Supersonic Turbine Research Articles

Investigation on the integrated characteristics of n-decane/air rotating detonation combustor and supersonic turbine

In this study, the two-dimensional numerical simulations are conducted to study the flow field characteristics, turbine performance and loss mechanism of integrated system of rotating detonation combustor and supersonic turbine under two different directions of detonation wave propagation. The results indicate that the propagation direction of RDW affects the incident angle between OSW and guide vanes, resulting in different operating modes for aligned and unaligned modes. OSW in the turbine cascade undergoes the leading-edge shock, blade surface reflection and trailing edge diffraction. The backward-propagating rake-type shock envelope is formed due to leading-edge shock of the rotor. In misaligned mode, the stator has higher damping of both pressure and temperature. The stator significantly improves the circumferential uniformity of both velocity and pressure, particularly when operating in aligned mode. The total pressure loss in aligned mode is less, therefore the turbine achieves higher rim work and efficiency. The viscosity is one of the main sources of loss in the flow field. The reflected shock waves at the leading edge of the rotor and in the stator cascade are the primary factors contributing to the leading-edge vortices on the stator vanes.

Design and analysis of a single-stage supersonic turbine with partial admission

Supersonic impulse turbines are commonly used to harvest waste heat from systems based on organic Rankine cycles, automotive applications, and gas-generator liquid rocket engines. This research aims to design and analyze a single-stage supersonic turbine with a rated power of approximately 440 KW. We present a comprehensive preliminary design roadmap followed by design parameters optimization. The preliminary sizing is carried out using one-dimensional flow analysis, while the blade geometrical design is performed using the two-dimensional method of characteristics. The turbine performance is then analyzed through numerical simulations on three-dimensional models. Comprehensive parametric analyses are conducted to examine the turbine performance sensitivity to changes in partial admission (PA), mean radius, and turbine blade height using three-dimensional sliding mesh techniques (SMT). The current analysis simulates the full turbine to capture PA effects, rather than focusing on a single periodic sector as is common in turbomachinery simulation practices. Different PA scenarios are investigated, emphasizing a single rotor passage over a complete rotor cycle in a partially admitted turbine, and examining its charging and discharging mechanisms. Flow tracking over such a blade shows that the turbine rotor exhibits compressor-like behavior over most of the unadmitted zone, contributing to efficiency drops and severe transients.

Analysis of coupling supersonic turbine stage with rotating detonation combustor under different turbine parameters

The rotating detonation gas turbine boasts several advantageous characteristics, including self-pressurization, minimal entropy increase, high thermal efficiency, and a high thrust-to-weight ratio. These features make it a promising candidate for the next generation of aerospace power devices. This study investigates the flow characteristics and performance attributes of the supersonic turbine stage coupled with a rotating detonation chamber based on different rotational speeds, blade ratio of rotor and stator, and hub radius, and uses DMD (Dynamic Mode Decomposition) to analyze the main modes of the supersonic turbine. The research identifies distinct wave modes in both the stator and rotor blades of the supersonic turbine. In the stator blades, the modes include waveless contact modes, opposite side λ-wave oblique shock wave modes, and derived modes from opposite side λ-wave oblique shock waves. In the rotor blades, the modes encompass non-separation modes (waveless modes, asymmetric dual λ-wave modes, oblique cut wave modes, and single-side λ-wave modes on the pressure surface), leading-edge stator excitation single separation bubble modes, saddle-type separation modes, and dual separation bubble modes. The efficiency and power of the supersonic turbine stage increase with the rotational speed. The turbine efficiency is highest when the blade ratio is 10:7. Under multi-wave mode conditions, detonation waves demonstrate adaptive self-regulation abilities, ultimately achieving equilibrium in speed, detonation wave height, and spacing. The dominant modes in the stator blades and combustion chamber are governed by detonation waves and slip lines, while in rotor blades, the primary mode is the flow-directional vortex mode governed by detonation waves and slip lines at a frequency lower than their propagation frequency. This research holds certain reference significance for understanding the internal flow characteristics of rotating detonation gas turbines, as well as discussing the patterns of efficiency, power, and load characteristics, thereby offering valuable insights for the practical application of rotating detonation gas turbines.

