate expressions for the characteristic roots and times are obtained by first identifying a small quantity in the coefficients of the characteristic equation, and then expanding in a perturbation series about the origin of the plane of engine possibilities, for which the solution is always known. These expressions agree very well with the exact solutions over a wide range of altitudes and thrust laws. The period of the oscillatory translational mode (phugoid) is found to be independent to first order of the thrust law, generalizing results found by previous investigators for specific thrust laws. The results apply to the speed range from hypersonic to orbital.

Effects of an in-flight thrust reverser on the stability and control characteristics of a single-engine fighter airplane model

NASA Technical Reports Server (NTRS)

Mercer, C. E.; Maiden, D. L.

1972-01-01

The changes in thrust minus drag performance as well as longitudinal and directional stability and control characteristics of a single-engine jet aircraft attributable to an in-flight thrust reverser of the blocker-deflector door type were investigated in a 16-foot transonic wind tunnel. The longitudinal and directional stability data are presented. Test conditions simulated landing approach conditions as well as high speed maneuvering such as may be required for combat or steep descent from high altitude.

LEO-to-GEO low thrust chemical propulsion

NASA Technical Reports Server (NTRS)

Shoji, J. M.

1980-01-01

One approach being considered for transporting large space structures from low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) is the use of low thrust chemical propulsion systems. A variety of chemical rocket engine cycles evaluated for this application for oxygen/hydrogen and oxygen/hydrocarbon propellants (oxygen/methane and oxygen/RF-1) are discussed. These cycles include conventional propellant turbine drives, turboalternator/electric motor pump drive, and fuel cell/electric motor pump drive as well as pressure fed engines. Thrust chamber cooling analysis results are presented for regenerative/radiation and film/radiation cooling.

Parametric analysis of a down-scaled turbo jet engine suitable for drone and UAV propulsion

NASA Astrophysics Data System (ADS)

Wessley, G. Jims John; Chauhan, Swati

2018-04-01

This paper presents a detailed study on the need for downscaling gas turbine engines for UAV and drone propulsion. Also, the procedure for downscaling and the parametric analysis of a downscaled engine using Gas Turbine Simulation Program software GSP 11 is presented. The need for identifying a micro gas turbine engine in the thrust range of 0.13 to 4.45 kN to power UAVs and drones weighing in the range of 4.5 to 25 kg is considered and in order to meet the requirement a parametric analysis on the scaled down Allison J33-A-35 Turbojet engine is performed. It is evident from the analysis that the thrust developed by the scaled engine and the Thrust Specific Fuel Consumption TSFC depends on pressure ratio, mass flow rate of air and Mach number. A scaling factor of 0.195 corresponding to air mass flow rate of 7.69 kg/s produces a thrust in the range of 4.57 to 5.6 kN while operating at a Mach number of 0.3 within the altitude of 5000 to 9000 m. The thermal and overall efficiency of the scaled engine is found to be 67% and 75% respectively for a pressure ratio of 2. The outcomes of this analysis form a strong base for further analysis, design and fabrication of micro gas turbine engines to propel future UAVs and drones.

Aircraft emissions of methane and nitrous oxide during the alternative aviation fuel experiment.

PubMed

Santoni, Gregory W; Lee, Ben H; Wood, Ezra C; Herndon, Scott C; Miake-Lye, Richard C; Wofsy, Steven C; McManus, J Barry; Nelson, David D; Zahniser, Mark S

2011-08-15

Given the predicted growth of aviation and the recent developments of alternative aviation fuels, quantifying methane (CH(4)) and nitrous oxide (N(2)O) emission ratios for various aircraft engines and fuels can help constrain projected impacts of aviation on the Earth's radiative balance. Fuel-based emission indices for CH(4) and N(2)O were quantified from CFM56-2C1 engines aboard the NASA DC-8 aircraft during the first Alternative Aviation Fuel Experiment (AAFEX-I) in 2009. The measurements of JP-8 fuel combustion products indicate that at low thrust engine states (idle and taxi, or 4% and 7% maximum rated thrusts, respectively) the engines emit both CH(4) and N(2)O at a mean ± 1σ rate of 170 ± 160 mg CH(4) (kg Fuel)(-1) and 110 ± 50 mg N(2)O (kg Fuel)(-1), respectively. At higher thrust levels corresponding to greater fuel flow and higher engine temperatures, CH(4) concentrations in engine exhaust were lower than ambient concentrations. Average emission indices for JP-8 fuel combusted at engine thrusts between 30% and 100% of maximum rating were -54 ± 33 mg CH(4) (kg Fuel)(-1) and 32 ± 18 mg N(2)O (kg Fuel)(-1), where the negative sign indicates consumption of atmospheric CH(4) in the engine. Emission factors for the synthetic Fischer-Tropsch fuels were statistically indistinguishable from those for JP-8.

High Thrust-to-Power Annular Engine Technology

NASA Technical Reports Server (NTRS)

Patterson, Michael J.; Thomas, Robert E.; Crofton, Mark W.; Young, Jason A.; Foster, John E.

2015-01-01

Gridded ion engines have the highest efficiency and total impulse of any mature electric propulsion technology, and have been successfully implemented for primary propulsion in both geocentric and heliocentric environments with excellent ground/in-space correlation of performance. However, they have not been optimized to maximize thrust-to-power, an important parameter for Earth orbit transfer applications. This publication discusses technology development work intended to maximize this parameter. These activities include investigating the capabilities of a non-conventional design approach, the annular engine, which has the potential of exceeding the thrust-to-power of other EP technologies. This publication discusses the status of this work, including the fabrication and initial tests of a large-area annular engine. This work is being conducted in collaboration among NASA Glenn Research Center, The Aerospace Corporation, and the University of Michigan.

High Thrust-to-Power Annular Engine Technology

NASA Technical Reports Server (NTRS)

Patterson, Michael; Thomas, Robert; Crofton, Mark; Young, Jason A.; Foster, John E.

2015-01-01

Gridded ion engines have the highest efficiency and total impulse of any mature electric propulsion technology, and have been successfully implemented for primary propulsion in both geocentric and heliocentric environments with excellent ground-in-space correlation of performance. However, they have not been optimized to maximize thrust-to-power, an important parameter for Earth orbit transfer applications. This publication discusses technology development work intended to maximize this parameter. These activities include investigating the capabilities of a non-conventional design approach, the annular engine, which has the potential of exceeding the thrust-to-power of other EP technologies. This publication discusses the status of this work, including the fabrication and initial tests of a large-area annular engine. This work is being conducted in collaboration among NASA Glenn Research Center, The Aerospace Corporation, and the University of Michigan.

In-flight thrust determination on a real-time basis

NASA Technical Reports Server (NTRS)

Ray, R. J.; Carpenter, T.; Sandlin, T.

1984-01-01

A real time computer program was implemented on a F-15 jet fighter to monitor in-flight engine performance of a Digital Electronic Engine Controlled (DEES) F-100 engine. The application of two gas generator methods to calculate in-flight thrust real time is described. A comparison was made between the actual results and those predicted by an engine model simulation. The percent difference between the two methods was compared to the predicted uncertainty based on instrumentation and model uncertainty and agreed closely with the results found during altitude facility testing. Data was obtained from acceleration runs of various altitudes at maximum power settings with and without afterburner. Real time in-flight thrust measurement was a major advancement to flight test productivity and was accomplished with no loss in accuracy over previous post flight methods.

Fuel-Efficient Descent and Landing Guidance Logic for a Safe Lunar Touchdown

NASA Technical Reports Server (NTRS)

Lee, Allan Y.

2011-01-01

The landing of a crewed lunar lander on the surface of the Moon will be the climax of any Moon mission. At touchdown, the landing mechanism must absorb the load imparted on the lander due to the vertical component of the lander's touchdown velocity. Also, a large horizontal velocity must be avoided because it could cause the lander to tip over, risking the life of the crew. To be conservative, the worst-case lander's touchdown velocity is always assumed in designing the landing mechanism, making it very heavy. Fuel-optimal guidance algorithms for soft planetary landing have been studied extensively. In most of these studies, the lander is constrained to touchdown with zero velocity. With bounds imposed on the magnitude of the engine thrust, the optimal control solutions typically have a "bang-bang" thrust profile: the thrust magnitude "bangs" instantaneously between its maximum and minimum magnitudes. But the descent engine might not be able to throttle between its extremes instantaneously. There is also a concern about the acceptability of "bang-bang" control to the crew. In our study, the optimal control of a lander is formulated with a cost function that penalizes both the touchdown velocity and the fuel cost of the descent engine. In this formulation, there is not a requirement to achieve a zero touchdown velocity. Only a touchdown velocity that is consistent with the capability of the landing gear design is required. Also, since the nominal throttle level for the terminal descent sub-phase is well below the peak engine thrust, no bound on the engine thrust is used in our formulated problem. Instead of bangbang type solution, the optimal thrust generated is a continuous function of time. With this formulation, we can easily derive analytical expressions for the optimal thrust vector, touchdown velocity components, and other system variables. These expressions provide insights into the "physics" of the optimal landing and terminal descent maneuver. These insights could help engineers to achieve a better "balance" between the conflicting needs of achieving a safe touchdown velocity, a low-weight landing mechanism, low engine fuel cost, and other design goals. In comparing the computed optimal control results with the preflight landing trajectory design of the Apollo-11 mission, we noted interesting similarities between the two missions.

Loss estimates for a Puente Hills blind-thrust earthquake in Los Angeles, California

USGS Publications Warehouse

Field, E.H.; Seligson, H.A.; Gupta, N.; Gupta, V.; Jordan, T.H.; Campbell, K.W.

2005-01-01

Based on OpenSHA and HAZUS-MH, we present loss estimates for an earthquake rupture on the recently identified Puente Hills blind-thrust fault beneath Los Angeles. Given a range of possible magnitudes and ground motion models, and presuming a full fault rupture, we estimate the total economic loss to be between $82 and $252 billion. This range is not only considerably higher than a previous estimate of $69 billion, but also implies the event would be the costliest disaster in U.S. history. The analysis has also provided the following predictions: 3,000-18,000 fatalities, 142,000-735,000 displaced households, 42,000-211,000 in need of short-term public shelter, and 30,000-99,000 tons of debris generated. Finally, we show that the choice of ground motion model can be more influential than the earthquake magnitude, and that reducing this epistemic uncertainty (e.g., via model improvement and/or rejection) could reduce the uncertainty of the loss estimates by up to a factor of two. We note that a full Puente Hills fault rupture is a rare event (once every ???3,000 years), and that other seismic sources pose significant risk as well. ?? 2005, Earthquake Engineering Research Institute.

Test Results for a Non-toxic, Dual Thrust Reaction Control Engine

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.; Turpin, Alicia A.

2005-01-01

A non-toxic, dual thrust reaction control engine (RCE) was successfully tested over a broad range of operating conditions at the Aerojet Sacramento facility. The RCE utilized LOX/Ethanol propellants; and was tested in steady state and pulsing modes at 25-lbf thrust (vernier) and at 870-lbf thrust (primary). Steady state vernier tests vaned chamber pressure (Pc) from 0.78 to 5.96 psia, and mixture ratio (MR) from 0.73 to 1.82, while primary steady state tests vaned Pc from 103 to 179 psia and MR from 1.33 to 1.76. Pulsing tests explored EPW from 0.080 to 10 seconds and DC from 5 to 50 percent at both thrust levels. Vernier testing accumulated a total of 6,670 seconds of firing time, and 7,215 pulses, and primary testing accumulated a total of 2,060 seconds of firing time and 3,646 pulses.

Investigation of two-dimensional wedge exhaust nozzles for advanced aircraft

NASA Technical Reports Server (NTRS)

Maiden, D. L.; Petit, J. E.

1975-01-01

Two-dimensional wedge nozzle performance characteristics were investigated in a series of wind-tunnel tests. An isolated single-engine/nozzle model was used to study the effects of internal expansion area ratio, aftbody cowl boattail angle, and wedge length. An integrated twin-engine/nozzle model, tested with and without empenage surfaces, included cruise, acceleration, thrust vectoring and thrust reversing nozzle operating modes. Results indicate that the thrust-minus-aftbody drag performance of the twin two-dimensional nozzle integration is significantly higher, for speeds greater than Mach 0.8, than the performance achieved with twin axisymmetric nozzle installations. Significant jet-induced lift was obtained on an aft-mounted lifting surface using a cambered wedge center body to vector thrust. The thrust reversing capabilities of reverser panels installed on the two-dimensional wedge center body were very effective for static or in-flight operation.

Orbit Transfer Rocket Engine Technology - 7.5K-LB Thrust Rocket Engine Preliminary Design

DTIC Science & Technology

1993-10-15

AND SPACE ADMINISTRATION October, 1993 r W NASA-Lewis Research Center Cleveland, Ohio 44135 94-08572 Contract Nc. NAS3-23773 Task B.7 and D.5 4I3’OA4 3 ...APPROACH 1 4.0 SUMMARY OF ACCOMPLISHMENTS 2 5.0 TECHNICAL DISCUSSIONS 3 6.0 PROGRAM WORK PLAN 5 6.1 Engine Analysis 5 6.2 Component Analysis 15 6.2.1...FIGURES Page Figure 1 Advanced Engine Studv Logic Diagram 4 Figure 2 Design Point Engine Pertormance at Full Thrust & MR = 6.0 7 Figure 3 Off-Design

Development of circulation control technology for powered-lift STOL aircraft

NASA Technical Reports Server (NTRS)

Englar, Robert J.

1987-01-01

The flow entraining capabilities of the Circulation Control Wing high lift system were employed to provide an even stronger STOL potential when synergistically combined with upper surface mounted engines. The resulting configurations generate very high supercirculation lift in addition to a vertical component of the pneumatically deflected engine thrust. A series of small scale wind tunnel tests and full scale static thrust deflection tests are discussed which provide a sufficient data base performance. These tests results show thrust deflections of greater than 90 deg produced pneumatically by nonmoving aerodynamic surfaces, and the ability to maintain constant high lift while varying the propulsive force from high thrust recovery required for short takeoff to high drag generation required for short low speed landings.

Closeup view of the aft fuselage of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up view of the aft fuselage of the Orbiter Discovery looking at the thrust structure that supports the Space Shuttle Main Engines (SSMEs). In this view, SSME number two position is on the left and SSME number three position is on the right. The thrust structure transfers the forces produce by the engines into and through the airframe of the orbiter. The thrust structure includes the SSMEs load reaction truss structure, engine interface fittings and the hydraulic-actuator support structure. The propellant feed lines are the plugged and capped orifices within the engine bays. Note that SSME position two is rotated ninety degrees from position three and one. This was needed to enable enough clearance for the engines to fit and gimbal. Note in engine bay three is a clear view of the actuators that control the gambling of that engine. This view was taken from a service platform in the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Development Status of the NASA MC-1 (Fastrac) Engine

NASA Technical Reports Server (NTRS)

Ballard, Richard O.; Olive, Tim; Turner, James E. (Technical Monitor)

2000-01-01

The MC-1 (formerly known as the Fastrac 60K) Engine is being developed for the X-34 technology demonstrator vehicle. It is a pump-fed liquid rocket engine with fixed thrust operating at one rated power level of 60,000 lbf vacuum thrust using a 15:1 area ratio nozzle (slightly higher for the 30:1 flight nozzle). Engine system development testing of the MC-1 has been ongoing since 24 Oct 1998. To date, 48 tests have been conducted on three engines using three separate test stands. This paper will provide some details of the engine, the tests conducted, and the lessons learned to date.

Nonlinear dynamic simulation of single- and multi-spool core engines

NASA Technical Reports Server (NTRS)

Schobeiri, T.; Lippke, C.; Abouelkheir, M.

1993-01-01

In this paper a new computational method for accurate simulation of the nonlinear dynamic behavior of single- and multi-spool core engines, turbofan engines, and power generation gas turbine engines is presented. In order to perform the simulation, a modularly structured computer code has been developed which includes individual mathematical modules representing various engine components. The generic structure of the code enables the dynamic simulation of arbitrary engine configurations ranging from single-spool thrust generation to multi-spool thrust/power generation engines under adverse dynamic operating conditions. For precise simulation of turbine and compressor components, row-by-row calculation procedures were implemented that account for the specific turbine and compressor cascade and blade geometry and characteristics. The dynamic behavior of the subject engine is calculated by solving a number of systems of partial differential equations, which describe the unsteady behavior of the individual components. In order to ensure the capability, accuracy, robustness, and reliability of the code, comprehensive critical performance assessment and validation tests were performed. As representatives, three different transient cases with single- and multi-spool thrust and power generation engines were simulated. The transient cases range from operating with a prescribed fuel schedule, to extreme load changes, to generator and turbine shut down.

REAR DETAIL OF RIGHT ENGINE AND WING. THRUST REVERSER REMAINS ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

REAR DETAIL OF RIGHT ENGINE AND WING. THRUST REVERSER REMAINS OPEN. MECHANICS JONI BAINE (R) AND BILL THEODORE(L) OPEN FLAP CARRIAGE ACCESS WITH AN IMPACT GUN. THEY WILL CHECK TRANSMISSION FLUID AND OIL THE JACK SCREW. AT FAR LEFT UTILITY MECHANICS BEGIN BODY POLISHING. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

Computational analysis of liquid hypergolic propellant rocket engines

NASA Technical Reports Server (NTRS)

Krishnan, A.; Przekwas, A. J.; Gross, K. W.

1992-01-01

The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

NASA Technical Reports Server (NTRS)

Smith, J. A.; Stechman, R. C.

1981-01-01

A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

Advanced Space Transportation Program (ASTP)

NASA Image and Video Library

1997-08-07

This double exposure depicts Marshall Space Flight Center's (MSFC) Test Stand 116 hosting a 60K Bantam Fastrac thrust chamber assembly test. The lower right exposure shows the engine firing in the test stand while the center exposure reveals workers monitoring the test in the interior block house of the test facility. The thrust chamber assembly is only part of the Fastrac engine project to build a low-cost engine for the X-34, an alternate light-weight unmarned launch vehicle. Both the nozzle and the engine for Fastrac are being manufactured at MSFC.

Research Technology

NASA Image and Video Library

1998-09-16

A team of engineers at Marshall Space Flight Center (MSFC) has designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket that produces lower thrust but has better thrust efficiency than the chemical combustion engines. This segmented array of mirrors is the solar concentrator test stand at MSFC for firing the thermal propulsion engines. The 144 mirrors are combined to form an 18-foot diameter array concentrator. The mirror segments are aluminum hexagons that have the reflective surface cut into it by a diamond turning machine, which is developed by MSFC Space Optics Manufacturing Technology Center.

Saturn Apollo Program

NASA Image and Video Library

1968-01-01

A complete F-1 engine assembly is shown in this photograph. Designed and developed by Rocketdye under the direction of the Marshall Space Flight Center, the engine measured 19-feet tall by 12.5 feet at the nozzle exit, and each engine produced a 1,500,000-pound thrust using liquid oxygen and kerosene as the propellant. A cluster of five F-1 engines was mounted on the Saturn V S-IC (first) stage and burned 15 tons of liquid oxygen and kerosene each second to produce 7,500,000 pounds of thrust.

Liquid rocket booster study. Volume 2, book 4, appendices 6-8: Reports of Rocketdyne, Pratt and Whitney, and TRW

NASA Technical Reports Server (NTRS)

1988-01-01

For the pressure fed engines, detailed trade studies were conducted defining engine features such as thrust vector control methods, thrust chamber construction, etc. This was followed by engine design layouts and booster propulsion configuration layouts. For the pump fed engines parametric performance and weight data was generated for both O2/H2 and O2/RP-1 engines. Subsequent studies resulted in the selection of both LOX/RP-1 and O2/H2 propellants for the pump fed engines. More detailed analysis of the selected LOX/RP-1 and O2/H2 engines was conducted during the final phase of the study.

Project-based introduction to aerospace engineering course: A model rocket

NASA Astrophysics Data System (ADS)

Jayaram, Sanjay; Boyer, Lawrence; George, John; Ravindra, K.; Mitchell, Kyle

2010-05-01

In this paper, a model rocket project suitable for sophomore aerospace engineering students is described. This project encompasses elements of drag estimation, thrust determination and analysis using digital data acquisition, statistical analysis of data, computer aided drafting, programming, team work and written communication skills. The student built rockets are launched in the university baseball field with the objective of carrying a specific amount of payload so that the rocket achieves a specific altitude before the parachute is deployed. During the course of the project, the students are introduced to real-world engineering practice through written report submission of their designs. Over the years, the project has proven to enhance the learning objectives, yet cost effective and has provided good outcome measures.

Structures-propulsion interactions and requirements. [large space structures

NASA Technical Reports Server (NTRS)

Coyner, J. V.

1982-01-01

The effects of low-thrust primary propulsion system characteristics on the mass, area, and orbit transfer characteristics of large space systems (LSS) were determined. Three general structural classes of LSS were considered, each with a broad range of diameters and nonstructural surface densities. While transferring the deployed structure from LEO and to GEO, an acceleration range of 0.02 to 0.1 g's was found to maximize deliverable payload based on structural mass impact. After propulsion system parametric analyses considering four propellant combinations produced values for available payload mass, length and volume, a thrust level range which maximizes deliverable LSS diameter was determined corresponding to a structure and propulsion vehicle. The engine start and/or shutdown thrust transients on the last orbit transfer (apogee) burn can impose transient loads which would be greater than the steady-state loads at the burnout acceleration. The effect of the engine thrust transients on the LSS was determined from the dynamic models upon which various engine ramps were imposed.

Nuclear Thermal Rocket (NTR) Propulsion and Power Systems for Outer Planetary Exploration Missions

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Cataldo, R. L.

2001-01-01

The high specific impulse (I (sub sp)) and engine thrust generated using liquid hydrogen (LH2)-cooled Nuclear Thermal Rocket (NTR) propulsion makes them attractive for upper stage applications for difficult robotic science missions to the outer planets. Besides high (I (sub sp)) and thrust, NTR engines can also be designed for "bimodal" operation allowing substantial amounts of electrical power (10's of kWe ) to be generated for onboard spacecraft systems and high data rate communications with Earth during the course of the mission. Two possible options for using the NTR are examined here. A high performance injection stage utilizing a single 15 klbf thrust engine can inject large payloads to the outer planets using a 20 t-class launch vehicle when operated in an "expendable mode". A smaller bimodal NTR stage generating approx. 1 klbf of thrust and 20 to 40 kWe for electric propulsion can deliver approx. 100 kg using lower cost launch vehicles. Additional information is contained in the original extended abstract.

Low Thrust, Deep Throttling, US/CIS Integrated NTRE

NASA Astrophysics Data System (ADS)

Culver, Donald W.; Kolganov, Vyacheslav; Rochow, Richard F.

