Engine Thrust Research Articles

Measurement report: In-flight and ground-based measurements of nitrogen oxide emissions from latest-generation jet engines and 100 % sustainable aviation fuel

Abstract. Nitrogen oxides, emitted from air traffic, are of concern due to their impact on climate by changing atmospheric ozone and methane levels. Using the DLR research aircraft Falcon, total reactive nitrogen (NOy) in-flight measurements were carried out at high altitudes to characterize emissions in the fresh aircraft exhaust from the latest-generation Rolls-Royce Trent XWB-84 engine aboard the long-range Airbus A350-941 aircraft during the ECLIF3 (Emission and CLimate Impact of alternative Fuels 3) experiment. The impact of different engine thrust settings, monitored in terms of combustor inlet temperature, pressure and engine fuel flow, was tested for two different fuel types: Jet A-1 and, for the first time, a 100 % sustainable aviation fuel (SAF) under similar atmospheric conditions. In addition, a range of combustor temperatures and an additional blended SAF were tested during ground-based emission measurements. For the data measured during ECLIF3, we confirm that the NOx emission index increases with increasing combustion temperature, pressure and fuel flow. We find that as expected, the fuel type has no measurable effect on the NOx emission index. These measurements are used to compare to cruise NOx emission index estimates from three engine emission prediction methods. Our measurements thus help to understand the ground to cruise correlation of current engine emission prediction methods while serving as input for climate modelling and extending the extremely sparse data set on in-flight aircraft nitrogen oxide emissions to newer engine generations.

Development and research testing of a low thrust on gaseous-propellant rocket engine chamber

The article presents preliminary results of the development and research testing of a rocket engine chamber with a thrust of 100 N operating on fuel components: gaseous oxygen-gaseous hydrogen, designed to be used as the main engine of a small upper stage for launching a payload weighing up to 150 kg into target orbits. The main features of the engine chamber are its fabrication by selective laser melting from 12X18N10T steel powder and its regenerative cooling with gaseous hydrogen. Calculation and experiments confirmed the possibility of regenerative cooling of the chamber in the nominal mode. In addition, the current results of the development of a technique for optical recording of the process of structural material removal from the chamber during tests using different optical filters are presented. It is shown that there is a correlation between the brightness of the obtained frames and the hydrogen flow rate. It is also shown that the afterburning of the removed material particles, in contrast to high thrust engines, occurs mainly in the tail of the jet.

Improved modelling of low-pressure rotor speed in commercial turbofan engines: A comprehensive analysis of machine learning approaches

Accurate engine thrust modelling offers notable opportunities for reducing fuel consumption, mitigating emission effects, managing air traffic, and optimizing the preliminary design of aircraft. This study focuses on the development of machine learning models to predict the low-pressure rotor speed (N1) parameter, which is closely related to the thrust of turbofan engines, for commercial aircraft during the cruise phase. The analysis utilizes a Flight Data Recorder (FDR) dataset from 1086 actual flights involving twin-engine, narrow-body, short-to-medium haul commercial aircraft equipped with CFM56-7B turbofan engines. To achieve accurate N1 parameter predictions, the study employs machine learning models, including Deep Learning (DL), Random Forest (RF), Gradient Boosted Machines (GBM), and Artificial Neural Networks (ANN). These models are evaluated using statistically significant indicators, demonstrating that machine learning models yield highly accurate results. Among the models tested, the DL model provides the most precise estimates of the N1 parameter. This study not only advances the understanding of engine thrust modelling through machine learning but also provides practical insights for the aerospace industry. The findings underline the potential of machine learning techniques in delivering superior prediction accuracy, which can be integrated into real-time flight management systems, ultimately contributing to more sustainable and efficient aviation operations.