Mach number effects on shock-boundary layer interactions over curved surfaces of supersonic turbine cascades

Mach number effects on shock-boundary layer interactions over curved surfaces of supersonic turbine cascades

Thermodynamic analysis of a novel precooled supersonic turbine engine based on aircraft/engine integrated optimal design

Aircraft and precooled engine efficient matching work is a key issue in the thermal cycle design of precooled engines, to solve this problem, an aircraft/precooled turbine engine integrated design method considering the fuel precooling impact mechanism is newly proposed. The impact of n-decane, ammonia, and hydrogen fuel precooling on aircraft design parameters and engine thrust requirements is revealed. Furthermore, a parameter optimal design method for fuel precooled engine is carried out to satisfy the multi-objective performance requirements of high specific thrust, low fuel consumption, high exergy efficiency, and low weight. The simulation results show that the n-decane fuel precooled engine reduces the takeoff weight of the aircraft by 4.23% and the fuel load by 11.6%, which has a better comprehensive performance than that of ammonia and hydrogen. Ammonia fuel has the maximum precooling heat sink and performance improvement space. After multi-objective optimization, it increases the maximum specific thrust by 7.3%, reduces fuel consumption by 2.2%, and improves the engine exergy efficiency by 2.7%, with only a 1.3% increase in rotating component weight. Ammonia and n-decane dual fuel precooling can achieve complementary advantages of high combustion heat value and high precooling heat sink, making it an ideal precooled engine scheme.

Numerical study on the integration of supersonic turbine guide vanes and three-dimensional hydrogen/air rotating detonation combustor

Hydrogen/air rotating detonation turbine engine is expected to become a new generation of aerospace power plant because of its compact structure, high cycle thermal efficiency, and superior thrust performance. It can also reduce fuel consumption, save energy, and reduce carbon emissions. However, the highly unsteady oscillation characteristics of the outlet flow of the rotating detonation combustor make it difficult to integrate the supersonic turbine with the rotating detonation combustor. In this paper, the supersonic turbine guide vanes are designed by the method of characteristics and Bessel parameterization and are integrated with three-dimensional hydrogen/air rotating detonation combustors for numerical studies. The effects of aligned mode and misaligned mode on the coupling of supersonic turbine guide vanes and rotating detonation combustor are discussed carefully. The results show that the supersonic turbine guide vanes can make the rotating detonation wave change from a single-wave mode to a double-wave alternating strength and weak propagation mode. It can effectively suppress the oscillation of the combustion chamber outlet airflow. In the aligned mode, the peak pressure at the outlet of the supersonic turbine is about 70% lower than that at the cascade inlet, the pressure oscillation amplitude is reduced by 93.33%, and the temperature amplitude is reduced by 23.81%; the average total pressure loss coefficient of the cascade is 11.63%. In the misaligned mode, compared with the cascade inlet, the peak value of the pressure signal at the cascade outlet decreases by about 50%, while the pressure oscillation amplitude decreases by about 33.33%, and the temperature oscillation amplitude decreases by 11.11%; the average total pressure loss coefficient of the cascade is 4.83%. The supersonic turbine guide vanes have a better suppression effect on the oscillation signal in the aligned mode, but the relative total pressure loss is relatively large. This is because that the oblique shock wave, channel shock wave, and supersonic turbine guide vanes interact to generate more complex wave system and secondary flow in the aligned mode. These features provide important reference information for the coupling of supersonic turbines and rotating detonation combustors.