1994-07-01

In 1993 our international team performed a follow-on ``Nuclear Thermal Rocket Engine (NTRE) Extended Life Feasibility Assessment'' study for the Nuclear Propulsion Office (NPO) at NASAs Lewis Research Center. The main purpose of this study was to complete the 1992 study matrix to assess NTRE designs at thrust levels of 22.5, 11.3, and 6.8 tonnes, using Commonwealth of Independent States (CIS) reactor technology. An additional Aerojet goal was to continue improving the NTRE concept we had generated. Deep throttling, mission performance optimized engine design parametrics, and reliability/cost enhancing engine system simplifications were studied, because they seem to be the last three basic design improvements sorely needed by post-NERVA NTRE. Deep throttling improves engine life by eliminating damaging thermal and mechanical shocks caused by after-cooling with pulsed coolant flow. Alternately, it improves mission performance with steady flow after-cooling by minimizing reactor over-cooling. Deep throttling also provides a practical transition from high pressures and powers of the high thrust power cycle to the low pressures and powers of our electric power generating mode. Two deep throttling designs are discussed; a workable system that was studied and a simplified system that is recommended for future study. Mission-optimized engine thrust/weight (T/W) and Isp predictions are included along with system flow schemes and concept sketches.

Breadboard RL10-2B low-thrust operating mode (second iteration) test report

NASA Technical Reports Server (NTRS)

Kanic, Paul G.; Kaldor, Raymond B.; Watkins, Pia M.

1988-01-01

Cryogenic rocket engines requiring a cooling process to thermally condition the engine to operating temperature can be made more efficient if cooling propellants can be burned. Tank head idle and pumped idle modes can be used to burn propellants employed for cooling, thereby providing useful thrust. Such idle modes required the use of a heat exchanger to vaporize oxygen prior to injection into the combustion chamber. During December 1988, Pratt and Whitney conducted a series of engine hot firing demonstrating the operation of two new, previously untested oxidizer heat exchanger designs. The program was a second iteration of previous low thrust testing conducted in 1984, during which a first-generation heat exchanger design was used. Although operation was demonstrated at tank head idle and pumped idle, the engine experienced instability when propellants could not be supplied to the heat exchanger at design conditions.

Quantifying the air quality-CO2 tradeoff potential for airports

NASA Astrophysics Data System (ADS)

Ashok, Akshay; Dedoussi, Irene C.; Yim, Steve H. L.; Balakrishnan, Hamsa; Barrett, Steven R. H.

2014-12-01

Aircraft movements on the airport surface are responsible for CO2 emissions that contribute to climate change and other emissions that affect air quality and human health. While the potential for optimizing aircraft surface movements to minimize CO2 emissions has been assessed, the implications of CO2 emissions minimization for air quality have not been quantified. In this paper, we identify conditions in which there is a tradeoff between CO2 emissions and population exposure to O3 and secondary PM2.5 - i.e. where decreasing fuel burn (which is directly proportional to CO2 emissions) results in increased exposure. Fuel burn and emissions are estimated as a function of thrust setting for five common gas turbine engines at 34 US airports. Regional air quality impacts, which are dominated by ozone and secondary PM2.5, are computed as a function of airport location and time using the adjoint of the GEOS-Chem chemistry-transport model. Tradeoffs between CO2 emissions and population exposure to PM2.5 and O3 occur between 2-18% and 5-60% of the year, respectively, depending on airport location, engine type, and thrust setting. The total duration of tradeoff conditions is 5-12 times longer at maximum thrust operations (typical for takeoff) relative to 4% thrust operations (typical for taxiing). Per kilogram of additional fuel burn at constant thrust setting during tradeoff conditions, reductions in population exposure to PM2.5 and O3 are 6-13% and 32-1060% of the annual average (positive) population exposure per kilogram fuel burn, where the ranges encompass the medians over the 34 airports. For fuel burn increases due to thrust increases (i.e. for constant operating time), reductions in both PM2.5 and O3 exposure are 1.5-6.4 times larger in magnitude than those due to increasing fuel burn at constant thrust (i.e. increasing operating time). Airports with relatively high population exposure reduction potentials - which occur due to a combination of high duration and magnitude of tradeoff conditions - are identified. Our results are the first to quantify the extent of the tradeoff between CO2 emissions and air quality impacts at airports. This raises the possibility of reducing the air quality impacts of airports beyond minimizing fuel burn and/or optimizing for minimum net environmental impact.

Adaptive Failure Compensation for Aircraft Flight Control Using Engine Differentials: Regulation

NASA Technical Reports Server (NTRS)

Yu, Liu; Xidong, Tang; Gang, Tao; Joshi, Suresh M.

2005-01-01

The problem of using engine thrust differentials to compensate for rudder and aileron failures in aircraft flight control is addressed in this paper in a new framework. A nonlinear aircraft model that incorporates engine di erentials in the dynamic equations is employed and linearized to describe the aircraft s longitudinal and lateral motion. In this model two engine thrusts of an aircraft can be adjusted independently so as to provide the control flexibility for rudder or aileron failure compensation. A direct adaptive compensation scheme for asymptotic regulation is developed to handle uncertain actuator failures in the linearized system. A design condition is specified to characterize the system redundancy needed for failure compensation. The adaptive regulation control scheme is applied to the linearized model of a large transport aircraft in which the longitudinal and lateral motions are coupled as the result of using engine thrust differentials. Simulation results are presented to demonstrate the effectiveness of the adaptive compensation scheme.

Performance improvements of a highly integrated digital electronic control system for an F-15 airplane

NASA Technical Reports Server (NTRS)

Putnam, T. W.; Burcham, F. W., Jr.; Andries, M. G.; Kelly, J. B.

1985-01-01

The NASA highly integrated digital electronic control (HIDEC) program is structured to conduct flight research into the benefits of integrating an aircraft flight control system with the engine control system. A brief description of the HIDEC system installed on an F-15 aircraft is provided. The adaptive engine control system (ADECS) mode is described in detail, together with simulation results and analyses that show the significant excess thrust improvements achievable with the ADECS mode. It was found that this increased thrust capability is accompanied by reduced fan stall margin and can be realized during flight conditions where engine face distortion is low. The results of analyses and simulations also show that engine thrust response is improved and that fuel consumption can be reduced. Although the performance benefits that accrue because of airframe and engine control integration are being demonstrated on an F-15 aircraft, the principles are applicable to advanced aircraft such as the advanced tactical fighter and advanced tactical aircraft.

EC94-42645-9

NASA Image and Video Library

1994-06-27

The modified F-18 High Alpha Research Vehicle (HARV) carries out air flow studies on a flight from the Dryden Flight Research Center, Edwards, California. Using oil, researchers were able to track the air flow across the wing at different speeds and angles of attack. A thrust vectoring system had been installed on the engines' exhaust nozzles for the high angle of attack research program. The thrust vectoring system, linked to the aircraft's flight control system, moves a set of three paddles on each engine to redirect thrust for directional control and increased maneuverability at angles of attack at up to 70 degrees.

Partitioning of convergence in Northwest Sub-Himalaya: estimation of late Quaternary uplift and convergence rates across the Kangra reentrant, North India

NASA Astrophysics Data System (ADS)

Thakur, V. C.; Joshi, M.; Sahoo, D.; Suresh, N.; Jayangondapermal, R.; Singh, A.

2014-06-01

The Kangra reentrant constitutes a ~ 80-km-wide zone of fold-thrust belt made of Cenozoic strata of the foreland basin in NW Sub-Himalaya. Earlier workers estimated the total long-term shortening rate of 14 ± 2 mm/year by balanced cross-section between the Main Boundary Thrust and the Himalayan Frontal Thrust. Geologically estimated rate is nearly consistent with the GPS-derived slip rate of 14 ± 1 mm/year. There are active faults developed within 4-8 km depth of the Sub-Himalayan fold-thrust belt of the reentrant. Dating the strath surfaces of the abandoned fluvial terraces and fans above the thrust faults, the uplift (bedrock incision) rates are computed. The dips of thrust faults are measured in field and from available seismic (depth) profiles. From the acquired data, late Quaternary shortening rates on the Jawalamukhi Thrust (JT), the Soan Thrust (ST) and the Himalayan Frontal Thrust (HFT) are estimated. The shortening rates on the JT are 3.5-4.2 mm/year over a period 32-30 ka. The ST yields a shortening rate of 3.0 mm/year for 29 ka. The corresponding shortening and slip rates estimated on the HFT are 6.0 and 6.9 mm/year during a period 42 ka. On the back thrust of Janauri Anticline, the shortening and slip rates are 2.0 and 2.2 mm/year, respectively, for the same period. The results constrained the shortening to be distributed largely across a 50-km-wide zone between the JT and the HFT. The emergence of surface rupture of a great and mega earthquakes recorded on the reactivated HFT implies ≥100 km width of the rupture. The ruptures of large earthquakes, like the 1905 Kangra and 2005 Kashmir, remained restricted to the hinterland. The present study indicates that the high magnitude earthquakes can occur between the locking line and the active thrusts.

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

NASA Technical Reports Server (NTRS)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hulka, James; Yang, Virog

2009-01-01

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Emergency Multiengine Aircraft System for Lateral Control Using Differential Thrust Control of Wing Engines

NASA Technical Reports Server (NTRS)

Burken, John J. (Inventor); Burcham, Frank W., Jr. (Inventor); Bull, John (Inventor)

2000-01-01

Development of an emergency flight control system is disclosed for lateral control using only differential engine thrust modulation of multiengine aircraft is currently underway. The multiengine has at least two engines laterally displaced to the left and right from the axis of the aircraft. In response to a heading angle command psi(sub c) is to be tracked. By continually sensing the heading angle psi of the aircraft and computing a heading error signal psi(sub e) as a function of the difference between the heading angle command psi(sub c) and the sensed heading angle psi, a track control signal is developed with compensation as a function of sensed bank angle phi. Bank angle rate phi, or roll rate p, yaw rate tau, and true velocity produce an aircraft thrust control signal ATC(sub psi(L,R)). The thrust control signal is differentially applied to the left and right engines, with equal amplitude and opposite sign, such that a negative sign is applied to the control signal on the side of the aircraft. A turn is required to reduce the error signal until the heading feedback reduces the error to zero.

CVD Rhenium Engines for Solar-Thermal Propulsion Systems

NASA Technical Reports Server (NTRS)

Williams, Brian E.; Fortini, Arthur J.; Tuffias, Robert H.; Duffy, Andrew J.; Tucker, Stephen P.

1999-01-01

Solar-thermal upper-stage propulsion systems have the potential to provide specific impulse approaching 900 seconds, with 760 seconds already demonstrated in ground testing. Such performance levels offer a 100% increase in payload capability compared to state-of-the-art chemical upper-stage systems, at lower cost. Although alternatives such as electric propulsion offer even greater performance, the 6- to 18- month orbital transfer time is a far greater deviation from the state of the art than the one to two months required for solar propulsion. Rhenium metal is the only material that is capable of withstanding the predicted thermal, mechanical, and chemical environment of a solar-thermal propulsion device. Chemical vapor deposition (CVD) is the most well-established and cost-effective process for the fabrication of complex rhenium structures. CVD rhenium engines have been successfully constructed for the Air Force ISUS program (bimodal thrust/electricity) and the NASA Shooting Star program (thrust only), as well as under an Air Force SBIR project (thrust only). The bimodal engine represents a more long-term and versatile approach to solar-thermal propulsion, while the thrust-only engines provide a potentially lower weight/lower cost and more near-term replacement for current upper-stage propulsion systems.

Determination of PM mass emissions from an aircraft turbine engine using particle effective density

NASA Astrophysics Data System (ADS)

Durdina, L.; Brem, B. T.; Abegglen, M.; Lobo, P.; Rindlisbacher, T.; Thomson, K. A.; Smallwood, G. J.; Hagen, D. E.; Sierau, B.; Wang, J.

2014-12-01

Inventories of particulate matter (PM) emissions from civil aviation and air quality models need to be validated using up-to-date measurement data corrected for sampling artifacts. We compared the measured black carbon (BC) mass and the total PM mass determined from particle size distributions (PSD) and effective density for a commercial turbofan engine CFM56-7B26/3. The effective density was then used to calculate the PM mass losses in the sampling system. The effective density was determined using a differential mobility analyzer and a centrifugal particle mass analyzer, and increased from engine idle to take-off by up to 60%. The determined mass-mobility exponents ranged from 2.37 to 2.64. The mean effective density determined by weighting the effective density distributions by PM volume was within 10% of the unit density (1000 kg/m3) that is widely assumed in aircraft PM studies. We found ratios close to unity between the PM mass determined by the integrated PSD method and the real-time BC mass measurements. The integrated PSD method achieved higher precision at ultra-low PM concentrations at which current mass instruments reach their detection limit. The line loss model predicted ∼60% PM mass loss at engine idle, decreasing to ∼27% at high thrust. Replacing the effective density distributions with unit density lead to comparable estimates that were within 20% and 5% at engine idle and high thrust, respectively. These results could be used for the development of a robust method for sampling loss correction of the future PM emissions database from commercial aircraft engines.

The 30-centimeter ion thrust subsystem design manual

NASA Technical Reports Server (NTRS)

1979-01-01

The principal characteristics of the 30-centimeter ion propulsion thrust subsystem technology that was developed to satisfy the propulsion needs of future planetary and early orbital missions are described. Functional requirements and descriptions, interface and performance requirements, and physical characteristics of the hardware are described at the thrust subsystem, BIMOD engine system, and component level.

Spectroscopy-based thrust sensor for high-speed gaseous flows

NASA Technical Reports Server (NTRS)

Hanson, Ronald K. (Inventor)

1993-01-01

A system and method for non-intrusively obtaining the thrust value of combustion by-products of a jet engine is disclosed herein. The system includes laser elements for inducing absorption for use in determining the axial velocity and density of the jet flow stream and elements for calculating the thrust value therefrom.

An experimental investigation of thrust vectoring two-dimensional convergent-divergent nozzles installed in a twin-engine fighter model at high angles of attack

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Mason, Mary L.; Leavitt, Laurence D.

1990-01-01

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine thrust vectoring capability of subscale 2-D convergent-divergent exhaust nozzles installed on a twin engine general research fighter model. Pitch thrust vectoring was accomplished by downward rotation of nozzle upper and lower flaps. The effects of nozzle sidewall cutback were studied for both unvectored and pitch vectored nozzles. A single cutback sidewall was employed for yaw thrust vectoring. This investigation was conducted at Mach numbers ranging from 0 to 1.20 and at angles of attack from -2 to 35 deg. High pressure air was used to simulate jet exhaust and provide values of nozzle pressure ratio up to 9.

Over-the-wing model thrust reverser noise tests

NASA Technical Reports Server (NTRS)

Goodykoontz, J.; Gutierrez, O.

1977-01-01

Static acoustic tests were conducted on a 1/12 scale model over-the-wing target type thrust reverser. The model configuration simulates a design that is applicable to the over-the-wing short-haul advanced technology engine. Aerodynamic screening tests of a variety of reverser designs identified configurations that satisfied a reverse thrust requirement of 35 percent of forward thrust at a nozzle pressure ratio of 1.29. The variations in the reverser configuration included, blocker door angle, blocker door lip angle and shape, and side skirt shape. Acoustic data are presented and compared for the various configurations. The model data scaled to a single full size engine show that peak free field perceived noise (PN) levels at a 152.4 meter sideline distance range from 98 to 104 PNdb.

Parametric scramjet analysis

NASA Astrophysics Data System (ADS)

Choi, Jongseong

The performance of a hypersonic flight vehicle will depend on existing materials and fuels; this work presents the performance of the ideal scramjet engine for three different combustion chamber materials and three different candidate fuels. Engine performance is explored by parametric cycle analysis for the ideal scramjet as a function of material maximum service temperature and the lower heating value of jet engine fuels. The thermodynamic analysis is based on the Brayton cycle as similarly employed in describing the performance of the ramjet, turbojet, and fanjet ideal engines. The objective of this work is to explore material operating temperatures and fuel possibilities for the combustion chamber of a scramjet propulsion system to show how they relate to scramjet performance and the seven scramjet engine parameters: specific thrust, fuel-to-air ratio, thrust-specific fuel consumption, thermal efficiency, propulsive efficiency, overall efficiency, and thrust flux. The information presented in this work has not been done by others in the scientific literature. This work yields simple algebraic equations for scramjet performance which are similar to that of the ideal ramjet, ideal turbojet and ideal turbofan engines.

Modified RS2101 rocket engine study program

NASA Technical Reports Server (NTRS)

1971-01-01

The purpose of the program is to perform design studies and analyses to determine the effects of incorporating a 60:1 expansion area ratio nozzle extension, extended firing time, and modified operating conditions and environments on the MM'71 rocket engine assembly. An injector-to-thrust chamber seal study was conducted to define potential solutions for leakage past this joint. The results and recommendations evolving from the engine thermal analyses, the injector-to-thrust chamber seal studies, and the nozzle extension joint stress analyses are presented.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accomodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part I-Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2-Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25.800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3-Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine, two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5,000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC material in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust Engines

NASA Technical Reports Server (NTRS)

Halliwell, Ian

2001-01-01

The exoskeletal engine concept is one in which the shafts and disks are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Omission of the shafts and disks leads to an open channel at the engine centerline. This has immense potential for reduced jet noise and for the accommodation of an alternative form of thruster for use in a combined cycle. The use of ceramic composite materials has the potential for significantly reduced weight as well as higher working temperatures without cooling air. The exoskeletal configuration is also a natural stepping-stone to complete counter-rotating turbomachinery. Ultimately this will lead to reductions in weight, length, parts count and improved efficiency. The feasibility studies are in three parts. Part 1: Systems and Component Requirements addressed the mechanical aspects of components from a functionality perspective. This effort laid the groundwork for preliminary design studies. Although important, it is not felt to be particularly original, and has therefore not been included in the current overview. Part 2: Preliminary Design Studies turned to some of the cycle and performance issues inherent in an exoskeletal configuration and some initial attempts at preliminary design of turbomachinery were described. Twin-spoon and single-spool 25,800-lbf-thrust turbofans were used as reference vehicles in a mid-size commercial subsonic category in addition to a single-spool 5,000-lbf-thrust turbofan that represented a general aviation application. The exoskeletal engine, with its open centerline, has tremendous potential for noise suppression and some preliminary analysis was done which began to quantify the benefits. Part 3: Additional Preliminary Design Studies revisited the design of single-spool 25,800-lbf-thrust turbofan configurations, but in addition to the original FPR = 1.6 and BPR = 5.1 reference engine. two additional configurations used FPR = 2.4 and BPR = 3.0 and FPR = 3.2 and BPR = 2.0 were investigated. The single-spool 5.000-lbf-thrust turbofan was refined and the small engine study was extended to include a 2,000-lbf-thrust turbojet. More attention was paid to optimizing the turbomachinery. Turbine cooling flows were eliminated, in keeping with the use of uncooled CMC materials in exoskeletal engines. The turbine performance parameters moved much closer to the nominal target values, demonstrating the great benefits to the cycle of uncooled turbines.

Low Thrust Orbital Maneuvers Using Ion Propulsion

NASA Astrophysics Data System (ADS)

Ramesh, Eric

2011-10-01

Low-thrust maneuver options, such as electric propulsion, offer specific challenges within mission-level Modeling, Simulation, and Analysis (MS&A) tools. This project seeks to transition techniques for simulating low-thrust maneuvers from detailed engineering level simulations such as AGI's Satellite ToolKit (STK) Astrogator to mission level simulations such as the System Effectiveness Analysis Simulation (SEAS). Our project goals are as follows: A) Assess different low-thrust options to achieve various orbital changes; B) Compare such approaches to more conventional, high-thrust profiles; C) Compare computational cost and accuracy of various approaches to calculate and simulate low-thrust maneuvers; D) Recommend methods for implementing low-thrust maneuvers in high-level mission simulations; E) prototype recommended solutions.

Implementation of the Orbital Maneuvering Systems Engine and Thrust Vector Control for the European Service Module

NASA Technical Reports Server (NTRS)

Millard, Jon

2014-01-01

The European Space Agency (ESA) has entered into a partnership with the National Aeronautics and Space Administration (NASA) to develop and provide the Service Module (SM) for the Orion Multipurpose Crew Vehicle (MPCV) Program. The European Service Module (ESM) will provide main engine thrust by utilizing the Space Shuttle Program Orbital Maneuvering System Engine (OMS-E). Thrust Vector Control (TVC) of the OMS-E will be provided by the Orbital Maneuvering System (OMS) TVC, also used during the Space Shuttle Program. NASA will be providing the OMS-E and OMS TVC to ESA as Government Furnished Equipment (GFE) to integrate into the ESM. This presentation will describe the OMS-E and OMS TVC and discuss the implementation of the hardware for the ESM.

Evaluation of a simplified gross thrust calculation technique using two prototype F100 turbofan engines in an altitude facility

NASA Technical Reports Server (NTRS)

Kurtenbach, F. J.

1979-01-01

The technique which relies on afterburner duct pressure measurements and empirical corrections to an ideal one dimensional flow analysis to determine thrust is presented. A comparison of the calculated and facility measured thrust values is reported. The simplified model with the engine manufacturer's gas generator model are compared. The evaluation was conducted over a range of Mach numbers from 0.80 to 2.00 and at altitudes from 4020 meters to 15,240 meters. The effects of variations in inlet total temperature from standard day conditions were explored. Engine conditions were varied from those normally scheduled for flight. The technique was found to be accurate to a twice standard deviation of 2.89 percent, with accuracy a strong function of afterburner duct pressure difference.

The effects of compressor seventh-stage bleed air extraction on performance of the F100-PW-220 afterburning turbofan engine

NASA Technical Reports Server (NTRS)

Evans, Alison B.

1991-01-01

A study was conducted to determine the effects of seventh-stage compressor bleed on the performance of the F100 afterburning turbofan engine. The effects of bleed on thrust, specific fuel consumption, fan turbine inlet temperature, bleed total pressure, and bleed total temperature were obtained from the engine manufacturer's status deck computer simulation. These effects were determined for power settings of intermediate, partial afterburning, and maximum afterburning for Mach numbers between 0.6 and 2.2 and for altitudes of 30,000, 40,000, and 50,000 ft. It was found that thrust loss and specific fuel consumption increase were approximately linear functions of bleed flow and, based on a percent-thrust change basis, were approximately independent of power setting.