Combinatorial Method for Quality Improvement of the Thrust Plate – A Case Study

Aim: This research aims to mitigate defects in the turning operation of thrust plates used in fighter jet fuel tank assemblies, thereby reducing the rejection rate and improving overall quality. This aligns with the aerospace industry's reliability goals. Background: The thrust plate is a critical component in fighter jet fuel tank assembly, transmitting engine thrust to the airframe. Quality compromises in this component can impair jet performance. It was observed that the thrust plate had a rejection rate of about 2.9% due to various defects. This real- world scenario underscores the importance of our study on the thrust plate and its potential impact on the aerospace industry. The rejection rate underscores its significance and potential for patent by quality improvement in turning of the thrust plate. Objective: The objective is to mitigate turning operation defects on the thrust plate to reduce rejection rates, aligning with aerospace industry reliability goals. Methods: Experimentation encompassed four pivotal factors: turning speed, feed rate, cutting depth, and tool inserts, implemented through Taguchi's Orthogonal Array technique. Grey Relational Analysis was utilized to optimize parameters in thrust plate turning. Specifically, this paper targeted the enhancement of its diameter, surface roughness, and tool life. Results: A single coefficient for the multiple responses, i.e., grey relational grade, has been determined, and optimum levels for the parameters have been identified. Confirmation experiments with the optimal factor level combination were carried out on a sample of thrust plates, and no rejections were observed. Conclusion: An experimental design based on Taguchi’s orthogonal array approach was used to conduct the experiments. The Grey Relational Analysis has been applied to analyze the experimental results and optimize the turning operation process parameters for the responses thrust plate diameter, tool life, and surface roughness. With this, the rejection of the thrust plate has been considerably reduced.

МЕТОДИКА ЕКСПЕРИМЕНТАЛЬНИХ ДОСЛІДЖЕНЬ СПРОЩЕНОЇ МОДЕЛІ БПЛА ТИПУ БІКОПТЕР

Modern aircraft are usually statically unstable objects. Due to this, increased maneuverability and controllability is achieved. One of the promising directions is the development and implementation of UAVs of the bicopter or convertible type. Their movement is provided by two screw electric drives. The non-identity of the engine parameters and the peculiarities of the movement of UAVs of this type cause special requirements for the control systems. To ensure sustainability, modeling and conducting research with the help of various tools is necessary. The most available models of unstable movements are pendulum devices of various types, which allow studying stabilization processes. The purpose of the work is to use a computer stabilization system of a simplified bicopter UAV model by changing the thrust of the engines. The tasks of the work are: development of a computer stabilization system for the angle of pitch or roll, creation and verification of stabilization algorithms on a laboratory bench using an ARDUINO controller, research of the functional capabilities of angular stabilization circuits. The object of the study: a computer control system of an aircraft that performs the task of stabilizing a UAV. Subject of research: processes of stabilization of the angular position of a simplified bicopter model. When performing the work, a control law using angular and angular velocity feedback was chosen. A variant of aircraft stabilization when using engine thrust control is proposed. The control law of the computer stabilization system, which is implemented on the ARDUINO controller, was selected and tested. The applied value of the work is the further use of the obtained results in modern samples of bicopters.

Numerical study on flow field characteristics and performance of a Mach 10 internal injection kerosene-fueled oblique detonation engine

Oblique detonation engines (ODE) have significant potential for hypersonic propulsion, yet there is a paucity of research investigating internal injection ODEs. In this study, a numerical simulation of the internal flow field of a Mach 10 internal injection kerosene-fueled ODE is conducted. The fuel mixing, pre-combustion, and combustor wave structure in the flow field are analyzed in a situation closer to the real flow field conditions. Further studies have demonstrated that alterations to the upper wall initiation position of nozzle can influence the separation zone, flow field stability and the engine performance. An upper wall initiation position that is too far forward will increase the separation zone area and reduce the engine thrust. Conversely, an upper wall initiation position that is too far back will lead to flow field destabilization and eventually thermal choking. Finally, the effects of increasing the equivalent ratio on the flow field structure and engine performance for a certain configuration are analyzed. The results demonstrate that when the equivalent ratio is elevated, an increase in either the bottom or top incoming flow equivalent ratio results in a transformation of the wave structure within the combustor due to the presence of the incoming boundary layer and the subsonic zone. A large-scale separation zone will form at the bottom of the combustor, resulting in a reduction in nozzle thrust. However, the wedge drag is reduced more, thereby increasing the engine's specific impulse.

