Aerojet Rocketdyne Research Articles

Propulsion Alternatives for Mars Transportation Architectures

Recently, NASA has pushed for returning humans to the moon, with in-situ resource utilization being the key capability to provide sustainability. One of the potential future developments could be a propellant depot in lunar distant retrograde orbit. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville Nuclear Thermal Propulsion (NTP) engine models, this research analyzed the impacts of using chemical and engines as well as the liquid-oxygen (LOX) Augmented Nuclear Thermal Rocket (LANTR) engines for these missions and compared their performances to the reference hydrogen-based NTP (H-NTP) engines all the while assuming a propellant depot at lunar distant retrograde orbit. For a human mission to Mars originating in the lunar distant retrograde parking orbit, the LANTR engines will offer better overall performance than H-NTP engines with a predicted 55.6% decrease in propellant volume, 39% decrease in vehicle dry mass, and 50% decrease in the number of aggregation launches. This is due to LANTR’s 22% higher specific impulse than conventional chemical propulsion systems, three times higher density than pure hydrogen, and 440% higher thrust than the baseline H-NTP engines. However, these benefits come at the cost of the propellant mass, which is 32.4% higher for the conjunction class mission and 106.7% higher for the opposition class mission than the baseline H-NTP system.

Impacts of In Situ Alternative Propellant on Nuclear Thermal Propulsion Mars Vehicle Architectures

Recently, NASA has pushed for returning humans to the moon sustainably with In situ resource utilization being the central focus. In its permanently shadowed regions, the moon has an abundance of water and ammonia, which are two potential alternative propellants to hydrogen in nuclear thermal propulsion (NTP) engines. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville NTP engine models, this research analyzed the impacts of using water or ammonia as propellants on mission architectures. For a human mission to Mars originating in the lunar distant retrograde parking orbit, when comparing the baseline hydrogen vehicles to vehicles using water or ammonia, the effects of increased propellant density by an order of magnitude despite a 62% decrease in specific impulse will still require an average of 53% of the launches for vehicle assembly, 59% of the dry mass, 25% of the propellant volume for the conjunction-class mission, and 50% of the propellant volume for the opposition-class mission when comparing against the baseline hydrogen vehicle. Due to the oxidative nature of water, which results in 29% of the engine life compared with that of hydrogen and ammonia, ammonia outperforms water in terms of the total number of missions per engine block by an average factor of 6.8. Because ammonia’s specific impulse is 40.6% that of hydrogen, the propellant mass is increased by a factor of 2.9 and longer burn times result in an average of a 10% decrease in the total number of missions that the hydrogen system can achieve. Since the changes to the vehicle architecture are more extreme and in favor of ammonia than the total mission capability, for reusable vehicles using NTP technology ferrying humans to Mars while utilizing in situ propellants from the moon, this research recommends ammonia as the propellant of choice.

Seeded Hydrogen in Mars Transfer Vehicles Using Nuclear Thermal Propulsion Engines

Using the NASA Design Reference Architecture, an Aerojet Rocketdyne vehicle architecture, and a University of Alabama in Huntsville nuclear thermal propulsion engine model, the performance of a Mars transfer vehicle using hydrogen propellant seeded with argon was analyzed. Seeded hydrogen, up to a maximum seed mass concentration of 55.85%, increased engine and vehicle performance by reducing pressure losses, decreasing reactor power, and increasing the overall while assuming constant propellant volume and vehicle dry mass. The tradeoff for using seeded hydrogen was lowered specific impulse and increased net propellant mass, resulting in increased vehicle wetted mass. Vehicle performance monotonically increased with seed mass concentration and provided a best-case 32-day reduction in round-trip transit time to Mars versus pure hydrogen.

LEU NTP Engine System Trades and Mission Options

The future of human exploration missions to Mars is dependent on solutions to the technology challenges being worked on by the National Aeronautics and Space Administration (NASA) and industry. One of the key architecture technologies involves propulsion that can transport the human crew from Earth orbit to other planets and back to Earth with the lowest risk to crew and the mission. Nuclear thermal propulsion (NTP) is a proven technology that provides the performance required to enable benefits in greater payload mass, shorter transit time, wider launch windows, and rapid mission aborts due to its high specific impulse and high thrust. Aerojet Rocketdyne (AR) has stayed engaged for several decades in working NTP engine systems and has worked with NASA recently to perform an extensive study on using low-enriched uranium NTP engine systems for a Mars campaign involving crewed missions from the 2030s through the 2050s. Aerojet Rocketdyne has used a consistent set of NASA ground rules and they are constantly updated as NASA adjusts its sights on obtaining a path to Mars, now via the Lunar Operations Platform-Gateway. Building on NASA’s work, AR has assessed NTP as the high-thrust propulsion option to transport the crew by looking at how it can provide more mission capability than chemical or other propulsion systems. The impacts of the NTP engine system on the Mars transfer vehicle configuration have been assessed via several trade studies since 2016, including thrust size, number of engine systems, liquid hydrogen stage size, reaction control system sizing, propellant losses, NASA Space Launch System (SLS) payload fairing size impact, and aggregation orbit. An AR study activity in 2018 included examining NTP stages derived from Mars crew mission elements to deliver extremely large cargo via multiple launches or directly off the NASA SLS. This paper provides an update on the results of the ongoing engine system and mission trade studies.