Coupling study of supersonic turbine stage and two-dimensional hydrogen/air rotating detonation combustor

The hydrogen/air rotating detonation turbine engine has the advantages of self-supercharging, small entropy increase, high thermal efficiency, high thrust-to-weight ratio, low fuel consumption, and low carbon emissions. However, the high-frequency and high-speed oscillation characteristics of the outflow of the rotating detonation combustor pose challenges for the turbine to extract work efficiently. In this study, a supersonic turbine stage designed using the Python code of the method of characteristics coupled with a two-dimensional rotating detonation combustor is numerically investigated. The propagation characteristics of the detonation wave in the aligned mode and misaligned mode and interaction with the supersonic turbine stage are carefully discussed. The results show that the coupling of the supersonic turbine stage and the rotating detonation combustor will cause the detonation wave to change from a single-wave mode to a three-wave co-propagating mode. The Kelvin–Helmholtz instability of the slip line increases after passing through the induced oblique shock. In the multi-wave mode, the detonation wave is self-adaptive, and multiple detonation waves interact to automatically adjust the propagation velocity, intensity, and distance of each detonation wave, and finally achieve a dynamic balance. The supersonic turbine stage has good operating performance under the condition of rotating detonation flow, its power level can reach 110 kW, and the maximum stagnation adiabatic efficiency of the supersonic turbine stage can reach 86%. The supersonic turbine guide vanes can greatly reduce the oscillation amplitude of the incoming flow. In the aligned mode, the supersonic turbine guide vanes has a more obvious effect of suppressing the amplitude of the incoming flow. The total pressure loss of the supersonic turbine stage is smaller, and the supersonic turbine rotor can extract work more efficiently in the aligned mode. These findings provide a valuable reference for further research on the hydrogen/air rotating detonation turbine engine, ultimately leading to the practical application of an energy-saving, high-efficiency, and low-emission hydrogen/air rotating detonation turbine engine.

The Role of Endwall Shape Optimization in the Design of Supersonic Turbines for Rotating Detonation Engines

Abstract Rotating detonation engines (RDEs) are characterized by a thermodynamic cycle with an efficiency gain up to 15% at medium pressure ratios with respect to systems based on the conventional Joule–Bryton cycle. Multiple turbine designs can be considered, and this article deals with the supersonic inlet configuration. After having reviewed the main design steps of an exemplary RDE supersonic turbine, the article focuses on the considerable effects that endwall losses have on the performance of supersonic-inlet turbines and on the reasons why endwall contouring is strongly recommended for an efficient design. Parametric analyses, carried out by a novel in-house mean-line code validated against computational fluid dynamics (CFD), reveal that endwall friction losses contribute significantly to the overall stage loss. Endwall boundary layers also reduce the effective area, which can be critical for the self-starting capability of the supersonic channel. Therefore, a variable blade height geometry is necessary to extend the design space and guarantee a higher efficiency with respect to a constant-span configuration. The in-house CFD-based evolutionary shape optimization code was adapted to search for the optimal endwall shape for these unconventional machines. The optimal shape reduces shock losses and deviation angles and provides a significant gain in efficiency and work extraction. Finally, a novel technique is proposed to design the three-dimensional shape of the rotor based on the method of characteristics and tailored on the flow delivered by the stator.

Performance assessment for a novel supersonic turbine engine with variable geometry and fuel precooled: From feasibility, exergy, thermoeconomic perspectives

Aiming at the problem of turbine engine insufficient thrust caused by low efficiency at high Mach number, a three-axis variable geometry dual-mode turbine engine scheme based on ammonia and n-decane equivalent dual fuel precooling and adjustable low-pressure (LP) turbine guide vane is proposed, and parameters such as specific thrust price, relative thrust economic ratio, volumetric specific fuel consumption are defined to scientifically evaluate the feasibility of the scheme. The ammonia and n-decane dual fuel method is used to expand the efficient working range of the precooler. On the one hand, in the low speed section, the engine operates in the turbofan mode, adjusting the LP turbine guide vane can offset the adverse impact of the precooler low efficiency on the engine thrust performance under the low inlet total temperature, and improve the engine economy. On the other hand, the high speed section operates in the turbojet mode, and the equivalent fuel precooling combined with the LP turbine guide vane adjustment is conducive to improving the working efficiency and thrust performance of the turbojet engine. Feasibility, exergy and thermoeconomic analysis indicate that under different operating modes, due to different control plans, the influence of fuel ratio and heat transfer efficiency changes of dual fuel precooler on engine performance is opposite. For the typical climbing process with H = 0, Ma = 0 to H = 30 km, Ma = 5, this scheme can significantly increase the engine thrust, with a maximum increase of about 41%, but it increases the volume fuel consumption, reduces the engine exergy efficiency, exergy sustainability, and thermoeconomic under the turbofan mode. Therefore, the variable geometry turbine engine scheme based on dual fuel precooling and adjustable LP turbine guide vane has high application potential in supersonic commercial flight under the premise of slightly losing economy.

