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NASA NTRS1979

Two-Way Time Transfer via Geostationary Satellites NRC/NBS, NRC/USNO and NBS/USNO via Hermes and NRC/LPTF (France) via Symphonie

In most of the experiments, 1 pps pulses of the station atomic clocks were exchanged between the partners, and a cubic equation was fitted to the 1000 to 2000 second measurements. The equations were exchanged and substracted to obtain the time difference of the stations. The standard deviation in the fit of the equations varied, depending on conditions, from 1.5 ns to 16 ns. For the last month of the Hermes experiment, a 1 MHz signal was used, giving a standard deviation of 0.18 ns. The comparison of the time scales via satellite and via Loran-C (BIH Circular D) show clearly that some Loran-C links are very good, but that the NBS link varies by 1 micron s. Via the satellite the frequencies of the time scales can be compared with an accuracy of 2 x 10 to the minus 14 power.

Costain, C.↗

NASA NTRS1987

Hypersonic aerodynamic coefficients and convection flow estimation in a HERMES type project

An analysis of all the experimental and computation requirements of the HERMES project is presented. The discussion includes the influence of the upper atmosphere composition and possible chemical reactions, the information collected in Orbiter experiments, the possibility of utilizing plasma generators and CF4 in wind tunnel tests, the development of parabolized Navier-Stokes models and solutions, and the computing capacity requirements.

Pollak, C.↗

NASA NTRS1992

A multipurpose model of Hermes-Columbus docking mechanism

One of the foreseen missions of the HERMES spacevehicle is the servicing to the Columbus Free Flying Laboratory (MTFF). Docking between the two spacecraft is a critical operation in which the Docking Mechanism (DM) has a major role. In order to analyze and assess robustness of initially selected concepts and to identify suitable implementation solutions, through the investigation of main parameters involved in the docking functions, a multipurpose model of DM was developed and tested. This paper describes the main design features as well as the process of calibrating and testing.

Gonzalez-Vallejo, J. J.↗

NASA NTRS1992

Hermes CX-7: Air transport system design simulation

The Hermes CX-7 has been designed to service the overnight parcel package delivery needs of the cities of Aeroworld as determined in the G-Dome Enterprises market survey. The design optimization centers on the prime goal of servicing the needs of these cities as efficiently and profitably as possible. The greatest factors which affect the design of an aircraft for the mission outlined in the Request for Proposal are cost, construction feasibility and effectiveness of the design. Other influencing factors are given by the constraints of the market, including a maximum takeoff and landing distance of 60 feet, storage capability in a container of size 5 ft. x 3 ft. x 2 ft., cargo packages of 2 inch and 4 inch cubes, and ability to turn with a radius no larger than 60 feet. Safety considerations such as flying at or below Mach one (30 ft/s) and controllability and maintainability of the aircraft must also be designed into the aircraft. Another influential factor is the efficiency of the aircraft which involves optimizations and tradeoffs of such factors as weight, lifting surface sizing, structural redundancy, and material costs.

Amer, Brian↗

NASA NTRS1993

Mission activities planning for a Hermes mission by means of AI-technology

Mission Activities Planning is a complex task to be performed by mission control centers. AI technology can offer attractive solutions to the planning problem. This paper presents the use of a new AI-based Mission Planning System for crew activity planning. Based on a HERMES servicing mission to the COLUMBUS Man Tended Free Flyer (MTFF) with complex time and resource constraints, approximately 2000 activities with 50 different resources have been generated, processed, and planned with parametric variation of operationally sensitive parameters. The architecture, as well as the performance of the mission planning system, is discussed. An outlook to future planning scenarios, the requirements, and how a system like MARS can fulfill those requirements is given.

Pape, U.↗

NASA NTRS1955

Longitudinal Aerodynamic Characteristics and Effect of Rocket Jet on Drag of Models of the Hermes A-3A and A-3B Missiles in Free Flight at Mach Numbers From 0.6 to 2.0

