Experimental Investigation of Combustion and Propulsion ...

Defence Research and Recherche et développement Development Canada pour la défense Canada

Experimental Investigation of Combustion and Propulsion for Shock-Induced Combustion Ramjets A. Higgins Reactive Energetics Contract Scientific Authority: S. Murray, DRDC Suffield

The scientific or technical validity of this Contract Report is entirely the responsibility of the contractor and the contents do not necessarily have the approval or endorsement of Defence R&D Canada.

Defence R&D Canada

Contract Report

DRDC Suffield CR 2007-249

December 2006

Experimental Investigation of Combustion and Propulsion for Shock-Induced Combustion Ramjets A. Higgins Reactive Energetics 1678 Alexis Nihon Montreal QC H4R 2W2 Contract Number: W7702-05-R063/001/EDM Contract Scientific Authority: S. Murray (403-544-4729) The scientific or technical validity of this Contract Report is entirely the responsibility of the contractor and the contents do not necessarily have the approval or endorsement of Defence R&D Canada.

Defence R&D Canada – Suffield Contract Report DRDC Suffield CR 2007-249 December 2006

Experimental Investigation of Combustion and Propulsion for

Shock-Induced Combustion Ramjets

FINAL REPORT For work carried out under Contract No. W7702-05R063/001/EDM

Submitted by: Andrew Higgins

1678 Alexis Nihon, Montreal, Quebec H4R 2W2

Reactive Energetics

Submitted to: Dr. Stephen Murray Head/Threat Assessment Group Defence R&D Canada - Suffield PO Box 4000, Stn. Main Medicine Hat, AB T1A 8K6

01 December 2006

1

Experimental Investigation of Combustion and Propulsion for

Shock-Induced Combustion Ramjets

Executive Summary An investigation of the shock-induced-combustion ramjet propulsive cycle was carried out in the 38-mm-bore ram accelerator facility at the University of Washington. Titanium-alloy shcramjet projectiles were launched into reactive propellants at Mach numbers greater than 5.5 to determine if the combustion process could be shock initiated and stabilized, what levels of thrust can be generated, the reactivity of the projectile material in hypersonic flow, and to evaluate the efficacy of investigating the operating characteristics of hypersonic propulsive cycles using gun-launched projectiles. Experiments with four different projectile configurations were carried out in methane- and ethane-based propellants with and without carbon dioxide diluent. Positive acceleration was observed in CH4/O2/CO2 and C2H6/O2 propellants in the Mach range of 5.5-7 (1.7-2.1 km/s) for distances of up to 6 meters. In the majority of cases, the acceleration process was terminated by either unstart, cruise at constant velocity (i.e., thrust equal drag), or wave fall off. Sustained accelerations greater than 9000 g and average specific thrust 150 N*s/kg were achieved in these experiments. The hypersonic projectile drag coefficient (based on projected frontal area of the projectile) in the Mach range of 5 – 6 was determined to be CD = 0.09 from experiments in which the propellant did not ignite (wave fall-off).

Details of the experimental apparatus and procedures used in this investigation are described in the following. Key results of 18 pertinent hypersonic shcramjet experiments are presented and suggestions for follow on research are discussed. Velocity- and Mach-distance data plots and corresponding tube wall data from pressure and EM sensors for these experiments are provided in Appendices A and B, respectively. An abbreviated shot log having pertinent details of the each testing configuration and outcome for all experiments conducted in this program is provided in Appendix C.

2

Table of Contents Executive Summary................................................................... 1 Table of Contents ...................................................................... 2 Introduction................................................................................ 3 Experimental Apparatus and Procedure.................................... 3 Experimental Results................................................................. 5

Single Stage Experiment.................................................. 6 Projectile Geometry Variations....................................... 10 Carbon Dioxide Dilution Effects...................................... 12 Ethane Dilution Effects................................................... 14

Discussion ............................................................................... 17 Conclusions............................................................................. 19 Acknowledgments ................................................................... 19 Appendix A: Velocity-Distance and Mach-Distance Data....... 20 Appendix B: Pressure-Time Data ........................................... 30 Appendix C: Experiment Shot Log ......................................... 42

3

Introduction The shock-induced combustion ramjet (or “shcramjet”) potentially offers considerable advantages over the conventional scramjet. In particular, fuel injection on the vehicle forebody results in an effectively premixed mixture of fuel and air entering the combustor. This feature permits either shock-induced or detonative modes of combustion to be utilized in the combustor, which are much more rapid than the usual diffusion-controlled combustion of the conventional scramjet. Thus, the “shcramjet” may address one of the major problems of the conventional scramjet: unacceptably long combustor lengths (with corresponding unacceptably large drag losses). In addition, several studies have suggested that the oblique-detonation mode may have propulsive advantages over conventional scramjets, especially at the higher flight Mach numbers (Mach 12 to 16).

