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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Powered Descent Guidance with General Thrust-Pointing Constraints

The Powered Descent Guidance (PDG) algorithm and software for generating Mars pinpoint or precision landing guidance profiles has been enhanced to incorporate thrust-pointing constraints. Pointing constraints would typically be needed for onboard sensor and navigation systems that have specific field-of-view requirements to generate valid ground proximity and terrain-relative state measurements. The original PDG algorithm was designed to enforce both control and state constraints, including maximum and minimum thrust bounds, avoidance of the ground or descent within a glide slope cone, and maximum speed limits. The thrust-bound and thrust-pointing constraints within PDG are non-convex, which in general requires nonlinear optimization methods to generate solutions. The short duration of Mars powered descent requires guaranteed PDG convergence to a solution within a finite time; however, nonlinear optimization methods have no guarantees of convergence to the global optimal or convergence within finite computation time. A lossless convexification developed for the original PDG algorithm relaxed the non-convex thrust bound constraints. This relaxation was theoretically proven to provide valid and optimal solutions for the original, non-convex problem within a convex framework. As with the thrust bound constraint, a relaxation of the thrust-pointing constraint also provides a lossless convexification that ensures the enhanced relaxed PDG algorithm remains convex and retains validity for the original nonconvex problem. The enhanced PDG algorithm provides guidance profiles for pinpoint and precision landing that minimize fuel usage, minimize landing error to the target, and ensure satisfaction of all position and control constraints, including thrust bounds and now thrust-pointing constraints.

Carson, John M., III↗
NASA NTRS2021

Expected Thrust Fraction: Resilient Trajectory Design Applied to the Earth Return Orbiter

Electric propulsion is an enabling technology for NASA’s Mars Sample Return mission that affords an order of magnitude reduction in fuel. Unfortunately, elec- tric propulsion’s low-thrust capabilities require long thrusting arcs that make space- craft susceptible to disturbances resulting in missed thrust events. One method for designing trajectories resilient to such missed thrust events is expected thrust frac- tion, which embeds their stochastic nature into a time-varying duty cycle. This paper investigates the application of expected thrust fraction to the Earth Return Orbiter’s outbound Earth-to-Mars trajectory and compares it with trajectories de- signed using traditional methods to mitigate missed thrust events. Through the use of expected thrust fraction, a trajectory with a baseline 65.4% success rate is im- proved to an 86.6% success rate at the cost of only a 12 kg reduction in delivered mass. Additionally, when expected thrust fraction is used in conjunction with a terminal coast, a 96.0% success rate is achieved, which is higher than any single mitigation technique.

Laipert, Frank↗
NASA NTRS1962

DEVELOPMENT OF A LIGHT-WEIGHT REGENERATIVELY COOLED THRUST CHAMBER WITH BRAZED RIBS AND BRAZED OUTER WIRE TRAP

Design and fabrication procedures are discussed of a brazed-ribtype thrust chamber fabricated from stainless steel or inconel for a Sooo-lb thrust 20: 1 expansion-area-ratio regeneratively cooled liquidpropellant rocket engine which is designated as the JPL 6K Engine and operates at a chamber pressure of 150 psia. Hydrazine is used as the fuel and regenerative coolant and nitrogen tetroxide as the oxidizer. The nozzle-throat diameter is 5.42 in. The thrust chamber consists of a sheet-metal gas-side liner with longitudinal ribs spotwelded and brazed to the outer surface. Wire is wrapped over the ribs circumferentially, spot-welded to the ribs, and brazed. The liner, ribs, and wire form the coolant-passage walls. There are three basic advantages to this type of thrust chamber: light weight, ease of fabrication, and comparatively low cost. The total dry weight of a thrust chamber of this type and size made of 347 stainless steel is 46.2 lb, and the wet weight is 56.3 Ib. A total of 87 static engine tests were made on 19 of 23 thrust chambers fabricated by this technique. Four of the chambers tested had an expansion-area ratio of 20:1, and 15, which were shorter but had the same throat diameter, had an expansion-area ratio of 3.5:l. Accumulated test time was 1926 sec, including 558 sec regenerative cooling with hydrazine and 1368 sec cooling with water. One 20:l thrust chamber was subjected to 16 separate engine tests, with an accumulated test time of 399 sec, and was still in very good operating condition after the tests were completed. With water flowing through the coolant passages at 10 lb/sec, the measured nominal pressure drop between coolant manifolds of a 20:l thrust chamber wrapped with flat wire was 17.2 psi.