A free-flight investigation over a Mach number range from 0.6 to 2.0 has been conducted to determine the longitudinal aerodynamic characteristics and effect of rocket jet on zero-lift drag of 1/5-scale models of two ballistic-type missiles, the Hermes A-3A and A-3B. Models of both types of missiles exhibited very nearly linear normal forces and pitching moments over the angle-of-attack range of 8 deg to -4 deg and Mach number range tested. The centers of pressure for both missiles were not appreciably affected by Mach number over the subsonic range; however, between a Mach number of 1.02 and 1.50 the center of pressure for the A-3A model moved forward 0.34 caliber with increasing Mach number. At a trim angle-of-attack of approximately 30 deg, the A-3A model indicated a total drag coefficient 30% higher than the power-off zero-lift drag over the subsonic Mach number range and 10% higher over the supersonic range. Under the conditions of the present test, and excluding the effect of the jet on base drag, there was no indicated effect of the propulsive jet on the total drag of the A-3A model. The propulsive jet operating at a jet pressure ratio p(sub j)/p(sub o) of 0.8 caused approximately 100% increase in base drag over the Mach number range M = 0.6 to 1.0. This increase in base drag amounts to 15% of the total drag. An underexpanded jet operating at jet pressure ratios corresponding approximately to those of the full-scale missile caused a 22% reduction in base drag at M = 1.55 (p(sub j)/p(sub o) = 1.76) but indicated no change at M = 1.30 (p(sub j)/p(sub o) = 1.43). At M = 1.1 and p(sub j)/p(sub o) = 1.55, the jet caused a 50% increase in base drag.

Jackson, H. Herbert↗

NASA NTRS1954

Effect of Rocket-Motor Operation on the Drag of Three 1/5-Scale Hermes A-3A Models in Free Flight

Three 1/5-scale models of the Hermes A-3A missile have been flown to determine the effect of rocket-motor operation on the drag corresponding to various altitude and Mach number combinations. The flights covered a Mach number range from 0.5 to 1.8, and ratios of jet-exit static pressure to free-stream static pressure from 0.8 to 1.8. The results indicate that the power-on drag of the missile should be the same as the power-off drag at Mach number 1.3 and slightly less than the power-off drag at Mach number 1.55.

Jackson, H. Herbert↗

NASA NTRS1950

Tests of a Hermes A-2 Missile Body at Mach Number 4.04

Force tests on a proposed body shape of the Hermes A-2 missile with and without longitudinal spoilers were made at Mach number 4.04. Values of normal force coefficient, pitching-moment coefficient, and center-of-pressure position were obtained.

Ulmann, Edward F.↗

NASA NTRS1952

Wind-Tunnel Investigation of the Static Longitudinal Stability Characteristics of a 0.15-Scale Model of the Hermes A-1E2 Missile at High Subsonic Mach Numbers

The static longitudinal stability characteristics of a 0.15-scale model of the Hermes A-lE2 missile have been determined in the Langley high-speed 7- by 10-foot tunnel over a Mach number range of 0.50 to 0.98, corresponding to Reynolds numbers, based on body length, of 12.3 x 10(exp 6) to 17.1 x 10(exp 6). This paper presents results obtained with body alone and body-fins combinations at 0 degrees (one set of fins vertical and the other set horizontal) and 45 degree angle of roll. The results indicate that the addition of the fins to the body insures static longitudinal stability and provides essentially linear variations of the lift and pitching moment at small angles of attack throughout the Mach number range. The slopes of the lift and pitching-moment curves vary slightly with Mach number and show only small effects due to the angle of roll.

Alford, William J., Jr.↗

NASA NTRS1994

Hermes Global Orbiter: A Discovery Mission in Gestation

The hermes Global Orbiter (HGO) is a Discovery class mission under study, which is investigating the possibility of placing a small spacecraft in highly elliptical polar orbit about mercury. The purpose of the mission is to conduct observations of the planet's surface, atmosphere and magnetosphere.

spacecraft Mercury orbit payload Delta II Venus fl↗

NASA NTRS1952

Investigation in the Ames Supersonic Free-Flight Wind Tunnel of the Static Longitudinal Stability of the Hermes A-3B Missile at a Mach Number of 5.0

Models of the Hermes A-3B missile were tested in the Ames supersonic free-flight wind tunnel to determine the static-longitudinal-stability characteristics at a Mach number of 5.0 and a Reynolds number based on body length of 10 million. The results indicated that the model center of pressure was 45.3 percent of the body length aft of the nose and the lift-curve slope based on body frontal area was 0.064 per degree. Estimates indicated that the effect on these characteristics of aeroelastic twisting of the model fins was small but important if a precise location of center of pressure is required. A comparison of the test results with predictions based on available theory showed that the theory was useful only for rough estimates, The drag coefficient at zero lift, based on body frontal area, was found to be 0.155.