To date, no experimental demonstration of positive acceleration of a vehicle by shock-induced supersonic combustion has been reported in the open literature. This experimental study is an attempt to demonstrate the feasibility of the concept by firing conical projectiles into a tube of pre-mixed, high-pressure combustible gas and monitoring the pressure wave activity around the projectile and the projectile’s acceleration and velocity history.

Experimental Apparatus and Procedure The 38-mm-bore ram accelerator test section was configured to accelerate shcramjet projectiles with the thermally choked ram accelerator propulsive mode from an entrance velocity of ~1.1 km/s up to ~1.8 km/s through two stages, each 4-m-long. These stages, filled with CH4/O2/N2 and CH4/O2/H2 propellants to 50-60 bar, were necessary to augment the gas gun muzzle velocity to the point where shcramjet tests could be carried

Fig. 1 Titanium-alloy, 4-fin, shcramjet projectile with magnet and plug.

4

out at Mach ≈ 6 in CO2-diluted propellant. The hypersonic experiments were conducted in a 6-m-long third stage of the ram accelerator test section with various CH4- and C2H6-fueled propellants at fill pressure of 21 bar. Projectile time-distance data were determined from tube wall-mounted electromagnetic sensors (in-house EM probe design) and piezoelectric pressure transducers (PCB 119) monitored at a data acquisition sampling rate of 1 MHz. Projectile velocity is based on center-differencing the EM probe signals and the velocity of the lead pressure wave is determined from pressure transducer data.

Facility shake-down experiments were carried out with projectiles (left over from previous shot programs) fabricated from magnesium and aluminum alloys. These preliminary tests were used to establish gas-gun performance scaling, tune up instrumentation and validate propellant handling procedures, in addition to training student laboratory assistants in the pertinent experimental techniques. Details of the staging arrangements for these preliminary experiments are included in the shot log summary in Appendix C.

The projectile configuration most extensively tested during this research program was fabricated as one piece from titanium 6Al-4V alloy (referred to as Type ‘B’ or ‘D’ in the following). It was hollowed from the base primarily for mass reduction purposes and to accommodate the installation of a neodymium magnet, as shown in Fig. 1. Features to enhance operation at hypersonic velocity incorporated in this projectile design are a slightly rounded nose tip (0.01” radius), knife-edged fins at a 20˚ rake angle and 0.100” thickness, throat diameter of 1.14” (0.76 throat-to-tube diameter ratio), and a 2.80-inch-long body with a taper angle of 4.49˚. The hollow base allows pressure equalization to eliminate concerns of projectile collapse from external pressure loads. Ten experiments were carried out with projectiles of this external configuration. The center-of-mass of some these projectiles (Type ‘C’) was moved aft by inserting steel weights in the projectile base. These projectiles had a mass range of 106 – 118 g, depending on external geometry and whether or not a steel plug was used. A second batch of 10 projectiles was fabricated with similar characteristics, except that the fins were thickened to 0.150” and the magnet receptacle was placed deeper into the projectile to reduce its mass.

The hollow base resulted in a significant reduction of solid surface area for the tube sealing obturator to bear against, which led to launch failures due to the nominal two-piece obturator (polycarbonate) prematurely breaking up. A new three-piece obturator configuration (shown in Fig. 2) was developed that could handle the 40 MPa breech pressure used to launch the projectile to the requisite test section entrance velocity (~1.1 km/s) for the ram accelerator to be successfully started. The key face plate parameters were the number and diameters of the holes, material and thickness. Ultimately, aluminum alloy face plates of ~0.10” thickness with an outer ring of 12 holes at 7/32” diameter, inner ring of 6 holes at 1/8” diameter, and a center hole at 3/8” diameter were found to be reliable for launching these projectiles. Details of the various obturator configurations used in this study are included in the shot log summary in Appendix C.

5

Propellants were prepared by in-line mixing of multiple gas flows independently metered with electronic mass flow controllers. Due to irreproducible results when using the same mass flow controller for CO2 and N2, the CO2 (and later C2H6) gas flows were controlled with a sonic orifice (i.e., micro-metering valve). This procedure enabled CO2-diluted propellants to be formulated at fill pressures up to nearly 30 atm and C2H6-fueled propellants to 21 atm. Frequent calibrations insured that the propellant compositions were within 2% of the desired component ratios.

Two stages of thermally choked ram accelerator propulsion were used to accelerate the projectiles up to ~1.8 km/s for the hypersonic experiments. The fill pressures of these stages were nominally 50 – 60 atm, based on projectile mass. The composition of the first stage propellant was 2.6CH4+2O2+5.8N2 for all experiments. The composition of the second stage propellant was initially 6CH4+2O2+2H2 and later in the program changed to 5.5CH4+2O2+2H2 to enhance performance. These stages proved to be very reliable and incurred minimal projectile erosion over the Mach number range they were used (typically less than Mach 4.5). Pressure transducers were installed at the beginning and end of each of the two 4-m-long thermally choked stages (i.e., P1-P12 and P13-P22) to monitor how well they were operating.

