Adam Kozakiewicz

Preliminary aerodynamic assessment provides an efficient means of investigating internal flow organization in aircraft gas turbine combustors before more computationally demanding three-dimensional and reactive-flow simulations are undertaken. This study presents a preliminary CFD investigation of cold-flow phenomena in the annular combustion chamber of the F100-PW-229 low-bypass turbofan engine. A two-dimensional planar model was intentionally adopted to identify dominant flow structures and assess their response to representative engine operating conditions at reduced computational cost. Numerical simulations were performed in ANSYS Fluent using a pressure-based coupled solver and the k-ω SST turbulence model for idle, cruise, and maximum rotational speed conditions. Additional simulations were performed at cruise conditions for an altitude of 11 km to examine the influence of reduced ambient pressure and air density on the internal flow field. The results show that increasing engine rotational speed primarily changes the intensity of the aerodynamic field while preserving its dominant spatial organization. The maximum velocity at idle was approximately 55% lower than that at maximum rotational speed, while the corresponding value at cruise was 20.1% lower. Similarly, the maximum stagnation pressure at idle was approximately 67.1% lower than that at maximum rotational speed, while the difference between cruise and maximum speed was 24.2%. A distinct vortex structure was identified in the outer annular passage between the liner and casing, with its location shifting downstream as the operating condition changed. At 11 km altitude, the calculated pressure level was approximately 81% lower and the characteristic flow velocity approximately 9% lower than under corresponding ground-level conditions, while the dominant flow topology remained similar. These results highlight the role of combustor geometry and secondary-air distribution in governing global pressure redistribution, velocity development, and recirculation behavior. Within the scope of the adopted two-dimensional non-reacting formulation, the obtained flow patterns provide an initial aerodynamic reference for identifying regions of interest for subsequent high-fidelity simulations. The proposed approach is therefore intended as a computationally efficient preliminary engineering tool rather than a substitute for validated three-dimensional combustor modelling.

In this paper, we consider the solution of the initial-value problem in thermodiffusion in a solid body in three-dimensional space. Not only do we prove the behavior of the solution over time, but we identify some of its technical aspects as well. Stabilizing a thermodiffusion system in solids is essential for understanding the long-time behaviour of some materials, which are used in mechanical engineering; this is especially important for materials used in aviation, not only in civil aviation but in the air force as well.

The aim of the article is to review and analyze existing methods and design solutions of high-pressure turbine cooling systems used in fourth-generation engines. The need to increase the temperature of the working medium causes the development of cooling systems, which also targets minimizing the negative impact of cooling on turbine efficiency. The paper presents the idea of cooling and the most important heat transfer models connected to turbine blades. Attention was focused on analyzing two types of cooling, i.e., internal and external. In the case of internal cooling, the problem of directing the flow of the cooling stream was outlined. It was pointed out that it plays a major role in the area of heat transfer efficiency, which depends on the shape of the flow channel, as well as the Reynolds number for the cooling air. One of the latest solutions used for cooling high-pressure turbine blades, i.e., a system of Dimple Induced Secondary Flow, was presented. In the case of external cooling, attention was focused on the elements affecting the effectiveness of this solution, i.e. the geometry, arrangement, and also the angle of the cooling holes. Results on the comparative analysis of heat transfer coefficients for different configurations of cooling systems are summarized.

This study examines scientific and technical solutions designed to enhance thermodynamic processes in modern aircraft turbine engines by utilizing heat exchangers. A comprehensive literature review informed the development of a conceptual design for a turbofan engine incorporating both an intercooler and a recuperator. The research included an original parametric and constrained optimization analysis conducted for two engine configurations as follows: one intended for narrow-body and the other for wide-body aircraft. The study focused on achieving the required thrust while enhancing efficiency. Results indicate that integrating heat exchangers can significantly reduce specific fuel consumption (SFC) and/or increase engine power or thrust. Moreover, the recovery of residual heat from exhaust gases through recuperation contributes to improved overall energy efficiency. The study also explores a novel cryogenic design that utilizes liquid hydrogen for cooling the intercooler, recuperator, and turbine. Although not modeled directly, this concept demonstrates the potential to increase the bypass ratio, further reduce SFC, and lower NO_x emissions. These findings highlight the promise of combined intercooling and recuperation strategies for improving both economic and environmental performance, with optimal system parameters dependent on aircraft class. The research aligns with ongoing efforts in mechanical engineering and aviation to enhance turbine engine efficiency through innovative thermal management solutions.

