Mihnea Gall

Pressure-gain combustion (PGC) represents a promising alternative to conventional propulsion systems for interplanetary travel due to its key advantages, including higher thermodynamic efficiency, increased specific impulse, and more compact engine designs. However, to elevate this technology to a sufficient technology readiness level (TRL) for practical application, extensive experimental validation, particularly under vacuum conditions, is essential. This study focuses on the performance of a pulsed-detonation combustor (PDC) under near-vacuum conditions, with two primary objectives: to assess the combustor’s ignition capabilities and to characterize the shock wave behavior at the exit plane. To achieve these objectives, high-frequency pressure sensors are strategically positioned within both the vacuum chamber and the combustor prototype to capture the pressure cycles during operation, providing insights into pressure augmentation over a period of approximately 0.5 s. Additionally, the Schlieren visualization technique is employed to analyze and interpret the flow structures of the exhaust jet. The combination of these experimental methods enables a comprehensive understanding of the ignition dynamics and the development of shock waves, contributing valuable data to advance PGC technology for space-exploration applications.

Turbomachinery shock wave patterns occur as a natural result of operating at off-design points and are accountable for some of the loss in performance. In some cases, shock wave–boundary layer (SW-BLIs) interactions may even lead to map restrictions. The current paper refers to experimental findings on a transonic linear cascade specifically designed to mitigate shock waves using porous walls on the blades. Schlieren visualization reveals two phenomena: Firstly, the shock waves were dissipated in all bladed passages, as predicted by the CFD studies. Secondly, a lower-pressure wave pattern was observed upstream of the blades. It is this phenomenon that the paper reports and attempts to describe. Attempts to replicate this pattern using Reynolds-averaged Navier–Stokes (RANS) calculations indicate that the numerical method may be too dissipative to accurately capture it. The experimental campaign demonstrated a 4% increase in flow rate, accompanied by minimal variations in pressure and temperature, highlighting the potential of this approach for enhancing turbomachinery performance.

To alleviate the shock boundary layer interaction adverse effects, various active or passive flow control strategies have been investigated in the literature. This research sheds light on the behavior of perforated plates as passive flow control techniques applied to NACA0012 airfoils in cascade configurations. Two identical perforated plates with shallow cavities underneath are accommodated on the upper and lower surfaces of each airfoil in the cascade arrangement. Six different cascade arrangements, including a baseline configuration with no control applied, are additively manufactured, with different perforated plate orifice sizes in the range of 0.5–1.2 mm. A high-speed wind tunnel with Schlieren optical diagnosis and wall static pressure taps is used to investigate the changes in the shock waves pattern triggered by the perforated plates. Steady 3D density-based numerical simulations in Ansys FLUENT are conducted for further analysis and validation. In the cascade configuration, the perforated plates alter the shock structure, and the strong normal shock wave is replaced by a weaker X-type shock structure. Eventually, a 1% penalty in overall total pressure loss is induced by the perforated plates because of the negative loss balance between the reduced shock losses and the enhanced viscous losses. Further studies on perforated plate geometrical features are needed to improve this outcome in a cascade arrangement.

Combustion through detonation marks an important leap in efficiency over standard deflagration methods. This research introduces a Pulsed Detonation Combustor (PDC) model that uses Hydrogen as fuel and Oxygen as an oxidizer, specifically targeting carbon-free combustion efforts. The PDC aerodynamic features boost operating cycle frequency and facilitate Deflagration-to-Detonation Transition (DDT) within distances less than 200 mm by means of Hartmann–Sprenger resonators and cross-flow fuel/oxidizer injection. The achievement of quality mixing in a short-time filling process represents not only higher cycle operation but also enhanced performances. The scope of this paper is to assess the impact of different fuel injectors with different opening areas on the performances of the PDC. This assessment, expressed as a function of the Equivalence Ratio (ER), is conducted using two primary methods. Instantaneous static pressures are recorded and processed to extract the maximum and average cycle pressure and characterize the pressure augmentation. Thrust measurements obtained using a load cell are averaged over the detonation cycle to calculate the time-averaged thrust. The specific impulse is subsequently determined based on these thrust measurements and the corresponding mass flow data.

Understanding shock wave behavior in supersonic flow environments is critical for optimizing the aerodynamic performance of turbomachinery components. This study introduces a novel transonic linear cascade design, focusing on advanced blade manufacturing and experimental validation. Blades were 3D-printed using Inconel 625, enabling tight control over the geometry and surface quality, which were verified through extensive dimensional accuracy assessments and surface finish quality checks using coordinate measuring machines (CMMs). Numerical simulations were performed using Ansys CFX with an implicit pressure-based solver and high-order numerical schemes to accurately model the shock wave phenomena. To validate the simulations, experimental tests were conducted using Schlieren visualization, ensuring high fidelity in capturing the shock wave dynamics. A custom-designed test rig was commissioned to replicate the specific requirements of the cascade, enabling stable and repeatable testing conditions. Experiments were conducted at three different inlet pressures (0.7-bar, 0.8-bar, and 0.9-bar gauges) at a constant temperature of 21 °C. Results indicated that the shock wave intensity and position are highly sensitive to the inlet pressure, with higher pressures producing more intense and extensive shock waves. While the numerical simulations aligned broadly with the experimental observations, discrepancies at finer flow scales suggest the need for the further refinement of the computational models to capture detailed flow phenomena accurately.