The entrance velocity ranged from 1.7 – 1.9 km/s to the third stage test section containing the low sound speed propellant for hypersonic shcramjet experiments. Pressure was monitored at 80-cm-intervals throughout most of the third stage part of the test section (P23 through P37), whereas the EM probes in this stage were separated by 40 cm. The fill pressure for the third stage was nominally 21 atm and composed of C2H6+3.5O2+diluent propellant, where excess C2H6 and sometimes CH4 or H2 were added as diluents. Effects of third stage propellant variations are discussed in the results section and configuration details are included in shot summary log in Appendix C.

Experimental Results The results of a single stage ram accelerator experiment are presented below to serve as reference for other shcramjet tests. Three distinct groupings of experimental data are then discussed. In the first set, the operating characteristics of 4 different projectile

Fig. 2 Face plate, obturator, and disk for launching shcramjet projectiles.

6

configurations (labeled type A, B, C, D) in the same propellant (1.5CH4+2O2+5CO2) were examined. The second data set presents the results from tests using only the type D projectiles in CH4-O2 propellant with different levels of CO2 dilution. In the third data set, shcramjet operation in fuel-rich C2H6-O2 propellant was explored using projectiles from both the first and second batches of new projectiles (types B and D).

Single Stage Experiment

During the experiment shake-down phase of this program, starting failures would occur in which the projectile promptly unstarts, decelerates, and is caught in the projectile decelerator tube at velocities less than 500 m/s. Under these conditions the projectile sustains relatively little damage and is reusable for certain kinds of experiments (often the nose tip has to be “touched up” and the fin leading edges de-burred). Re-shooting projectiles for developing a new ram accelerator starting procedure is routinely done. The velocity- and Mach-distance data from an experiment with a previously shot projectile in a 14-m-long stage filled with 2.6CH4+2O2+5.8N2 propellant at 53 atm are shown in Fig. 3. The “pV-lab” and “pMach” data points indicate pressure wave velocity determined from time-distance differencing the pressure transducer data, whereas “V-lab” and “Mach” data are from EM probes tracking the magnet onboard the projectile. Unstart is readily evident where the wave speed suddenly deviates from the projectile velocity. The test section entrance velocity was ~1.13 km/s and the projectile accelerated up to ~2.0 km/s in a distance of ~10 m before it experienced an unstart. The acceleration profile is quite similar to those from experiments carried out with pristine projectiles in the same propellant; however, the peak velocity obtained by the re-shot projectile is ~10% lower than expected. These results indicate that the ram accelerator starting and operating characteristics over a relatively wide Mach number range (Mach 3 - 5) are not very sensitive to imperfections in projectile external condition arising from being caught at modest velocity.

7

The pressure-time data and corresponding EM probe signals from this single-stage experiment are shown in Fig. 4. The plot legend reference letters P and M indicate pressure and EM sensor data, respectively, and the number refers to the instrumentation station location. The first station before the entrance diaphragm to the ram accelerator test section is designated the “launch tube” or “lt”, there after the stations are numbered 1 through 42 (the launch tube pressure data are deleted from this Fig. 4 for clarity; however, they are included in the Appendix B data plots). In experiments where only 14 m of the test section are used, the last instrumentation station is number 37, which is

located 20 cm upstream of the exit diaphragm. The voltage amplitude of the pressure data is plotted as recorded, except where it has been scaled down for presentation purposes. The “scale factor” in the figure caption refers to the multiplier used in these cases.

The time-of-passage of the center of the magnet is generally associated with the first zero crossing of its signal, however, signal distortions arising from variations in sensor construction and potential ionization of the gas lead to velocity uncertainties of ~20 m/s. The amplitudes of the EM signals are adjusted in all plots for clarification purposes. As evident in Fig. 4, one advantage of placing the magnet in the nose cone (see Fig. 1) is that its signal arrives before that of the conical shock wave from a hypersonic projectile nose cone, providing irrefutable proof that the pressure waves are accelerating the projectile, not just a shock wave. Unstart is indicated by tube-wall data when the head of the pressure wave overtakes the EM signal, as shown by the P27 and M27 signals from this experiment. High-Mach unstart typically results in a strong over-driven detonation wave with extreme pressure rise. As the projectile decelerates, the velocity of the over-driven

HS1670 Single Stage w/ Re-shot Projectile

1100

1300

1500

1700

1900

2100

2300

2500

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h

V-lab (m/s)pV-lab (m/s)MachpMach

2.6CH4 + 2O2 + 5.8N253 atm

Fig. 3 Data from single stage experiment using re-shot Ti-alloy projectile.

