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Peng Ma, Shuangfeng Wang, Xiuzhen Wang
Journal of Propulsion and Power pp 1-8; https://doi.org/10.2514/1.b38462

Abstract:
Experiments have been performed to explore the combustion behaviors of spherical micron-sized aluminum (μAl) particles and liquid water for Al particle sizes in the range of 3.5–25 μm. The ignition of quasi-homogeneous μAl/water mixtures was successfully implemented by employing a novel ignition method, and self-sustained flame propagation was obtained in the mixtures over a broad range of fuel-equivalence ratio ϕ. The burning rates, flammability limits, and thermal structure of the propagating flame were determined. The combustion products were also analyzed. For the particle sizes considered, the burning rates were found to first increase and then decrease as ϕ increased, with the maximum values occurring at ϕ=1.7–2.0 and substantially lower than nano-Al/H2O mixtures. The dependence of the burning rate on particle size follows a power law, rb∼D−0.18, indicating that the reaction process of μAl and water is kinetically controlled. Base on the experimental observations, a simplified flame propagation model was developed to provide insight into the effects of particle size and equivalence ratio. Combustion product analyses revealed that Al residues increased as ϕ was further increased from 0.7, and the combustion efficiency of aluminum decreased accordingly.
Yongtae Yun, Juwon Kim, Sejin Kwon
Journal of Propulsion and Power pp 1-12; https://doi.org/10.2514/1.b38241

Abstract:
In this study, impact of the port diameter and length changes of a solid fuel on its performance parameters was investigated to propose design considerations for the solid fuel for hydrogen peroxide/high-density polyethylene hybrid rockets. A laboratory-scale hydrogen peroxide/high-density hybrid rocket was designed to perform a parametric study of the solid fuel through experiments. The ratio of the nozzle throat area to the fuel port area and ratio of the fuel length to the fuel port diameter were used to design the port diameter and length of the solid fuel. Altering the fuel port diameter affected the characteristic velocity efficiency, oxidizer mass flux, velocity-based erosive burning, oxidizer-to-fuel ratio, and regression rate. Altering the fuel length affected the oxidizer-to-fuel ratio and characteristic velocity efficiency and had little effect on the regression rate because it did not significantly contribute to the change in the oxidizer mass flux. No improvement was observed in the characteristic velocity efficiency after the ratio of the fuel length to the fuel port diameter reached 20. Finally, design considerations for the port diameter and length of the solid fuel for a hydrogen peroxide/high-density polyethylene hybrid rocket were suggested based on the experimental results considering the changes in the ratio of the nozzle throat area to the fuel port area and ratio of the fuel length to the fuel port diameter.
, Patrick Bätz
Journal of Propulsion and Power pp 1-13; https://doi.org/10.2514/1.b38349

Abstract:
Currently, several green propellants are under investigation to replace the toxic and carcinogenic monopropellant hydrazine (N2H4). Beside alternatives as ammonium dinitramide-based propellants, hydroxylammonium nitrate-based mixtures, or high concentrated hydrogen peroxide, mixtures of fuels with nitrous oxide (N2O) called HyNOx or NOFBX represent promising green propellants. Compared to classical N2H4, N2O/fuel mixtures offer a higher performance (Isp and c*), no or low toxicity, and low propellant costs. To investigate the handling and performance of a propellant mixture consisting of N2O and ethene, hot gas combustion tests with an experimental combustor were conducted. This paper summarizes the results of 134 combustion tests conducted with the premixed propellant injected in gaseous state. Calculated and measured performances (c* and ηc*) depending on mixture ratio, characteristic combustion chamber length L*, and chamber pressure are shown and discussed. Furthermore, the residence timescales of the propellant mixture inside the combustion chamber are normalized by the chemical timescale of the combustion process. Thus, combustion efficiencies depending on dimensionless Damköhler numbers are derived.
Momar Hughes, John Olsen
Journal of Propulsion and Power pp 1-13; https://doi.org/10.2514/1.b38393

Abstract:
This study investigates the application of an exhaust gas energy harvesting hybrid powertrain to three general aviation aircraft with reciprocating spark-ignition piston engines. The hybridization process involves downsizing the internal combustion engine to save weight and installation of an exhaust heat harvesting system composed of a harvester, a generator, lithium-ion traction battery, an electrical conversion system, and an electric motor. Novel calculations for a hybrid powertrain performance envelope are presented that estimate the upper limits to the degree of downsizing and mass of the harvester. An organic Rankine cycle (ORC) has been proposed as a potential compact, high-efficiency harvester. Five hydrofluroolefins (which have no ozone depletion potential and very low global warming potential) are considered as organic fluids for the ORC. It is found that R-1336mzz(Z) performs best in this application. Each ORC design parameter is investigated, and an optimized design with 14.3% thermal efficiency is found. At this efficiency, the internal combustion engine may be downsized up to 34.5%. The fuel burn reduction at this degree of downsizing may be as high as 12.7–14.5%.
Satoru Sawada, Keisuke Goto, Kazuki Ishihara, Akira Kawasaki, Ken Matsuoka, , Akiko Matsuo, Ikkoh Funaki
Journal of Propulsion and Power pp 1-12; https://doi.org/10.2514/1.b38374

Abstract:
A rotating detonation engine (RDE) generates a continuous thrust with one or more rotating detonation waves. Because of the velocity on the order of kilometers/second, the reaction zone is relatively small. Therefore, the RDE realizes a short combustion chamber length. However, the detonation waves induce an azimuthal motion of propellant, resulting in torque around the thrust axis. Because the motion does not contribute to the thrust, the torque is important in terms of performance loss. Herein, we conducted combustion tests with a six-axis force sensor to simultaneously measure 0.149±0.009 Nm torque and 48.1±0.9 N thrust. A comparison of detonation waves captured by high-speed camera revealed that the torque followed the direction and was offset when the waves existed in both of two directions simultaneously, which indicates the possibility of controlling the torque. Under a mass flow rate at 87±9 g/s and an equivalence ratio at 1.43±0.28, when the azimuthal component of shear force was 8.8±0.6% of the thrust, 0.77±0.10% of the total kinetic energy of the exit flow was distributed to the azimuthal component of velocity and did not contribute to the thrust. We therefore concluded that the effect of the azimuthal motion on the RDE’s performance was small.
Troy L. Messinger, Colin D. Hill, Declan T. Quinn, Craig T. Johansen
Journal of Propulsion and Power pp 1-8; https://doi.org/10.2514/1.b38330

Abstract:
Spectral analyses were performed on high-speed video of three nitrous-oxide/paraffin-based hybrid rocket launches and one static test fire of a nitrous/paraffin-based hybrid rocket motor. The imagery was taken to assess the combustion stability and identify any dominant instability modes. A spectral analysis of the image luminosity signals has shown distinct oscillatory modes in the plume flowfield. The plume oscillations have been compared and linked to existing hybrid combustion instabilities reported in the literature and theoretical predictions. The use of high-speed imagery has proven useful as a nonintrusive method of gathering high-frequency data in the analysis of hybrid combustion during launch. Analysis of high-speed imagery from a static test has revealed the time-varying aspect of the dominant oscillatory modes. The first longitudinal acoustic mode, identified in all data sets, has been used in a novel way to determine the characteristic velocity of the operating motor.
Zhuang Ma, Gaofeng Wang, Tao Cui, Yao Zheng
Journal of Propulsion and Power pp 1-10; https://doi.org/10.2514/1.b38410