Interrogation of possible imaging conditions for radiation sensitive metal organic frameworks in transmission electron microscopes

NASA Astrophysics Data System (ADS)

Patel, Harinkumar Rajendrabhai

One of the main area of research currently in air-breathing propulsion is increasing the fuel efficiency of engines. Increasing fuel efficiency of an air-breathing engine will be advantageous for civil transport as well as military aircraft. This objective can be achieved in several ways. Present design models are developed based on their uses: commercial transport, high range rescue aircraft, military aircraft. One of the main property of military aircraft is possessing high thrust but increasing fuel efficiency will also be advantageous resulting in more time in combat. Today's engine design operates best at their design point and has reduced thrust and high fuel consumption values in off-design. The adaptive cycle engine concept was introduced to overcome this problem. The adaptive cycle engine is a variable cycle engine concept equipped with an extra bypass (3rd bypass) stream. This engine varies the bypass ratio and the fan pressure ratio, the two main parameters affecting thrust and fuel consumption values of the engine. In cruise, more flow will flow through the third stream resulting in the high bypass engine giving lower fuel consumption. on the other hand, the engine will act as a low bypass engine producing more thrust by allowing more air to flow through core while in combat. The simulation of this engine was carried out using the Numerical Propulsion System Simulation (NPSS) software. The effect of the bypass ratio and the fan pressure ratio along with Mach number were studied. After the parametric variation study, the mixture configuration was also studied. Once the effect of the parameters were understood, the best design operating point configuration was selected and then the engine performance for off-design was calculated. Optimum values of bypass ratio and fan pressure ratio were also obtained for each altitude selected for off-design performance.

A first scramjet study

NASA Technical Reports Server (NTRS)

Emanuel, George

1989-01-01

A variety of related scramjet engine topics are examined. The flow is assumed to be 1-D, the gas is thermally and calorically perfect, and focus is on low hypersonic Mach numbers. The thrust and lift of an exposed half nozzle, which is used on the aerospace plane, is evaluated as well as a fully confined nozzle. A rough estimate of the drag of an aerospace plane is provided. Thermal effects and shock waves are next discussed. A parametric scramjet model is then presented based on the influence coefficient method, which evaluates the dominant scramjet processes. The independent parameters are the ratio of specific heats, a nondimensional heat addition parameter, and four Mach numbers. The total thrust generated by the combustor and nozzle is shown to be independent of the heat release distribution and the combustor exit Mach number, providing thermal choking is avoided. An operating condition for the combustor is found that maximizes the thrust. An alternative condition is explored when this optimum is no longer realistic. This condition provides a favorable pressure gradient and a reasonable area ratio for the combustor. Parametric results based on the model is provided.

Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems

NASA Astrophysics Data System (ADS)

Ryzhkov, V.; Lapshin, E.

2018-01-01

The paper presents an algorithm for designing liquid small thrust rocket engine, the process of which consists of five aggregated stages with feedback. Three stages of the algorithm provide engineering support for design, and two stages - the actual engine design. A distinctive feature of the proposed approach is a deep study of the main technical solutions at the stage of engineering analysis and interaction with the created knowledge (data) base, which accelerates the process and provides enhanced design quality. The using multifunctional graphic package Siemens NX allows to obtain the final product -rocket engine and a set of design documentation in a fairly short time; the engine design does not require a long experimental development.

Design and Demonstration of Emergency Control Modes for Enhanced Engine Performance

NASA Technical Reports Server (NTRS)

Liu, Yuan; Litt, Jonathan S.; Guo, Ten-Huei

2013-01-01

A design concept is presented for developing control modes that enhance aircraft engine performance during emergency flight scenarios. The benefits of increased engine performance to overall vehicle survivability during these situations may outweigh the accompanied elevated risk of engine failure. The objective involves building control logic that can consistently increase engine performance beyond designed maximum levels based on an allowable heightened probability of failure. This concept is applied to two previously developed control modes: an overthrust mode that increases maximum engine thrust output and a faster response mode that improves thrust response to dynamic throttle commands. This paper describes the redesign of these control modes and presents simulation results demonstrating both enhanced engine performance and robust maintenance of the desired elevated risk level.

Test and evaluation of the HIDEC engine uptrim algorithm

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemented into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

Reaction Control Engine for Space Launch Initiative

NASA Technical Reports Server (NTRS)

2002-01-01

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Space shuttle orbit maneuvering engine, reusable thrust chamber program. Task 6: Data dump hot fuel element investigation

NASA Technical Reports Server (NTRS)

Nurick, W. H.

1974-01-01

An evaluation of reusable thrust chambers for the space shuttle orbit maneuvering engine was conducted. Tests were conducted using subscale injector hot-fire procedures for the injector configurations designed for a regenerative cooled engine. The effect of operating conditions and fuel temperature on combustion chamber performance was determined. Specific objectives of the evaluation were to examine the optimum like-doublet element geometry for operation at conditions consistent with a fuel regeneratively cooled engine (hot fuel, 200 to 250 F) and the sensitivity of the triplet injector element to hot fuels.

The pasty propellant rocket engine development

NASA Astrophysics Data System (ADS)

Kukushkin, V. I.; Ivanchenko, A. N.

1993-06-01

The paper describes a newly developed pasty propellant rocket engine (PPRE) and the combustion process and presents results of performance tests. It is shown that, compared with liquid propellant rocket engines, the PPREs can regulate the thrust level within a wider range, are safer ecologically, and have better weight characteristics. Compared with solid propellant rocket engines, the PPREs may be produced with lower costs and more safely, are able to regulate thrust performance within a wider range, and are able to offer a greater scope for the variation of the formulation components and propellant characteristics. Diagrams of the PPRE are included.

Advanced engine study program

NASA Technical Reports Server (NTRS)

Masters, A. I.; Galler, D. E.; Denman, T. F.; Shied, R. A.; Black, J. R.; Fierstein, A. R.; Clark, G. L.; Branstrom, B. R.

1993-01-01

A design and analysis study was conducted to provide advanced engine descriptions and parametric data for space transfer vehicles. The study was based on an advanced oxygen/hydrogen engine in the 7,500 to 50,000 lbf thrust range. Emphasis was placed on defining requirements for high-performance engines capable of achieving reliable and versatile operation in a space environment. Four variations on the expander cycle were compared, and the advantages and disadvantages of each were assessed. Parametric weight, envelope, and performance data were generated over a range of 7,500 to 50,000 lb thrust and a wide range of chamber pressure and nozzle expansion ratio.

Space shuttle orbit maneuvering engine reusable thrust chamber

NASA Technical Reports Server (NTRS)

1972-01-01

A data dump is presented containing space shuttle orbiter maneuvering engine performance, weight, envelope, and interface pressure requirements for candidate propellant combinations (NTO/MMH, NTO50-50, LOX/MMH, LOX/50-50, LOX/N2H4, LOX/C3H8, and LOX/RP-1) and cooling concepts (regenerative and dump/film). These data are presented parametrically for the thrust, chamber pressure, nozzle expansion ratio, and engine mixture ratio ranges of interest. Also included is information describing sensitivity to system changes; reliability, maintainability and safety; development programs and associated critical technology areas; engine cost comparisons during development and operation; and ecological effects.

Research Technology

NASA Image and Video Library

2002-03-13

NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama, has begun a series of engine tests on the Reaction Control Engine developed by TRW Space and Electronics for NASA's Space Launch Initiative (SLI). SLI is a technology development effort aimed at improving the safety, reliability, and cost effectiveness of space travel for reusable launch vehicles. The engine in this photo, the first engine tested at MSFC that includes SLI technology, was tested for two seconds at a chamber pressure of 185 pounds per square inch absolute (psia). Propellants used were liquid oxygen as an oxidizer and liquid hydrogen as fuel. Designed to maneuver vehicles in orbit, the engine is used as an auxiliary propulsion system for docking, reentry, fine-pointing, and orbit transfer while the vehicle is in orbit. The Reaction Control Engine has two unique features. It uses nontoxic chemicals as propellants, which creates a safer environment with less maintenance and quicker turnaround time between missions, and it operates in dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The force of low level thrust allows the vehicle to fine-point maneuver and dock, while the force of the high level thrust is used for reentry, orbital transfer, and course positioning.

Rocketdyne - J-2 Saturn V 2nd and 3rd Stage Engine. Chapter 2, Appendix D

NASA Technical Reports Server (NTRS)

Coffman, Paul

2009-01-01

The J-2 engine was unique in many respects. Technology was not nearly as well-developed in oxygen/hydrogen engines at the start of the J-2 project. As a result, it experienced a number of "teething" problems. It was used in two stages on the Saturn V vehicle in the Apollo Program, as well as on the later Skylab and Apollo/Soyuz programs. In the Apollo Program, it was used on the S-II stage, which was the second stage of the Saturn V vehicle. There were five J-2 engines at the back end of the S-II Stage. In the S-IV-B stage, it was a single engine, but that single engine had to restart. The Apollo mission called for the entire vehicle to reach orbital velocity in low Earth orbit after the first firing of the Saturn-IV-B stage and, subsequently, to fire a second time to go on to the moon. The engine had to be man-rated (worthy of transporting humans). It had to have a high thrust rate and performance associated with oxygen/hydrogen engines, although there were some compromises there. It had to gimbal for thrust vector control. It was an open-cycle gas generator engine delivering up to 230,000 pounds of thrust.

Design and evaluation of thrust vectored nozzles using a multicomponent thrust stand

NASA Technical Reports Server (NTRS)

Carpenter, Thomas W.; Blattner, Ernest W.; Stagner, Robert E.; Contreras, Juanita; Lencioni, Dennis; Mcintosh, Greg

1990-01-01

Future aircraft with the capability of short takeoff and landing, and improved maneuverability especially in the post-stall flight regime will incorporate exhaust nozzles which can be thrust vectored. In order to conduct thrust vector research in the Mechanical Engineering Department at Cal Poly, a program was planned with two objectives; design and construct a multicomponent thrust stand for the specific purpose of measuring nozzle thrust vectors; and to provide quality low moisture air to the thrust stand for cold flow nozzle tests. The design and fabrication of the six-component thrust stand was completed. Detailed evaluation tests of the thrust stand will continue upon the receipt of one signal conditioning option (-702) for the Fluke Data Acquisition System. Preliminary design of thrust nozzles with air supply plenums were completed. The air supply was analyzed with regard to head loss. Initial flow visualization tests were conducted using dual water jets.

Phase 2 program on ground test of refanned JT8D turbofan engines and nacelles for the 727 airplane. Volume 1: Summary

NASA Technical Reports Server (NTRS)

1975-01-01

The propulsion performance, acoustic, structural, and systems changes to a 727-200 airplane retrofitted with a refan modification of the JT8D turbofan engine are evaluated. Model tests, design of certifiable airplane retrofit kit hardware, manufacture of test hardware, ground test of a current production JT8D engine, followed by test of the same engine modified to the refan configuration, detailed analyses of the retrofit impact on airplane airworthiness, performance, and noise, and a preliminary analysis of retrofit costs are included. Results indicate that the refan retrofit of the 727-200 would be certifiable and would result in a 6-to 8 EPNdb reduction in effective perceived noise level (EPNL) at the FAR 36 measuring points and an annoyance-weighted footprint area reduction of 68% to 83%. The installed refan engine is estimated to provide 14% greater takeoff thrust at zero velocity and 10% greater thrust at 100 kn (51.4 m/s). There would be an approximate 0.6% increase in cruise specific fuel consumption (SFC). The refan engine performance in conjunction with the increase in stalled weight results in a range reduction of approximately 15% over the unmodified airplane at the same brake release gross weight (BRGW), with a block fuel increase of 1.5% to 3%. With the particular model 727 that was studied, however, it is possible to operate the airplane (with minor structural modifications) at a higher BRGW and increase the range up to approximately 15% relative to the nonrefanned airplane (with equal or slightly increased noise levels). The JT8D refan engine also improves the limited-field range of the airplane.

Parameterization of a Conventional and Regenerated UHB Turbofan

NASA Astrophysics Data System (ADS)

Oliveira, Fábio; Brójo, Francisco

2015-09-01

The attempt to improve aircraft engines efficiency resulted in the evolution from turbojets to the first generation low bypass ratio turbofans. Today, high bypass ratio turbofans are the most traditional type of engine in commercial aviation. Following many years of technological developments and improvements, this type of engine has proved to be the most reliable facing the commercial aviation requirements. In search of more efficiency, the engine manufacturers tend to increase the bypass ratio leading to ultra-high bypass ratio (UHB) engines. Increased bypass ratio has clear benefits in terms of propulsion system like reducing the specific fuel consumption. This study is aimed at a parametric analysis of a UHB turbofan engine focused on short haul flights. Two cycle configurations (conventional and regenerated) were studied, and estimated values of their specific fuel consumption (TSFC) and specific thrust (Fs) were determined. Results demonstrate that the regenerated cycle may contribute towards a more economic and friendly aero engines in a higher range of bypass ratio.

A novel technique to neutralize the Yawing moment due to asymmetric thrust in a hybrid buoyant aircraft

NASA Astrophysics Data System (ADS)

Haque, Anwar U.; Asrar, Waqar; Omar, Ashraf A.; Sulaeman, Erwin; J. S Ali, Mohamed

2016-03-01

Dorsal fin is used in swimming animals like shark for the generation of thrust as well as to meet the requirement of the lateral stability. In the case of aircraft, rudders are normally used for the said requirement. In the present work, this nature inspired idea is explored for its application to neutralize the unavoidable asymmetric thrust produced by the twin engines of a hybrid buoyant aircraft. First, the estimation of asymmetric thrust is obtained with the help of analytical techniques for maximum thrust condition at 4 degree angle of attack. The moment generated by it is utilized for the sizing of a dorsal fin which looks similar to a tapered wing and is placed aft of the center of gravity. Wind tunnel testing at subsonic speed is carried out to explore the design features of this rotatable dorsal fin. It is found that a small rotation of 5 degree can generate the required moment. However, such rotation requires a complete pneumatic/electro-mechanical system and an alternative of it is to use a cambered airfoil for the dorsal fin installed at fixed location. Such a flow controlling device can also be used as an antenna mast, which is commonly installed out the fuselage of the aircraft for communication purposes. Moreover, by incorporating this technique, a pilot doesn't have to put an extra effort to make the aircraft stable in the presence of side wind.

Performance seeking control program overview

NASA Technical Reports Server (NTRS)

Orme, John S.

1995-01-01

The Performance Seeking Control (PSC) program evolved from a series of integrated propulsion-flight control research programs flown at NASA Dryden Flight Research Center (DFRC) on an F-15. The first of these was the Digital Electronic Engine Control (DEEC) program and provided digital engine controls suitable for integration. The DEEC and digital electronic flight control system of the NASA F-15 were ideally suited for integrated controls research. The Advanced Engine Control System (ADECS) program proved that integrated engine and aircraft control could improve overall system performance. The objective of the PSC program was to advance the technology for a fully integrated propulsion flight control system. Whereas ADECS provided single variable control for an average engine, PSC controlled multiple propulsion system variables while adapting to the measured engine performance. PSC was developed as a model-based, adaptive control algorithm and included four optimization modes: minimum fuel flow at constant thrust, minimum turbine temperature at constant thrust, maximum thrust, and minimum thrust. Subsonic and supersonic flight testing were conducted at NASA Dryden covering the four PSC optimization modes and over the full throttle range. Flight testing of the PSC algorithm, conducted in a series of five flight test phases, has been concluded at NASA Dryden covering all four of the PSC optimization modes. Over a three year period and five flight test phases 72 research flights were conducted. The primary objective of flight testing was to exercise each PSC optimization mode and quantify the resulting performance improvements.

Research Technology

NASA Image and Video Library

2002-03-11

Engineers at the Marshall Space Flight Center (MSFC) have begun a series of engine tests on a new breed of space propulsion: a Reaction Control Engine developed for the Space Launch Initiative (SLI). The engine, developed by TRW Space and Electronics of Redondo Beach, California, is an auxiliary propulsion engine designed to maneuver vehicles in orbit. It is used for docking, reentry, attitude control, and fine-pointing while the vehicle is in orbit. The engine uses nontoxic chemicals as propellants, a feature that creates a safer environment for ground operators, lowers cost, and increases efficiency with less maintenance and quicker turnaround time between missions. Testing includes 30 hot-firings. This photograph shows the first engine test performed at MSFC that includes SLI technology. Another unique feature of the Reaction Control Engine is that it operates at dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The low-level thrust of 25 pounds of force allows the vehicle to fine-point maneuver and dock while the high-level thrust of 1,000 pounds of force is used for reentry, orbit transfer, and coarse positioning. SLI is a NASA-wide research and development program, managed by the MSFC, designed to improve safety, reliability, and cost effectiveness of space travel for second generation reusable launch vehicles.

Advanced Control Considerations for Turbofan Engine Design

NASA Technical Reports Server (NTRS)

Connolly, Joseph W.; Csank, Jeffrey T.; Chicatelli, Amy

2016-01-01

This paper covers the application of a model-based engine control (MBEC) methodology featuring a self tuning on-board model for an aircraft turbofan engine simulation. The nonlinear engine model is capable of modeling realistic engine performance, allowing for a verification of the advanced control methodology over a wide range of operating points and life cycle conditions. The on-board model is a piece-wise linear model derived from the nonlinear engine model and updated using an optimal tuner Kalman Filter estimation routine, which enables the on-board model to self-tune to account for engine performance variations. MBEC is used here to show how advanced control architectures can improve efficiency during the design phase of a turbofan engine by reducing conservative operability margins. The operability margins that can be reduced, such as stall margin, can expand the engine design space and offer potential for efficiency improvements. Application of MBEC architecture to a nonlinear engine simulation is shown to reduce the thrust specific fuel consumption by approximately 1% over the baseline design, while maintaining safe operation of the engine across the flight envelope.

Feasibility of rotating fluidized bed reactor for rocket propulsion

NASA Technical Reports Server (NTRS)

Ludewig, H.; Manning, A. J.; Raseman, C. J.

1974-01-01

The rotating fluidized bed reactor concept is outlined, and its application to rocket propulsion is discussed. Experimental results obtained indicate that minimum fluidization correlations commonly in use for 1-g beds can also be applied to multiple-g beds. It was found that for a low thrust system (20,000 lbf) the fuel particle size and/or particle stress play a limiting role on performance. The superiority of U-233 as a fuel for this type of rocket engine is clearly demonstrated in the analysis. The maximum thrust/weight ratio for a 90,000N thrust engine was found to be approximately 65N/kg.

Low-thrust chemical propulsion system pump technology

NASA Technical Reports Server (NTRS)

Sabiers, R. L.; Siebenhaar, A.

1981-01-01

Candidate pump and driver systems for low thrust cargo orbit transfer vehicle engines which deliver large space structures to geosynchronous equatorial orbit and beyond are evaluated. The pumps operate to 68 atmospheres (1000 psi) discharge pressure and flowrates suited to cryogenic engines using either LOX/methane or LOX/hydrogen propellants in thrust ranges from 445 to 8900 N (100 to 2000 lb F). Analysis of the various pumps and drivers indicate that the low specific speed requirement will make high fluid efficiencies difficult to achieve. As such, multiple stages are required. In addition, all pumps require inducer stages. The most attractive main pumps are the multistage centrifugal pumps.

Electron microscopic study of soot particulate matter emissions from aircraft turbine engines.

PubMed

Liati, Anthi; Brem, Benjamin T; Durdina, Lukas; Vögtli, Melanie; Dasilva, Yadira Arroyo Rojas; Eggenschwiler, Panayotis Dimopoulos; Wang, Jing

2014-09-16

The microscopic characteristics of soot particulate matter (PM) in gas turbine exhaust are critical for an accurate assessment of the potential impacts of the aviation industry on the environment and human health. The morphology and internal structure of soot particles emitted from a CFM 56-7B26/3 turbofan engine were analyzed in an electron microscopic study, down to the nanoscale, for ∼ 100%, ∼ 65%, and ∼ 7% static engine thrust as a proxy for takeoff, cruising, and taxiing, respectively. Sampling was performed directly on transmission electron microscopy (TEM) grids with a state-of-the-art sampling system designed for nonvolatile particulate matter. The electron microscopy results reveal that ∼ 100% thrust produces the highest amount of soot, the highest soot particle volume, and the largest and most crystalline primary soot particles with the lowest oxidative reactivity. The opposite is the case for soot produced during taxiing, where primary soot particles are smallest and most reactive and the soot amount and volume are lowest. The microscopic characteristics of cruising condition soot resemble the ones of the ∼ 100% thrust conditions, but they are more moderate. Real time online measurements of number and mass concentration show also a clear correlation with engine thrust level, comparable with the TEM study. The results of the present work, in particular the small size of primary soot particles present in the exhaust (modes of 24, 20, and 13 nm in diameter for ∼ 100%, ∼ 65% and ∼ 7% engine thrust, respectively) could be a concern for human health and the environment and merit further study. This work further emphasizes the significance of the detailed morphological characteristics of soot for assessing environmental impacts.

Nanonewton thrust measurement of photon pressure propulsion using semiconductor laser

NASA Astrophysics Data System (ADS)

Iwami, K.; Akazawa, Taku; Ohtsuka, Tomohiro; Nishida, Hiroyuki; Umeda, Norihiro

2011-09-01

To evaluate the thrust produced by photon pressure emitted from a 100 W class continuous-wave semiconductor laser, a torsion-balance precise thrust stand is designed and tested. Photon emission propulsion using semiconductor light sources attract interests as a possible candidate for deep-space propellant-less propulsion and attitude control system. However, the thrust produced by photon emission as large as several ten nanonewtons requires precise thrust stand. A resonant method is adopted to enhance the sensitivity of the biflier torsional-spring thrust stand. The torsional spring constant and the resonant of the stand is 1.245 × 10-3 Nm/rad and 0.118 Hz, respectively. The experimental results showed good agreement with the theoretical estimation. The thrust efficiency for photon propulsion was also defined. A maximum thrust of 499 nN was produced by the laser with 208 W input power (75 W of optical output) corresponding to a thrust efficiency of 36.7%. The minimum detectable thrust of the stand was estimated to be 2.62 nN under oscillation at a frequency close to resonance.

Real-time in-flight engine performance and health monitoring techniques for flight research application

NASA Technical Reports Server (NTRS)

Ray, Ronald J.; Hicks, John W.; Wichman, Keith D.

1991-01-01

Procedures for real time evaluation of the inflight health and performance of gas turbine engines and related systems were developed to enhance flight test safety and productivity. These techniques include the monitoring of the engine, the engine control system, thrust vectoring control system health, and the detection of engine stalls. Real time performance techniques were developed for the determination and display of inflight thrust and for aeroperformance drag polars. These new methods were successfully shown on various research aircraft at NASA-Dryden. The capability of NASA's Western Aeronautical Test Range and the advanced data acquisition systems were key factors for implementation and real time display of these methods.

Internal performance predictions for Langley scramjet engine module

NASA Technical Reports Server (NTRS)

Pinckney, S. Z.