NASA Started a Propeller set on Board Voyager 1 After 37 Years of Break

If you try to start a car that has been in the garage for decades, you expect the engine not to respond. But a set of propellers onboard the NASA Voyager 1 spacecraft was launched on Wednesday, November 29, 37 years after its last use without any problems. Voyager 1 is the only man-made object that has arrived in interstellar space, being also the space probe created by NASA, which travels at the highest speed and is at the highest distance from Terra. The probe flies for 40 years and can change its position to keep its antenna pointing to the Terra using some small propellers operating in very short halves, in the order of milliseconds. NASA's Voyager team has been able to launch a set of back-up propellants that had not been in use since 1980. The test succeeds in extending Voyager 1's life to a minimum of 2-3 years. In 2014, NASA engineers noticed that Voyager's propellers used to change direction degraded. Over time, propellers end up working longer than normal to get the same effect on the direction of the probe. NASA experts have designed several working scenarios to solve the problem and concluded that it is best to use a series of back-up engines to control the probe's direction. These propellants had not been used for 37 years. NASA has been forced to search for decades old archives and use an obsolete programming language that no one uses to compile commands that have been transmitted by radio waves to the small computer on board to Voyager 1. The probe is more than 20 billion km from Terra. In the early years of the mission, Voyager 1 passed past Jupiter, Saturn and some of the satellites of these planets. In order to maintain the correct distance and orientation of on-board instruments, engineers used a series of Trajectory Correction Maneuvers (TCM) with dedicated, but identical size and functionality to those used for small flight corrections. These propellers used to correct the trajectory are placed on the back of the probe. After the encounter with Saturn, Voyager 1 did not need them, the last use being on November 8, 1980. These propellers had been used in a different way, meaning they were operating for long periods, not for very short-lived pulses. All engines on board Voyager were produced by Aerojet Rocketdyne, the same type of engine being installed on other spacecraft such as Cassini and Dawn. On November 28, Voyager engineers started the four TCM engines and tested their ability to steer the probe using 10 millisecond pulses. Researchers were then forced to wait for the test results to travel through space, in the form of radio waves, to be received after 19 h and 35 min by an antenna from Goldstone, California, part of NASA's Deep Space network.

Electrical Facility Effects on Hall Current Thrusters: Electron Termination Pathway Manipulation

A nonnegligible fraction of the charged particles in the Hall current thruster plume completes the electrical circuit through the conductive wall of a ground-based vacuum test facility. The resultant electrical circuit is different from the electrical circuit completed by the Hall current thruster in the onorbit environment. To understand the electrical circuit created in ground-based testing, this work examines the effect of an electrically biased metal plate, placed in the far-field plume of a Hall current thruster, on the plasma plume characteristics, the Hall current thruster thrust, and the electron termination pathways. An Aerojet Rocketdyne T-140 Hall current thruster is operated at 300 V and 10.3 A on xenon propellant. The operational neutral background pressure is , corrected for xenon. Two aluminum plates, one representative of the radial wall of the vacuum chamber and one representative of the axial wall of the vacuum chamber, are placed 2.3 m radially outward from the thruster centerline and 4.3 m axially downstream from the discharge channel exit plane, respectively. At each axial bias plate voltage, measurements of thrust, electrical characteristics of the Hall current thruster, thruster body electrical waveform, and radial–axial plate waveforms are recorded. A Langmuir probe, a Faraday probe, and an emissive probe are placed 1 m downstream of the Hall current thruster exit plane. The cathode-to-ground voltage and plasma potential behavior closely follow the trends observed from in-flight measurements of the Small Missions for Advanced Research in Technology-1 PPS-1350 Hall current thruster. This investigation experimentally quantifies the impact of the varying in-flight plasma plume conditions on Hall current thruster operation in a ground-based vacuum facility.

