Turbine Engine Performance Research Articles

Aeroengine transient performance simulation integrated with generic heat soakage and tip clearance model

AbstractThe simulations and assessment of transient performance of gas turbine engines during the conceptual and preliminary design stage may be conducted ignoring heat soakage and tip clearance variations due to lack of detailed geometrical and structural information. As a result, problems with transient performance stability may not be revealed correctly, and corresponding design iterations would be necessary and costly when those problems are revealed at a detailed design stage. To make an engine design more cost and time effective, it has become important to require better transient performance simulations during the conceptual and preliminary design stage considering all key impact factors such as fuel control schedule, rotor dynamics, inter-component volume effect as well as heat soakage and tip clearance variation effects. In this research, a novel transient performance simulation approach with generically simplified heat soakage and tip clearance models for major gas path components of gas turbine engines including compressors, turbines and combustors has been developed to support more realistic transient performance simulations of gas turbine engines at conceptual and preliminary design stages. Such heat soakage and tip clearance models only require thermodynamic design parameters as input, which is normally available during such design stages. The models have been implemented into in-house transient performance simulation software and applied to a model twin-spool turbojet engine to test their effectiveness. Comparisons between transient performance simulated with and without the heat soakage and tip clearance effects demonstrate that the results are promising. Although the introduced heat soakage and tip clearance models may not be as accurate as that using detailed component geometrical information, it is able to include the major heat soakage and tip clearance effects and make the transient performance simulations and analysis more realistic during conceptual and preliminary engine design stage.

Comparative analysis of turboshaft engines thermodynamic cycles calculation

A brief overview of domestic and foreign programs for calculating the thermodynamic cycle of the gas turbine engine operation is provided. The focus is on comparing GasTurb and Tyumen Motor Builders methods. Configuration files are provided for NASA Chemical Equilibrium with Applications program to calculate the thermodynamic properties of methane, air, and combustion products. The main provisions of the procedure for calculating thermal schemes of turboshaft engines are presented. Using the example of a three-shaft ship-type gas turbine engine UGT 15000, the potentials for increasing efficiency by 4-7% (abs.) and reducing CO2 emissions by 20% are shown due to the installation of an intercooler during compression of the working fluid. According to the results of comparison of GasTurb and Tyumen Motor Builders programs, the correspondence of the main parameters within the permissible measurement error is established. Based on published data on UGT 15000 (DG90), the parameters of the thermodynamic cycle were identified, including the polytropical efficiency of each node and cooling rates (ISO 2314). It has been shown that due to the isotermo-adiabatic cycle of V. V. Uvarov, already in the first approximation, without extreme parameters of the thermodynamic cycle, the characteristics of the V generation of the gas turbine engine can be implemented based on affordable and economical technologies of the II-IV generation (using equiaxed nickel superalloys, cyclone or convective-film cooling of high-pressure turbine vanes and blades). The Tyumen Motor Builders methodology program is open-source and has proven itself positively in the service-pack development for engines DG90/DN80/DU80 (III-IV generation) and the new ТМ16М2 engine construction.

A hybrid ant colony–particle swarm optimization method (ACOPSO) for aerospace propulsion systems

PurposeThe purpose of this paper is to introduce a hybrid, metaheuristic, multimodal and multi-objective optimization tool that is needed for aerospace propulsion engineering problems.Design/methodology/approachMulti-objective hybrid optimization code is integrated with various benchmark and test functions that are selected suitable to the difficulty level of the aero propulsion performance problems.FindingsAnt colony and particle swarm optimization (ACOPSO) has performed satisfactorily with benchmark problems.Research limitations/implicationsACOPSO is able to solve multi-objective and multimodal problems. Because every optimization problem has specific features, it is necessary to search their general behavior using other algorithms.Practical implicationsIn addition to the optimization solving, ACOPSO enables an alternative methodology for turbine engine performance calculations by using generic components maps. The user is flexible for searching various effects of component designs along with the compressor and turbine maps.Originality/valueA hybrid optimization code that has not been used before is introduced. It is targeted use is propulsion systems optimization and design such as Turboshaft or turbofan by preparing the necessary engine functions. A number of input parameters and objective functions can be modified accordingly.

