Supersonic Free-Jet Combustion in a Ramjet Burner

TLDR: In this paper, a new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described, where a variable combustor exit aperture is required, and the need for fuel staging to accommodate the combustion process is eliminated.

Abstract: A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is governed largely by the same physics as its classical counterpart. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode configurations. Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD) analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and 12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions between the free-jet and recirculating combustion chamber flows. Performance based on these CFD results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these interactions. These differences are quantified and discussed. Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation on combustion chamber pressure, flow structure, and performance. Calculations with constant temperature walls were also done to evaluate heat flux and overall heat loads. Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and combustion chamber pressurization. Also recommended is an examination of system-level aspects such as weight, thermal management and rocket integration as well as alternate geometries and variable geometry schemes.

Figures

Figure 14.—Performance over Mach range.

Figure 15.—Effect of throat area variation on static pressure ratio for Mach 8 flight condition.

Figure 7.—GASP results, static temperature contours.

Figure 20.—Heat load versus flight Mach number for various wall temperatures, compared to the estimates of Reference 7.

Figure 8.—Ethylene mass fraction versus axial distance for various flight conditions.

Figure 9.—Mass-averaged pressure distributions for various flight conditions.

Full text

Charles J. Trefny and Vance F. Dippold III
Glenn Research Center, Cleveland, Ohio
Supersonic Free-Jet Combustion in a Ramjet Burner
NASA/TM—2010-216932
November 2010
AIAA–2010–6643

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Charles J. Trefny and Vance F. Dippold III
Glenn Research Center, Cleveland, Ohio
Supersonic Free-Jet Combustion in a Ramjet Burner
NASA/TM—2010-216932
November 2010
AIAA–2010–6643
National Aeronautics and
Space Administration
Glenn Research Center
Cleveland, Ohio 44135
Prepared for the
46th Joint Propulsion Conference and Exhibit
cosponsored by AIAA, ASME, SAE, and ASEE
Nashville, Tennessee, July 25–28, 2010

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This work was sponsored by the Fundamental Aeronautics Program
at the NASA Glenn Research Center.
Level of Review: This material has been technically reviewed by technical management.

NASA/TM—2010-216932 1
Supersonic Free-Jet Combustion in a Ramjet Burner
Charles J. Trefny and Vance F. Dippold III
National Aeronautics and Space Administration
Glenn Research Center
Cleveland, Ohio 44135
Abstract
A new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number
range is described. Subsonic combustion mode is similar to that of a traditional ram combustor which
allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode
scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the
subsonic combustion chamber to a variable nozzle. The maximum flight Mach number of this scheme is
governed largely by the same physics as its classical counterpart. Although a variable combustor exit
aperture is required, the need for fuel staging to accommodate the combustion process is eliminated.
Local heating from shock-boundary-layer interactions on combustor walls is also eliminated. Given the
parallel nature of the present scheme, overall flowpath length is less than that of present dual-mode
configurations.
Cycle analysis was done to define the flowpath geometry for computational fluid dynamics (CFD)
analysis, and then to determine performance based on the CFD results. CFD results for Mach 5, 8, and
12 flight conditions indicate stable supersonic free-jet formation and nozzle reattachment, thereby
establishing the basic feasibility of the concept. These results also reveal the structure of, and interactions
between the free-jet and recirculating combustion chamber flows. Performance based on these CFD
results is slightly less than that of the constant-pressure-combustion cycle analysis primarily due to these
interactions. These differences are quantified and discussed.
Additional CFD results at the Mach 8 flight condition show the effects of nozzle throat area variation
on combustion chamber pressure, flow structure, and performance. Calculations with constant
temperature walls were also done to evaluate heat flux and overall heat loads.
Aspects of the concept that warrant further study are outlined. These include diffuser design, ramjet
operation, mode transition, loss mechanisms, and the effects of secondary flow for wall cooling and
combustion chamber pressurization. Also recommended is an examination of system-level aspects such as
weight, thermal management and rocket integration as well as alternate geometries and variable geometry
schemes.
Nomenclature
A Cross-sectional area
C
f
Friction coefficient
M Mach number
P Pressure
r Radial distance
x Axial distance
y
+
Nondimensional turbulent wall distance
Z Altitude
Subscripts
0 Freestream
1 Cylindrical inflow section exit station
2 Combustion chamber inlet station

Frequently asked questions

Q1. What are the contributions in this paper?

In this paper, a new dual-mode ramjet combustor concept intended for operation over a wide flight Mach number range is described. 

Q2. How much friction coefficient was used to account for the effect of shear layer formation on the required?

A combustor friction coefficient of 0.0025 was used to account approximately for the effect of shear layer formation on the required nozzle throat area. 

Q3. What can be done to reduce shock losses due to the periodic wave structure?

Shock losses due to the periodic wave structure can be reduced by better tailoring the mixing and combustion process, thereby mitigating the entry interaction. 

Q4. What can be done to reduce the viscous loss in the freejet?

Viscous losses may be reduced by reducing the combustion chamber wetted area, and the momentum transfer from the freejet to the recirculation zone. 

Q5. What are the factors that govern the combustion area ratio of a scramjet engine?

Other factors that must be considered include separation of boundary layers due to adverse pressure gradients, intense local heating at reattachment points and shock impingements, and fuel staging or variable geometry to accommodate the variation of combustion area ratio with freestream stagnation enthalpy. 

Q6. How much reduction in throat area is required between Mach 5 and 12?

Nozzle throat area variation requirements could also be relieved by a reduction in fuel-air ratio at the lower flight Mach numbers at the expense of net thrust. 

Q7. What was the surface area assumed for the calculations?

The surface area assumed for the calculations herein was a conical frustum extending from the diffuser exit to the nozzle throat. 

Q8. What is the way to evaluate the merit of the present concept?

thermal management, structural design, and weight must be considered in order to assess the overall merit of the present concept. 

Q9. How many percent of the thrust deficits were caused by shock and viscous losses?

Shock and viscous losses resulted in net thrust deficits of 8.6 percent at the Mach 8 flight condition and 24 percent at Mach 12. 

Q10. How many iterations did it take to converge the solutions?

Further solutions (i.e., nozzle throat area variation calculations, cooled-wall calculations) were then restarted from the baseline solutions to reduce computational costs, typically reconverging in 10,000 to 15,000 iterations. 

Q11. Where is the aft region of the scramjet flowpath?

a thermally-choked combustion process is established in the aft regions of the scramjet flowpath where the cross-sectional areas are greatest. 

Q12. What is the simplest way to extend the flight Mach number range of a scramjet?

In order to extend the operable flight Mach number range of the scramjet engine downward, toward the upper limit for turbojets, “dual-mode” operation was introduced by Curran, et al. in a 1972 patent (Ref. 4). 

Q13. What is the need for fuel staging to accommodate the combustion process?

Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. 

Q14. What is the effect of the free-jet combustor on the combustion chamber?

During this mode of operation, the propulsive stream is not in contact with the combustor walls, and equilibrates to the combustion chamber pressure. 

Q15. What is the difference between a traditional and a subsonic ram combu?

Subsonic combustion mode is similar to that of a traditional ram combustor which allows operation at higher efficiency, and to lower flight Mach numbers than current dual-mode scramjets.