8

detonation wave decreases until it reaches the Chapman-Jouguet detonation speed. If the conditions are such that a CJ wave can be stabilized, then the lead shock wave ceases to communicate with the projectile and propagates at CJ speed until it exits the test section. In this particular experiment, the detonation wave was still over-driven when it reached the exit of the test section.

9

Single Stage Pressure Data: HS1670(P = 53 atm, 2.6CH4+2O2+5.8N2)

1

2

3

4

5

6

7

8

9

10

11

12

13

50 100 150 200 250 300 350

Microseconds

Volts

P37

P35

P32

P30

M30

P28

M28

P27

M27

P25

M25

P23

M23

P22

M22

P13

M13

P12

M12

P1

M1

Fig. 4 Tube wall pressure data from single-stage experiment using Type B projectile. Data plotting scale factor for pressure transducers P25-35 = 3x, scale factor for all others = 1x.

10

Projectile Geometry Variations

Four different projectile configurations were fired into the same test propellant (1.5CH4+2O2+5CO2) at entrance velocities ranging from 1.7-1.8 km/s. The test propellant fill pressure was 21 atm in all cases. Projectile A is the “standard” 5-fin configuration used in many prior ram accelerator experiments (Fig. 5-left). It is characterized by a 10˚ nose cone, flat-faced fin leading edges at a 12.5˚ rake angle and 0.120” fin thickness. The throat diameter, aft body taper angle and base area are the same as all other projectiles used in this study. Projectiles B and C were of the first batch of ten new 4-finned projectiles having 12.5˚ nose cone, knife-edged fin leading edges at a 20˚ rake angle, and 0.100”-thick fins. Type B projectiles were modified to a C-type by pressing a 10 g plug in their base to move the projectile center-of-mass aft ~0.3” to just behind the point where the fin leading edge first comes into contact with the wall. Projectiles from the second batch of ten having the fins thickened to 0.150”, as shown in Fig. 5-right, and internal works for magnet installation moved forward by ~0.10” are referred to as type D. These changes shifted the center-of-mass aft by ~0.25” (relative to type B) and provided the same flow area profile over the body as that of the type A projectile. The motivations for these various modifications are discussed below.

The velocity-distance data from EM probes for experiments with the type A-D projectiles are shown in Fig. 6. The type A projectile (HS1664) entered the third stage propellant at ~1.76 km/s, accelerated throughout 6-m-long stage to a peak velocity of ~1.90 km/s at the exit of the test section, which corresponds to an average specific thrust off 68 N*s/kg at an average flight Mach number of M = 6.1. In the following experiment (HS1665), the lower mass (108 vs. 115 g) type B projectile was used with a similar staging configuration. This projectile entered the third stage propellant at ~1.79 km/s, accelerated for a distance of ~1.8 m up to a peak velocity of 1.85 km/s, and then experienced an unstart. The tube wall pressure data (see App. B) were not significantly different from those with the 5-fin type A projectile; however, the EM probe signals were extraordinarily strong throughout the third stage test section, indicating that the projectile was heavily canted toward the sensors. Since this projectile did not travel as far or

Fig. 5 External views of type A (left) and type D (right) projectile configurations.

11

accelerate up to as high of velocity as the prior experiment, the canting was assumed to be detrimental to operation under these conditions even though higher average specific thrust (100 N*s/kg at M = 6.1) was generated before the unstart.

Because of canting concerns, an attempt was made to make the projectile more stable by moving the center-of-mass back to behind the point where the leading edges of the fins first come into contact with the tube wall. A 10 g steel plugged pressed into the base of the projectile was found to be sufficient (converting a B to type C projectile). The symmetry of tube-wall data from the next experiment (HS1666) indicated that the type C projectile was better stabilized and it ran about 1 m deeper into the stage; however, it just weakly accelerated in the velocity range of 1.82-1.85 km/s for about 3 m before unstarting. The increased projectile mass (118 vs. 108 g) does not account for the level of acceleration decrease observed in the hypersonic velocity regime. Indeed, this more massive projectile accelerated at a higher level in the thermally choked stages than in the previous shot and entered the third stage at ~1.82 km/s.

Since both the type B and C projectiles experienced unstart at very nearly the same velocity (1.85 km/s) in third stage of the test section, it is possible that they had approached the gas dynamic limit (i.e., thrust = drag) for their external geometry and that the titanium alloy was not able to withstand more than a few milliseconds of operation under these conditions. The type A projectile reached a higher velocity without unstarting in the 6-m-long third stage; however, the pressure data from last transducer indicated that unstart was imminent. The key external variations between the type A and B projectiles are: cone angle (10º vs. 12.5º), number of fins (5 vs. 4) and thickness (0.120” vs. 0.100”), and fin leading edge geometry (flat vs. knife edged) and rake angle (12.5º vs. 20º). Of these variables, the cone angle is likely to be the most important.