Speed of Sound Measurements in Dense Siloxane D6 Vapor at Temperatures up to 645 K by Means of a New Prismatic Acoustic Resonator.

Estimating the speed of sound for the dense vapor phase of D6 (dodecamethylcyclohexasiloxane, C12H36O6Si6) is particularly relevant to the study of nonideal compressible fluid dynamics (NICFD), the gas dynamics of fluids whose properties depart significantly from those related by the ideal gas model. If molecular complexity is sufficiently large, dense vapor flows may exhibit so-called nonclassical gasdynamic effects, and D6 is a candidate for experimental studies aimed at proving for the first time the existence of these exotic phenomena. More in general, speed of sound measurements in the dense vapor phase are important for NICFD applications: for example, complex organic compounds are employed as working fluids in organic Rankine cycle power plants and the correct prediction of the dense-vapor speed of sound is of paramount importance for the design of the supersonic turbines equipping these systems. This type of measurements is challenging and thus rare, especially in the case of complex organic molecules, given the high temperature at which they must be performed, close to the temperature at which the molecule thermally decomposes. Therefore, a new prismatic resonator (the OVAR, organic vapor acoustic resonator) has been conceived, designed, and realized. It is suitable for speed of sound measurements in the vapor phase of organic compounds at temperatures up to approximately 670 K and pressures up to about 1.5 MPa. The speed of sound of D6 (97.4% pure) has been measured along eight isotherms between 555 and 645 K and from nearly saturated density to densities close to ideal gas conditions (compressibility factor Z ≈ 0.95). The estimated relative uncertainty of these sound speed measurements is 0.14%. In addition, corresponding density values have been obtained with an estimated uncertainty between 0.2 kg·m-3 and 1.2 kg·m-3.

Large-eddy simulation study of rotating detonation supersonic turbine nozzle generated by the method of characteristics under oscillating incoming flow

Rotating detonation turbine engine is receiving considerable attention due to its' high cycle efficiency, outstanding thrust characteristics, self-pressurization, and energy-saving attributes. Conventional turbines are inefficient (30%) under rotating detonation inflow conditions. In order to obtain the turbine operating efficiently under the condition of rotating detonation inflow, this paper uses the method of characteristics and Bessel parameterization to design the blade profile of the rotating detonation supersonic turbine. The Large Eddy Simulation is used to numerically study the flow field characteristics of the supersonic turbine blade designed by the method of characteristics. The study found that the rotating detonation supersonic turbine guide vane can effectively reduce the pressure oscillation amplitude of the incoming flow to 25% of the original amplitude, and the main frequency (10 kHz) of the incoming flow occupies the main part of the flow field frequency. Second, the morphological evolution of the shock waves attenuates the adverse pressure gradient on the suction surface. The separation area of the suction surface slowly oscillates and attenuates, and is eventually confined to a small region. The wake accelerates and dissipates under the squeezing jet of the dovetail wave and the intense shearing action, forming a small wake area. The attenuation of large-scale separation gradually reduces the separation loss and wake loss, and the convergence and interaction of shock waves and the wake vortex significantly enhance the proportion of entropy production in the shock region. From the pressure coefficient and is entropic Mach number distributions, it is found that the blade load is mainly concentrated in the tail, and is minimized when the flow field becomes stable. These features provide a reference for the design of rotating detonation supersonic turbines and a deeper understanding of the flow field characteristics of rotating detonation turbine engines.

Unsteadiness of shock-boundary layer interactions in a Mach 2.0 supersonic turbine cascade

Shock-boundary layer interactions in a supersonic turbine cascade are investigated using a wall-resolved large eddy simulation. Results reveal that near-wall streaky structures (red color) drive the motion of the suction side separation bubble (blue color), which in turn promotes oscillations of the reattachment shock (black color) and shear layer flapping.