Influence of laser absorptivity of CuCr0.8 substrate surface state on the characteristics of laser directed energy deposition inconel 718 single track

An effective method for fabricating copper-nickel bimetallic liquid rocket engine thrust chambers involves utilizing laser directed energy deposition (LDED) technology. However, the state of the substrate surface significantly impacts the LDED process. This study investigates the effects of various substrate treatments on LDED single tracks, using CuCr0.8 high-copper alloy as the substrate and Inconel 718 as the deposition material. The treatments include polishing, sandblasting, laser etching, and cold spraying. Substrate surface roughness, laser absorptivity, molten pool morphology, and microstructure were characterized, and the mechanisms of laser absorptivity change and the different LDED processes were analyzed. The results indicate that the laser-etched surface exhibits the worst surface roughness (Ra 15.20 ± 0.60 μm), the highest laser absorptivity(80.70% at 1080 nm wavelength), the largest deposition width (947.33 ± 29.85 μm), and the maximum number of fine grains among the four substrates. Additionally, the cold-sprayed surface shows the largest deposition depth (237.33 ± 39.04 μm), the minimum number of fine grains and a higher laser absorptivity (66.20% at 1080 nm wavelength). In situ observations of molten pool formation and flow during LDED was conducted using an in situ high-speed high-resolution imaging system. The mechanisms underlying the alteration in laser absorptivity primarily involve the "trapped light" effect and modifications to the surface material. This research is significant as it provides foundational insights for laser processing of highly reflective materials, offering important theoretical and practical implications for engineering applications.

Assessment of influencing factors on the flight performance of J-series fighter jets based on the entropy weight-TOPSIS model

To enhance the scientific approach to fighter jet design and improve performance in areas such as supersonic cruise, long-range strike, and high maneuverability, this paper first collects, organizes, and thoroughly analyzes relevant data on factors affecting the flight performance of existing J-series fighter jets, including indicators such as engine thrust, aspect ratio, root-to-tip ratio, and wing shape. Using the entropy weight-TOPSIS evaluation analysis model, the performance factors influencing J-series fighter jets are quantitatively analyzed and ranked, avoiding subjective assumptions. Ultimately, the flight performance scores of eight J-series fighter jets under ten influencing factors are obtained. The results show that the top four scoring fighter jets are J-31 > J-20 > J-10 > J-15. This outcome provides directional guidance for the future development of fighter jets, indicating that by selecting appropriate engines, root-to-tip ratios, aspect ratios, and tail fin types, fighter jets can achieve better flight performance. This has significant implications for enhancing national defense capabilities, promoting military technological progress, and fostering economic development.

Constrained Model Predictive Control for Generation Power Distribution on Aircraft Engines

Aiming at the increasing demand for electric energy in aircraft in the future, a multi-objective optimization aircraft engine constrained model predictive control method based on generation power distribution is proposed. Firstly, based on the aircraft engine component level model and the equilibrium manifold theory, the aircraft engine equilibrium manifold expansion model is established. Secondly, the influence of the power generation is modeled, and the influence of the low- and high-pressure shaft generators on the normal operation of the aircraft engine is studied and compared. The control variables such as fuel flow and total generation power are taken as the constraint conditions to design the constraint model predictive controller. Furthermore, the multi-objective grey wolf optimization algorithm is introduced to intelligently optimize the parameters of the designed controller. At last, the simulation based on the component level model shows that the high-pressure shaft generator has less influence on the state quantity, including engine thrust, than the low-pressure shaft generator. The proposed control method using the multi-objective gray wolf optimization (MOGWO) algorithm has rapid response and no steady-state error.

Rethinking Propellant-Optimal Powered Descent Guidance

The propellant-optimal powered descent solution requires a bang-bang thrust magnitude profile that switches instantaneously between the upper and lower bounds of the engine thrust. Implementing a bang-bang thrust pattern for guidance can be challenging for the engine and may cause practical and operational difficulties. Motivated by the observation that excellent propellant performance does not appear to be predicated on necessarily having a bang-bang thrust structure, a new powered descent guidance method is proposed in this work. The foundation lies in a propellant-optimal problem in which the thrust magnitude is parameterized by a user-prescribed continuous function of time. In particular, constant and linear functions are considered. The solution determines the optimal thrust direction, time of flight, and design parameters defining the thrust function. An indirect-method-based guidance algorithm is developed to solve this problem rapidly and reliably. Substantial evidence is provided to demonstrate that the solutions proposed in this paper in general yield a propellant consumption very close to that of the optimal bang-bang solution. Given its comparable propellant performance, implementability, and robustness, this paper makes a strong case that the proposed propellant-optimal powered descent guidance approach is preferred to the long-standing bang-bang guidance approach.