1978-01-01

A one dimensional theoretical method for the prediction of the internal performance of a scramjet engine is presented. The effects of changes in vehicle forebody flow parameters and characteristics on predicted thrust for the scramjet engine were evaluated using this method, and results are presented. A theoretical evaluation of the effects of changes in the scramjet engine's internal parameters is also presented. Theoretical internal performance predictions, in terms thrust coefficient and specific impulse, are provided for the scramjet engine for free stream Mach numbers of 5, 6, and 7 free stream dynamic pressure of 23,940 N/sq m forebody surface angles of 4.6 deg to 14.6 deg, and fuel equivalence ratio of 1.0.

Status of the NEXT Ion Engine Wear Test

NASA Technical Reports Server (NTRS)

Soulas, George C.; Domonkos, Matthew T.; Kamhawi, Hani; Patterson, Michael J.; Gardner, Michael M.

2003-01-01

The status of the NEXT 2000 hour wear test is presented. This test is being conducted with a 40 cm engineering model ion engine, designated EM1, at a beam current higher than listed on the NEXT throttle table. Pretest performance assessments demonstrated that EM1 satisfies all thruster performance requirements. As of 7/3/03, the ion engine has accumulated 406 hours of operation at a thruster input power of 6.9 kW. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with no indications of performance degradation. Images of the downstream discharge cathode, neutralizer, and accelerator aperture surfaces have exhibited no significant erosion to date.

Control technology for future aircraft propulsion systems

NASA Technical Reports Server (NTRS)

Zeller, J. R.; Szuch, J. R.; Merrill, W. C.; Lehtinen, B.; Soeder, J. F.

1984-01-01

The need for a more sophisticated engine control system is discussed. The improvements in better thrust-to-weight ratios demand the manipulation of more control inputs. New technological solutions to the engine control problem are practiced. The digital electronic engine control (DEEC) system is a step in the evolution to digital electronic engine control. Technology issues are addressed to ensure a growth in confidence in sophisticated electronic controls for aircraft turbine engines. The need of a control system architecture which permits propulsion controls to be functionally integrated with other aircraft systems is established. Areas of technology studied include: (1) control design methodology; (2) improved modeling and simulation methods; and (3) implementation technologies. Objectives, results and future thrusts are summarized.

Investigation of Water-spray Cooling of Turbine Blades in a Turbojet Engine

NASA Technical Reports Server (NTRS)

Freche, John C; Stelpflug, William J

1953-01-01

An analytical and experimental investigation was made with a J33-A-9 engine to determine the effectiveness of spray cooling as a means of increasing thrust by permitting engine operation at inlet-gas temperatures and speeds above rated. With the assumption of adequate spray cooling at a coolant-to-gas flow ratio of 3 percent, calculations for the sea-level static condition indicated a thrust may be achieved by engine operation at an inlet-gas temperature of 2000 degrees F and an overspeed of 10 percent. Of the water-injection configurations investigated experimentally, those located in the inner ring of the stator diaphragm provided the best cooling at rated engine speed.

Space Launch Initiative (SLI) Engine Test

NASA Technical Reports Server (NTRS)

2002-01-01

NASA's Marshall Space Flight Center (MSFC) in Huntsville, Alabama, has begun a series of engine tests on the Reaction Control Engine developed by TRW Space and Electronics for NASA's Space Launch Initiative (SLI). SLI is a technology development effort aimed at improving the safety, reliability, and cost effectiveness of space travel for reusable launch vehicles. The engine in this photo, the first engine tested at MSFC that includes SLI technology, was tested for two seconds at a chamber pressure of 185 pounds per square inch absolute (psia). Propellants used were liquid oxygen as an oxidizer and liquid hydrogen as fuel. Designed to maneuver vehicles in orbit, the engine is used as an auxiliary propulsion system for docking, reentry, fine-pointing, and orbit transfer while the vehicle is in orbit. The Reaction Control Engine has two unique features. It uses nontoxic chemicals as propellants, which creates a safer environment with less maintenance and quicker turnaround time between missions, and it operates in dual thrust modes, combining two engine functions into one engine. The engine operates at both 25 and 1,000 pounds of force, reducing overall propulsion weight and allowing vehicles to easily maneuver in space. The force of low level thrust allows the vehicle to fine-point maneuver and dock, while the force of the high level thrust is used for reentry, orbital transfer, and course positioning.

Comparison of Mars Aircraft Propulsion Systems

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.

2003-01-01

The propulsion system is a critical aspect of the performance and feasibility of a Mars aircraft. Propulsion system mass and performance greatly influence the aircraft s design and mission capabilities. Various propulsion systems were analyzed to estimate the system mass necessary for producing 35N of thrust within the Mars environment. Three main categories of propulsion systems were considered: electric systems, combustion engine systems and rocket systems. Also, the system masses were compared for mission durations of 1, 2, and 4 h.

Test and evaluation of the HIDEC engine uptrim algorithm. [Highly Integrated Digital Electronic Control for aircraft

NASA Technical Reports Server (NTRS)

Ray, R. J.; Myers, L. P.

1986-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. Performance improvements will result from an adaptive engine stall margin mode, a highly integrated mode that uses the airplane flight conditions and the resulting inlet distortion to continuously compute engine stall margin. When there is excessive stall margin, the engine is uptrimmed for more thrust by increasing engine pressure ratio (EPR). The EPR uptrim logic has been evaluated and implemente into computer simulations. Thrust improvements over 10 percent are predicted for subsonic flight conditions. The EPR uptrim was successfully demonstrated during engine ground tests. Test results verify model predictions at the conditions tested.

Reverse Flow Engine Core Having a Ducted Fan with Integrated Secondary Flow Blades

NASA Technical Reports Server (NTRS)

Kisska, Michael K. (Inventor); Princen, Norman H. (Inventor); Kuehn, Mark S. (Inventor); Cosentino, Gary B. (Inventor)

2014-01-01

Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Technical Excellence: A Requirement for Good Engineering

NASA Technical Reports Server (NTRS)

Gill, Paul S.; Vaughan, William W.

2008-01-01

Technical excellence is a requirement for good engineering. Technical excellence has many different ways of expressing itself within engineering. NASA has initiatives that address the enhancement of the Agency's technical excellence and thrust to maintain the associated high level of performance by the Agency on current programs/projects and as it moves into the Constellation Program and the return to the Moon with plans to visit Mars. This paper addresses some of the key initiatives associated with NASA's technical excellence thrust. Examples are provided to illustrate some results being achieved and plans to enhance these initiatives.

Orbit transfer vehicle engine study, phase A, extension 1: Volume 2: Study results

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1981-01-01

Because of the advantage of the Advanced Expander Cycle Engine brought out in initial studies, further design optimization and comparative analyses were undertaken. The major results and conclusion derived are summarized. The primary areas covered are (1) thrust chamber geometry optimization, (2) expander cycle optimization, (3) alternate low thrust capability, (4) safety and reliability, (5) development risk comparison, and (6) cost comparisons. All of the results obtained were used to baseline the initial design concept for the OTV Advanced Expander Cycle Engine Point Design Study.

Engine sizing and integration requirements for hypersonic airbreathing missile applications

NASA Astrophysics Data System (ADS)

Waltrup, P. J.; Billig, F. S.; Stockbridge, R. D.

1982-03-01

A procedure that provides a rational means for selecting an inlet/combustor configuration for a hypersonic airbreathing missile is presented. The particular problem that is addressed is the design of the sustained engine of a two stage missile that is constrained to be launched from a stowage volume that is either square or circular in cross section. The sustainer engine accelerates from a low altitude separation at Mach 4 and climbs to high altitude for cruise at Mach 8. The results show that a missile with an axisymmetric nose inlet provides a somewhat higher thrust capability and slightly better fuel efficiency than a chin type inlet. Aft entry inlets are shown to have a substantially lower thrust potential and lower engine efficiency. A criterion for determining the maximum contraction ratio of a fixed geometry inlet is established and applied to the exemplary missile designs. Combustor area ratio is examined and found to have a relatively small effect on engine performance for area ratios equal to or larger than that required to obtain maximum thrust at the take-over Mach number.

Oxygen-hydrogen thrusters for Space Station auxiliary propulsion systems

NASA Technical Reports Server (NTRS)

Berkman, D. K.

1984-01-01

The feasibility and technology requirements of a low-thrust, high-performance, long-life, gaseous oxygen (GO2)/gaseous hydrogen (GH2) thruster were examined. Candidate engine concepts for auxiliary propulsion systems for space station applications were identified. The low-thrust engine (5 to 100 lb sub f) requires significant departure from current applications of oxygen/hydrogen propulsion technology. Selection of the thrust chamber material and cooling method needed or long life poses a major challenge. The use of a chamber material requiring a minimum amount of cooling or the incorporation of regenerative cooling were the only choices available with the potential of achieving very high performance. The design selection for the injector/igniter, the design and fabrication of a regeneratively cooled copper chamber, and the design of a high-temperature rhenium chamber were documented and the performance and heat transfer results obtained from the test program conducted at JPL using the above engine components presented. Approximately 115 engine firings were conducted in the JPL vacuum test facility, using 100:1 expansion ratio nozzles. Engine mixture ratio and fuel-film cooling percentages were parametrically investigated for each test configuration.

Thrust control system design of ducted rockets

NASA Astrophysics Data System (ADS)

Chang, Juntao; Li, Bin; Bao, Wen; Niu, Wenyu; Yu, Daren

2011-07-01

The investigation of the thrust control system is aroused by the need for propulsion system of ducted rockets. Firstly the dynamic mathematical models of gas flow regulating system, pneumatic servo system and ducted rocket engine were established and analyzed. Then, to conquer the discussed problems of thrust control, the idea of information fusion was proposed to construct a new feedback variable. With this fused feedback variable, the thrust control system was designed. According to the simulation results, the introduction of the new fused feedback variable is valid in eliminating the contradiction between rapid response and stability for the thrust control system of ducted rockets.

14 CFR Appendix B to Part 121 - Airplane Flight Recorder Specification

Code of Federal Regulations, 2010 CFR

2010-01-01

... Transmitter Keying On-Off (Discrete) ±2° ±2% Thrust/Power on Each Engine Full Range Forward ±2° 1 (per engine) 0.2% 2 Trailing Edge Flap or Cockpit Control Selection Full Range or Each Discrete Position ±3° or... Discrete Position ±3° or as Pilot's Indicator 0.5 0.5% 2 Thrust Reverser Position Stowed, In Transit, and...

Orbital transfer vehicle engine technology: Baffled injector design, fabrication, and verification

NASA Technical Reports Server (NTRS)

Schneider, J. A.

1991-01-01

New technologies for space-based, reusable, throttleable, cryogenic orbit transfer propulsion are being evaluated. Supporting tasks for the design of a dual expander cycle engine thrust chamber design are documented. The purpose of the studies was to research the materials used in the thrust chamber design, the supporting fabrication methods necessary to complete the design, and the modification of the injector element for optimum injector/chamber compatibility.

Thrust Vector Control for Nuclear Thermal Rockets

NASA Technical Reports Server (NTRS)

Ensworth, Clinton B. F.

2013-01-01

Future space missions may use Nuclear Thermal Rocket (NTR) stages for human and cargo missions to Mars and other destinations. The vehicles are likely to require engine thrust vector control (TVC) to maintain desired flight trajectories. This paper explores requirements and concepts for TVC systems for representative NTR missions. Requirements for TVC systems were derived using 6 degree-of-freedom models of NTR vehicles. Various flight scenarios were evaluated to determine vehicle attitude control needs and to determine the applicability of TVC. Outputs from the models yielded key characteristics including engine gimbal angles, gimbal rates and gimbal actuator power. Additional factors such as engine thrust variability and engine thrust alignment errors were examined for impacts to gimbal requirements. Various technologies are surveyed for TVC systems for the NTR applications. A key factor in technology selection is the unique radiation environment present in NTR stages. Other considerations including mission duration and thermal environments influence the selection of optimal TVC technologies. Candidate technologies are compared to see which technologies, or combinations of technologies best fit the requirements for selected NTR missions. Representative TVC systems are proposed and key properties such as mass and power requirements are defined. The outputs from this effort can be used to refine NTR system sizing models, providing higher fidelity definition for TVC systems for future studies.

Twin jet shielding. [for aircraft noise reduction

NASA Technical Reports Server (NTRS)

Parthasarathy, S. P.; Cuffel, R. F.; Massier, P. F.

1979-01-01

For an over-the-wing/under-the-wing engine configuration on an airplane, the noise produced by the upper jet flow is partially reflected by the lower jet. An analysis has been performed which can be used to predict the distribution of perceived noise levels along the ground plane at take-off for an airplane which is designed to take advantage of the over/under shielding concept. Typical contours of PNL, the shielding benefit in the shadow zone, and the EPNL values at 3.5 nautical miles from brake release as well as EPNL values at sideline at 0.35 nautical miles have been calculated. This has been done for a range of flow parameters characteristic of engines producing inverted velocity profile jets suitable for use in a supersonic cruise vehicle. Reductions up to 6.0 EPNdB in community noise levels can be realized when the over engines are operated at higher thrust and the lower engines simultaneously operated with reduced thrust keeping the total thrust constant.

Tests and analysis of a vented D thrust deflecting nozzle on a turbofan engine. [conducted at the outdoor aerodynamic research facility of the Ames Research Center

NASA Technical Reports Server (NTRS)

Roseberg, E. W.

1982-01-01

The objectives were to: obtain nozzle performance characteristics in and out of ground effects; demonstrate the compatibility of the nozzle with a turbofan engine; obtain pressure and temperature distributions on the surface of the D vented nozzle; and establish a correlation of the nozzle performance between small scale and large scale models. The test nozzle was a boilerplate model of the MCAIR D vented nozzle configured for operation with a General Electric YTF-34-F5 turbofan engine. The nozzle was configured to provide: a thrust vectoring range of 0 to 115 deg; a yaw vectoring range of 0 to 10 deg; variable nozzle area control; and variable spacing between the core exit and nozzle entrance station. Compatibility between the YTF-34-T5 turbofan engine and the D vented nozzle was demonstrated. Velocity coefficients of 0.96 and greater were obtained for 90 deg of thrust vectoring. The nozzle walls remained cool during all test conditions.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

Workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, hoist the thrust structure assembly for the Saturn IB S-IB (first) stage. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

An inventory of aeronautical ground research facilities. Volume 2: Air breathing engine test facilities

NASA Technical Reports Server (NTRS)

Pirrello, C. J.; Hardin, R. D.; Heckart, M. V.; Brown, K. R.

1971-01-01

The inventory covers free jet and direct connect altitude cells, sea level static thrust stands, sea level test cells with ram air, and propulsion wind tunnels. Free jet altitude cells and propulsion wind tunnels are used for evaluation of complete inlet-engine-exhaust nozzle propulsion systems under simulated flight conditions. These facilities are similar in principal of operation and differ primarily in test section concept. The propulsion wind tunnel provides a closed test section and restrains the flow around the test specimen while the free jet is allowed to expand freely. A chamber of large diameter about the free jet is provided in which desired operating pressure levels may be maintained. Sea level test cells with ram air provide controlled, conditioned air directly to the engine face for performance evaluation at low altitude flight conditions. Direct connect altitude cells provide a means of performance evaluation at simulated conditions of Mach number and altitude with air supplied to the flight altitude conditions. Sea level static thrust stands simply provide an instrumented engine mounting for measuring thrust at zero airspeed. While all of these facilities are used for integrated engine testing, a few provide engine component test capability.

Design, analysis, and control of a large transport aircraft utilizing selective engine thrust as a backup system for the primary flight control. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Gerren, Donna S.

1995-01-01

A study has been conducted to determine the capability to control a very large transport airplane with engine thrust. This study consisted of the design of an 800-passenger airplane with a range of 5000 nautical miles design and evaluation of a flight control system, and design and piloted simulation evaluation of a thrust-only backup flight control system. Location of the four wing-mounted engines was varied to optimize the propulsive control capability, and the time constant of the engine response was studied. The goal was to provide level 1 flying qualities. The engine location and engine time constant did not have a large effect on the control capability. The airplane design did meet level 1 flying qualities based on frequencies, damping ratios, and time constants in the longitudinal and lateral-directional modes. Project pilots consistently rated the flying qualities as either level 1 or level 2 based on Cooper-Harper ratings. However, because of the limited control forces and moments, the airplane design fell short of meeting the time required to achieve a 30 deg bank and the time required to respond a control input.

Investigation of the effect of the ejector on the performance of the pulse detonation engine nozzle extension

NASA Astrophysics Data System (ADS)

Korobov, A. E.; Golovastov, S. V.

2015-11-01

Influence of an ejector nozzle extension on gas flow at a pulse detonation engine was investigated numerically and experimentally. Detonation formation was organized in stoichiometric hydrogen-oxygen mixture in cylindrical detonation tube. Cylindrical ejector was constructed and mounted at the open end of the tube. Thrust, air consumption and parameters of the detonation were measured in single and multiple regimes of operation. Axisymmetric model was used in numerical investigation. Equations of Navies-Stokes were solved using a finite-difference scheme Roe of second order of accuracy. Initial conditions were estimated on a base of experimental data. Numerical results were validated with experiments data.

Assessment of community noise for a medium-range airplane with open-rotor engines

NASA Astrophysics Data System (ADS)

Kopiev, V. F.; Shur, M. L.; Travin, A. K.; Belyaev, I. V.; Zamtfort, B. S.; Medvedev, Yu. V.

2017-11-01

Community noise of a hypothetical medium-range airplane equipped with open-rotor engines is assessed by numerical modeling of the aeroacoustic characteristics of an isolated open rotor with the simplest blade geometry. Various open-rotor configurations are considered at constant thrust, and the lowest-noise configuration is selected. A two-engine medium-range airplane at known thrust of bypass turbofan engines at different segments of the takeoff-landing trajectory is considered, after the replacement of those engines by the open-rotor engines. It is established that a medium-range airplane with two open-rotor engines meets the requirements of Chapter 4 of the ICAO standard with a significant margin. It is shown that airframe noise makes a significant contribution to the total noise of an airplane with open-rotor engines at landing.

UAV Mission Optimization through Hybrid-Electric Propulsion

NASA Astrophysics Data System (ADS)

Blackwelder, Philip Scott

Hybrid-electric powertrain leverages the superior range of petrol based systems with the quiet and emission free benefits of electric propulsion. The major caveat to hybrid-electric powertrain in an airplane is that it is inherently heavier than conventional petroleum powertrain due mostly to the low energy density of battery technology. The first goal of this research is to develop mission planning code to match powertrain components for a small-scale unmanned aerial vehicle (UAV) to complete a standard surveillance mission within a set of user input parameters. The second goal is to promote low acoustic profile loitering through mid-flight engine starting. The two means by which midmission engine starting will be addressed is through reverse thrust from the propeller and a servo actuated gear to couple and decouple the engine and motor. The mission planning code calculates the power required to complete a mission and assists the user in sourcing powertrain components including the propeller, motor, battery, motor controller, engine and fuel. Reverse thrust engine starting involves characterizing an off the shelf variable pitch propeller and using its torque coefficient to calculate the advance ratio required to provide sufficient torque and speed to start an engine. Geared engine starting works like the starter in a conventional automobile. A servo actuated gear will couple the motor to the engine to start it and decouple once the engine has started. Reverse thrust engine starting was unsuccessful due to limitations of available off the shelf variable pitch propellers. However, reverse thrust engine starting could be realized through a custom larger diameter propeller. Geared engine starting was a success, though the system was unable to run fully as intended. Due to counter-clockwise crank rotation of the engine and the right-hand threads on the crankshaft, cranking the engine resulted in the nut securing the engine starter gear to back off as the engine cranked. A second nut was added to secure the starter gear but at the expense of removing the engine drive pulley. Removing the engine pulley meant that the starter gear must remain engaged to transmit torque to the propeller shaft as opposed to the engine pulley. This issue can be resolved using different hardware, however changing the mounting hardware would require additional modifications to the associated component which time would not permit. Though battery technology still proves to be the main constraint of electrified powertrain, careful design and mission planning can help minimize the weight penalties incurred. The mission planning code complements previous research by comparing the weight penalties of a blended climb versus an engine only climb and selecting the lightest option. Though reverse thrust engine starting proved unsuccessful, the success of geared engine starting now allows the engine to be shut off during loiter reducing both acoustic profile and fuel consumption during loiter.

Performance deterioration based on simulated aerodynamic loads test, JT9D jet engine diagnostics program

NASA Technical Reports Server (NTRS)

Stromberg, W. J.

1981-01-01

An engine was specially prepared with extensive instrumentation to monitor performance, case temperatures, and clearance changes. A special loading device was used to apply known loads on the engine by the use of cables placed around the flight inlet. These loads simulated the estimated aerodynamic pressure distributions that occur on the inlet in various segments of a typical airplane flight. Test results indicate that the engine lost 1.3 percent in take-off thrust specific fuel consumption (TSFC) during the course of the test effort. Permanent clearance changes due to the loads accounted for 1.1 percent; increase in low pressure compressor airfoil roughness and thermal distortion in the high pressure turbine accounted for 0.2 percent. Pretest predicted performance loss due to clearance changes was 0.9 percent in TSFC. Therefore, the agreement between measurement and prediction is considered to be excellent.

The convertible engine: A dual-mode propulsion system

NASA Technical Reports Server (NTRS)

Mcardle, Jack G.

1988-01-01

A variable inlet guide vane (VIGV) convertible engine that could be used to power future high-speed rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open, fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed, fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip air flow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque. Previous mission analyses of a conceptual X-wing rotorcraft capable of 400-knot cruise speed were revised to account for more fan-tip churning power loss that was originally estimated. The calculations confirm that using convertible engines rather than separate life and cruise engines would result in a smaller, lighter craft with lower fuel use and direct operating cost.

Outdoor test stand performance of a convertible engine with variable inlet guide vanes for advanced rotorcraft propulsion

NASA Technical Reports Server (NTRS)

Mcardle, Jack G.

1986-01-01

A variable inlet guide van (VIGV) type convertible engine that could be used to power future high-speed rotorcraft was tested on an outdoor stand. The engine ran stably and smoothly in the turbofan, turboshaft, and dual (combined fan and shaft) power modes. In the turbofan mode with the VIGV open fuel consumption was comparable to that of a conventional turbofan engine. In the turboshaft mode with the VIGV closed fuel consumption was higher than that of present turboshaft engines because power was wasted in churning fan-tip airflow. In dynamic performance tests with a specially built digital engine control and using a waterbrake dynamometer for shaft load, the engine responded effectively to large steps in thrust command and shaft torque. Previous mission analyses of a conceptual X-wing rotorcraft capable of 400-knot cruise speed were revised to account for more fan-tip churning power loss than was originally estimated. The new calculations confirm that using convertible engines rather than separate lift and cruise engines would result in a smaller, lighter craft with lower fuel use and direct operating cost.