Effects of Front Plate Geometry on Brush Seal in Highly Swirling Environments of Gas Turbine

Advanced brush seal technology has a significant impact on the performance and efficiency of gas turbine engines. However, in highly inlet swirling environments, the bristles of a brush seal tend to circumferentially slip, which may lead to aerodynamic instability and seal failure. In this paper, seven different front plate geometries were proposed to reduce the impact of high inlet swirl on the bristle pack, and a three-dimensional porous medium model was carried out to simulate the brush seal flow characteristics. Comparisons of a plane front plate with a relief cavity, plane front plate with axial drilled holes, anti-“L”-type plate and their relative improved configurations on the pressure and flow fields as well as the leakage behavior were conducted. The results show that the holed front plate can effectively regulate and control the upstream flow pattern of the bristle pack, inducing the swirl flow to move radially inward, which results in decreased circumferential velocity component. The anti-“L” plate with both axial holes and one radial hole was observed to have the best effect on reducing the swirl of those investigated. The swirl velocity upstream the bristle pack can decline 50% compared to the baseline model with plane front plate, and the circumferential aerodynamic forces on the bristles, which scale with the swirl dynamic head, are reduced by a factor of 4. This could increase the bristle stability dramatically. Moreover, the front plate geometry does not influence the leakage performance significantly, and the application of the axial hole on the front plate will increase the leakage slightly by around 3.5%.

Relative Effects of Velocity- and Mixture-Coupling in a Thermoacoustically Unstable, Partially Premixed Flame

Abstract Combustion instability, which is the result of a coupling between combustor acoustic modes and unsteady flame heat release rate, is a severely limiting factor in the operability and performance of modern gas turbine engines. This coupling can occur through different pathways, such as flow-field fluctuations or equivalence ratio fluctuations. In realistic combustor systems, there are complex hydrodynamic and thermo-chemical processes involved, which can lead to multiple coupling pathways. In order to understand and predict the mechanisms that govern the onset of combustion instability in real gas turbine engines, we consider the influences that each of these coupling pathways can have on the stability and dynamics of a partially premixed, swirl-stabilized flame. In this study, we use a model gas turbine combustor with two concentric swirling nozzles of air, separated by a ring of fuel injectors, operating at an elevated pressure of 5 bar. The flow split between the two streams is systematically varied to observe the impact on the flow and flame dynamics. High-speed stereoscopic particle image velocimetry, OH planar laser-induced fluorescence, and acetone planar laser-induced fluorescence are used to obtain information about the velocity field, flame, and fuel-flow behavior, respectively. Depending on the flow conditions, a thermoacoustic oscillation mode or a hydrodynamic mode, identified as the precessing vortex core, is present. The focus of this study is to characterize the mixture coupling processes in this partially premixed flame as well as the impact that the velocity oscillations have on mixture coupling. Our results show that, for this combustor system, changing the flow split between the two concentric nozzles can alter the dominant harmonic oscillation modes in the system, which can significantly impact the dispersion of fuel into air, thereby modulating the local equivalence ratio of the flame. This insight can be used to design instability control mechanisms in real gas turbine engines.

Development of the system of energy and resource saving during the operation of the gas pumping unit

The object of research is the fuel gas system of a gas turbine engine. The study of the use of secondary energy resources of the gas-pumping unit at the compressor station of the main pipeline has been carried out. The work of a gas turbine engine, including the work of the fuel gas system, is considered. The main drawback of the fuel gas system is revealed – ineffective use of excess gas pressure. An informational analysis of the options that eliminate the identified drawback is carried out. It is shown that in order to eliminate the disadvantage, it is advisable to use a turbo-expander at the compressor station to utilize the excess pressure of the fuel gas. It is also shown that the operation of a fuel gas turboexpander to drive an additional air compressor as part of a gas turbine engine is impractical. An expander-generator set with the generation of additional electricity at the compressor station is recommended for use. Modeling the operation of the utilization system made it possible to recommend constructive proposals for its improvement. A schematic diagram of a system for the complex utilization of excess pressure of fuel gas and heat of combustion products from the operation of a gas turbine engine is proposed. The system of complex utilization includes parts-generator unit, heat exchanger for cooling process gas and heat exchanger for firing gas. Regenerative heating of fuel gas up to 250 °С reduces energy consumption for heating it up to the ignition temperature. A model of a robot installation of the type GPU 16/56-1.44 (Ukraine) is carried out. It is determined that when a component engine of the J-59 (Ukraine) type with a shaft power of 16 MW operates, it is possible to additionally receive 102 kW of electricity and save 64 m3/h of fuel gas. It is revealed that the subcooling of the process gas does not play a significant role in reducing energy consumption during its transportation. It is recommended to use the process gas to heat the cold fuel gas stream downstream of the turboexpander to positive temperatures. The integrated utilization system is not a simple connection of an expander-generator set and two heat exchangers along the flow of the fuel gas. As a result of its operation, a significant reduction in the consumption of fuel gas and electricity is achieved. The disadvantages that hinder the implementation of a comprehensive disposal system are identified. This is the use of equipment for generating electricity at a compressor station. It is uncharacteristic for the operation of the station and requires additional qualifications in service. It is also required that the characteristics of industrial expander-generator sets correspond to the fuel gas consumption of a gas turbine engine.

