THE PROBLEM OF HEAT ADDITION IN DUCTS

1946 ◽  
Vol 24f (4) ◽  
pp. 272-286 ◽  
Author(s):  
D. G. Samaras
Keyword(s):  
Mach Number ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Frictional Coefficient ◽  
Extreme Value ◽  
Heat Addition ◽  
Total Head ◽  
Head Pressure ◽  
Temperature And Pressure ◽  
Head Temperature

The one-dimensional problem of heat addition in a duct with friction losses is examined.The variation of the total head temperature and pressure ratio, as well as the static ones, have been obtained as functions of the Mach number and the frictional coefficient. It is shown that an extreme value of the total head temperature and pressure ratio is obtained when the Mach number equals unity. An extreme value of the static temperature is obtained at Mach numbers smaller or larger than unity, depending on the form of the duct.It is concluded that when the total head pressure losses are to be kept low, a diverging duct must be used. The frictional losses have a convergence effect on the flow. When the friction losses are disregarded, no total head pressure loss will result if the area variation of the duct is such as to ensure a constant Mach number.The results and conclusions found may be useful in the design of combustion chambers for gas turbines, propulsive ducts, and rockets.

scholarly journals 100s and 1000s Mw Open and Semi-Closed Cycle Gas Turbines for Base and Peak Operation

1974 ◽  
Author(s):  
V. V. Uvarov ◽  
V. S. Beknev ◽  
E. A. Manushin
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Simple Cycle ◽  
Inlet Temperature ◽  
Closed Cycle ◽  
Turbine Inlet Temperature ◽  
Unit Capacity ◽  
Temperature And Pressure ◽  
Different Levels

There are two different approaches to develop the gas turbines for power. One can get some megawatts by simple cycle or by more complex cycle units. Both units require very different levels of turbine inlet temperature and pressure ratio for the same unit capacity. Both approaches are discussed. These two approaches lead to different size and efficiencies of gas turbine units for power. Some features of the designing problems of such units are discussed.

GAS DYNAMIC TREATMENT OF EXOTHERMIC AND ENDOTHERMIC DISCONTINUITIES

1948 ◽  
Vol 26a (1) ◽  
pp. 1-21 ◽  
Author(s):  
D. G. Samaras
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Static Pressure ◽  
Density Ratio ◽  
Area Ratio ◽  
Compression Shock ◽  
Normal Velocity ◽  
Gas Dynamic ◽  
Total Head ◽  
Head Pressure

Endothermic and exothermic processes in gas dynamic flows taking place in a very narrow zone may be considered as discontinuities. The variation of the static and total head density ratio, pressure ratio, and temperature ratio as well as the angle of deviation, area ratio, and exit normal Mach number have been found as functions of the entry normal Mach number and of heat addition. In addition to these, some other useful quantities such as the area ratio parameter, the difference of the square of velocities, and the normal velocity product have been evaluated. It was found that, in a discontinuity, heat can be added until the exit normal Mach number reaches unity (choking). Depending on the entry normal Mach number, only a limited amount of heat can be added at the discontinuity. An exothermic discontinuity behaves as an expansion when the entry normal Mach number is subsonic, and it is accompanied by a drop in static pressure, density, and total head pressure. An exothermic discontinuity behaves as a compression shock wave when the entry normal Mach number is supersonic, and it is accompanied by an increase in static pressure and density and a decrease in total head pressure. An endothermic discontinuity behaves always as a compression shock wave, and it is accompanied by an increase in static density, pressure, and total head pressure. It is hoped that the results and conclusions found may be useful in a better understanding of many nearly discontinuous phenomena such as flame fronts, condensation and evaporation fronts, and other similar problems.

scholarly journals Feasibility of a Helium Closed-Cycle Gas Turbine for UAV Propulsion

2020 ◽  
Vol 11 (1) ◽  
pp. 28
Author(s):  
Emmanuel O. Osigwe ◽  
Arnold Gad-Briggs ◽  
Theoklis Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Low Noise ◽  
Closed Cycle ◽  
Propulsion Systems ◽  
Component Design ◽  
Readiness Level ◽  
Aerial Vehicle ◽  
Turbine Design