Study on the influence of hydrogen injector flow rate in hydrogen-oxygen engine thrust chamber

Study on the influence of hydrogen injector flow rate in hydrogen-oxygen engine thrust chamber

Features of achievement of limiting speed values of track tests of ballistic type aircraft

The development of high-speed ballistic aircraft with speeds exceeding 1000 m/s is currently a priority abroad and in Russia. The effectiveness of new such products is confirmed by track tests at the speed of their use. Test sites with rail tracks exist in almost all countries, for example in the USA there are more than 15 of them. Double-rail, monorail and various combinations thereof, differing in length, width of the rail pair, rails and the design of the track itself, including a sealed shell over the rail track to fill it with a lighter one environment. The longest track in the USA is Holloman AFB, located in New Mexico with a length of 15536 m. They have track ranges with different lengths and their own special design in England, France, Germany, Canada, Italy, Japan, India, China, Korea, Turkey and other countries, including African continent. High-speed range tests in Russia are carried out on the experimental installation “Rocket Rail Track 2500”, located on the territory of the FKP “GkNIPAS named after L. K. Safronov”. The experimental installation consists of a rail track placed on a special base, providing the necessary vertical profile of the track with sections of ascent and straight horizontal movement, as well as a technological descent section for braking moving technological equipment. The product under test is placed on a rocket track sled moving along rails on sliding supports. To accelerate the track carriage, solid fuel rocket engines are used, the thrust of which is selected based on ballistic calculations to achieve the required test speed. The length of the track plays an important role in achieving the maximum acceleration speeds of moving track equipment. The enormous aerodynamic drag, proportional to the square of the speed of movement of the carriage, when tested at high speeds, leads to the need to reduce the midsection and mass of the mobile unit. An increase in engine thrust leads to an increase in the weight and cost of track equipment, as well as to the need to increase the safety margin of sliding supports. However, an increase in test speed can be achieved by replacing the air medium with gases that have a significantly lower density, for example, helium. Track testing of new aircraft or their elements, although cheaper than flight testing, is quite expensive. In this regard, work on the theoretical assessment of replacing the medium from ambient air with helium, as well as with a mixture of helium and air at different concentrations in an indoor gallery on a track rail track, is a new, relevant and practically useful task. The work performed a numerical simulation of the problem of supersonic flow around a helium-air mixture at different volumetric ratios. Numerical values of aerodynamic resistance were obtained at a sled speed of 830 m/s. The results of numerical calculations of the motion dynamics of a 3D model of monorail track equipment, which are planned for use in conducting full-scale fire experiments, are presented.

Multi-objective optimization of the aerodynamic performance of butterfly-shaped film cooling holes in rocket thrust chamber

Abstract This study uses Multi-Island Genetic Algorithm (MIGA) and three-dimensional Computational Fluid Dynamics (CFD) software to optimize butterfly-shaped film cooling holes in the upper-stage rocket engine thrust chamber. The goal is to meet thermal protection and thrust requirements at high altitudes without re-ignition. To facilitate an all-encompassing worldwide search, the holes in the optimized design remain at set dimensions. Film continuity and stability at the nozzle outlet are greatly impacted by the hole structure. Inlet and divergence angles have little effect on thrust, according to regression research, but lip height (de) and outlet width (β) have a big impact on cold gas ejection, which affects cooling and thrust. Optimized results lead to a 20.49 K decrease in the monitoring section’s average wall temperature and a 52.8 N boost in thrust by reducing interference between supersonic airflow and extending film stability.