Denardo, Billy Pat↗

NASA NTRS2017

Performance, Stability, and Plume Characterization of the HERMeS Thruster with Boron Nitride Silica Composite Discharge Channel

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-3 (TDU-3) has been the subject of extensive technology maturation in preparation for flight system development. Detailed performance, stability, and plume characterization tests of the thruster were performed at NASA GRC's Vacuum Facility 5 (VF-5). The TDU-3 thruster implements a magnetic topology that is identical to TDU-1. The TDU-3 boron nitride silica composite discharge channel material is different than the TDU-1 heritage boron nitride discharge channel material. Performance and stability characterization of the TDU-3 thruster was performed at discharge voltages between 300V and 600V and at discharge currents between 5A and 21.8A. The thruster performance and stability were assessed for varying magnetic field strength, cathode flow fractions between 5% and 9%, varying harness inductance, and for reverse magnet polarity. Performance characterization test results indicate that the TDU-3 thruster performance is in family with the TDU-1 levels. TDU-3's thrust efficiency of 65% and specific impulse of 2,800sec at 600V and 12.5kW exceed performance levels of SOA Hall thrusters. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations (discharge current peak-to-peak and root mean square magnitudes), discharge current waveform power spectral density analysis, and maps of the current-voltage-magnetic field. Stability characterization test results indicate a stability profile similar to TDU-1. Finally, comparison of the TDU-1 and TDU-3 plume profiles found that there were negligible differences in the plasma plume characteristics between the TDU with heritage boron nitride versus the boron nitride silica composite discharge channel.

ion engines↗

NASA NTRS2018

Long Duration Wear Test of the NASA HERMeS Hall Thruster

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5-kW Hall thruster is the subject of extensive technology maturation by NASA GRC and JPL in preparation for development into a flight propulsion system. As part of this on-going effort, a series of three wear tests have been conducted to identify erosion phenomena and the accompanying failure modes as well as to validate service-life models for magnetically-shielded thrusters. This paper presents an overview and summary of the results obtained over the first 1715 hours of the third wear test, which has the overall goal of serving as a pathfinder to identify and correct design or facility issues prior to the flight qualification campaign. Overall, negligible changes in performance and stability are observed as a function of operating time as well as relative to previous wear tests. Erosion of the inner and outer front pole covers is shown to vary by 76-300 percent as a function of discharge voltage and by up to 40 percent as a function of magnetic field strength. Shifting the cathode position upstream relative to the pole covers is shown to reduce keeper erosion rates by 84 percent, which supports this approach for mitigating the elevated keeper wear observed during previous wear tests.

Frieman, Jason D.↗

NASA NTRS2018

Long Duration Wear Test of the NASA HERMeS Hall Thruster

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5-kW Hall thruster is the subject of extensive technology maturation by NASA GRC (Glenn Research Center) and JPL (Jet Propulsion Laboratory) in preparation for development into a flight propulsion system. As part of this on-going effort, a series of three wear tests have been conducted to identify erosion phenomena and the accompanying failure modes as well as to validate service-life models for magnetically-shielded thrusters. This paper presents an overview and summary of the results obtained over the first 1715 hours of the third wear test, which has the overall goal of serving as a pathfinder to identify and correct design or facility issues prior to the flight qualification campaign. Overall, negligible changes in performance and stability are observed as a function of operating time as well as relative to previous wear tests. Erosion of the inner and outer front pole covers is shown to vary by 76-300 percent as a function of discharge voltage and by up to 40 percent as a function of magnetic field strength. Shifting the cathode position upstream relative to the pole covers is shown to reduce keeper erosion rates by 84 percent, which supports this approach for mitigating the elevated keeper wear observed during previous wear tests.

Frieman, Jason D.↗

NASA NTRS2018

Long Duration Wear Test of the NASA HERMeS Hall Thruster

Frieman, Jason D.↗

NASA NTRS2019

HERMeS Thruster Magnetic Field Topology Optimization Study: Performance, Stability, and Wear Results

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The TDU-1 thruster implements a magnetically shielded field topology and has demonstrated the elimination of the discharge channel erosion as a life limiting mechanism. Extensive wear testing the TDU Hall thrusters has identified the thruster front pole covers as the next life limiting component. This effort aims to explore and investigate alternate magnetic field topologies to assess whether reductions in the front pole cover erosion can be attained while still maintaining very low erosion rates on the discharge channel walls. Four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the front pole covers were designed. Three of the four candidate magnetic field topologies (B1, B2, and B4) have been manufactured and were subjected to an extensive test campaign that includes laser induced fluorescence (LIF), performance, stability, wear, plume, and thermal characterization. In Phase I, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster's peak magnetic field does result in an upstream shift of the acceleration zone but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment results found at a normalized thruster magnetic field setting of 1, the thruster performance was very similar for all configurations. Discharge current waveforms ripple data indicated that configurations Bo, B1, and B2 has similar oscillatory profiles with the B2 configuration transitioning to higher oscillatory mode at 400V instead of the 450V observed for configurations B0 and B1. Configuration B4 ripple data indicates that the thruster was operating in a very oscillatory mode above 350V. Inner front pole cover erosion rates for configuration B1 were approximately 65% relative to B0 and the erosion rates for configuration B2 were 40% relative to B0. Analysis of the performance, stability, and wear test results of configurations B0, B1, B2, and B4 indicates that configuration B2 may present an alternative option for B0 because it mostly maintained the same performance and stability as configuration B0 but with 40% of the inner front pole cover erosion rate.