Projectile Variations

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

Proj D HS1679Proj C HS1666Proj B HS1665Proj A HS1664

2.6CH4+2O2+5.8N250 - 60 atm

5.5-6.0CH4+2O2+2H250 - 60 atm

1.5CH4+2O2+ 5 CO221 atm

Fig. 6 Velocity-distance data from 3-stage experiments using four different projectiles.

12

Time constraints mandated that the design for the second batch of ten projectiles for this program be finalized and immediately submitted to machine shop. Fabricating projectiles in the one-piece configuration (which is more economical than two-piece fabrication) using the external geometry of the type A projectile is possible, however, the 10º nose cone results in too much of a mass increase for the light gas gun to launch. Prior experiments have indicated the peak velocity of the projectile was the same regardless of the number of fins; however, the flow area profile of the aft body did affect the rate of acceleration at and above the CJ speed of the propellant. Thus a second batch of projectiles was fabricated with the same external parameters as the first (type B), except that the fins were thickened to 0.150” in order to provide the same aft body flow area profile as the 5-fin projectile (type A). The internal works for installing the magnet were moved deeper into the nose cone as a mass reduction measure. The combined effect of these two changes resulted in a greater stability margin by moving the center-of-mass aft ward ~0.25”. These projectiles are referred to as type D.

The favorable impact of type D projectile geometry is indicated by the velocity-distance data from HS1679 in Fig. 6. The pressure in the thermally choked stages was increased (from ~50 to 60 atm) and the excess fuel component in the second stage was reduced (from 6 to 5.5 CH4) to enable higher entrance velocities to the third stage test propellant. In this experiment, the projectile entered the third stage at 1.91 km/s and accelerated to 2.07 km/s over a distance of ~4 m prior to unstart. The improved performance of the type D projectile over that of type B is attributed primarily to the thicker fins and better stability margin. Nevertheless, there is a possibility that the type B and C projectiles would experience similar performance characteristics if they too were launched into the third stage test propellant at the same entrance velocity (i.e., 1.91 vs. 1.82 km/s). This issue remains an open question at this time.

Carbon Dioxide Dilution Effects

A series of experiments with the type D projectile were conducted in which the carbon dioxide molar level in a 1.5CH4+2O2+XCO2 propellant was varied in the range 2.8 < X < 8. The velocity-distance and Mach-distance data from these experiments are plotted in Figs. 7 and 8, respectively. In all of these experiments, the projectile entered the third stage test propellant in the velocity range of 1.87 – 1.93 km/s, corresponding to a Mach range of 6.0 – 6.3. The largest velocity increase (160 m/s) and peak velocity (2.07 km/s) were observed in the 5CO2 (HS1679) propellant (same shot discussed in previous section). More energetic propellants (2.8 and 4.1CO2) experienced unstart within 2 m or less after entering the third stage. The greatest distance (~5 m) of shcramjet projectile acceleration occurred in the propellant with 7CO2 (HS1673), where it reached ~1.98 km/s before unstart. In two firings into 6CO2 propellant (HS1682, HS1685), the projectile accelerated for more than 3 m and attained 1.96 km/s before unstart, demonstrating very reproducible results. At a dilution level of 8CO2, the projectile more or less just cruised at ~1.91 km/s (M = 6.6) for ~3 m before unstart, indicating that the thrust = drag gas dynamic limit may have been reached. It is not known whether this unstart was due to projectile erosion/failure, combusting boundary layer interactions arising from surface heating of the projectile, and/or other gas dynamic phenomena.

13

Fig. 7 Velocity data from experiments with various CO2 dilution levels in third stage propellant.

CO2 Dilution Variations (Projectile D)

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

) 8.0 CO2 HS1684

7.0 CO2 HS1683

6.0 CO2 HS1685

6.0 CO2 HS1682

5.0 CO2 HS1679

4.1 CO2 HS1681

2.8 CO2 HS1680

2.6CH4+2O2+5.8N255 - 60 atm

5.5CH4+2O2+2H255 - 60 atm

1.5CH4+2O2+ X CO221 atm

CO2 Diluent Variations (Projectile D)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

0 2 4 6 8 10 12 14

Distance (m)

Mac

h N

umbe

r

8.0 CO2 HS1684

7.0 CO2 HS1683

6.0 CO2 HS1685

6.0 CO2 HS1682

5.0 CO2 HS1679

4.1 CO2 HS1681

2.8 CO2 HS1680

2.6CH4+2O2+5.8N255 - 60 atm

5.5CH4+2O2+2H255 - 60 atm

1.5CH4+2O2+ X CO221 atm

Fig. 8 Mach number data from experiments with various CO2 dilution levels in third stage propellant.