Electrical Facility Effects on Hall-Effect-Thruster Cathode Coupling: Discharge Oscillations and Facility Coupling

The physical mechanisms that govern the electrical interaction between the Hall-effect-thruster electrical circuit and the conductive vacuum-facility walls are not fully understood. As a representative test bed, an Aerojet Rocketdyne T-140 Hall-effect thruster is operated at 3.05 kW and a xenon mass flow rate of with a vacuum facility operating neutral pressure of , corrected for xenon. Two electrical witness plates, representative of the facility chamber walls, are placed 2.3 m radially outward from thruster centerline and 4.3 m axially downstream from the thruster exit plane. The cathode is radially translated from 18.1 to 77.8 cm away from the thruster centerline. At each cathode position, the discharge current and the electrical waveform of the radial and axial plates are simultaneously measured. As the cathode radial position changes from 18.1 to 77.8 cm from the thruster centerline, the discharge-current oscillation frequency decreases between 17 and 35% for the electrically grounded thruster-body case, and between 15 and 23% for the electrically floating thruster-body case. The analysis of the electron current collected by the radial plate suggests that electrons directly sourced from the cathode impinge on the radial plate at large cathode positions. Overall, the results of this work show that the chamber walls act as an artificial electrical boundary condition that keeps the Hall-effect-thruster plume plasma potential to within certain bounds.

Electrical Facility Effects on Hall Thruster Cathode Coupling: Performance and Plume Properties

The impact of facility conductivity on Hall effect thruster cathode coupling is experimentally investigated. The 3.4 kW Aerojet Rocketdyne T-140 Hall effect thruster operating at a discharge voltage of 300 V, a discharge current of 10.3 A, and an anode flow rate of serves as a representative Hall effect thruster test bed. The nominal facility operating pressure during thruster operation is corrected for xenon. Two square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the conductive vacuum chamber at two locations with respect to the center of the thruster exit plane: 4.3 m axially downstream along the thruster centerline, and 2.3 m radially outward centered on the exit plane. The plates and body of the Hall effect thruster are configured in three distinct electrical configurations with corresponding measurements: 1) electrically grounded with measurements of currents to ground, 2) electrically isolated with measurements of floating voltages, and 3) isolated from ground but electrically connected with measurements of the current conducted between the plates. Measurements are taken as the radial position of the cathode is varied from 0 to 129 cm with respect to the nominal cathode location. Measurements of the current collected by the plates and thruster body indicate that cathode electrons preferentially travel to the thruster body, Hall effect thruster plume, and radial facility surfaces for cathode locations in the near field, midfield, and far field, respectively. These results indicate that cathode position alters the recombination pathways taken by cathode electrons in the Hall effect thruster circuit.

Patent Analysis of the Course of Aerojet's Business and the Key Technologies of Hypersonic

This study researched the patents published from 1942 to March of 2014 of Aerojet Company (including Aerojet Rocketdyne Company's patents) in USA, considering Aerojet's annals and other information, analyzed relationship of the patents and the course of Aerojet's business development, and the technology R&D process, and extracted the key hypersonic technologies of Aerojet Company, at last raised the conclusion.

Role of a Conducting Vacuum Chamber in the Hall Effect Thruster Electrical Circuit

The role of the electrically conductive vacuum chamber wall in the completion of the discharge circuit of a Hall effect thruster is experimentally investigated. The Aerojet Rocketdyne T-140 laboratory-model Hall effect thruster, operating at a discharge voltage of 300 V, a discharge current of 5.16 A, and an anode flow rate of , serves as a representative Hall effect thruster testbed. The nominal facility operating pressure during thruster firings is torr, corrected for xenon. Two , square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the stainless-steel vacuum chamber at two locations with respect to the center of the thruster exit plane: 4.3 m axially downstream along the thruster centerline and 2.3 m radially outward centered on the exit plane. The plates are configured in three distinct electrical configurations with corresponding measurements: 1) electrically grounded plates with measurements of currents to ground, 2) electrically isolated plates with measurements of floating voltages, and 3) isolated but electrically connected plates with measurements of the current conducted between them. The measurements are all taken simultaneously, with the discharge current oscillations of the thruster at a sampling frequency of 100 MHz. Measurements of the current conducted to ground in the electrically grounded configuration reveal that the axial and radial plates collect ion currents that are 13.6 and 10.7% of the discharge current, respectively; the collected current is coupled to the discharge current oscillations but is smaller in magnitude and phase delayed. In the electrically connected-plate configuration, 5.5% of the average discharge current is observed to flow from the axial plate to the radial plate, driven by a floating voltage difference of 7.6 V; this current is uncorrelated in time with the discharge current oscillations. These results indicate that the vacuum chamber conducts current and is a recombination site for a significant number of plume ions during Hall effect thruster operation.