Abstract:
The transition from the low-amplitude, aperiodic fluctuation to thermoacoustic instability is usually treated as a Hopf bifurcation, but the transition incorporates complex dynamical behaviors. Intermittent oscillations commonly occur in the transition process wherein chaotic states alternate with limit cycle oscillations and both the chaotic state and the limit cycle oscillation last with no fixed period. Intermittent oscillations that occur in the thermoacoustic system are demonstrated to be deterministic. A Koopman linearization procedure is implemented to model the intermittent combustion oscillations in the backward-facing step combustor. The nonlinear dynamics of intermittent combustion oscillations are approximated as a forced linear system in an orthogonal space, which is constructed by singular value decomposition of a delay coordinates matrix. The phase space of an intermittent combustion oscillation can be divided into linear and nonlinear regions by the magnitude of the nonlinear term in the forced linear system. As a result, intermittent combustion oscillations can be described as the transition between chaotic states and limit cycle oscillations, and the transition occurs when the nonlinear term is large.
Sudip Bhattrai, Liam P. McQuellin, Gaetano M. D. Currao, Andrew J. Neely, David R. Buttsworth
Journal of Propulsion and Power pp 1-14; https://doi.org/10.2514/1.b38348

Abstract:
The response and performance of an aeroelastic hypersonic intake was studied experimentally using fundamental geometry and structural boundary conditions. The experiments were conducted in a hypersonic wind tunnel at a Mach 5.85 condition. The most relevant deforming component was the compression ramp, which was treated as a cantilever surface to emulate the global deformation of the intake. Flowfield measurements were performed using pressure transducers, pressure-sensitive paints, and schlieren flow visualization. The dynamic structural response was measured using digital image correlation as well as feature tracking from the schlieren videos. A point measurement of total pressure in the isolator was taken using a pitot tube to quantify the effects of intake ramp deformation on the intake performance. The loss of total pressure in the isolator was found to correlate directly with the intake ramp deformation, with a direct transient correlation between the peak loss in total pressure recovery and the peak ramp deformation. While undergoing aeroelastic deformation, the total pressure recovery of the cantilever compliant ramp decreased by up to 20% from the baseline values. The analysis showed a strong coupling, along with hysteresis in dynamic response, between the intake structural deformation and the shock-wave/boundary-layer flowfield in the isolator.
Go Fujii, Yu Daimon, Katsumi Furukawa, Chihiro Inoue, Daijiro Shiraiwa, Nobuhiko Tanaka
Journal of Propulsion and Power pp 1-7; https://doi.org/10.2514/1.b38421

Abstract:
We conduct a high-speed visualization of coolant liquid film dynamics inside a 10N-class bipropellant thruster using monomethylhydrazine and a mixture of nitrogen tetroxide with approximately 3% nitric oxide as propellants at equivalent conditions with the flight model. The direct visualization of the liquid film inside a quartz glass chamber demonstrates the transient dynamics of film flow from ignition to cutoff for the first time. The scenario for the life of the film flow is found as follows: the coolant fuel jet impinges on the chamber wall being the liquid film; the friction force from the chamber wall rapidly decelerates the film as one-tenth of the injection velocity; next, the film moves sheared by the fast combustion gas downstream; and eventually the film completely evaporates by heat transfer from the hot combustion gas. The liquid film presents the typical ripple wave structure, covered by the velocity and thermal boundary layers. The effect of the mixture ratio is significant as the film flow rate increases at a small mixture ratio, leading to the longer film length.
Matthew T. Vernacchia, Kelly J. Mathesius, R. John Hansman
Journal of Propulsion and Power pp 1-13; https://doi.org/10.2514/1.b38104

Abstract:
Small, low-thrust, long-burn-time solid propellant rocket motors could provide propulsion for a new class of kilogram-scale, transonic, uncrewed aerial vehicles (UAVs). This paper investigates technological challenges of small, low-thrust solid rocket motors: slow-burn solid propellants, motors that have low thrust relative to their size (and thus have low chamber pressure), thermal protection for the motor case, and small nozzles that can withstand long burn times. Slow-burn propellants were developed using ammonium perchlorate and 0–20% oxamide (burn-rate suppressant), with burn rates of 1–4 mm⋅s−1 at 1 MPa. Using these propellants, a low-thrust motor successfully operated at a thrust/burn area ratio 10 times less than that of typical solid rocket motors. This kilogram-scale motor can provide 5–10 N of thrust for 1–3 min. An ablative thermal protection liner was tested in these firings, and a new ceramic-insulated nozzle was demonstrated. This paper shows that small, low-thrust solid motors are feasible and presents a baseline design for the integration of such a motor into a small UAV.
Daisuke Ichihara, Koichiro Oka, Ayumi Higo, Yusuke Nakamura, Kiyoshi Kinefuchi, Akihiro Sasoh
Journal of Propulsion and Power pp 1-4; https://doi.org/10.2514/1.b38480

Matteo Migliorini, ,
Journal of Propulsion and Power pp 1-13; https://doi.org/10.2514/1.b38127

Abstract:
Closer integration between the airframe and the propulsion system is expected for future aircraft to reduce fuel consumption, emissions, weight, and drag. The use of embedded or partially embedded propulsion systems will require the use of complex intakes. However, this can result in unsteady flow distortion, which can adversely affect the propulsion system efficiency and stability. Relative to conventional measurement systems, time-resolved particle image velocimetry provides sufficient spatial and temporal resolution to enable the development of new methods to assess unsteady flow distortion. This work proposes a novel analysis approach to assess the unsteady flow distortion. For an S-duct configuration, the method was successfully used to evaluate the perturbations of the indealized incidence angle. This example showed peaks up to ±30 deg incidence and a duration equivalent to the passing time of three blades. The introduction of a nonuniform total pressure profile at the S-duct inlet increased the probability of peak distortion events with higher magnitude. The method provides an estimate of the likelihood, magnitude, and duration of distortion events and is a new way to evaluate flow distortion that could induce instabilities for the propulsion system.
Simon Peterschmitt, Denis Packan
Journal of Propulsion and Power pp 1-10; https://doi.org/10.2514/1.b38156