14

As evident in Fig. 8, the projectile Mach number upon entering the third stage of these experiments does not vary much in the range of CO2 dilution levels investigated here. The CJ speed, however, varied by ~50% (1.25 < VCJ < 1.78 km/s). The influence of CJ speed is more apparent in Fig. 9, where the projectile velocity is normalized by CJ speed in the V/Vcj-distance plots. The projectile is accelerated up to ~95% CJ speed in the two thermally choked stages and makes a sudden transition to superdetonative velocity upon entrance to the third stage. In the most energetic propellant (2.8CO2) the entrance velocity is only 5% greater than CJ speed and the projectile unstarts within ~1 m. The highest average specific thrust (151 N*s/kg with average acceleration = 9100 g) was attained when entering at 20% greater than CJ speed in 4.1CO2 diluted propellant; however, as previously stated, this projectile only accelerated ~2 m before unstart. More stable operation was achieved when the third stage entrance velocity was higher than CJ speed by 30% or more. In these scenarios the projectile readily operated with a shcramjet-like propulsive cycle for 3 or more meters before unstart.

Ethane Dilution Effects

The fact that the 5-fin projectile (type A) accelerated for more than 6 meters in 1.5CH4+2O2+5CO2 propellant under conditions where the 4-fin types B and C projectiles only operated for 2 − 3 meters (see Fig. 6) indicated that the aerodynamic heating environment of Mach 6+ flight may be too severe for the relatively thin (0.100”) projectile fins. Another concern was that the methane-oxygen equivalence ratio (φ = 1.5)

CO2 Diluent Variations (Projectile D)

0.6

0.7

0.8

0.9

1.0

1.1

1.2

1.3

1.4

1.5

1.6

0 2 4 6 8 10 12 14

Distance (m)

V / V

cj

8.0 CO2 HS1684

7.0 CO2 HS1683

6.0 CO2 HS1685

6.0 CO2 HS1682

5.0 CO2 HS1679

4.1 CO2 HS1681

2.8 CO2 HS1680

2.6CH4+2O2+5.8N255 - 60 atm

5.5CH4+2O2+2H255 - 60 atm

1.5CH4+2O2+ X CO221 atm

Fig. 9 V/VCJ data from experiments with various CO2 dilution levels in third stage propellant.

15

was so close to stoichiometric that some degree of titanium burning may occur that could contribute to premature unstart. Thus it was decided to use the remaining type B projectiles in ethane-rich propellants formulated to have similar heat release per mass, CJ speed, and sound speed as the CH4-fueled propellant. Type D projectiles were also used; however, the effects of the small geometry differences seem to be negligible in this series of experiments and thus the results are discussed without distinguishing projectile type.

Since the stoichiometric ratio of ethane-oxygen is 1.0C2H6 : 3.5O2, propellants having more than 50−50 ratio of C2H6 to O2 are very fuel rich which minimizes the amount of oxygen available for metal combustion. In addition, due to its more complex molecular make up, the heat capacity of C2H6 is larger than that of CH4 and CO2 on a molar basis which makes it a very good diluent. These characteristics of ethane motivated experimentation with third stage ethane-oxygen propellants formulated to have sound speeds of ~300 m/s and CJ speeds of 1.5-1.7 km/s. Projectile entrance velocities of 1.8 – 1.9 km/s were used in this set of experiments and the resulting velocity-, Mach-, and V/VCJ-distance data are shown in Figs. 10, 11, and 12, respectively. The projectile mass was nominally either 110 or 117 g, depending on type of projectile (see App. C).

The first 2 test firings were with 4.3C2H6+1O2 and 3.0C2H6+1O2 propellants having theoretical CJ speeds of 1.50 and 1.62 km/s, respectively. In the 4.3C2H6 experiment the projectile entered the third stage at Mach 5.8, a velocity 20% higher than CJ speed whereas in the 3.0C2H6 experiment the projectile entered the third stage at Mach 5.5, an entrance velocity only 5% higher than CJ speed, as shown in Figs. 11 and 12. Remarkably, the combustion waves clearly fell off both of these projectiles (see pressure traces for HS1671 and HS1672 in App. B) and they smoothly decelerated at supersonic velocity (from about Mach 5.5 down to 5.1, see Fig 11) in the last four meters of the test section. Experiments with this behavior are labeled as WFO in the plot legend. The average drag forces (~25 kN) determined from the three WFO experiments shown here correspond to a drag coefficient of 0.09±0.01, based on projected frontal area of the projectile, in the Mach range of 5.3 to 5.7.

When the ethane content was further reduced to formulate 2C2H6+1O2 propellant (VCJ = 1.80 km/s), the combustion wave did not completely separate in the first 3 or so meters of the test section. Enough thrust was generated to offset the drag and allowing the projectile to “cruise” at relatively constant velocity. Eventually, the combustion wave clearly fell off the projectile in the last two meters of the test section and the projectile velocity decreased to the CJ speed of the propellant (see Fig. 12 and pressure data for HS1673 in Appendix B). It is unusual for a projectile to experience a wave fall off when traveling at velocities greater than the CJ speed of the propellant, as evident is the CO2-diluted series of experiments. In the last of the pure ethane-oxygen experiments (HS1677), the projectile was injected into 1C2H6+1O2 (VCJ = 2.24 km/s) with an entrance velocity of 1.9 km/s (Mach 6.1, V = 0.85VCJ). In this scenario the projectile promptly unstarted, which is to be expected when trying to operate in the thermally choked velocity regime with too energetic of propellant.