When selecting a design for an unmanned aerial vehicle, the choice of the propulsion system is vital in terms of mission requirements, sustainability, usability, noise, controllability, reliability and technology readiness level (TRL). This study analyses the various propulsion systems used in unmanned aerial vehicles (UAVs), paying particular focus on the closed-cycle propulsion systems. The study also investigates the feasibility of using helium closed-cycle gas turbines for UAV propulsion, highlighting the merits and demerits of helium closed-cycle gas turbines. Some of the advantages mentioned include high payload, low noise and high altitude mission ability; while the major drawbacks include a heat sink, nuclear hazard radiation and the shield weight. A preliminary assessment of the cycle showed that a pressure ratio of 4, turbine entry temperature (TET) of 800 °C and mass flow of 50 kg/s could be used to achieve a lightweight helium closed-cycle gas turbine design for UAV mission considering component design constraints.

Design and Analysis of a High Pitch to Chord Ratio Cascade Representative of Ducted Propfans

1991 ◽  
Author(s):  
A. Weber ◽  
W. Steinert ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
High Pitch ◽  
Integral Boundary ◽  
Pressure Loss Coefficient ◽  
Distribution Boundary ◽  
Very High

Efforts to reduce the specific fuel consumption of a modern aero engine focus in particular on increasing the by-pass ratio beyond the current level of around 5. One concept is the counterrotating shrouded propfan operating at low overall pressure ratio and having only very few fan blades of extremely high pitch/chord ratios. The relative inlet Mach numbers cover a range from 0.7 at the hub to 1.1 at the tip section of the first rotor. A propfan cascade was designed by taking into account two characteristic features of a propfan blade-blade section: • a very high pitch/chord ratio of s/c = 2.25 • an inlet Mach number of M1 = 0.90 which leads to transonic flow conditions inside the blade passage In the design process a profile generator and a quasi-3D Euler solver were used iteratively to optimize the profile Mach number distribution. Boundary layer behavior was checked with an integral boundary layer code. The cascade design was verified experimentally in the transonic cascade wind tunnel of DLR at Cologne. The extensive experimental results confirm the design goal of roughly 5 degree flow turning. A total pressure loss coefficient of less than 1.5% was measured at design conditions. This validates the very high efficiency level the propfan concept is calling for. A 2D Navier-Stokes flow analysis code yields good results in comparison to the experimental ones.

Numerical Modeling of Heat Transfer and Pressure Losses for Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

2007 ◽  
Author(s):  
Lamyaa A. El-Gabry
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Gas Turbine ◽  
Computational Study ◽  
Numerical Models ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Pressure Losses ◽  
Blade Tip ◽  
Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

Investigation of the 3D Unsteady Rotor Pressure Field in a HP Turbine Stage

2002 ◽  
Author(s):  
E. Valenti ◽  
J. Halama ◽  
R. De´nos ◽  
T. Arts
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Pressure Field ◽  
Pressure Ratio ◽  
Turbine Stage ◽  
Time Resolved ◽  
Mach Number Distribution ◽  
Rotor Surface ◽  
Exit Mach Number ◽  
Unsteady Pressure Measurements

This paper presents steady and unsteady pressure measurements at three span locations (15, 50 and 85%) on the rotor surface of a transonic turbine stage. The data are compared with the results of a 3D unsteady Euler stage calculation. The overall agreement between the measurements and the prediction is satisfactory. The effects of pressure ratio and Reynolds number are discussed. The rotor time-averaged Mach number distribution is very sensitive to the pressure ratio of the stage since the incidence of the flow changes as well as the rotor exit Mach number. The time-resolved pressure field is dominated by the vane trailing edge shock waves. The incidence and intensity of the shock strongly varies from hub to tip due to the radial equilibrium of the flow at the vane exit. The decrease of the pressure ratio attenuates significantly the amplitude of the fluctuations. An increase of the pressure ratio has less significant effect since the change in the vane exit Mach number is small. The effect of the Reynolds number is weak for both the time-averaged and the time-resolved rotor static pressure at mid-span, while it causes an increase of the pressure amplitudes at the two other spans.