Abstract:
The electron-cyclotron resonance thruster with magnetic nozzle relies on two successive energy transfer processes: first from electromagnetic energy to electron thermal energy, facilitated by a coupling structure; and second from electron thermal energy to ion directed kinetic energy, facilitated by a diverging magnetic field. The nature and geometry of the coupling structure are crucial to the first energy transfer process. This paper presents an experimental study of the performance of an electron-cyclotron resonance thruster with magnetic nozzle, equipped either with a waveguide-coupling structure or with a coaxial-coupling structure. The necessity of thrust balance measurements to perform such a comparison is demonstrated. The low coupling efficiency from microwave power to the plasma achieved by waveguide coupling is found to result in very large uncertainty with respect to the deposited power. A method to significantly reduce this uncertainty is proposed and implemented. Thrust balance measurements indicate 500 μN for the coaxial-coupled thruster and 240 μN for the waveguide-coupled thruster, both operated at 25 W of deposited microwave power and a mass flow rate of 98 μg/s of xenon. Electrostatic probe measurements reveal that this difference can be explained by a difference in ion energy. The results emphasize the critical role of the coupling structure, which may have been previously overlooked.
Daisuke Nakata, Kiyoshi Kinefuchi, Hitoshi Sakai, Suyalatu
Journal of Propulsion and Power pp 1-9; https://doi.org/10.2514/1.b38187

Abstract:
A high-efficiency concentric tubular-type resistojet with potential application to short-term orbit-raising maneuvers has been fabricated by 3-D printing and demonstrated. The propellant flows through multiple layers of cylindrical shells, with this structure also functioning as a single-piece heater. A 6-cm-high cylinder was realized with a wall thickness of 0.2 mm, using Inconel 718. A nodal thermal analysis was performed to identify the upper-limit current at a temperature limit of the wall material, and it was revealed that an outlet gas temperature of 871 K can be achieved with 77 A of current at 0.2 g/s of mass flow rate. The designed heater was combined with a boron nitride insulator and a stainless-steel housing, and thrust was measured in a vacuum chamber with nitrogen as the propellant. At a mass flow rate of 0.2 g/s and 75 A of current, an outlet temperature of 747 K, a specific impulse of 108 s, and a heater efficiency of 72% were achieved. These results with nitrogen propellant were used to predict the performance of a tungsten-made resistojet with a hydrogen propellant, and a specific impulse of over 700 s can be expected at a heater temperature of 2000 K.
Monique S. McClain, Aaron Afriat, Brandon J. Montano, Jeffrey F. Rhoads, I. Emre Gunduz, Steven F. Son
Journal of Propulsion and Power, Volume 37, pp 725-732; https://doi.org/10.2514/1.b38282

Abstract:
Typically, the burning surface of a composite solid propellant is controlled through grain geometry and formulation. However, combustion studies of grains constructed from different propellant formulations at fine scales (nominally 1 mm) are not readily accessible in open literature. With additive manufacturing, such configurations can be investigated easily. Propellants with a faster burning inner layer (enhanced with either 1 wt.% iron oxide or 5 wt.% nanoaluminum) were 3D printed between two layers of slower burning 85 wt.% ammonium perchlorate/hydroxyl-terminated polybutadiene propellant. The dynamic combustion behavior of the layered propellant was investigated at pressures ranging from 3.45 to 10.34 MPa. Overall, an increase in the burning surface area, without interlayer delamination, was observed. The driving force behind the propellant surface area increase was the difference in the burning rate between the layers. In addition, the nanoaluminum propellant layer had a more stable burning rate exponent than the cast nanoaluminum propellant. Overall, only a small addition of catalyzed propellant was needed to increase the burning rate of the bulk material. The results of this study lay the foundation for functionally grading propellant grains, which could tailor the thrust profile of solid rocket motors and gun propellants.
Philip M. Piper, Timothée L. Pourpoint
Journal of Propulsion and Power, Volume 37, pp 674-681; https://doi.org/10.2514/1.b38276

Abstract:
Fuel-film cooling is necessary to mitigate heat transfer in high-pressure oxygen-rich staged combustion engines. A 4.8 MPa (700 psia) axisymmetric kerosene–H2O2 combustor used fuel-film cooling to deposit carbonaceous material on removable metal samples. Posttest inspection of the samples revealed a two-layer structure, with a tenacious dense lower layer and a soot-like upper layer. Total deposit depth was measured using an optical profilometer and was repeatable between tests at the same conditions. Combustor conditions of fuel-film flow rate, bipropellant run time, fuel composition, chamber liner material, and chamber liner surface roughness were varied to determine their effects on total carbonaceous deposit depth as a function of position and time. Increasing the fuel-film flow rate by 40% resulted in similar deposit depths to lower fuel-film flow rates, but for longer axial lengths. Longer run times resulted in thicker deposits. The use of lower thermal conductivity chamber liners resulted in three to four times more deposits.
Philip M. Piper, Jason R. Gabl, Timothée L. Pourpoint, Timothy E. Dawson, Ranjan S. Mehta
Journal of Propulsion and Power, Volume 37, pp 759-768; https://doi.org/10.2514/1.b38275

Abstract:
Carbonaceous deposits can reduce heat flux in fuel-film-cooled kerosene rocket engines, but the deposition process at rocket conditions is poorly understood. Heat flux was measured in a 4.8 MPa (700 psia) fuel-film-cooled kerosene–hydrogen peroxide axisymmetric rocket combustor using 50 null point calorimeters and the Gauss–Newton algorithm with a constant Jacobian matrix. Carbon deposition caused a reduction in heat flux at axial positions furthest from the fuel-film injector, where deposits were later measured as thickest. Heat flux was reduced by a smaller amount in low-thermal-conductivity chamber liners when compared with high-thermal-conductivity copper liners. The roughness of the chamber-liner surface had little effect on quasi-steady-state heat flux but may have affected heat flux and carbon deposition during startup transients.
Puja Upadhyay, Khairul B. M. Q. Zaman
Journal of Propulsion and Power, Volume 37, pp 701-712; https://doi.org/10.2514/1.b38231

Abstract:
An experimental study is conducted to advance the understanding of flow physics associated with a boundary-layer ingesting, distributed propulsion system. The influence of incoming boundary-layer thickness on the performance of the system is examined. The propulsion model, integrated with electrical fans, is mounted on a flat plate and tested at subsonic speeds. Detailed characterization of the incoming boundary layer and the downstream flowfield is performed using hot-wire anemometry and static pressure measurements. Modification of the boundary layer is achieved by placing trip rods of two different diameters near the leading edge of the flat plate. Overall performance of the system is analyzed by estimating thrust, flow power, and input power to the fans. Ingestion of a thicker boundary layer is found to result in an increase in thrust and net flow power.
Darren C. Tinker, Marsalis P. Pullen, Robin J. Osborne, Robert W. Pitz
Journal of Propulsion and Power, Volume 37, pp 748-758; https://doi.org/10.2514/1.b38303

Abstract:
Spark discharges were parametrically examined for a cylindrical air-gap electrode configuration. The aim of this study was to elucidate spark discharge characteristics, arc penetration, and exhaust plume development to guide designers of relevant ignition devices. Spark gaps ranged from 0.5 to 2.3 mm, nominal pressures ranged from 150 to 2200 kPa, and two exciter types (bipolar and unipolar) were tested. Positive correlations were observed between the pressure–distance product and multiple dependent variables: breakdown voltage, energy discharged, and percentage of sparks quenched. Positive correlations were observed between the pressure–distance quotient and various other dependent variables: spark duration, channel resistance, and plume velocity. This study also discusses the effects of quenching on electrical measurements, how these effects are nontrivial, and the subtle irregularities in electrical results that are indicative of quenching.
Catherine A. M. Dillier, Erica D. Petersen, Thomas Sammet
Journal of Propulsion and Power, Volume 37, pp 693-700; https://doi.org/10.2514/1.b38173