Study of design methods and heat transfer characteristics of combined cooling structure for turbine blades with turbine inter-vane burner

The turbine inter-vane burner can greatly increase the engine thrust but brings about a harsh thermal environment to the turbine blades. Fuel-air combined cooling structure is an effective way to solve the thermal protection problem of the blades. There is a lack of research on the thermal protection of turbine blades with turbine inter-vane burner, especially systematic research on the fuel–air combined cooling structure. In this paper, the design methods of the fuel–air combined cooling structure of turbine blades with turbine inter-vane burner are proposed, and one-dimensional empirical formulas are used to design specific parameters. The heat transfer characteristics of the designed combined cooling structure are numerically investigated and analyzed without turbine inter-vane combustion, where the mainstream temperature varies from 818 K to 1414 K, the cold air temperature ranges from 500 K to 700 K, and the mass flow rate of cold air changes from 0.005 kg/s to 0.02 kg/s. The results show that in the combined cooling structure with embedded fuel channels, the heat absorbed by fuel is mostly influenced by cold air temperature compared to mainstream temperature and mass flow rate of cold air. The proportion of fuel heat absorption to overall heat absorption decreases from 0.25 to 0.13 when mainstream temperature increases from 818 K to 1418 K, as the increase of mainstream temperature has a minimal effect on the heat flux of the fuel channels, but can significantly increase the overall heat absorption. The proportion of fuel heat absorption to the overall heat absorption changes from 0.15 to 0.31 when cold air temperature increases from 500 K to 700 K, as the increase of cold air temperature enhances the heat exchange between the fuel channels and the cold air but decreases the heat exchange between the mainstream and the cold air. Besides, the flow and heat transfer performance of the combined cooling structure of turbine blades with turbine inter-vane combustion is numerically studied. The results show that the temperature of blade and fuel channels are maintained within the safe range with turbine inter-vane combustion, providing research support for subsequent studies on the fuel–air combined cooling structure of turbine blades with turbine inter-vane burner.

Simple model of turbine-based combined cycle propulsion system and smooth mode transition

Abstract Turbine-based combined cycle (TBCC) propulsion system is becoming one of the most promising propulsion systems for two-stage-to-orbit reusable launch vehicle. Mathematic model of this combined cycle engine is helpful for the basic understanding of performance analysis of this propulsion system. We developed mathematic model of TBCC propulsion system based on C++ platform in this paper. Firstly, turbojet engine was built on component level and ramjet engine was calculated through stream thrust function. Then, performance of turbine-based combined cycle propulsion system along a specific flight trajectory was investigated. According to the thrust of this combined cycle engine, mode transition point was suggested at Ma 2.5, which may achieve smooth mode transition from turbine mode to ramjet mode. Finally, mode transition based on smooth mass flow and smooth thrust criteria were studied. The thrust gap arises during smooth mass flow mode transition, particularly when the turbojet engine afterburner is powered off at the start of the mode transition, and it meets 34 % of the total thrust. Smooth thrust mode transition may be achieved by delaying the turbojet engine afterburner power off and injecting additional fuel into ramjet burner.

Application of semiempirical correlations for multidimensional heat transfer in rocket engine cooling channels

Heat transfer in cooling channels of liquid rocket engine thrust chambers is an intrinsically multidimensional problem due to the geometry of the channels, often rectangular in section, and the asymmetric distribution of heat input from the hot combustion gas. In this study, the possibility of addressing this problem with semiempirical correlations for the coolant heat transfer and multidimensional numerical simulations for the heat conduction in the wall is investigated. Using experimental data from a test campaign of a cooling channel representative of liquid rocket engine conditions, semiempirical correlations are derived. Both cases of constant and variable heat transfer coefficient along the perimeter of the channel section are considered. The resulting correlations are then used to evaluate their reliability in reproducing the experimental wall temperature data from the same test campaign. For comparison, a correlation taken from the open literature and obtained with experimental data of circular heated tubes is also considered. The results demonstrate that correlations based on a constant heat transfer coefficient fail to reproduce with sufficient accuracy the test cases characterized by higher wall temperatures, while correlations based on a variable heat transfer coefficient overestimate the wall temperature in the initial sections of the channel.