16

Ethane-Fueled Propellant (Projectile B, D)

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

WFO 1.5H2+2O2+5C2H6

Cruise 1.0CH4+2O2+3C2H6


Unstart 1.0C2H6+1O2

Cruise 2.0C2H6+1O2

WFO 3.0C2H6+1O2

WFO 4.3C2H6+1O2

2.6CH4+2O2+5.8N250 - 60 atm

5.5CH4+2O2+2H250 - 60 atm

X CH4+Y H2 + 2O2+ Z C2H621 atm

Fig. 10 Velocity-distance data from ethane propellant experiments.


3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

0 2 4 6 8 10 12 14

Distance (m)

Mac

h N

umbe

r

WFO 1.5H2+2O2+5C2H6



Unstart 1.0C2H6+1O2

Cruise 2.0C2H6+1O2

WFO 3.0C2H6+1O2

WFO 4.3C2H6+1O2

2.6CH4+2O2+5.8N250 - 60 atm

5.5CH4+2O2+2H250 - 60 atm


Fig. 11 Mach-distance data from ethane propellant experiments.

17

Experiments with other fuel species added to the ethane-oxygen mixture to reduce the reaction induction time without significantly raising the CJ speed were carried out. As seen in Fig. 10, the experimental results of adding methane and removing ethane (i.e., 1.5CH4+2O2+3C2H6, HS1675 and 1.0CH4+2O2+3C2H6, HS1676) were very similar to those with the 2.0C2H6+1.0O2 propellant in that the projectiles cruised throughout the test section at constant velocity and then decelerated after the wave fell off (see pressure data for HS1676, pressure data for HS1665 has been lost). Hydrogen addition was tried with sufficient ethane to keep the CJ speed below 1.8 km/s; however, the wave fell off the projectile within 2 meters after entering third stage (HS1678) in the manner seen in prior experiments.

Discussion Theoretical predictions of superdetonative ram accelerator operation with 1.5CH4+2O2+XCO2 propellant at 21 atm and similar projectile configuration indicate that the thrust = drag limit should be reached at velocities 40-50% greater than CJ speed. This indeed appears to be velocity ratio limit observed here in experiments using propellant dilution levels of 5 < X < 8 CO2 (see Fig. 9). Ideally the projectile should cruise throughout the remainder of the test section at constant velocity once this gas dynamic limit has been reached, which would correspond to hypersonic cruising conditions in a shcramjet engine application. The relatively small scale at which these experiments are carried out, however, certainly magnifies the impact of aerodynamic and/or combustion heating effects on the projectile. Thus, due to erosion effects, the

Fig. 12 V/Vcj-distance data from ethane propellant experiments.


0.6

0.7

0.8

0.9

1.0

1.1

1.2

1.3

0 2 4 6 8 10 12 14

Distance (m)

V / V

cj

WFO 1.5H2+2O2+5C2H6



Unstart 1.0C2H6+1O2

Cruise 2.0C2H6+1O2

WFO 3.0C2H6+1O2

WFO 4.3C2H6+1O2

2.6CH4+2O2+5.8N250 - 60 atm

5.5CH4+2O2+2H250 - 60 atm


18

ability to maintain constant velocity operation for long stretches in the test section may not be possible for titanium alloy projectiles. Even so, the demonstration of significant thrust for several meters confirms that shock-induced combustion can propel vehicles at hypersonic velocities.

Data in Fig. 12 indicate that sustained shcramjet operation at velocities greater than the propellant CJ speed was not demonstrated in any of the experiments with excess ethane present. Indeed, the subsequent wave fall-off behavior was not seen under any circumstances in the CH4/O2/CO2 propellant. It is possible that propellant combustion could not be initiated on the aft body of the projectile at the lower Mach number (5.5-6.1 vs. 6-6.6) and fraction of CJ speed (1.2Vcj max vs. 1.5Vcj max) of these experiments. Another possibility is that the tendency for metal combustion was completely suppressed which negatively impacted the ability for a supersonic combustion process to stabilize in a manner that would continuously accelerate the projectile. The observation of a distinct pressure wave falling behind the projectile in the ethane-fuel experiments implies that whatever thrust was seen may not even have been from oblique shock-induced combustion; i.e., carry over effects from the prior thermally choked stage may have had some influence. Experiments with higher entrance Mach numbers are necessary to gain better understanding shcramjet operating characteristics in ethane-rich propellant.