Pulsed Ejector Thrust Amplification Tested and Modeled

NASA Technical Reports Server (NTRS)

Wilson, Jack

2004-01-01

There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter.

Strength and stiffness of reinforced rectangular columns under biaxially eccentric thrust.

DOT National Transportation Integrated Search

1976-01-01

Compression tests on nine reinforced concrete rectangular columns subjected to : constant thrust and biaxially eccentric moments were conducted at the off-campus : research facility of The University of Texas, The Civil Engineering Structures : Labor...

The Negative Thrust and Torque of Several Full-scale Propellers and Their Application to Various Flight Problems

NASA Technical Reports Server (NTRS)

Hartman, Edwin P; Biermann, David

1938-01-01

Negative thrust and torque data for 2, 3, and 4-blade metal propellers having Clark y and R.A.F. 6 airfoil sections were obtained from tests in the NACA 20-foot tunnel. The propellers were mounted in front of a radial engine nacelle and the blade-angle settings covered in the tests ranged from l5 degrees to 90 degrees. One propeller was also tested at blade-angle settings of 0 degree, 5 degrees, and 10 degrees. A considerable portion of the report deals with the various applications of the negative thrust and torque to flight problems. A controllable propeller is shown to have a number of interesting, and perhaps valuable, uses within the negative thrust and torque range of operation. A small amount of engine-friction data is included to facilitate the application of the propeller data.

Optimal thrust level for orbit insertion

NASA Astrophysics Data System (ADS)

Cerf, Max

2017-07-01

The minimum-fuel orbital transfer is analyzed in the case of a launcher upper stage using a constantly thrusting engine. The thrust level is assumed to be constant and its value is optimized together with the thrust direction. A closed-loop solution for the thrust direction is derived from the extremal analysis for a planar orbital transfer. The optimal control problem reduces to two unknowns, namely the thrust level and the final time. Guessing and propagating the costates is no longer necessary and the optimal trajectory is easily found from a rough initialization. On the other hand the initial costates are assessed analytically from the initial conditions and they can be used as initial guess for transfers at different thrust levels. The method is exemplified on a launcher upper stage targeting a geostationary transfer orbit.

11. Engine room, horizontal corliss (engine #3). view in well ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

11. Engine room, horizontal corliss (engine #3). view in well showing close-up detail of thrust bearing, taken from southwest - East Boston Pumping Station, Chelsea Street at Chelsea Creek, Boston, Suffolk County, MA

Elimination of Intermediate-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Nesman, Tomas E.; Turner, Jim E. (Technical Monitor)

2001-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. The thrust chamber exhibited benign, yet marginally unstable combustion. The marginally unstable combustion was characterized by chamber pressure oscillations with large amplitudes and a frequency that was too low to be identified as acoustic or high-frequency combustion instability and too high to be identified as chug or low-frequency combustion instability. The source of the buzz or intermediate-frequency combustion instability was traced to the fuel venturi whose violently noisy cavitation caused resonance in the feedline downstream. Combustion was stabilized by increasing the throat diameter of the fuel venturi such that the cavitation would occur more quietly.

Low-thrust Isp sensitivity study

NASA Technical Reports Server (NTRS)

Schoenman, L.

1982-01-01

A comparison of the cooling requirements and attainable specific impulse performance of engines in the 445 to 4448N thrust class utilizing LOX/RP-1, LOX/Hydrogen and LOX/Methane propellants is presented. The unique design requirements for the regenerative cooling of low-thrust engines operating at high pressures (up to 6894 kPa) were explored analytically by comparing single cooling with the fuel and the oxidizer, and dual cooling with both the fuel and the oxidizer. The effects of coolant channel geometry, chamber length, and contraction ratio on the ability to provide proper cooling were evaluated, as was the resulting specific impulse. The results show that larger contraction ratios and smaller channels are highly desirable for certain propellant combinations.

Experimental Evaluation of Journal Bearing Stability and New Gas Bearing Material

NASA Technical Reports Server (NTRS)

Keith, Theo G., Jr.; Dimofte, Florin

2001-01-01

It has been estimated that the noise levels in aircraft engine transmissions can be reduced by as much as 10 dB through the use of journal bearings. The potential benefits of lower noise levels include reduced wear, longer gear life and enhanced comfort for passengers and crew. Based on this concept the journal-thrust wave bearing was analyzed and its performance was evaluated. Numerical codes, developed over the past 30 years by Dr. Dimofte, were used to predict the performance of the bearing. The wave bearing is a fluid film bearing and therefore was analyzed using the Reynolds pressure equation. The formulation includes turbulent flow concepts and possesses a viscosity-temperature correction. The centrifugal growth of the bearing diameter and the deformation of the bearing under gear loads were also incorporated into the code. An experimental rig was developed to test the journal-thrust wave bearing.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

Workmen secure a J-2 engine onto the S-IVB (second) stage thrust structure. As part of Marshall Space Center's "building block" approach to the Saturn development, the S-IVB was utilized in the Saturn IBC launch vehicle as a second stage and the Saturn V launch vehicle as a third stage. The booster, built for NASA by McDornell Douglas Corporation, was powered by a single J-2 engine, initially capable of 200,000 pounds of thrust.

Initiation of Long-Wave Instability of Vortex Pairs at Cruise Altitudes

NASA Technical Reports Server (NTRS)

Rossow, Vernon J.

2011-01-01

Previous studies have usually attributed the initiation of the long-wave instability of a vortex pair to turbulence in the atmosphere or in the wake of the aircraft. The purpose here is to show by use of observations and photographs of condensation trails shed by aircraft at cruise altitudes that another initiating mechanism is not only possible but is usually the mechanism that initiates the long-wave instability at cruise altitudes. The alternate initiating mechanism comes about when engine thrust is robust enough to form an array of circumferential vortices around each jet-engine-exhaust stream. In those cases, initiation begins when the vortex sheet shed by the wing has rolled up into a vortex pair and descended to the vicinity of the inside bottom of the combined shear-layer vortex arrays. It is the in-and-out (up and down) velocity field between sequential circumferential vortices near the bottom of the array that then impresses disturbance waves on the lift-generated vortex pair that initiate the long-wave instability. A time adjustment to the Crow and Bate estimate for vortex linking is then derived for cases when thrust-based linking occurs.

Optimal design and installation of ultra high bypass ratio turbofan nacelle

NASA Astrophysics Data System (ADS)

Savelyev, Andrey; Zlenko, Nikolay; Matyash, Evgeniy; Mikhaylov, Sergey; Shenkin, Andrey

2016-10-01

The paper is devoted to the problem of designing and optimizing the nacelle of turbojet bypass engine with high bypass ratio and high thrust. An optimization algorithm EGO based on development of surrogate models and the method for maximizing the probability of improving the objective function has been used. The designing methodology has been based on the numerical solution of the Reynolds equations system. Spalart-Allmaras turbulence model has been chosen for RANS closure. The effective thrust losses has been uses as an objective function in optimizing the engine nacelle. As a result of optimization, effective thrust has been increased by 1.5 %. The Blended wing body aircraft configuration has been studied as a possible application. Two variants of the engine layout arrangement have been considered. It has been shown that the power plant changes the pressure distribution on the aircraft surface. It results in essential diminishing the configuration lift-drag ratio.

NEXT Ion Engine 2000 Hour Wear Test Results

NASA Technical Reports Server (NTRS)

Soulas, George C.; Kamhawi, Hani; Patterson, Michael J.; Britton, Melissa A.; Frandina, Michael M.

2004-01-01

The results of the NEXT 2000 h wear test are presented. This test was conducted with a 40 cm engineering model ion engine, designated EM1, at a 3.52 A beam current and 1800 V beam power supply voltage. Performance tests, which were conducted over a throttling range of 1.1 to 6.9 kW throughout the wear test, demonstrated that EM1 satisfied all thruster performance requirements. The ion engine accumulated 2038 h of operation at a thruster input power of 6.9 kW, processing 43 kg of xenon. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The ion engine was also inspected following the test. This paper presents these findings.

Liquid Rocket Booster (LRB) for the Space Transportation System (STS) systems study. Appendix F: Performance and trajectory for ALS/LRB launch vehicles

NASA Technical Reports Server (NTRS)

1989-01-01

By simply combining two baseline pump-fed LOX/RP-1 Liquid Rocket Boosters (LRBs) with the Denver core, a launch vehicle (Option 1 Advanced Launch System (ALS)) is obtained that can perform both the 28.5 deg (ALS) mission and the polar orbit ALS mission. The Option 2 LRB was obtained by finding the optimum LOX/LH2 engine for the STS/LRB reference mission (70.5 K lb payload). Then this engine and booster were used to estimate ALS payload for the 28.5 deg inclination ALS mission. Previous studies indicated that the optimum number of STS/LRB engines is four. When the engine/booster sizing was performed, each engine had 478 K lb sea level thrust and the booster carried 625,000 lb of useable propellant. Two of these LRBs combined with the Denver core provided a launch vehicle that meets the payload requirements for both the ALS and STS reference missions. The Option 3 LRB uses common engines for the cores and boosters. The booster engines do not have the nozzle extension. These engines were sized as common ALS engines. An ALS launch vehicle that has six core engines and five engines per booster provides 109,100 lb payload for the 28.5 deg mission. Each of these LOX/LH2 LRBs carries 714,100 lb of useable propellant. It is estimated that the STS/LRB reference mission payload would be 75,900 lb.

Methods for determining the internal thrust of scramjet engine modules from experimental data

NASA Technical Reports Server (NTRS)

Voland, Randall T.

1990-01-01

Methods for calculating zero-fuel internal drag of scramjet engine modules from experimental measurements are presented. These methods include two control-volume approaches, and a pressure and skin-friction integration. The three calculation techniques are applied to experimental data taken during tests of a version of the NASA parametric scramjet. The methods agree to within seven percent of the mean value of zero-fuel internal drag even though several simplifying assumptions are made in the analysis. The mean zero-fuel internal drag coefficient for this particular engine is calculated to be 0.150. The zero-fuel internal drag coefficient when combined with the change in engine axial force with and without fuel defines the internal thrust of an engine.

Oil-Free Turbomachinery Research Enhanced by Thrust Bearing Test Capability

NASA Technical Reports Server (NTRS)

Bauman, Steven W.

2003-01-01

NASA Glenn Research Center s Oil-Free Turbomachinery research team is developing aircraft turbine engines that will not require an oil lubrication system. Oil systems are required today to lubricate rolling-element bearings used by the turbine and fan shafts. For the Oil-Free Turbomachinery concept, researchers combined the most advanced foil (air) bearings from industry with NASA-developed high-temperature solid lubricant technology. In 1999, the world s first Oil-Free turbocharger was demonstrated using these technologies. Now we are working with industry to demonstrate Oil-Free turbomachinery technology in a small business jet engine, the EJ-22 produced by Williams International and developed during Glenn s recently concluded General Aviation Propulsion (GAP) program. Eliminating the oil system in this engine will make it simpler, lighter (approximately 15 percent), more reliable, and less costly to purchase and maintain. Propulsion gas turbines will place high demands on foil air bearings, especially the thrust bearings. Up until now, the Oil-Free Turbomachinery research team only had the capability to test radial, journal bearings. This research has resulted in major improvements in the bearings performance, but journal bearings are cylindrical, and can only support radial shaft loads. To counteract axial thrust loads, thrust foil bearings, which are disk shaped, are required. Since relatively little research has been conducted on thrust foil air bearings, their performance lags behind that of journal bearings.

Preliminary flight results of an adaptive engine control system of an F-15 airplane

NASA Technical Reports Server (NTRS)

Myers, Lawrence P.; Walsh, Kevin R.

1987-01-01

Results of the flight demonstration of the adaptive engine control system (ADECS), an integrated flight and propulsion control system, are reported. The ADECS system provides additional engine thrust by increasing engine pressure ratio (EPR) at intermediate and afterburning power, with the amount of EPR uptrim modulated in accordance with the maneuver requirements, flight conditions, and engine information. As a result of EPR uptrimming, engine thrust has increased by as much as 10.5 percent, rate of climb has increased by 10 percent, and the time to climb from 10,000 to 40,000 ft has been reduced by 12.5 percent. Increases in acceleration of 9.3 and 13 percent have been obtained at intermediate and maximum power, respectively. No engine anomalies have been detected for EPR increases up to 12 percent.

Developments in REDES: The rocket engine design expert system

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) is being developed at the NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP, a nozzle design program named RAO, a regenerative cooling channel performance evaluation code named RTE, and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Developments in REDES: The Rocket Engine Design Expert System

NASA Technical Reports Server (NTRS)

Davidian, Kenneth O.

1990-01-01

The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

Perceived Noise Analysis for Offset Jets Applied to Commercial Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Henderson, Brenda S.; Berton, Jeffrey J.; Seidel, Jonathan A.

2016-01-01

A systems analysis was performed with experimental jet noise data, engine/aircraft performance codes and aircraft noise prediction codes to assess takeoff noise levels and mission range for conceptual supersonic commercial aircraft. A parametric study was done to identify viable engine cycles that meet NASAs N+2 goals for noise and performance. Model scale data from offset jets was used as input to the aircraft noise prediction code to determine the expected sound levels for the lateral certification point where jet noise dominates over all other noise sources. The noise predictions were used to determine the optimal orientation of the offset nozzles to minimize the noise at the lateral microphone location. An alternative takeoff procedure called programmed lapse rate was evaluated for noise reduction benefits. Results show there are two types of engines that provide acceptable range performance; one is a standard mixed-flow turbofan with a single-stage fan, and the other is a three-stream variable-cycle engine with a multi-stage fan. The engine with a single-stage fan has a lower specific thrust and is 8 to 10 EPNdB quieter for takeoff. Offset nozzles reduce the noise directed toward the thicker side of the outer flow stream, but have less benefit as the core nozzle pressure ratio is reduced and the bypass-to-core area ratio increases. At the systems level for a three-engine N+2 aircraft with full throttle takeoff, there is a 1.4 EPNdB margin to Chapter 3 noise regulations predicted for the lateral certification point (assuming jet noise dominates). With a 10 reduction in thrust just after takeoff rotation, the margin increases to 5.5 EPNdB. Margins to Chapter 4 and Chapter 14 levels will depend on the cumulative split between the three certification points, but it appears that low specific thrust engines with a 10 reduction in thrust (programmed lapse rate) can come close to meeting Chapter 14 noise levels. Further noise reduction is possible with additional reduction in takeoff thrust using programmed lapse rate, but studies are needed to investigate the practical limits for safety and takeoff regulations.

NASA Engineer Examines the Design of a Regeneratively-Cooled Rocket Engine

NASA Image and Video Library

1958-12-21

An engineer at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a drawing showing the assembly and details of a 20,000-pound thrust regeneratively cooled rocket engine. The engine was being designed for testing in Lewis’ new Rocket Engine Test Facility, which began operating in the fall of 1957. The facility was the largest high-energy test facility in the country that was capable of handling liquid hydrogen and other liquid chemical fuels. The facility’s use of subscale engines up to 20,000 pounds of thrust permitted a cost-effective method of testing engines under various conditions. The Rocket Engine Test Facility was critical to the development of the technology that led to the use of hydrogen as a rocket fuel and the development of lightweight, regeneratively-cooled, hydrogen-fueled rocket engines. Regeneratively-cooled engines use the cryogenic liquid hydrogen as both the propellant and the coolant to prevent the engine from burning up. The fuel was fed through rows of narrow tubes that surrounded the combustion chamber and nozzle before being ignited inside the combustion chamber. The tubes are visible in the liner sitting on the desk. At the time, Pratt and Whitney was designing a 20,000-pound thrust liquid-hydrogen rocket engine, the RL-10. Two RL-10s would be used to power the Centaur second-stage rocket in the 1960s. The successful development of the Centaur rocket and the upper stages of the Saturn V were largely credited to the work carried out Lewis.

10. Engine room, horizontal corliss (engine #3), view in well ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

10. Engine room, horizontal corliss (engine #3), view in well showing detail of thrust bearing and vertical shaft, taken from northwest - East Boston Pumping Station, Chelsea Street at Chelsea Creek, Boston, Suffolk County, MA

14 CFR 33.63 - Vibration.

Code of Federal Regulations, 2011 CFR

2011-01-01

... and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.63 Vibration. Each engine... range of rotational speeds and power/thrust, without inducing excessive stress in any engine part...

14 CFR 33.63 - Vibration.

Code of Federal Regulations, 2010 CFR

2010-01-01

... and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.63 Vibration. Each engine... range of rotational speeds and power/thrust, without inducing excessive stress in any engine part...

Adaptation Method for Overall and Local Performances of Gas Turbine Engine Model

NASA Astrophysics Data System (ADS)

Kim, Sangjo; Kim, Kuisoon; Son, Changmin

2018-04-01

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

Simulating Operation of a Large Turbofan Engine

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Frederick, Dean K.; DeCastro, Jonathan

2008-01-01

The Commercial Modular Aero- Propulsion System Simulation (C-MAPSS) is a computer program for simulating transient operation of a commercial turbofan engine that can generate as much as 90,000 lb (.0.4 MN) of thrust. It includes a power-management system that enables simulation of open- or closed-loop engine operation over a wide range of thrust levels throughout the full range of flight conditions. C-MAPSS provides the user with a set of tools for performing open- and closed-loop transient simulations and comparison of linear and non-linear models throughout its operating envelope, in an easy-to-use graphical environment.

A Visual Photographic Study of Cylinder Lubrication

NASA Technical Reports Server (NTRS)

Shaw, Milton C; Nussdorfer, Theodore

1946-01-01

A V-type engine provided with a glass cylinder was used to study visually the lubrication characteristics of an aircraft-type piston. Photographs and data were obtained with the engine motored at engine speeds up to 1000 r.p.m. and constant cylinder-head pressures of 0 and 50 pounds per square inch. A study was made of the orientation of the piston under various operating conditions, which indicated that the piston was inclined with the crown nearest the major-thrust cylinder face throughout the greater part of the cycle. The piston moved laterally in the cylinder under the influence of piston side thrust.

Turbofan forced mixer lobe flow modeling. Part 3: Application to augment engines

NASA Technical Reports Server (NTRS)

Barber, T.; Moore, G. C.; Blatt, J. R.

1988-01-01

Military engines frequently need large quantities of thrust for short periods of time. The addition of an augmentor can provide such thrust increases but with a penalty of increased duct length and engine weight. The addition of a forced mixer to the augmentor improves performance and reduces the penalty, as well as providing a method for siting the required flame holders. In this report two augmentor concepts are investigated: a swirl-mixer augmentor and a mixer-flameholder augmentor. Several designs for each concept are included and an experimental assessment of one of the swirl-mixer augmentors is presented.

Emergency Flight Control Using Computer-Controlled Thrust

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Fullerton, C. Gordon; Stewart, James F.; Gilyard, Glenn B.; Conley, Joseph A.

1995-01-01

Propulsion Controlled Aircraft (PCA) systems are digital electronic control systems undergoing development to provide limited maneuvering ability through variations of individual engine thrusts in multiple-engine airplanes. Provide landing capability when control surfaces inoperable. Incorporated on existing and future airplanes that include digital engine controls, digital flight controls, and digital data buses, adding no weight for additional hardware to airplane. Possible to handle total failure of hydraulic system, depending on how surfaces respond to loss of hydraulic pressure, and broken control cables or linkages. Future airplanes incorporate data from Global Positioning System for guidance to any suitable emergency runway in world.

Multiphysics Thrust Chamber Modeling for Nuclear Thermal Propulsion

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Cheng, Gary; Chen, Yen-Sen

2006-01-01

The objective of this effort is to develop an efficient and accurate thermo-fluid computational methodology to predict environments for a solid-core, nuclear thermal engine thrust chamber. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation. A two-pronged approach is employed in this effort: A detailed thermo-fluid analysis on a multi-channel flow element for mid-section corrosion investigation; and a global modeling of the thrust chamber to understand the effect of heat transfer on thrust performance. Preliminary results on both aspects are presented.

Fluid design studies of integrated modular engine system

NASA Technical Reports Server (NTRS)

Frankenfield, Bruce; Carek, Jerry

1993-01-01

A study was performed to develop a fluid system design and show the feasibility of constructing an integrated modular engine (IME) configuration, using an expander cycle engine. The primary design goal of the IME configuration was to improve the propulsion system reliability. The IME fluid system was designed as a single fault tolerant system, while minimizing the required fluid components. This study addresses the design of the high pressure manifolds, turbopumps and thrust chambers for the IME configuration. A physical layout drawing was made, which located each of the fluid system components, manifolds and thrust chambers. Finally, a comparison was made between the fluid system designs of an IME system and a non-network (clustered) engine system.

Boeing's CST-100 Launch Abort Engine Test

NASA Image and Video Library

2016-10-10

Boeing and Aerojet Rocketdyne have begun a series of developmental hot-fire tests with two launch abort engines similar to the ones that will be part of Boeing’s CST-100 Starliner service module, in the Mojave Desert in California. The engines, designed to maximize thrust build-up, while minimizing overshoot during start up, will be fired between half a second and 3 seconds each during the test campaign. If the Starliner’s four launch abort engines were used during an abort scenario, they would fire between 3 and 5.5. seconds, with enough thrust to get the spacecraft and its crew away from the rocket, before splashing down in the ocean under parachutes.

TF34 convertible engine control system design

NASA Technical Reports Server (NTRS)

Gilmore, D. R., Jr.

1984-01-01

The characteristics of the TF34 convertible engine, capable of producing shaft power, thrust, or a combination of both, is investigated with respect to the control system design, development, bench testing, and the anticipated transient response during engine testing at NASA. The modifications to the prototype standard TF34-GE-400 turbofan, made primarily in the fan section, consist of the variable inlet guide vanes and variable exit guide vanes. The control system was designed using classical frequency domain techniques and was based on the anticipated convertible/VTOL airframe requirements. The engine has been run in the fan mode and in the shaft mode, exhibiting a response of 0.14 second to a 5-percent thrust change.

Extended Operation of Turbojet Engine with Pentaborane

NASA Technical Reports Server (NTRS)

Useller, James W; Jones, William L

1957-01-01

A full-scale turbojet engine was operated with pentaborane fuel continuously for 22 minutes at conditions simulating flight at a Mach number of 0.8 at an altitude of 50,000 feet. This period of operation is approximately three times longer than previously reported operation times. Although the specific fuel consumption was reduced from 1.3 with JP-4 fuel to 0.98 with pentaborane, a 13.2-percent reduction in net thrust was also encountered. A portion of this thrust loss is potentially recoverable with proper design of the engine components. The boron oxide deposition and erosion processes within the engine approached an equilibrium condition after approximately 22 minutes of operation with pentaborane.