Abstract:
To further characterize the roles of aluminum and ammonium perchlorate (AP) characteristics in ammonium perchlorate/R45-M hydroxyl-terminated polybutadiene (AP/HTPB)–composite propellants, aluminized and nonaluminized formulations were tested at very high pressures, up to 68.9 MPa (10,000 psi). A total of nine formulations were examined: seven nonaluminized with varying AP distributions, concentrations, and average particle sizes, and two aluminized with varying aluminum concentrations. All nine propellants showed an exponent or “slope” break above 20 MPa (2900 psi) and a postbreak pressure exponent greater than one. Decreasing the AP particle size decreased both the characteristic pressure where the exponent break occurred (or P*) and the pressure exponent after the break. AP distribution also affected P*, whereas changes in AP concentration did not. The inclusion of aluminum lowered P* compared with the nonaluminized formulations. Additionally, increasing the aluminum concentration appeared to lower the postbreak pressure exponent. The present study provides one of the first fundamental, systematic studies on the exponent break feature of AP/HTPB-composite propellants and adds new data to the limited database available in the open literature on the AP-based, composite propellant burning rate exponent break at very high pressures, with emphasis on the effect of mixture variables and Al on the characteristic break pressure and the postbreak pressure exponent.
Yujun Leng, Nicole L. Key
Journal of Propulsion and Power, Volume 37, pp 682-692; https://doi.org/10.2514/1.b37765

Abstract:
A generalized flat plate cascade model is used in this study to predict the unsteady aerodynamics of a vibrating blade row with nonuniform blade spacing in subsonic compressible flow. The blade row is assumed to vibrate in an isolated family of blade-dominated modes. The effect of nonuniform blade spacing on compressor rotor flutter stability is demonstrated by case studies based on the geometric and flow conditions of a high-speed three-stage axial research compressor. The results show that nonuniform blade spacing can greatly alter the blades’ aerodynamic damping. At certain vibrational nodal diameters, some blades are destabilized so much that their aerodamping becomes negative. However, negative aerodamping of some blades do not necessarily lead to the instability of the whole blade row. A general multibladed system aeroelastic model is derived to study the effects of the nonuniform blade spacing on rotor stability through an eigenvalue approach. The aerodynamic influence coefficients matrix can be calculated using the generalized flat plate cascade model for a blade row with any user-specified blade spacing patterns. The case studies investigated in this paper show that alternating blade spacing and shifting only one blade position can slightly increase the stability of the least-stable eigenmode, whereas sinusoidal blade spacing has a slightly destabilizing effect. On the other hand, the eigenvectors of the least-stable mode for the nonuniformly spaced blade rows can be significantly different from the uniform blade spacing case.
Kan Xie, Xiaolong Yi, Jiang Chen, Dongfeng Yan
Journal of Propulsion and Power, Volume 37, pp 769-779; https://doi.org/10.2514/1.b37933

Abstract:
As the level of aerospace technology continues to advance, the requirements for aircraft engines are becoming more onerous. A new thermal-cycle propulsion-technology scheme is proposed based on the existing propulsion system to meet the needs of round-trip space and multitask flights that use hypersonic vehicles, where thermoelectric materials are introduced into the wall of the engine tail nozzle and the front of the aircraft. The residual heat at the tail nozzle of the engine and the aerodynamic heat generated during flight are converted into electrical energy and reused in the engine cycle. The results show that, compared with traditional engine schemes, the thermoelectric combined-power engine has a higher specific impulse and cycle thermal efficiency, but the performance of the required thermoelectric materials is higher. However, the performance of existing materials cannot achieve the required standards.
Matthew T. Vernacchia, Kelly J. Mathesius, R. J. Hansman
Journal of Propulsion and Power, Volume 37, pp 792-800; https://doi.org/10.2514/1.b38106

Abstract:
Low-thrust long-burn-time solid rocket motors may be useful as propulsion for small, fast, uncrewed aerial vehicles. These motors require a slow-burning propellant that can operate at unusually low chamber pressures (0.3–2 MPa). Slow-burn propellants were developed using ammonium perchlorate oxidizer and the burn rate suppressant oxamide. By varying the amount of oxamide (from 0 to 20%), burn rates from 4 to 1 mm⋅s−1 (at 1 MPa) were achieved. The adjustable burn rate allows a set of similar propellants to serve many aircraft and mission concepts. This work presents burn rate measurements (from both a strand burner and a research motor), minimum burn pressure measurements, and combustion chemical equilibrium simulations. A novel model of oxamide’s effect on burn rate is also presented, and it fits well to the experimental data. Finally, these propellant data and models are applied to select the propellant and chamber pressure for an example low-thrust solid rocket motor.
Ewan Fonda-Marsland, Graham T. Roberts, Charles N. Ryan, David Gibbon
Journal of Propulsion and Power, Volume 37, pp 713-724; https://doi.org/10.2514/1.b38083

Abstract:
Experimental testing of a number of novel additively manufactured monopropellant microthrusters was conducted under atmospheric conditions using 87.5% concentration hydrogen peroxide. The aim of this work was to select a specific catalyst bed geometry for the thruster system and to investigate more general methodologies for monopropellant packed catalyst bed optimization. Characteristic velocity efficiencies approaching 0.98 were demonstrated, and performance improved for smaller beds with low aspect ratios; although, these beds flooded at lower propellant flow rates. The onset of bed flooding was used to identify physical limits of propellant flow rate supported by the catalyst. The particular propellant–catalyst pairing limit was defined by a Damköhler number of 56, independent of the bed geometry, with thermal performance peaking for the high flow rates just before flooding occurred. It is suggested that this method is extensible to other monopropellant systems, although with further work required to confirm it is a more general effect beyond thrusters using hydrogen peroxide.
Michael R. Natisin, Henry L. Zamora, Zachary A. Holley, N. Ivan Arnold, Will A. McGehee, Michael R. Holmes,
Journal of Propulsion and Power, Volume 37, pp 650-659; https://doi.org/10.2514/1.b38160

Abstract:
The overall propulsion efficiency for ion-mode electrospray thrusters has been predicted to be as high as 90%; however, experimental measurements currently fall far short of these predictions. Further complicating this is that for passively fed electrospray thrusters, the mass flow rate, which is required to obtain the propulsion efficiency and specific impulse, is not directly controlled or measured, and so this parameter is typically estimated by assuming all mass loss is due to the emitted ion current. Presented here is a detailed investigation into the efficiencies associated with a porous-media-based electrospray thruster operated in the purely ionic regime using the ionic-liquid propellant 1-ethyl-3-methylimidazolium tetrafluoroborate in both positive and negative ion emission modes. Measurements of performance metrics that affect thruster efficiency are discussed, including the transmission, angular, polydispersive, energy, and mass utilization efficiencies, in order to determine their impact on overall efficiency. The overall propulsion efficiency and specific impulse are also calculated using a variety of methods to better investigate how the assumptions made to estimate the mass flow rate affect these parameters. These results suggest that the efficiency of these devices may primarily be limited by the presence of additional mass loss mechanisms other than ion emission occurring during thruster operation.
Samith Sirimanna, Balachandran Thanatheepan, Dongsu Lee, Shivang Agrawal, Yangxue Yu, Yuyao Wang, Aaron Anderson, Arijit Banerjee, Kiruba Haran
Journal of Propulsion and Power, Volume 37, pp 733-747; https://doi.org/10.2514/1.b38195