The article provides a detailed description of the design and key parameters of the GTM-120 turbine engine, along with flow calculations for this power unit at a selected operating point. These calculations were performed using GasTurb software, which is also capable of assessing nitrogen oxide emissions. Additionally, the article outlines the test stand and the methodology used to determine toxic compounds emitted by the engine. It analyzes the measured results of various exhaust components produced during the engine's operation, specifically focusing on the levels of nitrogen oxides and carbon oxides as influenced by rotational speed and temperature. Conducted studies have established a testing methodology for aviation turbine engines, based on the type of analyzer used. This methodology will facilitate the evaluation of other aviation turbine engines on stationary test stands.

This study investigates the aerodynamic performance of a fourth-generation normal shockwave inlet system, with a primary focus on minimizing pressure losses and ensuring uniform airflow distribution. A computational model was developed, incorporating a section of the fuselage along with the complete inlet duct. The model was discretized using a hybrid mesh approach to enhance numerical accuracy. The analysis was conducted at a flight altitude of 8000 m, encompassing 370 distinct cases defined by varying angles of attack and Mach numbers. This comprehensive parametric study yielded a dataset of 10,800 total pressure measurements across predefined sampling locations. Based on the obtained results, flow distortion coefficients in both circumferential (CDI) and radial directions (RDI) were systematically determined for each test case. The interdependencies between CDI, RDI, Mach number, and angle of attack (α) were analyzed and presented in a consolidated manner. In the second phase of the study, an artificial neural network (ANN) utilizing a Feed-Forward architecture was implemented to predict pressure distributions for intermediate flight conditions. The ANN was trained using the CFG algorithm, and the predictive accuracy was assessed through the determination coefficients computed by comparing ANN-based estimates with numerical simulation results. The findings demonstrate the efficacy of ANN-based modeling in enhancing the predictive capabilities of inlet flow dynamics, offering valuable insights for optimizing next-generation supersonic air intake systems.

The paper presents the projected emissions of combustion components over the next few years. The basic tasks that a modern combustion chamber must fulfill were defined. The process of hydrocarbon combustion in the theoretical and actual cases was analyzed. The assessment evaluates the effect of engine operating parameters such as rotational speed, thrust, temperature downstream the compressor and combustion on the formation of toxic combustion components. The paper also presents alternative fuels. One of the most promising aspects related to the matter is the use of sustainable aviation fuels - SAF. Alternative methods of powering aircraft engines, such as hydrogen or nuclear propulsion, were presented. An analysis on the latest combustion chamber design systems that allow to reduce the amount of exhaust gasses emitted into the atmosphere was conducted.

One of the biggest challenges of modern aviation is the development of technologies that reduce or eliminate emissions of harmful combustion components into the atmosphere. European authorities are imposing increasingly stringent emissions regulations. Therefore, new models of combustion chambers, new combustion methods, as well as new types of aviation fuels must be developed. This article presents the possibilities of using hydrogen propulsion in aviation. The reasons for conducting research on hydrogen propulsion are discussed, as well as the history of the introduction of hydrogen propulsion into aircraft engines. Problems that can be encountered in the production and storage of hydrogen are identified and explained. Proposals for the use of hydrogen combustion or the use of fuel cells to power turbine engines are also presented, and the economic aspect of this type of fuel is discussed.

The following paper evaluates the strength properties of the critical area of the turbine rotor stage of a jet engine. This is where the blade is attached to the rotor disk. Important aspects of this issue are characteristic geometrical dimensions and operating factors, e.g., changes in rotor speed. In the article, the strength properties of the rotor stage with a trapezoidal blade–disk connection were analyzed using parametric analysis considering fluid–structure interaction (FSI). Virtual models of the compressor rotor stages were developed, including parametric blade–disk connections and discrete models for FEM (finite element method) and CFD (computational fluid dynamics) analyses. The parametric analysis was aimed at assessing selected geometric parameters in terms of their impact on the maximum stresses in the blade root. The analysis performed shows that by changing the blade’s width and thickness angles at the root, we can influence the value of the maximum stresses in the rotor stage. It was noticed that some combinations of these two parameters may negatively affect the value of the stresses in the rim.