Sunmaster: An SEP cargo vehicle for Mars missions

NASA Technical Reports Server (NTRS)

Chiles, Aleasa; Fraser, Jennifer; Halsey, Andy; Honeycutt, David; Madden, Michael; Mcgough, Brian; Paulsen, David; Spear, Becky; Tarkenton, Lynne; Westley, Kevin

1991-01-01

Options are examined for an unmanned solar powered electric propulsion cargo vehicle for Mars missions. The 6 prime areas of study include: trajectory, propulsion system, power system, supporting structure, control system, and launch consideration. Optimization of the low thrust trajectory resulted in a total round trip mission time just under 4 years. The argon propelled electrostatic ion thruster system consists of seventeen 5 N engines and uses a specific impulse of 10,300 secs. At Earth, the system uses 13 engines to produce 60 N of thrust; at Mars, five engines are used, producing 25 N thrust. The thrust of the craft is varied between 60 N at Earth and 24 N at Mars due to reduced solar power available. Solar power is collected by a Fresnel lens concentrator system using a multistacked cell. This system provides 3.5 MW to the propulsion system after losses. Control and positioning to the craft are provided by a system of three double gimballed control moment gyros. Four shuttle 'C' launches will be used to transport the unassembled vehicle in modular units to low Earth orbit where it will be assembled using the Mobile Transporter of the Space Station Freedom.

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Burstadt, Paul L.

1984-01-01

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Critical engine system design characteristics for SSTO vehicles

NASA Astrophysics Data System (ADS)

Fanciullo, Thomas J.; Judd, D. C.; Obrien, C. J.

1992-02-01

Engine system design characteristics are summarized for typical vertical take-off and landing (VTOL) and vertical take-off and horizontal landing (VTHL) Strategic Defense Initiative Organization (SDIO) single stage to orbit (SSTO) vehicles utilizing plug nozzle configurations. Power cycle selection trades involved the unique modular platelet engine (MPE) with the use of (1) LO2 and LH2 at fixed and variable mixture ratios, (2) LO2 and propane or RP-1, and (3) dual fuels (LO2 with LH2 and C3H8). The number of thrust cells and modules were optimized. Dual chamber bell and a cluster of conventional bell nozzle configurations were examined for comparison with the plug configuration. Thrust modulation (throttling) was selected for thrust vector control. Installed thrust ratings were established to provide an additional 20 percent overthrust capability for engine out operation. Turbopumps were designed to operate at subcritical speeds to facilitate a wide range of throttling and long life. A unique dual spool arrangement with hydrostatic bearings was selected for the LH2 turbopump. Controls and health monitoring with expert systems for diagnostics are critical subsystems to ensure minimum maintenance and supportability for a less than seven day turnaround. The use of an idle mode start, in conjunction with automated health condition monitoring, allows the rocket propulsion system to operate reliably in the manner of present day aircraft propulsion.

Magnetohydrodynamic Augmentation of Pulse Detonation Rocket Engines (Preprint)

DTIC Science & Technology

2010-09-28

augmentation of the thrust . Ejectors typically transfer energy between streams through shear stress between separate flow streams, where a portion of the...the opportunity to extract energy and apply it to a separate stream where the net thrust can be increased. With MHD augmentation , such as in the Pulse...with the PDRIME for separate or additional thrust augmentation . Results show potential performance gains under many flight and operating conditions

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In one of the initial assembly steps for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, position the thrust structure. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Saturn Apollo Program

NASA Image and Video Library

1969-01-01

In one of the initial assembly steps for the Saturn IB launch vehicle's S-IB (first) stage, workers at the Michoud Assembly Facility (MAF) near New Orleans, Louisiana, complete the thrust structure. Developed by the Marshall Space Flight Center and built by the Chrysler Corporation at Michoud Assembly Facility (MAF), the S-IB utilized eight H-1 engines and each produced 200,000 pounds of thrust, a combined thrust of 1,600,000 pounds.

Performance of a RBCC Engine in Rocket-Operation

NASA Astrophysics Data System (ADS)

Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

Affordable Development and Demonstration of a Small NTR engine and Stage: A Preliminary NASA, DOE, and Industry Assessment

NASA Technical Reports Server (NTRS)

Borowski, S. K.; Sefcik, R. J.; Fittje, J. E.; McCurdy, D. R.; Qualls, A. L.; Schnitzler, B. G; Werner, J.; Weitzberg, A.; Joyner, C. R.

2015-01-01

In FY'11, Nuclear Thermal Propulsion (NTP) was identified as a key propulsion option under the Advanced In-Space Propulsion (AISP) component of NASA's Exploration Technology Development and Demonstration (ETDD) program A strategy was outlined by GRC and NASA HQ that included 2 key elements -"Foundational Technology Development" followed by specific "Technology Demonstration" projects. The "Technology Demonstration "element proposed ground technology demonstration (GTD) testing in the early 2020's, followed by a flight technology demonstration (FTD) mission by approx. 2025. In order to reduce development costs, the demonstration projects would focus on developing a small, low thrust (approx. 7.5 -16.5 klb(f)) engine that utilizes a "common" fuel element design scalable to the higher thrust (approx. 25 klb(f)) engines used in NASA's Mars DRA 5.0 study(NASA-SP-2009-566). Besides reducing development costs and allowing utilization of existing, flight proven engine hard-ware (e.g., hydrogen pumps and nozzles), small, lower thrust ground and flight demonstration engines can validate the technology and offer improved capability -increased payloads and decreased transit times -valued for robotic science missions identified in NASA's Decadal Study.

Dual throat engine design for a SSTO launch vehicle

NASA Technical Reports Server (NTRS)

Obrien, C. J.; Salmon, J. W.

1980-01-01

A propulsion system analysis of a dual fuel, dual throat engine for launch vehicle application was conducted. Basic dual throat engine characterization data are presented to allow vehicle optimization studies to be conducted. A preliminary baseline engine system was defined. Dual throat engine performance, envelope, and weight parametric data were generated over the parametric range of thrust from 890 to 8896 KN (200K to 2M lb-force), chamber pressure from 6.89 million to 34.5 million N/sq m (1000 to 5000 psia) thrust ratio from 1.2 to 5, and a range of mixture ratios for the two tripropellant combinations: LO2/RP-1 + LH2 and LO2/LCH4 + LH2. The results of the study indicate that the dual fuel dual throat engine is a viable single stage to orbit candidate.

Progress with variable cycle engines

NASA Technical Reports Server (NTRS)

Westmoreland, J. S.

1980-01-01

The evaluation of components of an advanced propulsion system for a future supersonic cruise vehicle is discussed. These components, a high performance duct burner for thrust augmentation and a low jet noise coannular exhaust nozzle, are part of the variable stream control engine. An experimental test program involving both isolated component and complete engine tests was conducted for the high performance, low emissions duct burner with excellent results. Nozzle model tests were completed which substantiate the inherent jet noise benefit associated with the unique velocity profile possible of a coannular exhaust nozzle system on a variable stream control engine. Additional nozzle model performance tests have established high thrust efficiency levels at takeoff and supersonic cruise for this nozzle system. Large scale testing of these two critical components is conducted using an F100 engine as the testbed for simulating the variable stream control engine.

Integrated flight/propulsion control - Adaptive engine control system mode

NASA Technical Reports Server (NTRS)

Yonke, W. A.; Terrell, L. A.; Meyers, L. P.

1985-01-01

The adaptive engine control system mode (ADECS) which is developed and tested on an F-15 aircraft with PW1128 engines, using the NASA sponsored highly integrated digital electronic control program, is examined. The operation of the ADECS mode, as well as the basic control logic, the avionic architecture, and the airframe/engine interface are described. By increasing engine pressure ratio (EPR) additional thrust is obtained at intermediate power and above. To modulate the amount of EPR uptrim and to prevent engine stall, information from the flight control system is used. The performance benefits, anticipated from control integration are shown for a range of flight conditions and power settings. It is found that at higher altitudes, the ADECS mode can increase thrust as much as 12 percent, which is used for improved acceleration, improved turn rate, or sustained turn angle.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This chart is an illustration of J-2 Engine characteristics. A cluster of five J-2 engines powered the Saturn V S-II (second) stage with each engine providing a thrust of 200,000 pounds. A single J-2 engine powered the S-IVB stage, the Saturn IB second stage, and the Saturn V third stage. The engine was uprated to provide 230,000 pounds of thrust for the fourth Apollo Saturn V flight and subsequent missions. Burning liquid hydrogen as fuel and using liquid oxygen as the oxidizer, the cluster of five J-2 engines for the S-II stage burned over one ton of propellant per second, during about 6 1/2 minutes of operation, to take the vehicle to an altitude of about 108 miles and a speed of near orbital velocity, about 17,400 miles per hour.

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James

2008-01-01

The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

The design and operating characteristics of an advanced 30-kW ammonia arcjet engine

NASA Technical Reports Server (NTRS)

Deininger, William D.; Pivirotto, Thomas J.; Brophy, John R.

1987-01-01

Experimental investigations were conducted to evaluate the effects of a contoured nozzle and modified cathode shape on ammonia arcjet engine performance. The contoured nozzle performance data were compared to the performance data of an arcjet which had a 38-deg included-angle, conical nozzle. Thrust improvements of up to 10 percent were demonstrated which corresponded to 3 percent improvements in specific impulse and 10 percent improvements in thrust efficiency. Performance characterizations for the modified cathode tip were conducted with the contoured nozzle arcjet. A uniform 15 percent decrease in arc voltage was demonstrated over a mass flow range of 0.175 to 0.350 g/s. A 4 percent improvement in thrust efficiency was noted at 22.0 kW.

Thrust performance of a variable-geometry, divergent exhaust nozzle on a turbojet engine at altitude

NASA Technical Reports Server (NTRS)

Straight, D. M.; Collom, R. R.

1983-01-01

A variable geometry, low aspect ratio, nonaxisymmetric, two dimensional, convergent-divergent exhaust nozzle was tested at simulated altitude on a turbojet engine to obtain baseline axial, dry thrust performance over wide ranges of operating nozzle pressure ratios, throat areas, and internal expansion area ratios. The thrust data showed good agreement with theory and scale model test results after the data were corrected for seal leakage and coolant losses. Wall static pressure profile data were also obtained and compared with one dimensional theory and scale model data. The pressure data indicate greater three dimensional flow effects in the full scale tests than with models. The leakage and coolant penalties were substantial, and the method to determine them is included.

DC-9 flight demonstration program with refanned JT8D engines. Volume 4: Flyover noise

NASA Technical Reports Server (NTRS)

1975-01-01

Flyover noise tests were conducted to determine the noise reductions achievable by modifying the engines and nacelles of DC-9-30 airplanes. The two stage fan of the JT8D-9 engine was replaced with a larger diameter, single stage fan and sound absorbing materials were incorporated in the engines and nacelles. The noise levels were determined to be 95.3 EPNdB at the sideline, 96.2 EPNdB for a full thrust takeoff, 87.5 EPNdB for takeoff with thrust cutback, and 97.4 EPNdB for landing approach. The noise reductions relative to the hardwall JT8D-9 were 8.2 EPNdB for takeoff with cutback and 8.7 EPNdB for landing. The 90 EPNdB noise contour areas were reduced by 40% for missions requiring maximum design takeoff and landing weights. For typical mission weights, the reductions were 19% for full thrust takeoff and 34% for takeoff with cutback. The 95 EPNdB contour areas were reduced by 50% for takeoff and 30% for takeoff with cutback for both missions.

Low-speed wind-tunnel investigation of a large scale advanced arrow-wing supersonic transport configuration with engines mounted above wing for upper-surface blowing

NASA Technical Reports Server (NTRS)

Shivers, J. P.; Mclemore, H. C.; Coe, P. L., Jr.

1976-01-01

Tests have been conducted in a full scale tunnel to determine the low speed aerodynamic characteristics of a large scale advanced arrow wing supersonic transport configuration with engines mounted above the wing for upper surface blowing. Tests were made over an angle of attack range of -10 deg to 32 deg, sideslip angles of + or - 5 deg, and a Reynolds number range of 3,530,000 to 7,330,000. Configuration variables included trailing edge flap deflection, engine jet nozzle angle, engine thrust coefficient, engine out operation, and asymmetrical trailing edge boundary layer control for providing roll trim. Downwash measurements at the tail were obtained for different thrust coefficients, tail heights, and at two fuselage stations.

Saturn Apollo Program

NASA Image and Video Library

1960-01-01

This chart provides the vital statistics for the F-1 rocket engine. Developed by Rocketdyne, under the direction of the Marshall Space Flight Center, the F-1 engine was utilized in a cluster of five engines to propel the Saturn V's first stage, the S-IC. Liquid oxygen and kerosene were used as its propellant. Initially rated at 1,500,000 pounds of thrust, the engine was later uprated to 1,522,000 pounds of thrust after the third Saturn V launch (Apollo 8, the first marned Saturn V mission) in December 1968. The cluster of five F-1 engines burned over 15 tons of propellant per second, during its two and one-half minutes of operation, to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour.

Highly integrated digital engine control system on an F-15 airplane

NASA Technical Reports Server (NTRS)

Burcham, F. W., Jr.; Haering, E. A., Jr.

1984-01-01

The Highly Integrated Digital Electronic Control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine/airframe control systems. This system is being used on the F-15 airplane. An integrated flightpath management mode and an integrated adaptive engine stall margin mode are implemented into the system. The adaptive stall margin mode is a highly integrated mode in which the airplane flight conditions, the resulting inlet distortion, and the engine stall margin are continuously computed; the excess stall margin is used to uptrim the engine for more thrust. The integrated flightpath management mode optimizes the flightpath and throttle setting to reach a desired flight condition. The increase in thrust and the improvement in airplane performance is discussed.

Applications of the DOE/NASA wind turbine engineering information system

NASA Technical Reports Server (NTRS)

Neustadter, H. E.; Spera, D. A.

1981-01-01

A statistical analysis of data obtained from the Technology and Engineering Information Systems was made. The systems analyzed consist of the following elements: (1) sensors which measure critical parameters (e.g., wind speed and direction, output power, blade loads and component vibrations); (2) remote multiplexing units (RMUs) on each wind turbine which frequency-modulate, multiplex and transmit sensor outputs; (3) on-site instrumentation to record, process and display the sensor output; and (4) statistical analysis of data. Two examples of the capabilities of these systems are presented. The first illustrates the standardized format for application of statistical analysis to each directly measured parameter. The second shows the use of a model to estimate the variability of the rotor thrust loading, which is a derived parameter.

A Demonstration of a Retrofit Architecture for Intelligent Control and Diagnostics of a Turbofan Engine

NASA Technical Reports Server (NTRS)

Litt, Jonathan S.; Turso, James A.; Shah, Neerav; Sowers, T. Shane; Owen, A. Karl

2005-01-01

A retrofit architecture for intelligent turbofan engine control and diagnostics that changes the fan speed command to maintain thrust is proposed and its demonstration in a piloted flight simulator is described. The objective of the implementation is to increase the level of autonomy of the propulsion system, thereby reducing pilot workload in the presence of anomalies and engine degradation due to wear. The main functions of the architecture are to diagnose the cause of changes in the engine s operation, warning the pilot if necessary, and to adjust the outer loop control reference signal in response to the changes. This requires that the retrofit control architecture contain the capability to determine the changed relationship between fan speed and thrust, and the intelligence to recognize the cause of the change in order to correct it or warn the pilot. The proposed retrofit architecture is able to determine the fan speed setting through recognition of the degradation level of the engine, and it is able to identify specific faults and warn the pilot. In the flight simulator it was demonstrated that when degradation is introduced into an engine with standard fan speed control, the pilot needs to take corrective action to maintain heading. Utilizing the intelligent retrofit control architecture, the engine thrust is automatically adjusted to its expected value, eliminating yaw without pilot intervention.

Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program

NASA Technical Reports Server (NTRS)

Asbury, Scott C.; Yetter, Jeffrey A.

2000-01-01

The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9%-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems.

Rapid prototype fabrication processes for high-performance thrust cells

NASA Technical Reports Server (NTRS)

Hunt, K.; Chwiedor, T.; Diab, J.; Williams, R.

1994-01-01

The Thrust Cell Technologies Program (Air Force Phillips Laboratory Contract No. F04611-92-C-0050) is currently being performed by Rocketdyne to demonstrate advanced materials and fabrication technologies which can be utilized to produce low-cost, high-performance thrust cells for launch and space transportation rocket engines. Under Phase 2 of the Thrust Cell Technologies Program (TCTP), rapid prototyping and investment casting techniques are being employed to fabricate a 12,000-lbf thrust class combustion chamber for delivery and hot-fire testing at Phillips Lab. The integrated process of investment casting directly from rapid prototype patterns dramatically reduces design-to-delivery cycle time, and greatly enhances design flexibility over conventionally processed cast or machined parts.

A digital controller for variable thrust liquid rocket engines

NASA Astrophysics Data System (ADS)

Feng, X.; Zhang, Y. L.; Chen, Q. Z.

1993-06-01

The paper describes the design and development of a built-in digital controller (BDC) for the variable thrust liquid rocket engine (VTLRE). Particular attention is given to the function requirements of the BDC, the hardware and software configuration, and the testing process, as well as to the VTLRE real-time computer simulation system used for the development of the BDC. A diagram of the VLTRE control system is presented as well as block diagrams illustrating the hardware and software configuration of the BDC.

Saturn Apollo Program

NASA Image and Video Library

1964-11-01

The Saturn I S-IV stage (second stage) assembly for the SA-9 mission underwent the weight and balance test in the hanger building at Cape Canaveral. The S-IV stage had six RL-10 engines which used liquid hydrogen and liquid oxygen as its propellants arranged in a circle. Each RL-10 engine produced a thrust of 15,000 pounds, a total combined thrust of 90,000 pounds. The SA-9 mission was the first Saturn with operational payload Pegasus I, meteoroid detection satellite, and launched on February 16, 1965.

The Mission Defines the Cycle: Turbojet, Turbofan and Variable Cycle Engines for High Speed Propulsion

DTIC Science & Technology

2010-09-01

RTO-EN-AVT-185 2 - 1 The Mission Defines the Cycle: Turbojet, Turbofan and Variable Cycle Engines for High Speed Propulsion Joachim Kurzke...following turbine parts 1 %. With T4=2000K the amounts of cooling air are 10% and 6% respectively. Burner pressure ratio is taken into account with 0.97 and...Figure 2 . Figure 3 shows specific thrust (i.e. thrust per unit of air flow) and specific fuel consumption SFC for three altitude / Mach number

Coal-Fired Rocket Engine

NASA Technical Reports Server (NTRS)

Anderson, Floyd A.

1987-01-01

Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

14 CFR 25.121 - Climb: One-engine-inoperative.

Code of Federal Regulations, 2014 CFR

2014-01-01

... engine inoperative and the remaining engines at the power or thrust available when retraction of the... retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun...

14 CFR 25.121 - Climb: One-engine-inoperative.

Code of Federal Regulations, 2011 CFR

2011-01-01

... engine inoperative and the remaining engines at the power or thrust available when retraction of the... retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun...

14 CFR 25.121 - Climb: One-engine-inoperative.

Code of Federal Regulations, 2012 CFR

2012-01-01

... engine inoperative and the remaining engines at the power or thrust available when retraction of the... retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun...

14 CFR 25.121 - Climb: One-engine-inoperative.

Code of Federal Regulations, 2010 CFR

2010-01-01

... engine inoperative and the remaining engines at the power or thrust available when retraction of the... retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun...

14 CFR 25.121 - Climb: One-engine-inoperative.

Code of Federal Regulations, 2013 CFR

2013-01-01

... engine inoperative and the remaining engines at the power or thrust available when retraction of the... retracted; and (2) The weight equal to the weight existing when retraction of the landing gear is begun...

Charged particle measurements on a 30-CM diameter mercury ion engine thrust beam

NASA Technical Reports Server (NTRS)

Sellen, J. M., Jr.; Komatsu, G. K.; Hoffmaster, D. K.; Kemp, R. F.

1974-01-01

Measurements of both thrust ions and charge exchange ions were made in the beam of a 30 centimeter diameter electron bombardment mercury ion thruster. A qualitative model is presented which describes magnitudes of charge exchange ion formation and motions of these ions in the weak electric field structure of the neutralized thrust beam plasma. Areas of agreement and discrepancy between observed and modeled charge exchange properties are discussed.

Aircraft dual-shaft jet engine with indirect action fuel flow controller

NASA Astrophysics Data System (ADS)

Tudosie, Alexandru-Nicolae

2017-06-01

The paper deals with an aircraft single-jet engine's control system, based on a fuel flow controller. Considering the engine as controlled object and its thrust the most important operation effect, from the multitude of engine's parameters only its rotational speed n is measurable and proportional to its thrust, so engine's speed has become the most important controlled parameter. Engine's control system is based on fuel injection Qi dosage, while the output is engine's speed n. Based on embedded system's main parts' mathematical models, the author has described the system by its block diagram with transfer functions; furthermore, some Simulink-Matlab simulations are performed, concerning embedded system quality (its output parameters time behavior) and, meanwhile, some conclusions concerning engine's parameters mutual influences are revealed. Quantitative determinations are based on author's previous research results and contributions, as well as on existing models (taken from technical literature). The method can be extended for any multi-spool engine, single- or twin-jet.

MHD thrust vectoring of a rocket engine

NASA Astrophysics Data System (ADS)

Labaune, Julien; Packan, Denis; Tholin, Fabien; Chemartin, Laurent; Stillace, Thierry; Masson, Frederic

2016-09-01

In this work, the possibility to use MagnetoHydroDynamics (MHD) to vectorize the thrust of a solid propellant rocket engine exhaust is investigated. Using a magnetic field for vectoring offers a mass gain and a reusability advantage compared to standard gimbaled, elastomer-joint systems. Analytical and numerical models were used to evaluate the flow deviation with a 1 Tesla magnetic field inside the nozzle. The fluid flow in the resistive MHD approximation is calculated using the KRONOS code from ONERA, coupling the hypersonic CFD platform CEDRE and the electrical code SATURNE from EDF. A critical parameter of these simulations is the electrical conductivity, which was evaluated using a set of equilibrium calculations with 25 species. Two models were used: local thermodynamic equilibrium and frozen flow. In both cases, chlorine captures a large fraction of free electrons, limiting the electrical conductivity to a value inadequate for thrust vectoring applications. However, when using chlorine-free propergols with 1% in mass of alkali, an MHD thrust vectoring of several degrees was obtained.

SSME/side loads analysis for flight configuration, revision A. [structural analysis of space shuttle main engine under side load excitation

NASA Technical Reports Server (NTRS)

Holland, W.

1974-01-01

This document describes the dynamic loads analysis accomplished for the Space Shuttle Main Engine (SSME) considering the side load excitation associated with transient flow separation on the engine bell during ground ignition. The results contained herein pertain only to the flight configuration. A Monte Carlo procedure was employed to select the input variables describing the side load excitation and the loads were statistically combined. This revision includes an active thrust vector control system representation and updated orbiter thrust structure stiffness characteristics. No future revisions are planned but may be necessary as system definition and input parameters change.