Abstract:
Electric aircraft propulsion is a growing research area that looks into achieving propulsion through fully electric or hybrid electric systems while achieving low CO2 emissions. The system-level benefit gained by different electric and hybrid-electric propulsion schemes depends heavily on the performance of system-level components in the electric drive-train, including the electric motor, gear box, motor drive, protection systems, as well as the thermal management system. When comparing motor topologies, it is important to understand performance measures such as efficiency and specific power on a drive system level. Many different motor types have been qualitatively compared and can be found in the literature. To guide appropriate component selection, this paper presents details of a quantitative study for a given electric propulsion drive system. A Pareto optimal front for a notional drive system of a 1.5 MW electrical propulsor with different motor types is generated and compared. An optimization algorithm coupled with an electromagnetic finite element analysis software tool was used to optimize the induction motor, switched reluctance motor, wound rotor synchronous motor, permanent magnet synchronous motor (PMSM), slotless PMSM, permanent-magnet-assisted synchronous reluctance motor, brushless DC motor, and brushless doubly fed reluctance motor types for efficiency and specific power. Overall advantages considering system-level efficiency, specific power, and a few other key metrics such as origin of losses, cooling complexity, manufacturing tolerance, and fault tolerance are discussed. This gives an indication of the relative performance of different motor types and confirms the overall advantage of PM motor topologies in aircraft propulsion.
Alexis J. Harroun, Stephen D. Heister, Joseph H. Ruf
Journal of Propulsion and Power, Volume 37, pp 660-673; https://doi.org/10.2514/1.b38244

Abstract:
A computational and experimental study was conducted on nozzle geometries for rocket application rotating detonation engines (RDEs). Three geometries, including a nozzleless blunt body typically employed in RDE combustor hot-fire testing and two aerospike nozzles, were investigated. Simulations of the exhaust flow of a rotating high-frequency, high-pressure ratio wave based on rocket RDE test results were related to comparable constant-pressure conditions. Computational and experimental results showed the high momentum added by the highest-pressure detonation products influences the exhaust plume differently than a comparable steady flowfield fed by the same average product gas flow rate. In particular, the RDE exhaust flow tended to enhance entrainment on the nozzleless blunt body recirculation region and delay flow separation on nozzle expansion surfaces due to overexpansion compared to a constant-pressure engine. Results have important ramifications for isolating RDE combustor performance from nozzle effects and must be considered for future design of nozzle geometries to exploit the high-frequency, high-pressure ratio outflow of an RDE.
V. I. Yazhini, Balusamy Kathiravan, T. M. Muruganandam,
Journal of Propulsion and Power, Volume 37, pp 780-791; https://doi.org/10.2514/1.b38217

Abstract:
Experiments have been carried out to investigate the effect of cowl length variation on performance characteristics of a single expansion ramp nozzle. The performance parameters were estimated for cowl lengths of 0, 25, 50, 75, and 100% with respect to the horizontal length of the ramp. Experiments were conducted for different nozzle pressure ratios ranging between 1.5 and 9. The wall static pressure distribution data were measured from the tests to estimate the various performance parameters, such as axial thrust, normal force, gross thrust, thrust vectoring angle, and coefficient of pitching moment. High-speed schlieren imaging was used to visualize the flow separation and shock patterns and to measure the jet width. The flow was separated from the ramp wall up to a nozzle pressure ratio of 3 for all cowl cases. The shorter cowl length delays the downstream movement of shock-induced boundary separation inside the nozzle as compared to the longer cowl. The cowl trailing-edge flow was more underexpanded than the ramp tip flow. As cowl length increases, the increased restriction results in higher axial thrust and also increases the normal force. The pitching moment and thrust vectoring were dominated by normal force. Overall, as the nozzle pressure ratio increases, the axial force and jet width increase, whereas the normal force and the pitching moment increase up to a certain level and then decrease. As the cowl length increases, the axial thrust, normal thrust, pitching moment, and thrust vector angle increase, while the jet width decreases.
Patrick K. Dubois, Céderick Landry, Dominik Thibault, Jean-Sébastien Plante, Mathieu Picard, Benoît Picard
Journal of Propulsion and Power pp 1-8; https://doi.org/10.2514/1.b38004

Abstract:
Distributed aircraft propulsion has renewed the interest in power-dense, high-efficiency power packs. Ceramic turbomachinery could be a major enabler, although no successful design has been achieved in microturbine rotors. Rotor blade loading is tensile and a hurdle for successful conversion to ceramics. The inside-out ceramic turbine (ICT) rotor uses the superior compressive properties of monolithic ceramics by supporting ceramic blades against a structural composite rotating shroud. This enables low stress levels throughout the blade, increasing reliability and extending service life. An experimental demonstration of two ICT designs was conducted with 15-kW scale prototypes to identify critical issues: design A, a flexible hub that clamps blades against the structural shroud and design B, a sliding-blade configuration that allows free displacement of the blade. The flexible-hub design was tested up to 1000°C. Rotor integrity was preserved, but local blade cracking occurred. The sliding-blade design was successfully tested up to 1100°C for over 1 hour at a tip speed of 350 m/s with no issue. Tensile loading at the ceramic/metallic interfaces remains the key challenge to address. Reducing friction should overcome blade cracking and allow the proposed ICT to reach the targeted temperature of 1275°C and tip speed of 425 m/s.
Darren Dehesa, Shyam Menon, Sean Brown, Christopher Hagen
Journal of Propulsion and Power pp 1-13; https://doi.org/10.2514/1.b38261

Abstract:
Hybrid–electric powertrains offer the potential for performance improvements in unmanned aerial systems. However, for small unmanned aerial systems, potential gains in range and endurance can depend significantly on the aircraft flight profile and powertrain control logic. Subsequently, these impact the performance of individual powertrain components. This study uses dynamic simulations of an unmanned aerial system (UAS) with different powertrain control logic approaches to evaluate the performance of a series hybrid–electric powertrain. Component models generated using lookup table approaches and model parameterization are combined to generate a dynamic system model of the unmanned aerial system. The performance of the powertrain is evaluated for three representative mission profiles. Fuel consumption and battery state of charge form two metrics that are used to evaluate the performance of a baseline controller against an ideal operating line strategy. The ideal operating line strategy, which uses a performance map obtained by engine characterization on a specialized dynamometer, produces an average fuel economy improvement ranging from 0.5–2.0 g/km for a 30-min-long mission profile. The study demonstrates the need to consider a dynamic analysis aided by detailed component performance maps and a robust control strategy in evaluating hybridization approaches for UAS powertrains.
Mehmet Kahraman, Ibrahim Ozkol, M. Arif Karabeyoglu
Journal of Propulsion and Power pp 1-12; https://doi.org/10.2514/1.b38417