Thrust Chamber↗
NASA NTRS1975

Direct thrust measurement of a 30-cm ion thruster

A direct thrust measurement of a 30-cm diameter ion thruster was accomplished by means of a laser interferometer thrust stand. The thruster was supported in a pendulum manner by three 3.65-m long wires. Electrical power was provided by means of 18 mercury filled pots. A movable 23-button planar probe rake was used to determine thrust loss due to ion beam divergence. Values of thrust, thrust loss due to ion beam divergence, and thrust loss due to multiple ionization were measured for ion beam currents ranging from 0.5 A to 2.5 A. Measured thrust values indicate an accuracy of approximately 1% and are in good agreement with thrust values calculated by indirect measurements.

Banks, B.↗
NASA NTRS1975

Direct thrust measurement of a 30-cm ion thruster

A direct thrust measurement of a 30-cm diameter ion thruster was accomplished by means of a laser interferometer thrust stand. The thruster was supported in a pendulum manner by three 3.65-m long wires. Electrical power was provided by means of 18 mercury filled pots. A movable 23-button planar probe rake was used to determine thrust loss due to ion beam divergence. Values of thrust, thrust loss due to ion beam divergence, and thrust loss due to multiple ionization were measured for ion beam currents ranging from 0.5 to 2.5 A. Measured thrust values indicate an accuracy of approximately 1% and are in good agreement with thrust values calculated by indirect measurements.

Banks, B.↗
NASA NTRS1989

Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices has been designed and tested. The thrust stand was specifically tailored to the needs of a 0.1 to 0.25 MW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. The paper gives a detailed description of the thrust stand design and operation with a 0.1 MW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability.

Haag, Thomas W.↗
NASA NTRS1989

Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices was designed and tested. The thrust stand was specifically tailored to the needs of a 100 to 250 kW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. A detailed description is given for the thrust stand design and operation with a 100 kW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability.

Haag, Thomas W.↗
NASA NTRS1994

Static performance investigation of a skewed-throat multiaxis thrust-vectoring nozzle concept

The static performance of a jet exhaust nozzle which achieves multiaxis thrust vectoring by physically skewing the geometric throat has been characterized in the static test facility of the 16-Foot Transonic Tunnel at NASA Langley Research Center. The nozzle has an asymmetric internal geometry defined by four surfaces: a convergent-divergent upper surface with its ridge perpendicular to the nozzle centerline, a convergent-divergent lower surface with its ridge skewed relative to the nozzle centerline, an outwardly deflected sidewall, and a straight sidewall. The primary goal of the concept is to provide efficient yaw thrust vectoring by forcing the sonic plane (nozzle throat) to form at a yaw angle defined by the skewed ridge of the lower surface contour. A secondary goal is to provide multiaxis thrust vectoring by combining the skewed-throat yaw-vectoring concept with upper and lower pitch flap deflections. The geometric parameters varied in this investigation included lower surface ridge skew angle, nozzle expansion ratio (divergence angle), aspect ratio, pitch flap deflection angle, and sidewall deflection angle. Nozzle pressure ratio was varied from 2 to a high of 11.5 for some configurations. The results of the investigation indicate that efficient, substantial multiaxis thrust vectoring was achieved by the skewed-throat nozzle concept. However, certain control surface deflections destabilized the internal flow field, which resulted in substantial shifts in the position and orientation of the sonic plane and had an adverse effect on thrust-vectoring and weight flow characteristics. By increasing the expansion ratio, the location of the sonic plane was stabilized. The asymmetric design resulted in interdependent pitch and yaw thrust vectoring as well as nonzero thrust-vector angles with undeflected control surfaces. By skewing the ridges of both the upper and lower surface contours, the interdependency between pitch and yaw thrust vectoring may be eliminated and the location of the sonic plane may be further stabilized.

Wing, David J.↗
NASA NTRS1995

Maximum thrust mode evaluation

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.