Preliminary Flight Results of a Fly-by-throttle Emergency Flight Control System on an F-15 Airplane

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Maine, Trindel A.; Fullerton, C. Gordon; Wells, Edward A.

1993-01-01

A multi-engine aircraft, with some or all of the flight control system inoperative, may use engine thrust for control. NASA Dryden has conducted a study of the capability and techniques for this emergency flight control method for the F-15 airplane. With an augmented control system, engine thrust, along with appropriate feedback parameters, is used to control flightpath and bank angle. Extensive simulation studies were followed by flight tests. The principles of throttles only control, the F-15 airplane, the augmented system, and the flight results including actual landings with throttles-only control are discussed.

Static noise tests on modified augmentor wing jet STOL research aircraft

NASA Technical Reports Server (NTRS)

Cook, G. R.; Lilley, B. F.

1981-01-01

Noise measurements were made to determine if recent modifications made to the bifurcated jetpipe to increase engine thrust had at the same time reduced the noise level. The noise field was measured by a 6-microphone array positioned on a 30.5m (100 ft) sideline between 90 and 150 degrees from the left engine inlet. Noise levels were recorded at three flap angles over a range of engine thrust settings from flight idle to emergency power and plotted in one-third octave band spectra. Little attenuation was observed at maximum power, but significant attenuation was achieved at approach and cruise power levels.

14 CFR 33.66 - Bleed air system.

Code of Federal Regulations, 2010 CFR

2010-01-01

... Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.66 Bleed air system. The engine must supply bleed air without adverse effect on the engine, excluding reduced thrust or power...

14 CFR 33.66 - Bleed air system.

Code of Federal Regulations, 2011 CFR

2011-01-01

... Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.66 Bleed air system. The engine must supply bleed air without adverse effect on the engine, excluding reduced thrust or power...

Neural Network and Regression Methods Demonstrated in the Design Optimization of a Subsonic Aircraft

NASA Technical Reports Server (NTRS)

Hopkins, Dale A.; Lavelle, Thomas M.; Patnaik, Surya

2003-01-01

The neural network and regression methods of NASA Glenn Research Center s COMETBOARDS design optimization testbed were used to generate approximate analysis and design models for a subsonic aircraft operating at Mach 0.85 cruise speed. The analytical model is defined by nine design variables: wing aspect ratio, engine thrust, wing area, sweep angle, chord-thickness ratio, turbine temperature, pressure ratio, bypass ratio, fan pressure; and eight response parameters: weight, landing velocity, takeoff and landing field lengths, approach thrust, overall efficiency, and compressor pressure and temperature. The variables were adjusted to optimally balance the engines to the airframe. The solution strategy included a sensitivity model and the soft analysis model. Researchers generated the sensitivity model by training the approximators to predict an optimum design. The trained neural network predicted all response variables, within 5-percent error. This was reduced to 1 percent by the regression method. The soft analysis model was developed to replace aircraft analysis as the reanalyzer in design optimization. Soft models have been generated for a neural network method, a regression method, and a hybrid method obtained by combining the approximators. The performance of the models is graphed for aircraft weight versus thrust as well as for wing area and turbine temperature. The regression method followed the analytical solution with little error. The neural network exhibited 5-percent maximum error over all parameters. Performance of the hybrid method was intermediate in comparison to the individual approximators. Error in the response variable is smaller than that shown in the figure because of a distortion scale factor. The overall performance of the approximators was considered to be satisfactory because aircraft analysis with NASA Langley Research Center s FLOPS (Flight Optimization System) code is a synthesis of diverse disciplines: weight estimation, aerodynamic analysis, engine cycle analysis, propulsion data interpolation, mission performance, airfield length for landing and takeoff, noise footprint, and others.

Effect of Operating Frequency on PDE Driven Ejector Thrust Performance

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pal, Sibtosh; Landry, K.; Shehadeh, R.; Bouvet, N.; Lee, S.-Y.

2005-01-01

Results of an on-going study of pulse detonation engine driven ejectors are presented and discussed. The experiments were conducted using a pulse detonation engine (PDE) designed to operate at frequencies up to 50 Hz. The PDE used in these experiments utilizes an equi-molar mixture of oxygen and nitrogen as the oxidizer, and ethylene (C2H4) as the fuel, with the propellant mixture having an equivalence ratio of one. A line of sight laser absorption technique was used to determine the time needed for proper filling of the tube. Thrust measurements were made using an integrated spring damper system coupled with a linear variable displacement transducer. The baseline thrust of the PDE was first measured at each desired frequency and agrees with experimental and modeling results found in the literature. Thrust augmentation measurements were then made for constant diameter ejectors. The ejectors had varying lengths, and two different inlet geometries were tested for each ejector configuration. The parameter space for the study included PDE operation frequency, ejector length, overlap distance and the radius of curvature for the ejector inlets. For the studied experimental matrix, the results showed a maximum thrust augmentation of 106% at an operational frequency of 30 Hz.

The Experimental Study about the Effect of Operating Conditions on Multi-tube Pulse Detonation Engine Performance

NASA Astrophysics Data System (ADS)

Kim, Jung-Min; Han, Hyung-Seok; Choi, Jeong-Yeol

2018-04-01

This study examines a multi-tube pulse detonation engine (PDE) which has a type of constant volume combustion. We designed and made the multi-tube PDE and then conducted an experiment in various operating frequencies and equivalence ratios. First, experiments with operating frequencies of 40, 80, 120, 160, and 200 Hz resulted in an average thrust and specific impulse 23.14 N and 42.34 s. The next experiment resulted in the equivalence ratio varying from 0.81 to 1.38, which resulted in an average thrust and specific impulse 22.36 N and 40.11 s. The average detonation velocity was 8% lower than that calculated according to C-J theory. The incidence ratios of the detonation wave were stable with the exception of the operating frequency of 200 Hz. However, at 200 Hz, the incidence ratio was less than 50%. We assumed that a low fill fraction occurred for this problem. The thrust of the PDE increased with the operating frequency. However, the thrust increase was at a lower rate than in previous studies, because of a lost thrust output result from the slow response time of the load cell amplifier.

Subsonic flight test evaluation of a performance seeking control algorithm on an F-15 airplane

NASA Technical Reports Server (NTRS)

Gilyard, Glenn B.; Orme, John S.

1992-01-01

The subsonic flight test evaluation phase of the NASA F-15 (powered by F 100 engines) performance seeking control program was completed for single-engine operation at part- and military-power settings. The subsonic performance seeking control algorithm optimizes the quasi-steady-state performance of the propulsion system for three modes of operation. The minimum fuel flow mode minimizes fuel consumption. The minimum thrust mode maximizes thrust at military power. Decreases in thrust-specific fuel consumption of 1 to 2 percent were measured in the minimum fuel flow mode; these fuel savings are significant, especially for supersonic cruise aircraft. Decreases of up to approximately 100 degree R in fan turbine inlet temperature were measured in the minimum temperature mode. Temperature reductions of this magnitude would more than double turbine life if inlet temperature was the only life factor. Measured thrust increases of up to approximately 15 percent in the maximum thrust mode cause substantial increases in aircraft acceleration. The system dynamics of the closed-loop algorithm operation were good. The subsonic flight phase has validated the performance seeking control technology, which can significantly benefit the next generation of fighter and transport aircraft.

78 FR 63015 - Exhaust Emissions Standards for New Aircraft Gas Turbine Engines and Identification Plate for...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-10-23

... Turbine Engines and Identification Plate for Aircraft Engines AGENCY: Federal Aviation Administration (FAA... regulatory requirements for aircraft turbofan or turbojet engines with rated thrusts greater than 26.7... standards for certain turbine engine powered airplanes to incorporate the standards promulgated by the...

Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.

A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injectionmore » system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.« less

Emergency Control Aircraft System Using Thrust Modulation

NASA Technical Reports Server (NTRS)

Burken, John J. (Inventor); Burcham, Frank W., Jr. (Inventor)

2000-01-01

A digital longitudinal Aircraft Propulsion Control (APC system of a multiengine aircraft is provided by engine thrust modulation in response to comparing an input flightpath angle signal (gamma)c from a pilot thumbwheel. or an ILS system with a sensed flightpath angle y to produce an error signal (gamma)e that is then integrated (with reasonable limits) to generate a drift correction signal to be added to the error signal (gamma)e after first subtracting a lowpass filtered velocity signal Vel(sub f) for phugoid damping. The output error signal is multiplied by a constant to produce an aircraft thrust control signal ATC of suitable amplitude to drive a throttle servo for all engines. each of which includes its own full-authority digital engine control (FADEC) computer. An alternative APC system omits sensed flightpath angle feedback and instead controls the flightpath angle by feedback of the lowpass filtered velocity signal Vel(sub f) which also inherently provides phugoid damping. The feature of drift compensation is retained.

Saturn Apollo Program

NASA Image and Video Library

1967-01-01

This is a cutaway illustration of the Saturn V launch vehicle with callouts of the major components. The Saturn V is the largest and most powerful launch vehicle developed in the United States. It was a three stage rocket, 363 feet in height, used for sending American astronauts to the moon and for placing the Skylab in Earth orbit. The Saturn V was designed to perform Earth orbital missions through the use of the first two stages, while all three stages were used for lunar expeditions. The S-IC stage (first stage) was powered by five F- engines, which burned kerosene and liquid oxygen to produce more than 7,500,000 pounds of thrust. The S-II (second) stage was powered by five J-2 engines, that burned liquid hydrogen and liquid oxygen and produced 1,150,000 pounds thrust. The S-IVB (third) stage used one J-2 engine, producing 230,000 pounds of thrust, with a re-start capability. The Marshall Space Flight Center and its contractors designed, developed, and assembled the Saturn V launch vehicle stages.

Propulsion technology needs for advanced space transportation systems. [orbit maneuvering engine (space shuttle), space shuttle boosters

NASA Technical Reports Server (NTRS)

Gregory, J. W.

1975-01-01

Plans are formulated for chemical propulsion technology programs to meet the needs of advanced space transportation systems from 1980 to the year 2000. The many possible vehicle applications are reviewed and cataloged to isolate the common threads of primary propulsion technology that satisfies near term requirements in the first decade and at the same time establish the technology groundwork for various potential far term applications in the second decade. Thrust classes of primary propulsion engines that are apparent include: (1) 5,000 to 30,000 pounds thrust for upper stages and space maneuvering; and (2) large booster engines of over 250,000 pounds thrust. Major classes of propulsion systems and the important subdivisions of each class are identified. The relative importance of each class is discussed in terms of the number of potential applications, the likelihood of that application materializing, and the criticality of the technology needed. Specific technology programs are described and scheduled to fulfill the anticipated primary propulsion technology requirements.

Mathematical model of marine diesel engine simulator for a new methodology of self propulsion tests

NASA Astrophysics Data System (ADS)

Izzuddin, Nur; Sunarsih, Priyanto, Agoes

2015-05-01

As a vessel operates in the open seas, a marine diesel engine simulator whose engine rotation is controlled to transmit through propeller shaft is a new methodology for the self propulsion tests to track the fuel saving in a real time. Considering the circumstance, this paper presents the real time of marine diesel engine simulator system to track the real performance of a ship through a computer-simulated model. A mathematical model of marine diesel engine and the propeller are used in the simulation to estimate fuel rate, engine rotating speed, thrust and torque of the propeller thus achieve the target vessel's speed. The input and output are a real time control system of fuel saving rate and propeller rotating speed representing the marine diesel engine characteristics. The self-propulsion tests in calm waters were conducted using a vessel model to validate the marine diesel engine simulator. The simulator then was used to evaluate the fuel saving by employing a new mathematical model of turbochargers for the marine diesel engine simulator. The control system developed will be beneficial for users as to analyze different condition of vessel's speed to obtain better characteristics and hence optimize the fuel saving rate.

Saturn Apollo Program

NASA Image and Video Library

1964-01-01

This close-up view of the F-1 engine for the Saturn V S-IC (first) stage shows the engine's complexity, and also its large size as it dwarfs the technician. Developed by Rocketdyne, under the direction of the Marshall Space Flight Center, the F-1 engine was utilized in a cluster of five engines to propel the Saturn V's first stage, the S-IC. Liquid oxygen and kerosene were used as its propellant. Initially rated at 1,500,000 pounds of thrust, the engine was later uprated to 1,522,000 pounds of thrust after the third Saturn V launch (Apollo 8, the first marned Saturn V mission) in December 1968. The cluster of five F-1 engines burned over 15 tons of propellant per second, during its two and one-half minutes of operation, to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour.

Saturn Apollo Program

NASA Image and Video Library

1963-01-01

A close-up view of the F-1 Engine for the Saturn V S-IC (first) stage depicts the complexity of the engine. Developed by Rocketdyne under the direction of the Marshall Space Flight Center, the F-1 engine was utilized in a cluster of five engines to propel the Saturn V's first stage, the S-IC. Liquid oxygen and kerosene were used as its propellant. Initially rated at 1,500,000 pounds of thrust, the engine was later uprated to 1,522,000 pounds of thrust after the third Saturn V launch (Apollo 8, the first marned Saturn V mission) in December 1968. The cluster of five F-1 engines burned over 15 tons of propellant per second, during its two and one-half minutes of operation, to take the vehicle to a height of about 36 miles and to a speed of about 6,000 miles per hour.

SST Technology Follow-on Program - Phase I, Performance Evaluation of an SST Noise Suppressor Nozzle System. Volume 1. Suppressed Mode.

DTIC Science & Technology

ACOUSTIC INSULATION, *TURBOJET EXHAUST NOZZLES, *JET ENGINE NOISE, REDUCTION, JET TRANSPORT AIRCRAFT, THRUST AUGMENTATION , SUPERSONIC NOZZLES, DUCT...INLETS, CONVERGENT DIVERGENT NOZZLES, SUBSONIC FLOW, SUPERSONIC FLOW, SUPPRESSORS, TURBOJET INLETS, BAFFLES, JET PUMPS, THRUST , DRAG, TEMPERATURE

Upper Stage Flight Experiment 10K Engine Design and Test Results

NASA Technical Reports Server (NTRS)

Ross, R.; Morgan, D.; Crockett, D.; Martinez, L.; Anderson, W.; McNeal, C.

2000-01-01

A 10,000 lbf thrust chamber was developed for the Upper Stage Flight Experiment (USFE). This thrust chamber uses hydrogen peroxide/JP-8 oxidizer/fuel combination. The thrust chamber comprises an oxidizer dome and manifold, catalyst bed assembly, fuel injector, and chamber/nozzle assembly. Testing of the engine was done at NASA's Stennis Space Center (SSC) to verify its performance and life for future upper stage or Reusable Launch Vehicle applications. Various combinations of silver screen catalyst beds, fuel injectors, and combustion chambers were tested. Results of the tests showed high C* efficiencies (97% - 100%) and vacuum specific impulses of 275 - 298 seconds. With fuel film cooling, heating rates were low enough that the silica/quartz phenolic throat experienced minimal erosion. Mission derived requirements were met, along with a perfect safety record.

Manual Manipulation of Engine Throttles for Emergency Flight Control

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Fullerton, C. Gordon; Maine, Trindel A.

2004-01-01

If normal aircraft flight controls are lost, emergency flight control may be attempted using only engines thrust. Collective thrust is used to control flightpath, and differential thrust is used to control bank angle. Flight test and simulation results on many airplanes have shown that pilot manipulation of throttles is usually adequate to maintain up-and-away flight, but is most often not capable of providing safe landings. There are techniques that will improve control and increase the chances of a survivable landing. This paper reviews the principles of throttles-only control (TOC), a history of accidents or incidents in which some or all flight controls were lost, manual TOC results for a wide range of airplanes from simulation and flight, and suggested techniques for flying with throttles only and making a survivable landing.

Experimental Study of Propulsion Performance by Single-Pulse Rotating Detonation with Gaseous Fuels-Oxygen Mixtures

NASA Astrophysics Data System (ADS)

Toshimitsu, Kazuhiko; Hara, Kosei; Mikajiri, Shuuto; Takiguchi, Naoki

2016-12-01

A rotating detonation engine (RDE) is one of candidates of aerospace engines for supersonic cruse, which is better for propulsion system than a pulse detonation engine (PDE) from the view of continuous thrust and simple structure. The propulsion performance of a proto-type RDE and a PDE by single pulse explosion with methane-oxygen is investigated. Furthermore, the performance of the RDE with acetylene-oxygen gas mixtures is investigated. Its impulse is estimated through ballistic pendulum method with maximum displacement and damping ratio. The comparison of specific impulses of the mixture gases at atmospheric pressure is shown. The specific impulses of the RDE and the PDE are almost same with methane-oxygen gas. Furthermore, the fuel-base specific impulse of the RDE with acetylene-oxygen gas is about over twice as large as one of methane-oxygen, and its maximum specific impulse is 1100 seconds.

Status on the Verification of Combustion Stability for the J-2X Engine Thrust Chamber Assembly

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hinerman, Tim; Kenny, R. Jeremy; Hulka, Jim; Barnett, Greg; Dodd, Fred; Martin, Tom

2013-01-01

Development is underway of the J -2X engine, a liquid oxygen/liquid hydrogen rocket engine for use on the Space Launch System. The Engine E10001 began hot fire testing in June 2011 and testing will continue with subsequent engines. The J -2X engine main combustion chamber contains both acoustic cavities and baffles. These stability aids are intended to dampen the acoustics in the main combustion chamber. Verification of the engine thrust chamber stability is determined primarily by examining experimental data using a dynamic stability rating technique; however, additional requirements were included to guard against any spontaneous instability or rough combustion. Startup and shutdown chug oscillations are also characterized for this engine. This paper details the stability requirements and verification including low and high frequency dynamics, a discussion on sensor selection and sensor port dynamics, and the process developed to assess combustion stability. A status on the stability results is also provided and discussed.

Highly integrated digital engine control system on an F-15 airplane

NASA Technical Reports Server (NTRS)

Burcham, F. W., Jr.; Haering, E. A., Jr.

1984-01-01

The highly integrated digital electronic control (HIDEC) program will demonstrate and evaluate the improvements in performance and mission effectiveness that result from integrated engine-airframe control systems. This system is being used on the F-15 airplane at the Dryden Flight Research Facility of NASA Ames Research Center. An integrated flightpath management mode and an integrated adaptive engine stall margin mode are being implemented into the system. The adaptive stall margin mode is a highly integrated mode in which the airplane flight conditions, the resulting inlet distortion, and the engine stall margin are continuously computed; the excess stall margin is used to uptrim the engine for more thrust. The integrated flightpath management mode optimizes the flightpath and throttle setting to reach a desired flight condition. The increase in thrust and the improvement in airplane performance is discussed in this paper.

ESCORT: A Pratt & Whitney nuclear thermal propulsion and power system for manned mars missions

NASA Astrophysics Data System (ADS)

Feller, Gerald J.; Joyner, Russell

1999-01-01

The purpose of this paper is to describe the conceptual design of an upgrade to the Pratt & Whitney ESCORT nuclear thermal rocket engine. The ESCORT is a bimodal engine capable of supporting a wide range of vehicle propulsive and electrical power requirements. The ESCORT engine is powered by a fast-spectrum beryllium-reflected CERMET-fueled nuclear reactor. In propulsive mode, the reactor is used to heat hot hydrogen to approximately 2700 K which is expanded through a converging/diverging nozzle to generate thrust. Heat pickup in the nozzle and the radial beryllium reflectors is used to drive the turbomachinery in the ESCORT expander cycle. In electrical mode, the reactor is used to heat a mixture of helium and xenon to drive a closed-loop Brayton cycle in order to generate electrical energy. This closed loop system has the additional function of a decay heat removal system after the propulsive mode operation is discontinued. The original ESCORT design was capable of delivering 4448.2 N (1000 lbf) of thrust at a vacuum impulse level of approximately 900 s. Design Reference Mission requirements (DRM) from NASA Johnson Space Center and NASA Lewis Research Center studies in 1997 and 1998 have detailed upgraded requirements for potential manned Mars missions. The current NASA DRM requires a nuclear thermal propulsion system capable of delivering total mission requirements of 200170 N (45000 lbf) thrust and 50 kWe of spacecraft electrical power. This is met assuming three engines capable of each delivering 66723 N (15000 lbf) of vacuum thrust and 25 kWe of electrical power. The individual engine requirements were developed assuming three out of three engine reliability for propulsion and two out of three engine reliability for spacecraft electrical power. The approximate target vacuum impulse is 925 s. The Pratt & Whitney ESCORT concept was upgraded to meet these requirements. The hexagonal prismatic fuel elements were modified to address the uprated power requirements while maintaining the peak fuel temperature below the 2880 K limit for W-UO2 CERMET fuels. A system integrated performance methodology was developed to assess the sensitivity to weight, thrust and impulse to the DRM requirements. Propellant tanks, shielding, and Brayton cycle power conversion unit requirements were included in this evaluation.

Wind-tunnel investigation of the powered low-speed longitudinal aerodynamics of the Vectored-Engine-Over (VEO) wing fighter configuration

NASA Technical Reports Server (NTRS)

Paulson, J. W.; Whitten, P. D.; Stumpfl, S. C.

1982-01-01

A wind-tunnel investigation incorporating both static and wind-on testing was conducted in the Langley 4- by 7-Meter Tunnel to determine the effects of vectored thrust along with spanwise blowing on the low-speed aerodynamics of an advanced fighter configuration. Data were obtained over a large range of thrust coefficients corresponding to takeoff and landing thrust settings for many nozzle configurations. The complete set of static thrust data and the complete set of longitudinal aerodynamic data obtained in the investigation are presented. These data are intended for reference purposes and, therefore, are presented without analysis or comment. The analysis of the thrust-induced effects found in the investigation are not discussed.

Entropy-Based Performance Analysis of Jet Engines; Methodology and Application to a Generic Single-Spool Turbojet

NASA Astrophysics Data System (ADS)

Abbas, Mohammad

Recently developed methodology that provides the direct assessment of traditional thrust-based performance of aerospace vehicles in terms of entropy generation (i.e., exergy destruction) is modified for stand-alone jet engines. This methodology is applied to a specific single-spool turbojet engine configuration. A generic compressor performance map along with modeled engine component performance characterizations are utilized in order to provide comprehensive traditional engine performance results (engine thrust, mass capture, and RPM), for on and off-design engine operation. Details of exergy losses in engine components, across the entire engine, and in the engine wake are provided and the engine performance losses associated with their losses are discussed. Results are provided across the engine operating envelope as defined by operational ranges of flight Mach number, altitude, and fuel throttle setting. The exergy destruction that occurs in the engine wake is shown to be dominant with respect to other losses, including all exergy losses that occur inside the engine. Specifically, the ratio of the exergy destruction rate in the wake to the exergy destruction rate inside the engine itself ranges from 1 to 2.5 across the operational envelope of the modeled engine.