Effect of Recirculation Device With Counter Swirl Vane on Performance of High Pressure Ratio Centrifugal Compressor

2011 ◽  
Author(s):  
Hideaki Tamaki
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Negative Slope ◽  
Tip Leakage Flow ◽  
Recirculation Flow ◽  
Operating Range ◽  
High Pressure Ratio

Centrifugal compressors used for turbochargers need to achieve a wide operating range. The author has developed a high pressure ratio centrifugal compressor with pressure ratio 5.7 for a marine use turbocharger. In order to enhance operating range, two different types of recirculation devices were applied. One is a conventional recirculation device. The other is a new one. The conventional recirculation device consists of an upstream slot, bleed slot and the annular cavity which connects both slots. The new recirculation device has vanes installed in the cavity. These vanes were designed to provide recirculation flow with negative preswirl at the impeller inlet, a swirl counterwise to the impeller rotational direction. The benefits of the application of both of the recirculation devices were ensured. The new device in particular, shifted surge line to a lower flow rate compared to the conventional device. This paper discusses how the new recirculation device affects the flow field in the above transonic centrifugal compressor by using steady 3-D calculations. Since the conventional recirculation device injects the flow with positive preswirl at the impeller inlet, the major difference between the conventional and new recirculation device is the direction of preswirl that the recirculation flow brings to the impeller inlet. This study focuses on two effects which preswirl of the recirculation flow will generate. (1) Additional work transfer from impeller to fluid. (2) Increase or decrease of relative Mach number. Negative preswirl increases work transfer from the impeller to fluid as the flow rate reduces. It increases negative slope on pressure ratio characteristics. Hence the recirculation flow with negative preswirl will contribute to stability of the compressor. Negative preswirl also increases the relative Mach number at the impeller inlet. It moves shock downstream compared to the conventional recirculation device. It leads to the suppression of the extension of blockage due to the interaction of shock with tip leakage flow.

scholarly journals Methanol Dissociative Intercooling in Gas Turbines

1982 ◽  
Author(s):  
M. F. Bardon ◽  
J. A. C. Fortin
Keyword(s):  
Carbon Monoxide ◽  
Theoretical Analysis ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Unit Mass ◽  
Cycle Model ◽  
Optimum Pressure ◽  
Heating Value ◽  
Compressor Work

This paper examines the possibility of injecting methanol into the compressor of a gas turbine, then dissociating it to carbon monoxide and hydrogen so as to cool the air and reduce the work of compression, while simultaneously increasing the fuel’s heating value. A theoretical analysis shows that there is a net reduction in compressor work resulting from this dissociative intercooling effect. Furthermore, by means of a computer cycle model, the effects of dissociation on efficiency and work per unit mass of airflow are predicted for both regenerated and unregenerated gas turbines. The effect on optimum pressure ratio is examined and practical difficulties likely to be encountered with such a system are discussed.

scholarly journals Industrial Gas Turbine Uprates: Tips, Tricks and Traps

1997 ◽  
Author(s):  
T. L. Ragland
Keyword(s):  
Gas Turbine ◽  
Output Power ◽  
Gas Turbines ◽  
Customer Value ◽  
Low Cost ◽  
Pressure Ratio ◽  
Original Design ◽  
Advantages And Disadvantages ◽  
Industrial Gas Turbine ◽  
Cycle Efficiency

After industrial gas turbines have been in production for some amount of time, there is often an opportunity to improve or “uprate” the engine’s output power or cycle efficiency or both. In most cases, the manufacturer would like to provide these uprates without compromising the proven reliability and durability of the product. Further, the manufacturer would like the development of this “Uprate” to be low cost, low risk and result in an improvement in “customer value” over that of the original design. This paper describes several options available for enhancing the performance of an existing industrial gas turbine engine and discusses the implications for each option. Advantages and disadvantages of each option are given along with considerations that should be taken into account in selecting one option over another. Specific options discussed include dimensional scaling, improving component efficiencies, increasing massflow, compressor zero staging, increasing firing temperature (thermal uprate), adding a recuperator, increasing cycle pressure ratio, and converting to a single shaft design. The implications on output power, cycle efficiency, off-design performance engine life or time between overhaul (TBO), engine cost, development time and cost, auxiliary requirements and product support issues are discussed. Several examples are provided where these options have been successfully implemented in industrial gas turbine engines.

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