Orme, John S.↗
NASA NTRS1995

PPT Thrust Stand

A torsional-type thrust stand has been designed and built to test Pulsed Plasma Thrusters (PPT's) in both single shot and repetitive operating modes. Using this stand, momentum per pulse was determined strictly as a function of thrust stand deflection, spring constant, and natural frequency. No empirical corrections were required. The accuracy of the method was verified using a swinging impact pendulum. Momentum transfer data between the thrust stand and the pendulum were consistent to within 1%. Following initial calibrations, the stand was used to test a Lincoln Experimental Satellite (LES-8/9) thruster. The LES-8/9 system had a mass of approximately 7.5 kg, with a nominal thrust to weight ratio of 1.3 x 10(exp -5). A total of 34 single shot thruster pulses were individually measured. The average impulse bit per pulse was 266 microN-s, which was slightly less than the value of 300 microN-s published in previous reports on this device. Repetitive pulse measurements were performed similar to ordinary steady-state thrust measurements. The thruster was operated for 30 minutes at a repetition rate of 132 pulses per minute and yielded an average thrust of 573 microN. Using average thrust, the average impulse bit per pulse was estimated to be 260 microN-s, which was in agreement with the single shot data. Zero drift during the repetitive pulse test was found to be approximately 1% of the measured thrust.

Haag, Thomas W.↗
NASA NTRS2004

An Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold: first, to add to the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. Second, to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and predicting what thrust augmentation values may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (having no mean component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the various measurements made. These include both thrust and Digital Particle Imaging Velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called Formation Number serves as a useful dimensionless number by which to characterize the jet and predict performance.

Paxson, Daniel E.↗
NASA NTRS2007

Experimental Investigation of Unsteady Thrust Augmentation Using a Speaker-Driven Jet

An experimental investigation is described in which a simple speaker-driven jet was used as a pulsed thrust source (driver) for an ejector configuration. The objectives of the investigation were twofold. The first was to expand the experimental body of evidence showing that an unsteady thrust source, combined with a properly sized ejector generally yields higher thrust augmentation values than a similarly sized, steady driver of equivalent thrust. The second objective was to identify characteristics of the unsteady driver that may be useful for sizing ejectors, and for predicting the thrust augmentation levels that may be achieved. The speaker-driven jet provided a convenient source for the investigation because it is entirely unsteady (i.e., it has no mean velocity component) and because relevant parameters such as frequency, time-averaged thrust, and diameter are easily variable. The experimental setup will be described, as will the two main measurements techniques employed. These are thrust and digital particle imaging velocimetry of the driver. It will be shown that thrust augmentation values as high as 1.8 were obtained, that the diameter of the best ejector scaled with the dimensions of the emitted vortex, and that the so-called formation time serves as a useful dimensionless parameter by which to characterize the jet and predict performance.

Paxson, Daniel E.↗
NASA NTRS2018

Uncertainty in Inverted Pendulum Thrust Measurements

An uncertainty analysis of a common configuration of electric propulsion thrust stand is presented. The analysis applies to inverted pendulum thrust stands operating in a null-coil configuration with in-situ calibration. Several sources of bias and precision uncertainty are discussed, propagated, and combined to form conservative estimates of the relative and absolute thrust uncertainties. A case study of the NASA Glenn Research Center Vacuum Facility 6 thrust stand is presented. For the thruster investigated, the uncertainty was estimated to be +/-6.9 mN over the entire span of thrust. This uncertainty represents a maximum instrument bias introduced by the thrust stand. The paper does not address repeatability of actual thrust measurements, as this is generally beyond the influence of the thrust stand and can be dependent on a large number of factors.

Mackey, J.↗
NASA NTRS2018

Uncertainty in Inverted Pendulum Thrust Measurements

An uncertainty analysis of a common configuration of electric propulsion thrust stand is presented. The analysis applies to inverted pendulum thrust stands operating in a null-coil configuration with in-situ calibration. Several sources of bias and precision uncertainty are discussed, propagated, and combined to form conservative estimates of the relative and absolute thrust uncertainties. A case study of the NASA Glenn Research Center Vacuum Facility 6 thrust stand is presented. For the thruster investigated, the uncertainty was estimated to be ±6.9 mN over the entire span of thrust. This uncertainty represents a maximum instrument bias introduced by the thrust stand. The paper does not address repeatability of actual thrust measurements, as this is generally beyond the influence of the thrust stand and can be dependent on a large number of factors.