Aero-engine direct thrust control based on nonlinear model predictive control with composite predictive model

Abstract A novel NMPC (Nonlinear Model Predictive Control) based on composite predictive model is proposed and applied to direct engine thrust control. To improve the real-time of NMPC, an adaptive composite model based on SVM (State Variable Model), KF (Kalman Filter), and CLM (Component Level Model) is proposed as predictive model. The correction theory is adopted to establish a full envelope adaptive on-board predictive dynamic model and reduce the data storage of predictive model. At each sampling time, the CLM is calculated only once in the proposed NMPC, instead of many times in the popular NMPC based on EKF (extended Kalman filler). Therefore, the proposed NMPC has better real-time performance than the popular one. The simulations that consist of the proposed NMPC, the popular NMPC based on EKF, and the traditional controller PID are conducted. The simulations demonstrate that the proposed NMPC not only has greatly better real time performance than popular NMPC, but also has faster response speed than traditional controller PID.

Research on performance seeking control of turbofan engine in minimum hot spot temperature mode

Abstract The uneven temperature distribution at the combustion chamber outlet seriously affects the working life of the engine. In order to reduce the heat spot temperature at the combustion chamber outlet, a performance optimization control method of the engine minimum heat spot temperature pattern is proposed. Firstly, based on CFD method, the temperature distribution characteristics of combustion chamber outlet under different working conditions were obtained, and a component-level model of turbofan engine was established to characterize the heat spot temperature at combustion chamber outlet. Secondly, the high precision and high real-time engine on-board model is established by deep neural network. Compared with the component-level model, the average relative error of each performance parameter is less than 0.3 %, and the real-time performance is improved by 12 times. Finally, based on the feasible sequential quadratic programming algorithm, the performance optimization control of the minimum hot spot temperature model in the typical flight envelope is simulated and verified. The simulation results show that under the condition of ensuring the safe and stable operation of the engine and constant thrust, the heat spot temperature at the combustion chamber outlet decreases by 21 K maximum. Compared with the conventional minimum turbine front temperature optimization mode, the minimum heat spot temperature mode significantly reduces the heat spot temperature at the combustion chamber outlet.

Experimental and numerical investigation of the effects of porous bleed systems on a model scramjet inlet under high backpressure conditions

In this study, we conducted experimental and numerical investigations to assess the effects of backpressure and configurations of porous bleeds on a scramjet inlet's performance at Mach 5. Experiments in a high-enthalpy wind tunnel, using a backpressure controller and a porous bleed with 1 mm holes and 10 mm lateral spacing, were paired with Reynolds-averaged Navier-Stokes simulations to explore the mitigation of flow separation and the prevention of inlet unstart under various backpressure conditions. The simulations provided insights into how backpressure levels and porous bleed geometries impact internal flow structures, as well as the critical backpressure ratio at which inlet unstart occurs. Additional simulations varied the bleed system's hole diameter and spacing, identifying configurations that optimize aerodynamic performance by enhancing total pressure recovery and Mach number at the isolator exit, while reducing bleed mass flow rates. The study also quantified the impact of these configurations on engine thrust, determining that certain bleed system designs significantly improve thrust by efficiently suppressing the interaction between the core flow and corner recirculation zones. This study examined the internal three-dimensional flow structure characteristics under high back pressure and provided insights into porous bleed configuration capable of efficiently suppressing inlet unstart.

Central cone film cooling scheme of turbofan engine based on multi-fidelity simulation

To improve the computational efficiency of multi-fidelity simulation, an improved fully coupled method was proposed, which was applied to the coupled simulation of a 3-D model of an exhaust system and a 0-D model of a turbofan engine. Different from conventional approach, by “truncating” the engine model, the CFD calculation of the exhaust system can be decoupled from the solution of nonlinear equations in the engine model, and an outer iteration equation set was constructed to solve the multi-fidelity model. It was found that the improved fully coupled method has better computational efficiency than traditional methods while maintaining equivalent computational results. Then, the improved fully coupled method was applied to the design of the central cone film cooling system, the influence of cooling gas bleed mechanism, the hole open area ratio of the central cone and hole density, hole direction, hole shape on the infrared characteristics was investigated. The impact of central cone film cooling on the overall performance of the engine was quantitatively evaluated. Combined with various favorable factors, when maintaining the HP rotor speed unchanged, the integrated infrared radiation intensity of the central cone under supersonic cruise conditions is reduced by 93 % compared to the uncooled state, while the engine thrust increases by 0.3 % and fuel consumption rate increases by 0.18 %. The multi-fidelity simulation method proposed in this paper and the conclusions obtained are of great significance for the structural design of low-infrared exhaust systems.