Abstract:
Low fuel regression rate is one of the major drawbacks of hybrid rocket motors. In this study, a novel injector concept is proposed to provide a substantial enhancement in the fuel regression rate. A tubular injector, Distributed Tube Injector (DTI), is inserted in the center of the cylindrical fuel port in order to tailor the oxidizer flow introduced into the combustion chamber with the desired combination of radial, tangential, or axial components of velocity. This concept has been tested using a 500 N thrust class hybrid rocket motor, which uses a paraffin-based fuel and supercharged N2O (L) as the oxidizer. As a result of 30 hot firings conducted in the test program, it is determined that the DTI configuration provides regression rates up to 3.9 times higher than the regression rates obtained using fore-end injector commonly employed in conventional hybrid rockets. Using the motor test data, a comprehensive nondimensional and scalable regression rate relation has been established. This nondimensional regression rate equation can be used to design the internal ballistic configuration of hybrid rocket motors using the novel injector.
Florin Saceleanu, Lauren LeSergent, Yiqi Zhang, Victoria Kerr, , Catalin F. Petre, Pascal Beland
Journal of Propulsion and Power pp 1-10; https://doi.org/10.2514/1.b38150

Abstract:
This experimental study investigates low power (3 W) laser ignition and the subsequent burning characteristics of consolidated Al/CuxO nanoparticle pellets and loose nanolaminate flakes. The effects of laser absorption, packing density (bulk porosity), and material properties on ignition are quantified, while simplified numerical models are used to examine their predominance. It is found that ignition of consolidated Al/CuO pellets is controlled by their effective thermal conductivity, such that ignition delay (approximately 1 to 100 ms) is minimized in the most porous specimen. However, the burning rate of these pellets is a compromise between the availability of reactive interfaces and the porosity allowing for gas expansion. On the other hand, ignition of loose Al/Cu2O nanolaminate flakes is predominately controlled by power absorption in the laser irradiated layer made of Al or Cu2O, especially when the thickness of the Cu2O layer is smaller than 60 nm. Ignition delay (approximately 0.1 to 1 ms) is tunable by adjusting the thickness of the top Cu2O layer, which absorbs more radiation compared to a top Al layer. Moreover, the burning rate of the nanolaminate samples is a compromise between the layer thickness for power absorption and the equivalence ratio that controls the extent of Al/CuxO reactions.
Chengjin Huang, Jianling Li, Mu Li, Ting Si, Cha Xiong, Wei Fan
Journal of Propulsion and Power pp 1-9; https://doi.org/10.2514/1.b37878

Abstract:
It has been urgently required to develop ionic liquid electrospray thruster (ILET) for rapidly expanding micro-/nanosatellites. In this work, ILETs with high-density array emitters are fabricated to quantitatively study the electrospray emission process of three typical ionic liquids (ILs). Onset voltage, current, voltage–current characteristics, and ratio of positive to negative currents are analyzed, and the effect of the properties of ILs on ILET is revealed. The onset voltage of the ILET is measured to be as low as 740 V, and the collected current is up to 100 μA. The results show that the viscosity of ILs has a great influence on the onset voltage and the emitted current. The inconsistency between theoretical and experimental results for the onset voltage of different ILs is explained in detail. Taking viscous stress into account, the estimated results agree well with the experimental ones. The analysis shows that the dynamic effect will be significant when the viscosity is high. In addition, the voltage–current characteristics and the ratio of positive to negative current are studied. The research is valuable to get insight into generative mechanism of IL electrospray and to provide a better design of both ILs and ILET.
Alicia Benhidjeb-Carayon, Jason R. Gabl, Benjamin E. Whitehead, Timothée L. Pourpoint
Journal of Propulsion and Power pp 1-11; https://doi.org/10.2514/1.b38266

Abstract:
Combined with a common fuel binder, solid hypergols can simplify the overall complexity of hybrid rocket engines, as the fuel grain can be ignited and reignited without any external power source or external fluid. Also, with the hypergolic additive embedded in the binder, the flame zone could be placed at the surface of the grain itself, thereby providing heat to the grain to maximize regression rate and promote combustion sustainability. The objective of this study was to demonstrate hypergolic ignition and successive relights of a 2-in motor grain configuration, with a paraffin-based fuel and MON-3 (3 wt % nitric oxide in nitrogen tetroxide) as the oxidizer. With sodium amide and potassium bis(trimethylsilyl)amide as solid hypergolic additives, the study focused on the influence of the additive type, loading, location, and format on the grain ignition delay and combustion sustainability. Hypergolic ignition was achieved with grain configurations composed of a front segment with 90 wt % additive and the main grain with loadings of additive from 40 down to 0 wt %. Two-s single burn tests provided regression rate estimates for the different grain combinations. Grain ignition delays varied between ∼20 and ∼250 ms, depending on the grain configuration, with C* efficiencies between 64% and 98%.
Scott J. Hall, Benjamin A. Jorns, Alec D. Gallimore, Hani Kamhawi, Wensheng Huang
Journal of Propulsion and Power pp 1-14; https://doi.org/10.2514/1.b38081

Abstract:
The plasma plume properties of a three-channel 100-kW-class nested Hall thruster were measured on xenon propellant for total powers up to 80 kW. The thruster was throttled through all seven available channel combinations for conditions spanning 300 to 500 V discharge voltage and three discharge current densities. A plasma diagnostics array, which included a Wien filter spectrometer, a retarding potential analyzer, and a planar Langmuir probe, was placed in the far-field plume of the thruster and used to measure the beam ion charge state, the ion energy distribution function, and the local plasma potential. These data were used to calculate thruster phenomenological efficiencies. These efficiencies are compared across the discharge voltage and channel combination, and they are compared to similar results from the NASA-300M single-channel high-power Hall thruster. An estimate of cross-channel ingestion, which is a phenomenon in the nested configuration that may improve thruster efficiency and that will be present in space, is calculated, and the results for mass utilization efficiency are corrected for this effect. These plasma diagnostic results are discussed in the context of the state of the art, as well as in that of the viability and potential benefits of the nested channel thruster configuration.
Edward Canepa, Davide Lengani, Alessandro Nilberto, Daniele Petronio, Daniele Simoni, Francesco Bertini, Simone Rosa Taddei
Journal of Propulsion and Power pp 1-12; https://doi.org/10.2514/1.b38259