Design and optimization of supersonic turbines for detonation combustors

Detonation-based engines offer a potential surge in efficiency for compact thermal power systems. However, these cycles require ad-hoc components adapted to the high outlet velocity from the detonation combustors. This paper presents the design methodology of turbine stages suitable for supersonic inlet conditions and provides a detailed analysis of optimized turbine geometries. A reduced-order solver examines the supersonic blade rows’ functional design space, quantifies the turbine’s non-isentropic performance, and budgets the turbine loss for different optimized leading-edge designs and chord to pitch ratios. The shock-wave interactions were identified as the predominant contributor to turbine losses, and optimal pitch-chord ratios were determined for various inlet Mach numbers. Finally, with this tool, the specific-power output for a wide range of design configurations was computed; and the metal angle that ensures flow starting and maximizes power extraction was calculated. The detailed numerical study describes the flow interactions in a supersonic turbine and offers new correlations to guide the design of future supersonic turbines.

Effect of Centrifugal Force on Gas Flow in a Supersonic Turbine

Abstract To clarify the loss mechanism in supersonic turbines, the authors experimentally and numerically studied the flow condition between the rotor blades of a supersonic turbine for liquid rocket engines. The entrance Mach number was 1.94, and the turning angle of the blades was 120 deg. Shock waves were created at the leading edge of the blade. The Mach number and total pressure decreased in the passage between the blades, where the flow condition was restricted by the centrifugal force. Such degradation, which has also been reported in past studies, is an inherent feature of high-speed, large-turning-angle blades. Here, it was found that a convergent–divergent configuration of the passage between the blades suppressed the performance degradation of the supersonic turbine.

Design and Parametric Analysis of a Supersonic Turbine for Rotating Detonation Engine Applications

Pressure gain combustion is a promising alternative to conventional gas turbine technologies and within this class the Rotating Detonation Engine has the greatest potential. The Fickett–Jacobs cycle can theoretically increase the efficiency by 15% for medium pressure ratios, but the combustion chamber delivers a strongly non-uniform flow; in these conditions, conventionally designed turbines are inadequate with an efficiency below 30%. In this paper, an original mean-line code was developed to perform an advanced preliminary design of a supersonic turbine; self-starting capability of the supersonic channel has been verified through Kantrowitz and Donaldson theory; the design of the supersonic profile was carried out employing the Method of Characteristics; an accurate evaluation of the aerodynamic losses has been achieved by considering shock waves, profile, and mixing losses. Afterwards, an automated Computational Fluid Dynamics (CFD) based optimization process was developed to find the optimal loading condition that minimizes losses while delivering a sufficiently uniform flow at outlet. Finally, a novel parametric analysis was performed considering the effect of inlet angle, Mach number, reaction degree, peripheral velocity, and blade height ratio on the turbine stage performance. This analysis has revealed for the first time, in authors knowledge, that this type of machines can achieve efficiencies over 70%.

Design of a supersonic turbine for the organic Rankine cycle system

Design of a supersonic turbine for the organic Rankine cycle system

Nonlinear model predictive control of organic Rankine cycles for automotive waste heat recovery: Is it worth the effort?

Using organic Rankine cycles (ORC) for waste heat recovery in vehicles promises significant reductions in fuel consumption. Controlling the organic Rankine cycle, however, is difficult due to the highly transient exhaust gas conditions. To tackle this issue, nonlinear model predictive control (NMPC) has been proposed and approximate NMPC solutions have been investigated to reduce computational demand. Herein, we compare (i) an idealized economic NMPC (eNMPC) scheme as a benchmark to (ii) a NMPC enforcing minimal superheat and (iii) a PI controller with dynamic feed-forward term (PI-ff) in a control case study with highly transient disturbances. We show that, for an ORC system with supersonic turbine, the economic control problem can be reduced to a single-input single-output superheat tracking problem combined with a decoupled steady-state real-time optimization (RTO) of turbine operation, assuming an idealized condenser. Our results indicate that the NMPC enforcing minimal superheat provides good control performance with negligible losses in average power compared to the full solution of the economic NMPC problem and that even PI-ff only results in marginal losses in average power compared to the model-based controllers.