Mackey, Jonathan A.↗
NASA NTRS2018

Uncertainty in Inverted Pendulum Thrust Measurements

An uncertainty analysis of a common configuration of electric propulsion thrust stand is presented. The analysis applies to inverted pendulum thrust stands operating in a null-coil configuration with in-situ calibration. Several sources of bias and precision uncertainty are discussed, propagated, and combined to form conservative estimates of the relative and absolute thrust uncertainties. A case study of the NASA Glenn Research Center Vacuum Facility 6 thrust stand is presented. For the thruster investigated, the uncertainty was estimated to be ±6.9mN over the entire span of thrust. This uncertainty represents a maximum instrument bias introduced by the thrust stand. The paper does not address repeatability of actual thrust measurements, as this is generally beyond the influence of the thrust stand and can be dependent on a large number of factors.

Mackey, Jonathan A.↗
NASA NTRS2020

Using Reinforcement Learning to Design Missed Thrust Resilient Trajectories

From ion thrusters to solar sails, spacecraft continue to adopt new and more efficient forms of propulsion. As these low-thrust propulsion meth- ods have become more prevalent, new challenges have arisen. Depending on the mission, low-thrust propulsion elements may need to thrust con- tinuously for days/months. During these thrusting periods, external fac- tors, such as a micro-meteoroid impact or a software glitch, may cause the spacecraft to prematurely cease its thrust stage. Half of all deep space missions enter a safe mode where they cannot thrust every four months. These missed thrust events can result in the complete loss of a space- craft for time-dependent trajectories like planetary rendezvous. This paper demonstrates how neural networks, trained using reinforcement learning, can autonomously correct for missed thrust events during an interplanetary trajectory.

Laipert, Frank E.↗
DOE OSTI2025

Carbon Storage in Fold‐and‐Thrust Belts: An Overlooked Gigatonne Storage Opportunity

This study presents numerical investigations of the trapping characteristics of fold-and-thrust belt structures, defining three carbon capture and storage (CCS) play types that could be used to store commercial volumes (millions of tonnes) of CO 2 . Specifically, we present simulations of CO 2 storage in three fold-and-thrust belt models comprising a thrust-ramp, duplex, and thrust-fold geometry. To constrain these play types in realistic geology, each model is based on a study site, including a novel investigation of a greenfield saline reservoir in Virginia, USA, being considered for commercial carbon storage and two well-characterized petroleum fields: the Wilburton field in Oklahoma, USA, and the Incahuasi field in Bolivia. Our results provide insight into several key parameters, such as the long-term security of injected CO 2 in these geologies and injection strategies for maximizing storage efficiency while reducing pressure-related risk. These results improve the understanding of CCS in fold-and-thrust belt storage sites globally by describing general storage parameters that may be applied to site-specific projects. We find that thrust-ramp geometries may securely trap CO 2 through solubility and hydrodynamic trapping under suitable reservoir conditions, duplex structures may store some quantities of CO 2 but are pressure-constrained, and that thrust-ramp structures may store large quantities of CO 2 by maximizing fetch volume, which simultaneously lowers geomechanical risk by reducing pressure buildup along zones of weakness.

42 ENGINEERING↗
NASA NTRS1977

Thrust-augmented vortex attenuation

An experimental investigation was conducted to determine the vortex attenuating effect of engine thrust. Tests were made using a 0.03-scale model of the Boeing 747 transport aircraft as a vortex generating model. A Learjet-class probe model was used to measure the vortex induced rolling moment at a scale separation distance of 1.63 km. These tests were conducted at a lift coefficient of 1.4 at a model velocity of 30.48 m/s. The data presented indicate that engine thrust is effective as a vortex attenuating device when the engines are operated at high thrust levels and are positioned to direct the high energy engine wake into the core of the vortex. The greatest thrust vortex attenuation was obtained by operating the inboard engine thrust reversers at one-quarter thrust and the outboard engines at maximum forward thrust.

Patterson, J. C., Jr.↗