Tunable diode laser absorption spectroscopy (TDLAS) sensors for simultaneous detection of temperature, carbon monoxide (CO), and carbon dioxide (CO2) concentrations have been developed and demonstrated at the exit of a scramjet combustor in the ACT-II facility. Short scramjet combustor flow-through times, combined with elevated temperatures characteristic of the high enthalpy flight environment, create unfavorable conditions for complete reaction progress to combustion products. The implementation of these sensors addressed these challenges by quantitatively characterizing combustion progress. Different fueling schemes and equivalence ratios were tested to understand the effects of injection location and core flow penetration on combustion performance. Along with laser absorption measurements, simultaneous static pressure and high-speed broadband imaging were used for more comprehensive characterization. Results indicated that more robust combustion was achievable through near-cavity fueling as opposed to upstream injection. Greater differences in CO and CO2-inferred temperatures existed in cases of nonuniform measurement conditions.
The increasing complexity of electrified propulsion systems for aviation requires scalable and generalizable models to support design activities exploring the potential of new enabling technologies. Traditional computational frameworks rely on predefined architectures and hybridization factors, limiting flexibility when introducing novel powertrain configurations. This paper proposes a computational framework for modeling propulsion systems based on graph theory, offering a topological representation capable of describing virtually any acyclic powertrain. In the proposed approach, system components are represented as vertices connected by directed edges that define power flows. The resulting acyclic directed graphs enable consistent and modular descriptions of systems with multiple energy sources, including fuel, battery, and hydrogen-based configurations. Thus, the computational framework defined in this work can be applied to any propulsion system presented in the literature without major changes. Furthermore, new parameters are introduced to represent the contribution of the power sources to each propulsive line, which provides the designer with greater control over the distribution to explore alternatives in the power distribution that have been neglected in previous works. Finally, the method is not prone to singularities common in conventional formulations.
In highly loaded transonic compressors, strong pulsation appears in the flowfield before stall onset. However, the connection between strong pulsation and stall triggering remains unclear. This study adjusts the pulsation intensity before stall in a transonic compressor by modifying the axial spacing between the inlet guide vane and the downstream rotor and investigates its impact on the compressor’s stability. Numerical simulations are conducted to obtain the stability boundaries and stall evolution processes for two configurations with different interaction intensities. The results show that strong pulsation before stall can alter the compressor’s stable operating range. During the stall evolution process, the amplitude of the unsteadiness composed of components with frequencies below 0.5 times the blade pass frequency gradually increases. This corresponds to the decrease in the oscillation frequency and the increase in the oscillation range of the suction-side leading-edge separation bubble at the blade tip, ultimately triggering stall. The difference in the stability boundaries of the two configurations is attributed to the pressure pulsations induced by inlet-guide-vane–rotor interaction, which modify the development of low-frequency oscillations of the blade-tip suction-side separation bubble.
This article describes a framework to scale hydrogen fully premixed swirled flames between laboratory experiment and turbofan’s combustor. The framework combines three major elements: 1) turbomachinery calculations, 2) combustor performance maps, and 3) flame stabilization and matching conditions for scaling. First, the operating point of the engine is determined with a thermodynamics-based analysis applied to all components of the engine: fan, compressor, combustor, and turbine. Second, the method to obtain and compute the combustor’s operating point in the fully premixed mode, coined the combustor performance map, is applied. The method includes six requirements: the thermal power for the engine, mass, momentum, energy conservation budgets in the combustor, a revisited compressor/combustor/turbine power balance for fully premixed combustion, and a coupling relationship. The results are synthesized as combustor performance maps that are functions of flame equivalence ratio [Formula: see text] and injector flow split [Formula: see text]. The flame stagnation temperature, [Formula: see text] species mass fraction, and emission index are reported on these maps. The flow split and flame equivalence ratio of the operating point are deduced from these requirements. Third, a theoretical method to scale lean fully premixed swirled flames between the laboratory experiment and the jet engine at the determined conditions is described based on the criterion for stabilization.
This study investigates the aeroelastic characteristics of a transonic linear cascade using an experimental approach to understand the influence of compressibility, incidence angle, interblade phase angle, and reduced frequency on flutter behavior. Vibration-related failure remains a major concern in modern turbomachinery, as thin and lightweight blades operating in transonic flow regimes are highly susceptible to aeroelastic instabilities. Controlled torsional oscillations are imposed on the blades ([Formula: see text] and [Formula: see text]) adjacent to the stationary central blade [Formula: see text] to study the effect of phase and frequency variation, while the central blade, instrumented with a fast-response load cell, measures the unsteady aerodynamic forces induced by the neighboring blade motion. The results indicate that phase variation can significantly increase instability when unsteady aerodynamic forces synchronize with blade motion, leading to reduced aerodynamic stability. In the transonic regime, passage shock further triggers the onset of choke-type flutter at low incidence angles and specific phase conditions. At a very high incidence, stall flutter helps the stability due to the interaction of upstream shock motion and the separation layer. However, frequency variation modifies temporal synchronization between blade passages; unlike phase variation, it does not influence blade stability significantly. The experimental results demonstrate that variations in incidence, reduced frequency, and phase synchronization determine the stability limits and flutter characteristics of turbomachinery blade rows.
