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Centrifugal Nuclear Thermal Propulsion is a second-generation propulsion concept currently being researched to succeed traditional solid core nuclear thermal propulsion. Current efforts across the space industry are focusing on development of technologies for enabling manned flight to Mars and scientific missions to the outer planets. To enable these types of mission's direct injection trajectories are needed to decrease exposure to astronauts and scientific payloads to the harsh environment of space. Current efforts in nuclear thermal propulsion have shown great promise in higher specific impulses ∼900 s, but even these levels of specific impulse fall short for manned Mars missions with durations less than two years or direct injection to orbit of the outer ice giants Uranus and Neptune without the use of gravity assists. High performance nuclear thermal propulsion has been proposed using bubble through nuclear reactors to reach specific impulses as high as 1800 s, however significant challenges exist to prove the feasibility of such a system. One of the major challenges is the fluid mechanics within the core of the engine since little is known about uranium in the liquid phase. An experimental study has been conducted to better understand the fluid dynamics to better inform the thermodynamic and nucleonic models. Using the data from the experiment, high-fidelity simulations have been developed to predict the fluid dynamics of the uranium fuel. This information can then be further fed into thermodynamics models in development to predict core temperatures and system behaviour. The validated models show that small bubbles reach thermal equilibrium, and high centrifugal forces and surface tension assist in confining the molten uranium within the reactor core. Injector methods suggest feasible configurations to encourage bubble flow while minimizing backflow of the propellant, motivating further study of Centrifugal Nuclear Thermal Propulsion based systems. • Galinstan (Ga67In20.5Sn12.5) was determined to be a suitable uranium simulant for bubble formation. • High Speed Radiography was used to determine bubble characteristics using machine learning algorithm. • Hydrogen bubbles in uranium rapidly reach local equilibrium with diameters of 0.266 mm and reaching velocities of 16 m/s. • To date all research indicates that a centrifugal nuclear thermal rocket is feasible. [ABSTRACT FROM AUTHOR]

The non-equilibrium spectra of shock-heated nitrogen in the UV-visible range are investigated in a detonation driven shock tube. Spectral identification indicates that the main contributors to the spectra are molecular radiation of N 2 (2+) and N 2 + (1-). Vibrational and rotational temperatures of N 2 (C) and N 2 + (B) are determined through spectral fitting. The post-shock spectra of pure nitrogen are obtained for shock velocities ranging from 6.82–9.00 km/s and P 0 = 30–200 Pa at the shock arrival time. Differences between vibrational and rotational temperatures of N 2 (C) and N 2 + (B) are observed, suggesting that nitrogen is in a non-equilibrium state at shock arrival moment. Time-resolved spectra are obtained within varying delay times of 0– 2. 0 μ s at P 0 = 200 Pa and shock velocity of V sw = 6.60 km/s. Time-resolved temperatures are determined to illustrate the time-resolved non-equilibrium characteristics of nitrogen at high temperature. It is observed that the non-equilibrium characteristics of nitrogen gradually weaken with increasing time and the thermal equilibrium was not obtained within 2. 0 μ s. Finally, the time-resolved temperatures are compared with CEA (Chemical Equilibrium with Applications) prediction, revealing that the difference in temperature between the experiment and the CEA equilibrium calculation decreases as time increases. • The radiative spectra of nitrogen are obtained in shock tube with shock velocity of 6.5–9.0 km/s. • The shock velocity-dependent non-equilibrium characteristics of nitrogen are revealed. • The time-resolved non-equilibrium characteristics in the post-shock region are studied. • The non-equilibrium effects of N 2 (C) and N 2 +(B) are obtained and analyzed. [ABSTRACT FROM AUTHOR]

Shock/shock interactions (SSIs) often lead to high thermal loads. To understand the influences of thermochemical non-equilibrium effects on the Type III SSI, four gas models based on the assumption of thermal perfect gas (TPG), thermal equilibrium chemical non-equilibrium gas (TECNG), thermal non-equilibrium chemical frozen gas (TNCFG), and thermochemical non-equilibrium gas (TCNEG) are employed to simulate the Type III SSI at Mach 10. The intersection of incident shock and bow shock is determined by dimensionless intercept of 0.05 and 0.1. It is found that the chemical non-equilibrium effects significantly lift up the impingement position of shear layer by reducing the standoff distance of bow shock. As a result, the peaks of wall pressure and heat flux calculated by TECNG and TCNEG model are higher than TPG and TNCFG model, respectively. Type IIIa SSI occur in the flows calculated by the TECNG and TCNEG model for the case with a dimensionless intercept of 0.1. The peaks of wall pressure and heat flux calculated by TECNG model are both over four times higher than those of TPG model. The thermal non-equilibrium effects slightly increase the standoff distance of bow shock and lower the impingement position of shear layer. In addition, the thermal non-equilibrium reduces the angle to the horizontal direction of shear layer by increasing the specific heat ratio. • Type III shock/shock interaction at Mach 10 are simulated by four gas models. • The influences of non-equilibrium effects on flow fields are analyzed. • The chemical non-equilibrium effects significantly increase the peak values of heat flux and pressure. • The mechanisms of non-equilibrium effects on peak heat flux and pressure are revealed. [ABSTRACT FROM AUTHOR]