Effects of heat input on metallurgical behavior in HAZ of multi-pass and multi-layer welded IN-939 superalloy

The hot components of a gas turbine are susceptible to damage in the high-temperature environment of turbine engine operation. Given that these components are relatively costly to manufacture, they are often repaired than replaced when damaged. Fusion welding is an economical method for repairing the damaged components of a gas turbine. This research examines the roles of heat input, pass number and layer number on the intergranular liquation cracking of the Inconel-939 (IN-939) precipitate nickel base superalloy during tungsten arc welding. Several specimens were welded with IN-625 filler alloy under argon gas by following the Taguchi method and an L8 array. These specimens were then characterized via metallographic investigations and hardness measurements. Results show that, during welding, the IN-939 heat-affected zone (HAZ) is acutely prone to cracks that propagate along the grain boundaries. Moreover, layer number, heat input and pass number significantly influence the microstructure and liquation cracking of HAZs with impact percentages of 72.37%, 22.17% and 4.84%, respectively.

Mathematical model of brush seals for gas turbine engines: A nonlinear analytical solution

Presented herein is a mathematical model employing differential equations formulation for brush seals used in gas turbine engines. These components are used to seal the bearing chamber from the environment and reduce the loss of lubricant in the atmosphere, ensuring a MTBR long enough to have required the change the seals only during the engine overhaul operation. The model assumes a single curved bristle loop in the form of a curved-bridge beam subjected to the influences of complex external loads (static and dynamic). Further, a model for clustered bristles is proposed. Specifically, the static forces acting on the curved-bridge beam include the weight of the oil capillary attached to the beam, the weight of the beam itself, the capillary force developed between the surfaces of the bristles in the brush and the temperature gradient. The dynamic forces include the leakage oil pressure and the rotation of the shaft. This complex loading induces a nonlinear large deflection on the curved-bridge beam. Also, the temperature gradient present on the bristles during the gas turbine engine operation generates a change in the geometry of the beam and in the magnitude of the forces acting on the bristles modeled as beams. In the present model, the weights are assumed as uniformly distributed forces on the surface of the beam while the capillary forces and the force generated by the rotating shaft are considered to be non-uniform. The equation expressing the curvature of the beam under general loading force is developed and one can choose the appropriate method of solving the generated differential equation after the expression of the general force is defined. Hence, the ordinary differential equation describing the nonlinear large deflection of the curved-bridge beam will be derived using general nonlinear elasticity theory.

Devising a method for calculating the turboshaft gas turbine engine performance involving a blade-by-blade description of the multi-stage compressor in a two-dimensional setting

The design and adjustment of modern gas turbine engines significantly rely on the use of numerical research methods. This paper reports a method devised for calculating the thermogasdynamic parameters and characteristics of a turboshaft gas turbine engine. The special feature of a given method is a two-dimensional blade-by-blade description of the compressor in the engine system. Underlying the calculation method is a nonlinear mathematical model that makes it possible to describe the established processes occurring in individual nodes and in the engine in general. To build a mathematical model, a modular principle was chosen, involving the construction of a system of interrelated and coordinated models of nodes and their elements. The approach used in modeling a two-dimensional flow in the compressor makes it possible to estimate by calculation a significant number of parameters that characterize its operation. With the help of the reported method, it is possible to estimate the effect of changing the geometric parameters of the compressor height on the characteristics of the engine. To take into consideration the influence of variable modes of air intake or overflow in various cross-sections along the compressor tract, to determine the effect of the input radial unevenness on the parameters of the compressor and engine in general. To verify the method described, the calculation of thermogasdynamic parameters and throttle characteristics of a single-stage turboshaft gas turbine engine with a 12-stage axial compressor was performed. Comparison of the calculation results with experimental data showed satisfactory convergence. Thus, the standard deviation of the calculation results from the experimental data is 0.45 % for the compressor characteristics, 0.4 % for power, and 0.15 % for specific fuel consumption. Development and improvement of methods for calculating the parameters and characteristics of gas turbine engines make it possible to improve the quality of design and competitiveness of locally-made aircraft engines.