Abstract:
Particle image velocimetry measurements have been carried out in a low-pressure turbine cascade operating under unsteady inflow to deeply investigate reduced frequency and flow coefficient effects on flow dynamics, and, consequently, on loss generation in the boundary layer and in the core flow region. Two independent measuring setups have been used for the purpose. The first one captured a large view of the entire blade passage, thus allowing the observation of the incoming wakes and related large-scale vortices developing in the core flow region. The second setup was instead focused on the rear part of the blade suction side to analyze the boundary layer development and to observe the mechanisms dominating the wake–boundary- layer interaction. Tests were performed for four flow cases, varying the reduced frequency and the flow coefficient independently. Proper orthogonal decomposition has been applied to quantify the turbulent kinetic energy production in the core flow, due to wake dilatation and distortion, and in the boundary-layer region. Upstream wake migration and boundary-layer-related losses are consequently quantified from particle image velocimetry data and compared with total pressure measurements for the different combinations of the inflow parameters, providing a clear view of the different loss sources affecting the unsteady operation of low-pressure turbine cascades.
Scott J. Hall, Benjamin A. Jorns, Sarah E. Cusson, Alec D. Gallimore, Hani Kamhawi, Peter Y. Peterson, Thomas W. Haag, Jonathan A. Mackey, Matthew J. Baird, James H. Gilland
Journal of Propulsion and Power pp 1-11; https://doi.org/10.2514/1.b38080

Abstract:
The performance of a three-channel, 100-kW nested Hall thruster was evaluated on xenon propellant for total powers up to 102 kW. The thruster demonstrated stable operation in all seven available channel combinations at discharge voltages from 300 to 500 V and three different current densities. The resulting test matrix contained forty-six unique conditions ranging from 5 to 102 kW total power and 16 to 247 A discharge current. At each operating condition, thrust, specific impulse, and efficiency were characterized. All seven channel combinations showed similar performance at a given discharge voltage and current density. The largest thrust recorded was 5.4±0.1 N at 99 kW/400 V discharge voltage. Total efficiency and specific impulse ranged from 0.54 to 0.67±0.03 and from 1800 to 2650 s ±60 s, respectively. Discharge current oscillations were also characterized with peak-to-peak values and with high-speed camera analysis, which provide insight into how the discharge channels oscillate and how those oscillations are affected by the presence of other operating channels. These results are discussed in the context of differences between single- and multichannel operation, as are the implications for the general viability of nested Hall thruster technology for future mission applications.
Journal of Propulsion and Power pp 1-9; https://doi.org/10.2514/1.b38346

Abstract:
A reversed gas-feed configuration for Hall thrusters is proposed and studied numerically. As an alternative to standard direct injection, we investigate the effect of injecting the propellant near the channel exit and directing it backwards, towards the anode. The resulting neutral density and average velocity fields are studied with the direct simulation Monte Carlo method for both cold and warm anode conditions. The residence time of neutral particles inside the channel and in the ionization region is computed using a test-particle Monte Carlo method. The computations indicate that the reversed injection allows for increasing the mass utilization efficiency of a standard feed configuration from 2–5% to a maximum of 20–30% depending on the initial efficiency.
Chihiro Inoue, Yuki Oishi, Yu Daimon, Go Fujii, Kaname Kawatsu
Journal of Propulsion and Power pp 1-8; https://doi.org/10.2514/1.b38310

Abstract:
We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination, and nozzle expansion ratio. The theoretical formulation deduces a framework for a quantitative noncombustion test or cold-flow test to predict the performance indices. We simply employ water and dyed water as simulant liquids and then measure the local ratios of mixture and flow rate using an absorbance spectrometer. Density ratio mismatch between hypergolic propellants and water can be reasonably convertible. Combined with chemical equilibrium analysis, we obtain characteristic velocity and specific impulse across wide injection mixture ratios. The validity of the quantitative water-flow diagnostic is evidenced by comparing the results with those of corresponding combustion tests using nitrogen tetroxide and monomethylhydrazine as the propellants under several injector unlike-doublet and triplet configurations, showing that the mixing states of bipropellant thrusters under combustion can be reproduced using the water-flow diagnostic.
Antonio Rubino, Piero Colonna, Matteo Pini
Journal of Propulsion and Power pp 1-9; https://doi.org/10.2514/1.b37920

Abstract:
The lack of established optimal design guidelines for turbomachinery operating in the nonideal flow regime (e.g., organic Rankine cycle turbines, CO2 compressors, compressors for refrigeration systems) demands for effective and efficient automated design methods. Past research work focused on gradient-free methods applied to computational fluid-dynamic simulations. The application of the adjoint method is a cost-effective alternative as it enables gradient-based optimization irrespective of the number of design variables. This paper presents the application of a fully turbulent unsteady adjoint method for the automated design of multirow turbomachinery partly operating in the nonideal flow regime. The method therefore allows for the solution of constrained unsteady fluid-dynamic optimization problems, in which the thermodynamic properties of the working fluid need to be modeled by means of complex equations of state. The optimal designs computed with unsteady simulations obtained with the harmonic balance method are then compared with optimal design resulting from mixing-plane simulations. The method is applied to the optimization of 1) a two-dimensional turbine cascade subject to time-varying inlet conditions, and 2) a two-dimensional turbine stage of an organic Rankine cycle power system. The results demonstrate the importance of computing fluid properties using accurate thermodynamic models and of using unsteady simulations for shape optimization of these machines.
Prashanth Bangalore Venkatesh, Lars Osborne, Michael Fitzpatrick, Mesa Hollinbeck, Daudi Barnes
Journal of Propulsion and Power pp 1-10; https://doi.org/10.2514/1.b38326

Abstract:
This paper describes the testing and pulse performance analysis of a 45 N thruster, designated A45, using a 19.78% monomethylhydrazine–80.22% hydrazine fuel blend with mixed oxides of nitrogen containing 3% nitric oxide as the oxidizer. The thruster components are fabricated using additive manufacturing (AM), which enabled fine design features on injectors to achieve fast response times. The main focus of this work is to demonstrate precise and repeatable pulse performance and combustion stability and to compare thruster performance with theoretical predictions from NASA CEA across the chosen test matrix. As part of the pulse performance analysis, impulse produced by each pulse of a given 50 Hz pulse train is characterized, thereby illustrating that true pulse repeatability on such thrusters only starts with the third pulse. This is caused by the priming of the injector manifold at the start of a pulse train. Specific impulse as functions of pulse ON and OFF times are also characterized for a 50 Hz duty cycle and dry injector manifold operation. This method of per-pulse performance characterization aids guidance, navigation, and control teams in thrust/impulse control for attitude control. Overall, the specific impulse, characteristic velocity efficiency, and response times demonstrated by this additively manufactured thruster are similar to those fabricated by traditional methods and support wider use of AM in the aerospace industry.
Giuseppe Gallo, Stefano Mungiguerra, Raffaele Savino
Journal of Propulsion and Power pp 1-17; https://doi.org/10.2514/1.b38333