Amplification of Operational Uncertainty Induced by Nonideal Flows in Supersonic Turbine Cascades

Abstract In high-temperature transcritical organic Rankine cycles (ORCs), the expansion process may take place in the neighborhood of the thermodynamic critical point. In this region, many organic fluids feature a value of the fundamental derivative of gas dynamics Γ that is less than unity. As a consequence, severe nonideal gas-dynamic effects can be possibly observed. Examples of these nonideal effects are the nonmonotonic variation of the Mach number along an isentropic expansion, oblique shocks featuring an increase of the Mach number, and a significant dependence of the flow field on the upstream total state. To tackle this latter nonideal effect, an uncertainty-quantification strategy combined with Reynolds-averaged flow simulations is devised to evaluate the turbine performance in presence of operational uncertainty. The results clearly indicate that a highly nonideal expansion process leads to an amplification of the operational uncertainty. Specifically, given an uncertainty in the order of 1% in cycle nominal conditions, the mass flow rate and cascade losses vary ±4% and ±0.75 percentage points, respectively. These variations are four and six times larger than those prompted by an ideal-like expansion process. The flow delivered to the first rotating cascade is severely altered as well, leading to local variations in the rotor incidence angle up to 10 deg. A decomposition of variance contributions reveals that the uncertainty in the upstream total temperature is mainly responsible for these variations. Finally, the understanding of the physical mechanism behind these changes allows us to generalize the present findings to other organic-fluid flows.

Numerical Simulation of Shock Wave-Boundary Layer Interaction with Supersonic Film Cooling

As an effective external cooling method, film cooling can significantly reduce the wall temperature of hot end components of aero engines and large liquid rocket engines. With the development of supersonic turbine blade, liquid rocket engine thrust chamber and ramjet thermal protection technology, the impact of shock wave-boundary layer interaction on the cooling effectiveness of film cooling is a problem that must be considered in the engine heat transfer design. In this paper, the two-dimensional supersonic film cooling is simulated based on the plane slot model, and the blowing ratio is similar to the actual engine. Impact of wave structures on film cooling was studied, and flow loss was evaluated by total pressure and entropy parameter to investigate the film cooling efficiency and uniformity distribution. Physical mechanism of shock wave-boundary layer interaction and its impact on film form was revealed. It shows that the shock wave leads to dramatic changes in flow parameters. Boundary layer separation is found due to the shock, and significantly affects the cooling efficiency. The blowing ratio is an important factor affecting the cooling efficiency. With blowing ratio increasing, local film cooling effectiveness and average cooling efficiency are improved and cooling uniformity decreases. However, the shock wave will cause the temperature of the insulation wall to rise, resulting in decreased cooling efficiency.

Design, Optimization, and Analysis of Supersonic Radial Turbines

Abstract New compact engine architectures such as pressure gain combustion require ad hoc turbomachinery to ensure an adequate range of operation with high performance. A critical factor for supersonic turbines is to ensure the starting of the flow passages, which limits the flow turning and airfoil thickness. Radial outflow turbines inherently increase the cross section along the flow path, which holds great potential for high turning of supersonic flow with a low stage number and guarantees a compact design. First, the preliminary design space is described. Afterward a differential evolution multi-objective optimization with 12 geometrical design parameters is deducted. With the design tool autoblade 10.1, 768 geometries were generated and hub, shroud, and blade camber line were designed by means of Bezier curves. Outlet radius, passage height, and axial location of the outlet were design variables as well. Structured meshes with around 3.7 × 106 cells per passage were generated. Steady three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) simulations, enclosed by the k-omega shear stress transport turbulence model were solved by the commercial solver CFD++. The geometry was optimized toward low entropy and high-power output. To prove the functionality of the new turbine concept and optimization, a full wheel unsteady RANS simulation of the optimized geometry exposed to a nozzled rotating detonation combustor (RDC) has been performed and the advantageous flow patterns of the optimization were also observed during transient operation.