Analytical STEM Investigations of Nozzle Vane Surfaces after Turbine Engine Operation

The scope of this work includes detailed microstructural investigations using high-resolution scanning transmission electron microscopy of the oxide scales and deposits formed on nozzle ring vanes made of IN713LC alloy during the operation of an auxiliary power unit under real conditions. The paper presents the differences of oxidation processes between the suction and the pressure sides of the vanes. It has been shown that in the pressure side of the vane, where a greater flow of exhaust gases from the combustion chamber is present, the oxidation process is accompanied by a deposition of among others (i.a.) Fe, Mg, Ca, Na, and P, while on the suction side of the vane, the thermal stresses and the mechanical loading most likely lead to a cracking of the oxide scale. A segregation of Cr, Ni, Ta, and Ti to the boundaries of α-Al2O3 grains is observed, which may indicate their diffusion from a metal to a gas atmosphere.

CFD를 이용한 항공용 마이크로 가스터빈 엔진 시뮬레이션

Turbojet-type micro gas turbine engine has been widely adopted for an aero-propulsion system of small-sized aircraft due to its simplicity and prominent thrust-to-weight ratio. The number of applications using micro gas turbine engine continues to grow with the development of various type in aerial vehicles. To adopt an appropriate micro gas turbine engine for a specific aero application, an evaluation of the engine performance is very important. A thermodynamic cycle analysis is an usual method for the gas turbine engine performance evaluation. Establishing the thermodynamic cycle model of turbojet engine is relatively simple and simulation time is the shortest compared to other methods. However, details of engine characteristics would not be captured with the cycle model and the result might be incorrect if analysis condition were apart from component’s operational range. These deficiencies could be compensated by taking the advantage of CFD analysis, which is capable of simulating the details of flow characteristics inside engine components such as stall, choke and shock. These flow phenomena can deteriorate engine performance or can cause severe engine failure. This paper demonstrates performance analysis of micro gas turbine engine using CFD analysis. Major components in a micro gas turbine engine, such as compressor, combustor, turbine and exhaust nozzle, were modeled and linked together to establish entire engine model. Fortran sub-routine was newly developed and embedded to CFD code, which is manipulating the fuel flow into combustor to match compressor and turbine torques. The sub-routine continues to run until it finds matching point of each component at designated rotating speed. Using this CFD model, not only engine cycle performance, but also component performance can be evaluated with a single calculation.

Effect of Laser Treatment on the Surface Roughness of Multilayer Plasma Sprayed Thermal Barrier Coating System

Thermal barrier coatings (TBCs) are used in advanced engines working at higher temperatures. Higher efficiency and performance of gas turbine engines will require careful selection of TBCs. In this study, Ni22Cr10Al1.0Y (Amdry 9625) bond coat and two types of top coat including ceria stabilized zirconia (CSZ) ZrO2-24CeO2-2.5Y2O3) and yttria stabilized zirconia (YSZ) ZrO2-8Y2O3 were deposited on IN 625 by air plasma spraying (APS). The thickness of the duplex ceramic coat based on zirconia was in the range between 350 to 400 µm. The effect of high power Yb:YAG solid state laser at different laser parameters on feature, microstructure and roughness of plasma sprayed and laser sealed coating of multilayer ceria stabilized zirconia/ yttria stabilized zirconia was investigated. Surface roughness has been reduced significantly after laser sealing. The effect of laser process parameters carried out using Taguchi’s L16 orthogonal array design. Minimum roughness can be obtained at moderate power density and longer interaction time with sufficient specific energy to produce complete melting of coating. Characterization and analysis of results was achieved by employing scanning electron microscopy (SEM) , (EDS) and image J analysis. It was found from the results, there were significant improvements in the performance of plasma sprayed coatings after laser sealing due to the reduction of surface coating defects.

