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Detonation gas turbines: during the early days of gas turbine development in the 1930's, getting combustors to work efficiently took a lot of ingenuity and effort.

Both inventors of the aviation gas turbine--Hans von Ohain in Germany and Frank Whittle in England--first resorted to hydrogen fuel with its more forgiving high flame speed, before they solved liquid fuel combustion problems. The first industrial gas turbine, Brown Boveri's 1939 Neuchatel Switzerland 4 MW unit, had a single, very large and long combustor, reflective of the nascent state of liquid fuel combustion technology at the time [1].

By contrast, modern gas turbine combustors are now compact, robust, tolerant of a wide variety of fuels, and provide the highest combustion intensities (rate of energy released per unit volume, as high as 75,000 Btu/s [ft.sup.3]). They heat gas path flow in a near constant pressure process, to thermodynamically approximate the "energy in" isobaric part of a Brayton cycle.

Currently, all flame processes in gas turbines fall within the combustion category of deflagration [2]. This is the term describing subsonic combustion propagating through heat transfer. In a typical gas turbine combustor this subsonic heat addition leads to near constant pressure process in the gas path flow.

In contrast to deflagration, detonation combustion involves a supersonic flow, with the chemical reaction front accelerating, driving a shock wave system in its advancement. Detonations, usually associated with explosives and explosions, produce very high pressures and high velocities. Thus, if harnessed in a gas turbine combustor, detonation could reduce the need for some expensive compressor and turbine hardware, lighten engine weight and possibly increase gas turbine power output.

In recent years there have been growing research activities to utilize detonation to combust fuel/air mixtures in gas turbine combustors. In the 1990s detonation based power concepts began with pulse detonation engines (PDEs), and have now moved into the continuous detonation mode, termed rotating detonation engines (RDEs). Currently, RDE experimental and computational activities are taking place at the U.S. Naval Research Laboratory [3], the U.S. Air Force Research Laboratory [4], the University of Texas at Arlington [5] and here at the University of Connecticut [6].

The RDE concept for a detonation combustor consists of a concentric circular-tube annulus, where a premixed fuel/air mixture is injected axially at the annulus entrance. Once initiated, a detonation propagates at supersonic speeds circumferentially around the annulus, moving downstream in an axial direction with at first rapidly rising pressures and then falling, as the swirling flow approaches the exit of the annulus. Challenges in the design included high heat transfer loads to the annular surfaces and decelerating the exiting flow to subsonic velocities, while minimizing aerodynamic losses.

A possible application of RDE to a gas turbine is shown in Fig. 1, taken from Nordeen [6]. Figure la) is a sketch of a conventional twin-spool gas turbine with various stations labeled. Figure lb) is the sketch of an equivalent engine where the entire high pressure spool and conventional combustor have been replaced with an RDE combustor. Conceptually, the pressure rise of the high compressor in la) is brought about by the supersonic shock system in the RDE combustor in Fig. lb). The Fig. la) high turbine is also eliminated, since there is no high compressor to be driven. Thus, the RDE design in Fig. lb) has one less spool, resulting in less weight and a shorter engine.

Figure 2, taken from Schwer and Kailasanath [3], is a pressure P (normalized by ambient pressure [P.sub.ref]) volume v (normalized by specific volume [v.sub.ref]) diagram for an ideal fluid particle moving through each of the gas turbines in Fig. 1. Positions on each Pv curve are indicated by corresponding station numbers in Fig. 1.

The two-spool gas turbine in Fig. la) is represented as the Brayton cycle with an operating compressor pressure ratio of 10, between positions 2.0 and 3.0.

The single spool RDE gas turbine of Fig. lb) is represented by a detonation cycle which accounts for the supersonic features of the heat addition, starting at station 2.5'. This is at the exit of the low pressure compressor (LPC) which has an operating compressor pressure ratio of 2. (The shape of the detonation cycle is similar to that of a constant volume heat addition Humphrey cycle [7].) The exit of the RDE combustor is at station 4.5'.

As many of us learned in our thermo courses, the area enclosed by a cycle on a Pv diagram is a measure of the work out of the device, for quasi-static processes. In Fig. 2, the detonation cycle area exceeds that of the Brayton cycle. The exceeding area has been labeled as additional work by Schwer and Kailasanath, providing the promise of an increased output for a RDE gas turbine.

Continued research and development by the RDE technical community is needed to see if the promise of improved performance and downsized turbomachinery for a detonation cycle is real. However, any skeptics should remember the travail and tribulation (alluded to at our beginning here) early gas turbine developers experienced to perfect the deflagration combustors of today.


[1.] Langston, Lee S., 2010, "Visiting the Museum of the World's First Gas Turbine Power plant", Global Gas Turbine News, April, p. 3.

[2.] Lefebvre, Arthur H., 1983, Gas Turbine Combustion, McGraw-Hill, p. 33.

[3.] Schwer, D.A., and Kailasanath, K., 2011, "Rotating Detonation-Wave Engines", 2011 NRL Review, pp. 90-94.

[4.] Tellefsen, Jonathan R., King, Paul I., Schauer, Frederick R., and Hoke, John L, 2012, "Analysis of an RDE with Convergent Nozzle in Preparation for Turbine Integration", American Institute of Aeronautics and Astronautics, paper AIAA 2012-0773, pp. 1-11.

[5.] Braum, Eric M., Lu, Frank K., Wilson, Donald R., and Camberos, Jose A., 2013, "Airbreathing Rotating Detonation Wave Engine Cycle Analysis", Aerospace Science and Technology, Vol. 27, pp. 201-208.

[6.] Nordeen, Craig A., 2013 "Thermodynamics of a Rotating Detonation Engine", Ph.D. Thesis, University of Connecticut.

[7.] Heiser, William H., and Pratt, David T, 2002 "Thermodynamic Cycle Analysis of Pulse Detonation Engines", Journal of Propulsion and Power, Vol. 18, No. 1, pp. 68-76.

By Dr. Lee S. Langston, Professor Emeritus of Mechanical Engineering, University of Connecticut

Lee Langston is former editor of the ASME Journal of Engineering for Gas Turbines and Power and has served on the ASME IGTI Board as both Chair and Treasurer.
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Title Annotation:AS THE TURBINE TURNS ...
Author:Langston, Lee S.
Publication:Mechanical Engineering-CIME
Geographic Code:1USA
Date:Dec 1, 2013
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