Detonation combustion represents a significant advancement in efficiency over traditional deflagration methods. This paper presents a Pulsed Detonation Combustor (PDC) model that is designed with an aerodynamic mixing chamber featuring Hartmann–Sprenger resonators and crossflow injection. This design enhances operational cycle frequency and enables sustained detonation over short distances (below 200 mm). The PDC’s performance was evaluated through a comprehensive full-factorial experimental campaign, incorporating four factors with four discrete levels each. Testing was conducted using both hydrogen/air and hydrogen/oxygen mixtures, highlighting the PDC’s potential as a carbon-free combustion chamber suitable for both air-breathing and space-based propulsion systems. One advantage is the versatility of our PDC breadboard, which lies in its applicability to both terrestrial and in-space applications, such as interplanetary travel or trajectory corrections. Thrust measurements were recorded using a load cell and time-averaged thrust levels were determined over the detonation cycle and are reported herein, together with the specific impulse. The results underscore the PDC’s promise as an efficient propulsion technology for future aerospace applications.

A pulsed detonation chamber (PDC) equipped with Hartmann–Sprenger resonators has been designed and tested for both Hydrogen/air and Hydrogen/Oxygen mixtures. A full-factorial experimental campaign employing four factors with four levels each has been carried out for both mixtures. Instantaneous static pressure has been measured at two locations on the exhaust pipe of the PDC, and the signal has been processed to extract the average and maximum cycle pressures and the operating frequency of the spark plug. The PDC has been shown to be able to reach sustained detonation cycles over a length below 200 mm, measured from the spark plug to the first pressure sensor. The optimal regimes for both air and Oxygen operation have been determined, and the influence of the four factors on the responses is discussed.

This research investigated a passive flow control technique to mitigate the adverse effects of shock wave–boundary layer interaction on a NACA 0012 airfoil. A perforated plate with a strategically positioned cavity beneath the shock wave anchoring spot was employed. Airfoils with perforated plates of varying orifice sizes (ranging from 0.5 to 1.2 mm) were constructed using various manufacturing techniques. Experimental analysis utilized an “Eiffel”-type open wind tunnel and a Z-type Schlieren system for flow visualization, along with static pressure measurements obtained from the bottom wall. Empirical observations were compared with steady 3D density-based numerical simulations conducted in Ansys FLUENT for comprehensive analysis and validation. The implementation of the perforated plate induced a significant alteration in shock structure, transforming it from a strong normal shock wave into a large lambda-type shock. The passive control case exhibited a 0.2% improvement in total pressure loss and attributed to the perforated plate’s capability to diminish the intensity of the shock wave anchored above. Significant fluctuations in shear stress were introduced by the perforated plate, with lower stress observed in the plate area due to flow detachment from cavity blowing. Balancing shock and viscous losses proved crucial for achieving a favorable outcome with this passive flow control method.

In connection with subsonic jet noise production, especially regarding the hot jet from a micro turbojet engine, we encountered a lack of recent high-resolution data in the literature describing the flow field using experimental validation through optical diagnoses. The objective of this paper is to examine and compare the influence on shear layers of the exhaust plug nozzle of a micro turbojet engine with and without chevrons mounted, using a high-speed camera used in Schlieren-type optical system diagnosis. Three different operating regimes are examined for both the baseline configuration and the configuration with 16 triangular-shaped chevrons. In conjunction with the image captures, the sound pressure level was recorded with the help of a microphone placed perpendicular to the flow, 0.4 m from the exhaust of the nozzle which was further processed. In quantitative terms, we found that the OASPL decreases by more than 1% when the engine is operating at higher regimes. Moreover, we found that the average exhaust jet angle, which is a measure of the quality of the fluid mixing layer is increased by 5% with respect to the baseline nozzle. By using the “darkest pixel” technique in Schlieren imaging, we can verify experimentally, for all working regimes, the theory that asserts that subsonic jet noise is a consequence of fine-scale homogeneous turbulence. Additionally, the potential novelty lies in the specific observations related to consistent dispersion of fine-scale eddies and how the presence of chevrons amplifies this uniformity within the turbulent field.

This paper presents the development of high-specific-speed mixed flow/centrifugal compressor vaned diffusers. Specifically, the design of a test rig that will make the visualization of shock waves on diffuser vanes manageable is addressed in the current study. In this particular case, linearization of an existing state-of-the-art compressor stage was used. For the computational modeling, a series of RANS analyses were conducted to examine the flow characteristics of the two cases explored: a complete transonic cascade and an idealized periodic passage. The distinct behavior exhibited by each vane passage within the entire cascade offers the opportunity to analyze the shockwave structures across a mass flow range of ±9% around the design point. Overall, the pressure coefficient distributions and flow field patterns appear to align with the single-passage conditions, although there are some minor lateral wall influences, particularly in the first passage close to the suction lateral wall. However, because of the nature of the flow, which is characterized by high velocity and density differences near the vanes, the equivalent mass flow per individual passage was difficult to estimate. This may also be attributed to the small endwall axial vortices; nonetheless, for the purposes of this paper, this was of little consequence.