Challenges in Qualifying Additive Manufacturing for Turbine Components: A Review

Gas turbine engine advancements have been enabled by innovation in materials and manufacturing technologies. The evolution of additive manufacturing (AM) has changed the face of direct digital technologies for the rapid production of models, prototypes, and functional parts including repair and maintenance for turbine engines. Metal 3D printing is poised to be an enabler for the next industrial revolution in enabling advancements in turbine engine performance. While there has been tremendous efforts on research and development in utilizing this versatile technology, the number of qualified metallic parts that are running in the engine has not been commensurate with the applied research and development efforts and the benefits that the AM technology offers. This review addresses some of the key technical issues that are currently limiting the prolific usage of AM as a successful vehicle for accelerated progress in gas turbine engines.

ANALISIS HUMAN ERROR PADA OPERATOR ENGINE ROOM PT. ADEI PLANTATION & INDUSTRY DENGAN METODE HUMAN ERROR ASSESMENT AND REDUCTION TECHNIQUE (HEART)

PT. Adei Plantation & Industry is a plantation company that operates on plantation and palm oil industries. CPO production of palm oil is needed for consumption, which is run by power stations and for distribution (engine room) involving operators. Since january to june 2020, there is a sudden shutdown in the work of research for the analysis of human errors in the engine room using HEART method. The HEART method is used to evaluate the possibility of human error in all the work and to count the erros in various mays. Data management shows that there are 4 task operators that are common to the plant turbine engines and 6 task forces on the genertor. The highest HEP rating on the turbin engine was 0,6528 with the greatest error potential for task 3, the turbine engine operation. For the highest HEP value on the generator engine is 3,1759 with an error potential in task 5.

Lowest emission sustainable aviation biofuels as the potential replacement for the Jet-A fuels

Purpose The purpose of this study is to predict the performance and emission characteristics of micro gas turbine engines powered by alternate fuels. The micro gas turbine engine performance, combustion and emission characteristics are analyzed for the jet fuel with different additives. Design/methodology/approach The experimental investigation was carried out with Jet A-1 fuel on the gas turbine engines at different load conditions. The primary blends of the Jet A-1 fuels are from canola and solid waste pyrolysis oil. Then the ultrasonication of highly concentrated multiwall carbon nanotubes is carried with the primary blends of canola (Jet-A fuel 70%, canola 20% and 10% ethanol) and P20E (Jet-A 70% fuel, 20% PO and 10% ethanol). Findings The consumption of the fuel is appreciable with the blends at a very high static thrust. The 39% reduction in thrust specific fuel consumption associated with a 32% enhance in static thrust with P20E blend among different fuel blends. Moreover, due to the increase in ethanol concentration in the blends PO20E and C20E lead to a 22% rise in thermal efficiency and a 9% increase in higher oxygen content is observed. Practical implications The gas turbine engine emits very low emission of gases such as CO, CO2 and NOx by using the fuel blends, which typically reduces the fossil fuel usage limits with reduced pollutants. Originality/value The emission of the gas turbine engines is further optimized with the addition of hydrogen in Jet-A fuel. That is leading to high specific fuel exergy and owing to the lower carbon content in the hydrogen fuel when compared with that of the fossil fuels used in gas turbine engines. Therefore, the usage of hydrogen with nanofluids was so promising based on the results obtained for replacing fossil fuels.

Micromachining porous alumina ceramic for high quality trimming of turbine blade cores via double femtosecond laser scanning

Turbine blade cores are made of porous alumina ceramic and determine the molding accuracy of the cavity of turbine blades, which strongly affect thermal diffusion performance and service life of turbine engines. To get a high quality ceramic core, accurate trimming for a preliminarily processed core is needed and therefore, micromachining porous alumina ceramic, which differs from general alumina substrates, is crucial. This paper dealt with a processing technology for the special material via double femtosecond laser scanning. The materials ablation threshold was firstly determined through parameter fitting and then this material was machined at a combination of different laser processing parameters. Considering the produced debris blocks the lasers further propagation into the material, double femtosecond laser scanning was newly proposed and experimentally verified with the comparison of gas jet assist and underwater laser processing ways. The removal profiles of the machined material were characterized in terms of cutting width, cutting depth, deviation of linearity and surface morphology, which exhibited high dependence on the femtosecond laser processing parameters. The optimal laser operating window was identified and high quality laser cutting of the porous alumina ceramic was demonstrated. The developed processing technology has potential application in trimming for ceramic casting cores. In addition, it might also give a novel view for high quality laser micromachining another materials.