Abstract:
This work presents a novel approach for the modeling of the entrainment in the numerical simulation of the internal ballistics of hybrid rocket engines with paraffin-based fuels. This model, coupled with a more sophisticated gas–surface interaction treatment, is an improvement of the model previously developed by some of the authors, which was based on some oversimplifying assumptions. Indeed, the old entrainment model was performed in a close range of averaged oxidizer mass flux, and it made the overall numerical model not scalable on different motor sizes. Therefore, a new correlation is introduced, which is based on the Reynolds number and takes into account the dependence of the entrained fraction on the shear stress exerted by the gas flow and the tube diameter. Firstly, the new and the old numerical models are compared in order to highlight the improvement obtained by the current efforts. Then, the model has been validated on experimental tests involving two different thrust class motors. Finally, the effect of the motor size on the fuel consumption is shown, thus revealing the crucial influence of the recirculating zone extension due to the oxidizer axial injection.
Çetin Ozan Alanyalioğlu, Yusuf Özyörük
Journal of Propulsion and Power, Volume 37, pp 528-543; https://doi.org/10.2514/1.b37839

Abstract:
Because of its excellent insulation capability, the usage of a silica-phenolic charring ablator as a nozzle liner is a common practice in the solid rocket motor industry. During the design of a solid rocket motor employing a silica-phenolic nozzle liner, it is desired to conduct an accurate analysis yielding in-depth thermal response and recession characteristics. As the interior ballistics and nozzle recession rate mutually interact, the best practice is to perform a coupled solution to both. Commonly used one-dimensional analysis tools with empirical approaches for estimation of convective heat transfer rate and blowing effect generally lack sought accuracy and do not model the transient shape-change phenomenon, which affects the nozzle performance. This Paper considers governing equations for charring, including pyrolysis gas injection and surface energy balance for melting ablation, along with a boundary condition governed by interior ballistics, and demonstrates a framework in which these equations are solved with governing equations for the nozzle flowfield in a coupled manner. Development and validation of a one-dimensional material response solver based on the same governing equations is also demonstrated. Also, results from a static firing test conducted with a small-scale ballistic evaluation motor employing a silica-phenolic nozzle insert are provided. Results from both investigations are compared and discussed.
Mahadevan Krishnan, Kent Frankovich, Jonathan A. Mackey
Journal of Propulsion and Power, Volume 37, pp 577-583; https://doi.org/10.2514/1.b38191

Abstract:
A torsional thrust stand, calibrated for impulse bits in the range of 0.1–0.5 mN⋅s, was used to measure impulse bits from a metal plasma thruster. Impulse data were obtained on roughly 1400 shots, with metal plasma thruster targets of molybdenum, niobium, palladium, aluminum, and carbon. Model predictions (based on a simple circuit model and published plasma parameters) were validated by data from the calibrated torsional thrust stand. Over a typical 50-shot firing sequence, the impulse bits measured showed a coefficient of variation of approximately 5%. This implies that individual impulse bits of about 0.4 mN⋅s can be imparted to a satellite with ±5% variation about the mean in a total impulse burst of 20 mN⋅s. For a 20 kg satellite, this would result in a velocity increment ΔV of just 1 mm/s, enabling very precise attitude control and fine positioning control of nano- and microsatellites. The metal plasma thruster uses solid metal propellant and hence requires no liquids, gases, flow valves, or flow controls and has no moving parts. Total impulse approximately 5000 (N⋅s)/liter provides orbit raising and drag compensation capability.
Ilaria De Dominicis, Sebastian Robens, Nina Wolfrum, Martin Lange, Volker Gümmer
Journal of Propulsion and Power, Volume 37, pp 615-624; https://doi.org/10.2514/1.b37909

Abstract:
This paper describes the analysis of the flowfield in a four-stage low-speed axial compressor at the design point investigated both numerically and experimentally, with particular emphasis on the impact of stator hub flow leakage in a section of the machine featuring a change of stator configuration. The goal of the work is to give insight into the effect of the stator hub configuration on the aerodynamic performance in terms of loss and efficiency of downstream blade rows as well as the resulting stage matching effects. The investigation is focused on aerodynamic behavior and stage interaction, which the stator hub configuration of a given stage induces in a downstream stage with respect to its performance and flowfield. Two shrouded and one cantilevered stator hub configurations, such as those commonly employed in industry, are investigated and discussed with respect to local flow phenomena and stagewise aerodynamic effects. The results show that the stator hub configuration of a given stage has a drastically high impact on the downstream rotor aerodynamic performance, whereas the downstream stator as well as the upstream rotor are only slightly influenced by the stator hub configuration.
Jacob Simmonds, Yevgeny Raitses
Journal of Propulsion and Power, Volume 37, pp 544-552; https://doi.org/10.2514/1.b37935

Abstract:
Conventional expressions and definitions describing performance of plasma thrusters, including the thrust, specific impulse, and the thruster efficiency, assume a steady-state plasma flow with a constant flow velocity. However, it is very common for these thrusters that the plasma exhibits unstable behavior resulting in time variations of the thrust and the exhaust velocity. For example, in Hall thrusters, the ionization instability leads to strong oscillations of the discharge current (so-called breathing oscillations), plasma density, ion energy, and as a result the ion flow. This paper revisits the formulation of the thrust and the thrust efficiency to account for time variations of the ion parameters, including the phase shift between the ion energy and the ion flow. For sinusoidal oscillations it was found that thrust can potentially change more than 20%. It is shown that, by modulating ion energy at specific amplitudes, thrust can be maximized in such regimes. Finally, an expression for the thruster efficiency of the modulating thruster is derived to show a mechanism for inefficiencies in such thrusters.
Mohammad A. Hossain, Ali Ameri, Jeffrey Bons
Journal of Propulsion and Power, Volume 37, pp 604-614; https://doi.org/10.2514/1.b37980

Abstract:
The Paper involves experimental and numerical heat transfer analyses of four pin-fin cooling configurations in a rectangular channel. Two design concepts were studied with two separate pin-fin (cylindrical and triangular) geometries. The first design includes a single-wall pin-fin configuration where the coolant flows through an array of full-length cylindrical (PF) and triangular pin fins (TF). The second configuration involves a double-wall pin-fin–jet configuration where the coolant jet impinges on the target wall, then flows through the array of partial-length cylindrical pin fins with jets (PFC) and partial-length triangular pin fins with jets (TFC). IR thermography was used to measure the steady state wall temperature of the target surface to estimate the overall cooling effectiveness ϕ. A transient heat transfer experiment was conducted to evaluate the internal heat transfer coefficient from the transient wall temperature. It is found that the double-wall pin-fin–jet (PFC and TFC) configurations show higher cooling effectiveness compared to the single-wall pin-fin (PF and TF) designs and the triangular pin fin (TF and TFC) showed a higher heat transfer augmentation compared to the cylindrical pin fin (PF and PFC) design. Numerical study later confirmed that the triangular pin fin creates a pair of streamwise vortices that augments the local heat transfer. The presence of the pin fin in the double-wall configuration prevents the development of a strong crossflow that negatively affects the heat transfer of the impinging jet. In addition, the triangular pin fin not only prevents the strong crossflow but redirects the coolant in the lateral direction that eventually contributes to the formation of a counterrotating vortex pair in the streamwise direction and thus augments local heat transfer and improves overall cooling effectiveness. Pressure drop measurement showed that the full-length triangular pin fin has a higher-pressure loss but the partial-length pin fin can lower the pressure loss considerably without compromising additional cooling performance.
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