Kamhawi, Hani↗

NASA NTRSDate not supplied

HERMeS Thruster Magnetic Field Topology Optimization Study: Plasma Plume Results

NASA’s Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5-kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. The magnetic optimization study aims to explore and investigate alternate magnetic field topologies. Four candidate magnetic field topologies that reduce the effectiveness of the shielding along the discharge channel walls with the intent to also reduce the erosion rates along the front pole covers were designed. Three of the four candidate magnetic field topologies (B1, B2, and B4) have been manufactured and were subjected to an extensive test campaign that included laser induced fluorescence (LIF), performance, stability, wear, and plume characterization. In Phase I, LIF measurements along the discharge chamber centerline found that upstream retraction of the thruster’s peak magnetic field does result in an upstream shift of the acceleration zone, but the magnitude of the shift does not correspond one-to-one to the shift in the location of the peak radial magnetic field magnitude. Phase II test segment results found that at a normalized thruster magnetic field setting of 1, the thruster performance and discharge current oscillations were very similar for configurations B0-B2. Also in Phase II abbreviated wear tests found that average inner front pole cover erosion rates for configuration B1 were approximately 65% of those at B0, and the average erosion rates for configuration B2 were approximately 35% of those at B0. This paper reports on the detailed plasma plume mapping results for configurations B0-B4. Faraday probe results indicate that, at 600 V operation, the plume divergence increased slightly for the B4 configuration when compared to B0-B2. Wien filter spectrometer results found that at 600 V, lower multiply-charged species fractions were attained as the thruster transitioned from configurations B0 to B4. Finally, retarding potential analyzer results found that at 600 V operation and at angles above 60,˚ significantly lower high-energy ion content was found for configurations B2 and B4 when compared to B0 and B1.

Electric propulsion,.Hall thruster, laser diagnost↗

NASA NTRS2018

Investigations of Pole Erosion Mechanisms in the 12.5 kW HERMeS Hall Thruster Using Numerical Simulations and Ion Velocity Measurements

When the Hall Effect Rocket with Magnetic Shielding (HERMeS) was tested at the NASA GRC’s VF-5 the erosion rate on the pole covers was found to be at least a factor of two higher at 300 V and 20.8 A than at other operating conditions in the 400 V to 600 V range. Simulations using our hydrodynamics code Hall2De do not predict increased erosion rates at 300 V but accurately compute erosion rates similar to the measurements at the other operating conditions. We investigate the source of the discrepancy between measurements and simulations at 300 V using a combination of numerical simulations and experimental measurements of the ion velocity fields in the acceleration region of the thruster. By examining previous simulations that predicted the measured erosion rates for other thrusters and operating conditions, we determine that sputtering of the pole surfaces by high energy ions is the most likely mechanism behind the erosion rates at 300 V. High energy ions sputter the pole surfaces when the acceleration region of the thruster is downstream of the pole surface plane. When the latter occurs, the curvature of the plasma potential contours at the edges of the channel accelerates a small fraction of the high energy ions radially. We find that neither our Hall2De simulations nor the experimental measurements produce ions of sufficiently high energy and flux to the inner pole cover to explain the measured erosion there. Only at the outer edge of the cover is where we find the ions needed to yield simulation results that are comparable to the erosion measurements. We argue that one possible mechanism for the higher erosion is high energy ions that graze the channel corner of the pole and become trapped in a sheath that develops downstream. Within the sheath, we find that the electric field is large enough to turn the ions towards the surface of the pole. A simplified calculation shows that the erosion rates produced by sputtering of trapped ions are similar to the measurements. We also propose that the presence of local oscillations and high-energy cathode ions may be sources of the enhanced erosion We conclude this article by proposing a combination of analysis, simulation, and experimental measurements that can be used to address the validity of these hypotheses.

Mikellides, Ioannis G.↗