To characterize the impact of thermal nonequilibrium conditions in shock-induced combustion in high Mach number scramjet engines, a comparative numerical investigation has been performed. The combustor configuration was based on the HyShot II scramjet while the total enthalpy is 3.3 MJ/kg which resembles Mach 8 flight condition. The improved delay detached eddy simulation with Park two-temperature approach in combination with the vibration-chemical coupling model was employed in the nonequilibrium simulations. It is shown that under thermal nonequilibrium inflow conditions, noticeable stratification in the thermal state exists in the circumferential direction both in the combustor section and exhaust nozzle. In the inner core where the fuel mixes with the mainstream, flow deceleration caused by the heat release leads to a sufficient transformation of energy between the vibrational and translational-rotational modes approaching the thermal equilibrium state. However, the τ v , m i x is quite large in the supersonic mainstream, whereby the thermal nonequilibrium dominates. For the case with higher T v in the inflow, the fuel jet's penetration height is lower while the local viscosity is higher which results in poor mixing between the fuel and air stream. Nevertheless, a 400 K increase in inflow T v gives a higher effective temperature, which promotes the dissociation reactions, shortens the ignition distance and yields an 8% increase in the overall combustion efficiency. With better combustion performance, an additional 6% deceleration in the supersonic mainstream and the energy exchange rate between vibrational and translational-rotational modes becomes almost four times faster at the combustor outlet, which promotes the restoration of the thermal equilibrium state. Furthermore, under higher inflow vibrational temperature, the shock-induced combustion in supersonic crossflow becomes less stable with noticeable fluctuations in combustion heat release and local static pressure. • Combustion dynamics for the Hyshot II scramjet engine under different thermal nonequilibrium inflow conditions were investigated numerically. • Noticeable stratification in thermal state in the circumferential direction was observed when vibrational nonequilibrium effect was considered. • Increase in vibrational temperature at combustor inlet by 400 K resulted in an 8% increase in the overall combustion efficiency. • The instability of shock-induced combustion was enhanced under higher inlet vibrational temperature. [ABSTRACT FROM AUTHOR]

Aero-thermal analysis of hypersonic vehicles is of fundamental interest for structural design and material selection. Therefore, we investigated the aerodynamic heating characteristics of a blunt cone made from stainless steel in a combustion-heated hypersonic wind tunnel with a free stream of Mach 6. During ground testing, the temperature rise of the model from room temperature to thermal equilibrium under the influence of hypersonic flow was measured with thermocouples, and the equilibrium temperature of the blunt section was 1130–1150 K. In addition, the transient thermal response of the model to the aerothermodynamic load was simulated by fluid-thermal coupled analysis and compared with the measurement data. The obtained results correspond well to typical thermocouple data, and the temperature predicts the measurement with an accuracy on the order of 8.8%. On this foundation, some factors affecting the radiation equilibrium temperature of the blunt cone are also discussed. The analysis indicates that emissivity has an important influence, while turbulence intensity at the nozzle inlet has minimal impact. • The effective experimental time reached 350 s, and the radiation equilibrium temperature of the cone section is obtained in 225 s. • Under the experimental conditions of Mach 6 and total temperature of 1712 K, the radiation equilibrium temperature of the cone section is 1130–1150 K. • The fluid-thermal coupling analysis method can effectively predict the thermal response process of the model. • The radiation equilibrium temperature in the blunt section is not sensitive to the change of the total temperature of the gas flow. • Surface emissivity has a great influence on the radiation equilibrium temperature, while the turbulence intensity gives the minimal impact. [ABSTRACT FROM AUTHOR]

Electric thrusters are considered to be the most promising propulsion system of choice for space gravitational wave detection. However, the requirement of high accuracy and drag-free control results in significant challenges in thrust measurements. To better understand the thrust noise characteristics, this work proposes a time-frequency-domain method for thrust noise analysis, i.e., a scalogram method attained through wavelet analysis. Unlike the conventional Fourier transform method, it can provide 3-D relationships for the time-frequency and thrust noise. A free-state thrust measurement experiment identified that the environmental temperature produces the dominant impact in the thrust noise. During a 70,000 s measurement, even without thruster operation, the thrust noise increased from 0.3 μN/Hz1/2 to 19 μN/Hz1/2 as the environmental temperature increased from −1 °C to 3 °C. The corresponding frequency of peak thrust noise for the entire period tends to be the same of 0.1 Hz. Additionally, a Hall thruster thrust measurement experiment was conducted to verify the method's time-varying analysis capability for thrust noise. The complete time-frequency thrust noise variation throughout the entire operating process of a Hall thruster was observed. The results at different stages of thruster operation indicate that thrust noise has the greatest influence 150–300 s after ignition. This is because the thrust stand will take hundreds of seconds to reach thermal equilibrium after ignition. Another finding during the Hall thruster operation is that the 0.1 Hz frequency corresponding to the peak thrust noise is highly consistent with the frequency of peak thrust noise in the free-state thrust experiment. This reveals the fact that the 0.1 Hz frequency should be the dominant frequency for thrust noise caused by temperature drift in our thrust measurement system. • A time-frequency-domain method for thrust noise analysis is developed for use with electric thrusters. • Time-varying and non-stationary features of thrust noise are analyzed using wavelet transforms. • The noise at 0.1 Hz is identified as the key frequency of thrust noise caused by zero drift. • Variations in the thrust noise during the entire operation process of a Hall thruster were obtained. [ABSTRACT FROM AUTHOR]