Non-Toxic Dual Thrust Reaction Control Engine Development for On-Orbit APS Applications

NASA Technical Reports Server (NTRS)

Robinson, Philip J.; Veith, Eric M.

2003-01-01

A non-toxic dual thrust proof-of-concept demonstration engine was successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the Next Generation Launch Technology (NGLT) program. The demonstration engine utilized the existing Kistler K-1 870 lbf LOX/Ethanol orbital maneuvering engine ( O m ) coupled with some special test equipment (STE) that enabled engine operation at 870 lbf in the primary mode and 25 lbf in the vernier mode. Ambient testing in primary mode varied mixture ratio (MR) from 1.28 to 1.71 and chamber pressure (P(c) from 110 to 181 psia, and evaluated electrical pulse widths (EPW) of 0.080, 0.100 and 0.250 seconds. Altitude testing in vernier mode explored igniter and thruster pulsing characteristics, long duration steady state operation (greater than 420 sec) and the impact of varying the percent fuel film cooling on vernier performance and chamber thermal response at low PC (4 psia). Data produced from the testing provided calibration of the performance and thermal models used in the design of the next version of the dual thrust Reaction Control Engine (RCE).

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2010 CFR

2010-01-01

....73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or... aircraft, without overtemperature, surge, stall, or other detrimental factors occurring to the engine...

14 CFR 33.73 - Power or thrust response.

Code of Federal Regulations, 2011 CFR

2011-01-01

....73 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.73 Power or... aircraft, without overtemperature, surge, stall, or other detrimental factors occurring to the engine...

Estimating Thruster Impulses From IMU and Doppler Data

NASA Technical Reports Server (NTRS)

Lisano, Michael E.; Kruizinga, Gerhard L.

2009-01-01

A computer program implements a thrust impulse measurement (TIM) filter, which processes data on changes in velocity and attitude of a spacecraft to estimate the small impulsive forces and torques exerted by the thrusters of the spacecraft reaction control system (RCS). The velocity-change data are obtained from line-of-sight-velocity data from Doppler measurements made from the Earth. The attitude-change data are the telemetered from an inertial measurement unit (IMU) aboard the spacecraft. The TIM filter estimates the threeaxis thrust vector for each RCS thruster, thereby enabling reduction of cumulative navigation error attributable to inaccurate prediction of thrust vectors. The filter has been augmented with a simple mathematical model to compensate for large temperature fluctuations in the spacecraft thruster catalyst bed in order to estimate thrust more accurately at deadbanding cold-firing levels. Also, rigorous consider-covariance estimation is applied in the TIM to account for the expected uncertainty in the moment of inertia and the location of the center of gravity of the spacecraft. The TIM filter was built with, and depends upon, a sigma-point consider-filter algorithm implemented in a Python-language computer program.

14 CFR 21.231 - Applicability.

Code of Federal Regulations, 2010 CFR

2010-01-01

... category rotorcraft; (4) Turbojet engines of not more than 1,000 pounds thrust; (5) Turbopropeller and reciprocating engines of not more than 500 brake horsepower; and (6) Propellers manufactured for use on engines covered by paragraph (a)(4) of this section; and (b) Issuing airworthiness approval tags for engines...

14 CFR 23.1143 - Engine controls.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine controls. 23.1143 Section 23.1143... Accessories § 23.1143 Engine controls. (a) There must be a separate power or thrust control for each engine... supercharger controls must be arranged to allow— (1) Separate control of each engine and each supercharger; and...

Research Technology

NASA Image and Video Library

1997-02-01

Researchers at the Marshall Space Flight Center (MSFC) have designed, fabricated, and tested the first solar thermal engine, a non-chemical rocket engine that produces lower thrust but has better thrust efficiency than a chemical combustion engine. This photograph shows components for the thermal propulsion engine being laid out prior to assembly. MSFC turned to solar thermal propulsion in the early 1990s due to its simplicity, safety, low cost, and commonality with other propulsion systems. As part of MSFC's Space Transportation Directorate, the Propulsion Research Center serves as a national resource for research of advanced, revolutionary propulsion technologies. The mission is to move the Nation's capabilities beyond the confines of conventional chemical propulsion into an era of aircraft-like access to Earth-orbit, rapid travel throughout the solar system, and exploration of interstellar space.

NASA Researcher Examines an Aircraft Model with a Four-Fan Thrust Reverser

NASA Image and Video Library

1972-03-21

National Aeronautics and Space Administration (NASA) researcher John Carpenter inspects an aircraft model with a four-fan thrust reverser which would be studied in the 9- by 15-Foot Low Speed Wind Tunnel at the Lewis Research Center. Thrust reversers were introduced in the 1950s as a means for slowing high-speed jet aircraft during landing. Engineers sought to apply the technology to Vertical and Short Takeoff and Landing (VSTOL) aircraft in the 1970s. The new designs would have to take into account shorter landing areas, noise levels, and decreased thrust levels. A balance was needed between the thrust reverser’s efficiency, its noise generation, and the engine’s power setting. This model underwent a series of four tests in the 9- by 15-foot tunnel during April and May 1974. The model, with a high-wing configuration and no tail, was equipped with four thrust-reverser engines. The investigations included static internal aerodynamic tests on a single fan/reverser, wind tunnel isolated fan/reverser thrust tests, installation effects on a four-fan airplane model in a wind tunnel, and single reverser acoustic tests. The 9-by 15 was built inside the return leg of the 8- by 6-Foot Supersonic Wind Tunnel in 1968. The facility generates airspeeds from 0 to 175 miles per hour to evaluate the aerodynamic performance and acoustic characteristics of nozzles, inlets, and propellers, and investigate hot gas re-ingestion of advanced VSTOL concepts. John Carpenter was a technician in the Wind Tunnels Service Section of the Test Installations Division.

SSME structural computer program development. Volume 3: BOPACE demonstration analysis of the SSME thrust chamber liner

NASA Technical Reports Server (NTRS)

Spring, A. H.

1973-01-01

The application of a structural computer program for analysis of a thrust chamber liner is discussed. Two objectives were accomplished as follows: (1) exercise of the full capabilities of the computer program and (2) definition of thermal and mechanical boundary conditions to reflect the emergency power level operating conditions for the SSME 47OK engine at a station just upstream of the thrust chamber throat. Creep information on the thrust chamber is presented as a reference curve of creep strain versus time for various temperatures. Contour plots of the effective plastic strain, effective stress, and effective creep strain are developed.

Preliminary flight test results of a fly-by-throttle emergency flight control system on an F-15 airplane

NASA Technical Reports Server (NTRS)

Burcham, Frank W., Jr.; Maine, Trindel A.; Fullerton, C. G.; Wells, Edward A.

1993-01-01

A multi-engine aircraft, with some or all of the flight control system inoperative, may use engine thrust for control. NASA Dryden has conducted a study of the capability and techniques for this emergency flight control method for the F-15 airplane. With an augmented control system, engine thrust, along with appropriate feedback parameters, is used to control flightpath and bank angle. Extensive simulation studies have been followed by flight tests. This paper discusses the principles of throttles-only control, the F-15 airplane, the augmented system, and the flight results including landing approaches with throttles-only control to within 10 ft of the ground.

Effect of a rough surface on the aerodynamic characteristics of a two-bladed wind-powered engine with cylindrical blades

NASA Astrophysics Data System (ADS)

Tanasheva, N. K.; Kunakbaev, T. O.; Dyusembaeva, A. N.; Shuyushbayeva, N. N.; Damekova, S. K.

2017-11-01

We have reported the results of experiments on determining the drag coefficient and the thrust coefficient of a two-bladed wind-powered engine based on the Magnus effect with rotating rough cylinders in the range of air flow velocity of 4-10 m/s (Re = 26800-90000) for a constant rotation number of a cylindrical blade about its own axis. The results show that an increase in the Reynolds number reduces the drag coefficient and the thrust coefficient. The extent of the influence of the relative roughness on the aerodynamic characteristics of the two-bladed wind-powered engine has been experimentally established.

Injection Principles from Combustion Studies in a 200-Pound-Thrust Rocket Engine Using Liquid Oxygen and Heptane

NASA Technical Reports Server (NTRS)

Heidmann, M. F.; Auble, C. M.

1955-01-01

The importance of atomizing and mixing liquid oxygen and heptane was studied in a 200-pound-thrust rocket engine. Ten injector elements were used with both steel and transparent chambers. Characteristic velocity was measured over a range of mixture ratios. Combustion gas-flow and luminosity patterns within the chamber were obtained by photographic methods. The results show that, for efficient combustion, the propellants should be both atomized and mixed. Heptane atomization controlled the combustion rate to a much larger extent than oxygen atomization. Induced mixing, however, was required to complete combustion in the smallest volume. For stable, high-efficiency combustion and smooth engine starts, mixing after atomization was most promising.

Divergence thrust loss calculations for convergent-divergent nozzles: Extensions to the classical case

NASA Technical Reports Server (NTRS)

Berton, Jeffrey J.

1991-01-01

The analytical derivations of the non-axial thrust divergence losses for convergent-divergent nozzles are described as well as how these calculations are embodied in the Navy/NASA engine computer program. The convergent-divergent geometries considered are simple classic axisymmetric nozzles, two dimensional rectangular nozzles, and axisymmetric and two dimensional plug nozzles. A simple, traditional, inviscid mathematical approach is used to deduce the influence of the ineffectual non-axial thrust as a function of the nozzle exit divergence angle.

Uprating the Frontal Thrust of a Spherical Gas-Dynamical Resonator-Pulse Amplifier

NASA Astrophysics Data System (ADS)

Bogdanov, V. I.; Khantalin, D. S.

2017-01-01

Calculations were carried out with application of current numerical methods and with the use of scientific-technical developments of a gas-dynamical resonator-thrust amplifier. The possibility of creating an exit device with a resonator for a small-size gas-turbine engine that in flight provides for thrust uprating by no less than 6% via gas mass attachment in the pulsating process is shown. In this way the size-mass characteristics of the exit device are preserved in the process.

VSCE technology definition study

NASA Technical Reports Server (NTRS)

Howlett, R. A.; Hunt, R. B.

1979-01-01

Refined design definition of the variable stream control engine (VSCE) concept for advanced supersonic transports is presented. Operating and performance features of the VSCE are discussed, including the engine components, thrust specific fuel consumption, weight, noise, and emission system. A preliminary engine design is presented.

Final Rule for Control of Air Pollution From Aircraft and Aircraft Engines; Emission Standards and Test Procedures

EPA Pesticide Factsheets

EPA adopted emission standards and related provisions for aircraft gas turbine engines with rated thrusts greater than 26.7 kilonewtons. These engines are used primarily on commercial passenger and freight aircraft.

77 FR 63268 - Airworthiness Directives; Airbus Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-10-16

... link bolts, center sway link bolts, and thrust link bolts on the forward engine mounts. This proposed... forward engine mounts. This condition, if left uncorrected, could result in a deterioration of the structural integrity of the front engine mount bolts [and possible damage to an engine or wing]. For the...

14 CFR 33.66 - Bleed air system.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Bleed air system. 33.66 Section 33.66... STANDARDS: AIRCRAFT ENGINES Design and Construction; Turbine Aircraft Engines § 33.66 Bleed air system. The engine must supply bleed air without adverse effect on the engine, excluding reduced thrust or power...

14 CFR 23.1143 - Engine controls.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Engine controls. 23.1143 Section 23.1143... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1143 Engine controls. (a) There must be a separate power or thrust control for each engine...

14 CFR 23.1143 - Engine controls.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Engine controls. 23.1143 Section 23.1143... STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Powerplant Powerplant Controls and Accessories § 23.1143 Engine controls. (a) There must be a separate power or thrust control for each engine...

The application of parameter estimation to flight measurements to obtain lateral-directional stability derivatives of an augmented jet-flap STOL airplane

NASA Technical Reports Server (NTRS)

Stephenson, J. D.

1983-01-01

Flight experiments with an augmented jet flap STOL aircraft provided data from which the lateral directional stability and control derivatives were calculated by applying a linear regression parameter estimation procedure. The tests, which were conducted with the jet flaps set at a 65 deg deflection, covered a large range of angles of attack and engine power settings. The effect of changing the angle of the jet thrust vector was also investigated. Test results are compared with stability derivatives that had been predicted. The roll damping derived from the tests was significantly larger than had been predicted, whereas the other derivatives were generally in agreement with the predictions. Results obtained using a maximum likelihood estimation procedure are compared with those from the linear regression solutions.

Surface rupture of the 2002 Denali fault, Alaska, earthquake and comparison with other strike-slip ruptures

USGS Publications Warehouse

Haeussler, Peter J.; Schwartz, D.P.; Dawson, T.E.; Stenner, Heidi D.; Lienkaemper, J.J.; Cinti, F.; Montone, Paola; Sherrod, B.; Craw, P.

2004-01-01

On 3 November 2002, an M7.9 earthquake produced 340 km of surface rupture on the Denali and two related faults in Alaska. The rupture proceeded from west to east and began with a 40-km-long break on a previously unknown thrust fault. Estimates of surface slip on this thrust are 3-6 m. Next came the principal surface break along ???218 km of the Denali fault. Right-lateral offsets averaged around 5 m and increased eastward to a maximum of nearly 9 m. The fault also ruptured beneath the trans-Alaska oil pipeline, which withstood almost 6 m of lateral offset. Finally, slip turned southeastward onto the Totschunda fault. Right-lateral offsets are up to 3 m, and the surface rupture is about 76 km long. This three-part rupture ranks among the longest strike-slip events of the past two centuries. The earthquake is typical when compared to other large earthquakes on major intracontinental strike-slip faults. ?? 2004, Earthquake Engineering Research Institute.

Thrust Measurements for a Pulse Detonation Engine Driven Ejector

NASA Technical Reports Server (NTRS)

Santoro, Robert J.; Pak, Sibtosh; Shehadeh, R.; Saretto, S. R.; Lee, S.-Y.

2005-01-01

Results of an experimental effort on pulse detonation driven ejectors aimed at probing different aspects of PDE ejector processes, are presented and discussed. The PDE was operated using ethylene as the fuel and an equimolar oxygen/nitrogen mixture as the oxidizer at an equivalence ratio of one. The thrust measurements for the PDE alone are in excellent agreement with experimental and modeling results reported in the literature and serve as a Baseline for the ejector studies. These thrust measurements were then used as a basis for quantifying thrust augmentation for various PDE/ejector setups using constant diameter ejector tubes and various detonation tube/ejector tube overlap distances. The results show that for the geometries studied here, a maximum thrust augmentation of 24% is achieved. The thrust augmentation results are complemented by shadowgraph imaging of the flowfield in the ejector tube inlet area and high frequency pressure transducer measurements along the length of the ejector tube.

Design and Optimization of Low-thrust Orbit Transfers Using Q-law and Evolutionary Algorithms

NASA Technical Reports Server (NTRS)

Lee, Seungwon; vonAllmen, Paul; Fink, Wolfgang; Petropoulos, Anastassios; Terrile, Richard

2005-01-01

Future space missions will depend more on low-thrust propulsion (such as ion engines) thanks to its high specific impulse. Yet, the design of low-thrust trajectories is complex and challenging. Third-body perturbations often dominate the thrust, and a significant change to the orbit requires a long duration of thrust. In order to guide the early design phases, we have developed an efficient and efficacious method to obtain approximate propellant and flight-time requirements (i.e., the Pareto front) for orbit transfers. A search for the Pareto-optimal trajectories is done in two levels: optimal thrust angles and locations are determined by Q-law, while the Q-law is optimized with two evolutionary algorithms: a genetic algorithm and a simulated-annealing-related algorithm. The examples considered are several types of orbit transfers around the Earth and the asteroid Vesta.

Performance characteristics of a wedge nozzle installed on an F-18 propulsion wind tunnel model

NASA Technical Reports Server (NTRS)

Petit, J. E.; Capone, F. J.

1979-01-01

The results of two-dimensional wedge non-axisymmetric nozzle (2D-AIN) tests to determine its performance relative to the baseline axisymmetric nozzle using an F-18 jet effects wind tunnel model are presented. Configurations and test conditions simulated forward thrust-minus drag, thrust vectoring/induced lift, and thrust reversing flight conditions from Mach .6 to 1.20 and attack angles up to 10 degrees. Results of the model test program indicate that non-axisymmetric nozzles can be installed on a twin engine fighter aircraft model with equivalent thrust minus drag performance as the baseline axisymmetric nozzles. Thrust vectoring capability of the non-axisymmetric nozzles provided significant jet-induced lift on the nozzle/aftbody and horizontal tail surfaces. Thrust reversing panels deployed from the 2D-AIN centerbody wedge were very effective for static and inflight operation

Fuel Burn Estimation Using Real Track Data

NASA Technical Reports Server (NTRS)

Chatterji, Gano B.

2011-01-01

A procedure for estimating fuel burned based on actual flight track data, and drag and fuel-flow models is described. The procedure consists of estimating aircraft and wind states, lift, drag and thrust. Fuel-flow for jet aircraft is determined in terms of thrust, true airspeed and altitude as prescribed by the Base of Aircraft Data fuel-flow model. This paper provides a theoretical foundation for computing fuel-flow with most of the information derived from actual flight data. The procedure does not require an explicit model of thrust and calibrated airspeed/Mach profile which are typically needed for trajectory synthesis. To validate the fuel computation method, flight test data provided by the Federal Aviation Administration were processed. Results from this method show that fuel consumed can be estimated within 1% of the actual fuel consumed in the flight test. Next, fuel consumption was estimated with simplified lift and thrust models. Results show negligible difference with respect to the full model without simplifications. An iterative takeoff weight estimation procedure is described for estimating fuel consumption, when takeoff weight is unavailable, and for establishing fuel consumption uncertainty bounds. Finally, the suitability of using radar-based position information for fuel estimation is examined. It is shown that fuel usage could be estimated within 5.4% of the actual value using positions reported in the Airline Situation Display to Industry data with simplified models and iterative takeoff weight computation.

Aerospike thrust chamber program. [cumulative damage and maintenance of structural members in hydrogen oxygen engines

NASA Technical Reports Server (NTRS)

Campbell, J., Jr.; Cobb, S. M.

1976-01-01

An existing, but damaged, 25,000-pound thrust, flightweight, oxygen/hydrogen aerospike rocket thrust chamber was disassembled and partially repaired. A description is presented of the aerospike chamber configuration and of the damage it had suffered. Techniques for aerospike thrust chamber repair were developed, and are described, covering repair procedures for lightweight tubular nozzles, titanium thrust structures, and copper channel combustors. Effort was terminated prior to completion of the repairs and conduct of a planned hot fire test program when it was found that the copper alloy walls of many of the thrust chamber's 24 combustors had been degraded in strength and ductility during the initial fabrication of the thrust chamber. The degradation is discussed and traced to a reaction between oxygen and/or oxides diffused into the copper alloy during fabrication processes and the hydrogen utilized as a brazing furnace atmosphere during the initial assembly operation on many of the combustors. The effects of the H2/O2 reaction within the copper alloy are described.

Experience with high performance V/STOL fighter projects at MBB

NASA Technical Reports Server (NTRS)

Aulehla, F.; Kissel, G. K.

1981-01-01

Flight control systems and aerodynamic aspects of experimental V/STOL aircraft are discussed. The VJ 101 C featured tilting engines for increased thrust, reheat for takeoff, simple translation, triangular decentralization of the engines for thrust modulation, and moderate ground effects. Two experimental aircraft were built, with and without reheat, capable of Mach 2 and Mach 1.04, respectively. The mechanical flight control system and tests are outlined, both for hover rig and flight configurations. Ground suction, acoustic and thermal loading, sodium silicate coatings to avoid ground corrosion, and recirculation are considered. Results of the follow-on project to the VJ 101 C, the AVS, which was developed by NASA, are reviewed, and it is noted that trends toward thrust-to-weight ratios exceeding one, in concert with low wing loading, favor the development of V/STOL aircraft.

Development and Validation of an NPSS Model of a Small Turbojet Engine

NASA Astrophysics Data System (ADS)

Vannoy, Stephen Michael

Recent studies have shown that integrated gas turbine engine (GT)/solid oxide fuel cell (SOFC) systems for combined propulsion and power on aircraft offer a promising method for more efficient onboard electrical power generation. However, it appears that nobody has actually attempted to construct a hybrid GT/SOFC prototype for combined propulsion and electrical power generation. This thesis contributes to this ambition by developing an experimentally validated thermodynamic model of a small gas turbine (˜230 N thrust) platform for a bench-scale GT/SOFC system. The thermodynamic model is implemented in a NASA-developed software environment called Numerical Propulsion System Simulation (NPSS). An indoor test facility was constructed to measure the engine's performance parameters: thrust, air flow rate, fuel flow rate, engine speed (RPM), and all axial stage stagnation temperatures and pressures. The NPSS model predictions are compared to the measured performance parameters for steady state engine operation.

NASA's Evolutionary Xenon Thruster (NEXT) Prototype Model 1R (PM1R) Ion Thruster and Propellant Management System Wear Test Results

NASA Technical Reports Server (NTRS)

VanNoord, Jonathan L.; Soulas, George C.; Sovey, James S.

2010-01-01

The results of the NEXT wear test are presented. This test was conducted with a 36-cm ion engine (designated PM1R) and an engineering model propellant management system. The thruster operated with beam extraction for a total of 1680 hr and processed 30.5 kg of xenon during the wear test, which included performance testing and some operation with an engineering model power processing unit. A total of 1312 hr was accumulated at full power, 277 hr at low power, and the remainder was at intermediate throttle levels. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The propellant management system performed without incident during the wear test. The ion engine and propellant management system were also inspected following the test with no indication of anomalous hardware degradation from operation.

Control Design for a Generic Commercial Aircraft Engine

NASA Technical Reports Server (NTRS)

Csank, Jeffrey; May, Ryan D.

2010-01-01

This paper describes the control algorithms and control design process for a generic commercial aircraft engine simulation of a 40,000 lb thrust class, two spool, high bypass ratio turbofan engine. The aircraft engine is a complex nonlinear system designed to operate over an extreme range of environmental conditions, at temperatures from approximately -60 to 120+ F, and at altitudes from below sea level to 40,000 ft, posing multiple control design constraints. The objective of this paper is to provide the reader an overview of the control design process, design considerations, and justifications as to why the particular architecture and limits have been chosen. The controller architecture contains a gain-scheduled Proportional Integral controller along with logic to protect the aircraft engine from exceeding any limits. Simulation results illustrate that the closed loop system meets the Federal Aviation Administration s thrust response requirements