High-Fidelity Finite Element Modeling and Analysis of Adaptive Gas Turbine Stator-Rotor Flow Interaction at Off-Design Conditions

ABSTRACTThe objective of this work is to computationally investigate the impact of an incidence-tolerant rotor blade concept on gas turbine engine performance under off-design conditions. When a gas turbine operates at an off-design condition such as hover flight or takeoff, large-scale flow separation can occur around turbine blades, which causes performance degradation, excessive noise, and critical loss of operability. To alleviate this shortcoming, a novel concept which articulates the rotating turbine blades simultaneous with the stator vanes is explored. We use a finite-element-based moving-domain computational fluid dynamics (CFD) framework to model a single high-pressure turbine stage. The rotor speeds investigated range from 100% down to 50% of the designed condition of 44,700 rpm. This study explores the limits of rotor blade articulation angles and determines the maximal performance benefits in terms of turbine output power and adiabatic efficiency. The results show articulating rotor blades can achieve an efficiency gain of 10% at off-design conditions thereby providing critical leap-ahead design capabilities for the U.S. Army Future Vertical Lift (FVL) program.

Characteristic and regulation method of parallel turbocompound engine with steam injection for waste heat recovery

In a parallel turbocompound engine, a power turbine (PT) is fitted in the exhaust pipe in parallel with the turbocharging turbine (CT) for waste heat recovery. In this study, steam is generated using the exhaust heat and injected at the turbines inlet to improve the turbine power and engine performance. Numerical model for the engine cycle is established using commercial software GT-POWER. The model is validated against the experimental results of the turbocompound engine and precisely predicts the engine performance. Based on the model, the impacts of the CT and PT sizes on the performance of the parallel turbocompound engine are evaluated firstly. It is found that variable geometry turbine (VGT) should be applied on PT to regulate the energy distribution between the two turbines at engine off-design condition. Then the influences of steam mass flow rate on the engine BSFC, power and efficiency are investigated. The optimum values of the steam mass flow rate and PT size is closely related to the exhaust temperature. At last, the characteristic and regulation pattern of the steam injected parallel turbocompound engines are compared with series turbocompound engine. In parallel turbcompound engine, the BSFC is decreased by 2.08%–3.28% with steam injection.

Effect of balance valve on an asymmetric twin-scroll turbine for heavy-duty diesel engine

A conventional asymmetric twin-scroll turbine with wastegate is capable of effectively tackling down the contradiction between fuel economy degradation and low nitrogen oxide emissions. However, as the engine speed has been rising at middle- and high-speed ranges, the pressure of small scroll inlet will be increasingly higher as compared with the intake pressure, thereby worsening fuel economy. In this study, a novel turbocharging technology of asymmetric twin-scroll turbine with a balance valve was first analyzed to more effectively balance the engine fuel economy and emission. The experiments on turbine test rig and engine performance were performed to explore the effects of balance valve on turbine performance, asymmetric ratio, exhaust gas recirculation rate, as well as engine performance. As the balance valve open degree was elevated, the turbine flow parameter was being extended, while the turbine efficiency was enhanced. Moreover, a lower asymmetric ratio could lead to a broader flow parameter range between that of partial admission and equal admission, thereby resulting in a broader regulating range of exhaust gas recirculation rate. In contrast with the asymmetric twin-scroll turbine with wastegate, the turbine running efficiency of asymmetric twin-scroll turbine with balance valve was enhanced by nearly 2%–11% at middle and high engine speed ranges, while the fuel economy was improved by nearly 1.5%–8%.

Experimental Analysis of an Ultra-Compact Combustor Powered Turbine Engine

Abstract The innovative ultra-compact combustor (UCC) is an alternative to traditional turbine engine combustors and has been shown to reduce the combustor length and offer potential improvements in combustion efficiency. Prior UCC configurations featured a circumferential combustion cavity positioned around the outside diameter (OD) of the engine. This configuration would be difficult to implement in a vehicle with a small, fixed diameter and had difficulty migrating the hot combustion products at the OD radially inward across an axial core flow to present a uniform temperature distribution to the first turbine stage. This study draws from preliminary computational analysis which enabled experimental testing of a new UCC configuration that featured a smaller diameter circumferential cavity that exhausted axially into a dilution zone positioned just upstream of the nozzle guide vanes. The combustor was sized as a replacement burner for the JetCat P90 RXi small-scale turbine engine and fit inside the engine casing. This combustor configuration achieved a 33% length reduction compared to the stock JetCat combustor and achieved comparable engine performance across a limited operating range. Self-sustained engine operation was achieved with a rotating compressor and turbine making this study the first to achieve operation of a UCC-powered turbine engine.