One of the key questions raised in this experimental program is whether the unstarts are caused by projectile erosion due to aerodynamic and/or propellant combustion heating effects at hypersonic Mach numbers or if there are other gas dynamic processes causing the combustion wave to be disgorged ahead of the projectile. An intriguing finding of the ethane-fuel experiments is that there were no unstarts at all in superdetonative velocity regime. Granted these tests were at somewhat lower Mach number, but do these results imply that combustion must be initiated for hypersonic unstart to occur? It is exigent to distinguish if unstarts are due to projectile erosion or combustion induced phenomena. Experiments with more refractory materials (e.g., nickel steel) and/or thermal insulation coatings and oxidizing barriers are very likely to establish if material properties are the main factor limiting the ability of the projectiles to cruise at hypersonic velocities greater than the CJ speed. If this is the case, then large-scale shcramjet engines with active cooling systems have great potential for atmospheric vehicles. Conversely, if the unstarts occur regardless of the projectile material due to gas dynamic phenomena, then much more extensive testing is required to determine if geometry and/or propellant composition tweaks can mitigate these effects. The fact that relatively rapid projectile proto-typing and experiment turn around can be carried out in gun-launched testing facilities indicate this approach is a very prudent risk reduction strategy to employ to support the development of shcramjet propulsion systems.

19

Conclusions Titanium-alloy projectiles were successfully launched into reactive propellants at 21 atm fill pressure with entrance Mach numbers ranging from 5.5 to 6.6 to investigate the operating characteristics of the shcramjet propulsive cycle. Results of experiments with different projectile configurations indicated aft body area profile and center-of-mass location significantly affect shcramjet projectile performance at hypersonic velocities. Positive acceleration was observed in the Mach range of 5.5 − 7 (1.7 − 2.1 km/s) for distances of up to 6 meters in 1.5CH4+2O2+XCO2 propellant with 2.8 < X < 8. In experiments where shcramjet operation was sustained for distances greater than 2 meters, average accelerations of 9000 g and average specific thrust 150 N*s/kg were demonstrated. Experiments with ethane-fueled propellant in the Mach number range of 5.5 – 6.1 found that projectiles would either experience a wave fall-off or cruise at relatively constant velocity without unstart. Significant projectile acceleration was not observed at superdetonative velocity in this propellant. The hypersonic projectile drag coefficient in the Mach range of 5 – 6 was determined to be CD ≈ 0.09 from experiments in which the combustion wave had clearly fallen off the projectile. Effective thrust in a shcramjet-like propulsive mode was unequivocally demonstrated; however, more experimentation is necessary to determine whether it was propellant combustion phenomena or the effects of aerodynamic heating that limit the peak velocity to which a projectile can be accelerated.

Acknowledgments The experimental work reported here was performed by Carl Knowlen with the assistance of Kyle Hughes, Viggo Hansen, Dan Balock, David Peters, Ryan Trescott, Xavier De Guillebon, Ninh Trinh, and Miguel Rodriquez with experiments and data processing is greatly appreciated.

20

Appendix A: Velocity-Distance and Mach-Distance Data for Experiments HS1664 – HS1685

21

HS1685

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h

V-lab (m/s)pV-labMachpMach

2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+6CO2

HS1684

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+8CO2

22

HS1683

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+7CO2

HS1682

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+6CO2

23

HS1681

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

8.0

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+4CO2

HS1680

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+2.8CO2

24

HS1679

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

2300

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h

V-lab (m/s)

pV-lab

Mach

pMach

2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 1.5CH4+2O2+5CO2

HS1678

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

2100

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h


2.6CH4+2O2+5.8N2 5.5CH4+2O2+2H2 5C2H6+2O2+1.5H2

25

HS1677

1100

1300

1500

1700

1900

2100

2300

2500

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

4.0

5.0

6.0

7.0

8.0

9.0

Mac

h


HS1676

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


26

HS1675

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h

V-lab (m/s)

Mach

HS1673

1100

1200

1300

1400

1500

1600

1700

1800

1900

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h


27

HS1672

1100

1200

1300

1400

1500

1600

1700

1800

1900

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h


HS1671

1100

1200

1300

1400

1500

1600

1700

1800

1900

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h


28

HS1670 Single Stage (Re-shot Projectile B)

1100

1300

1500

1700

1900

2100

2300

2500

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

Mac

h


2.6CH4 + 2O2 + 5.8N253 atm

HS1666 (Projectile C)

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4 + 2O2 + 5 8N2

6CH4 + 2O2 + 2H2 1.5CH4 + 2O2 + 5CO2

29

HS1665 (Projectile B)

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h


2.6CH4 + 2O2 + 5.8N2 6CH4 + 2O2 + 2N2 1.5CH4 + 2O2 + 5CO2

HS1664 (Projectile A)

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 2 4 6 8 10 12 14

Distance (m)

Velo

city

(m/s

)

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

Mac

h

V-lab (m/s)pV-lab (m/s)

MachpMach

1.5CH4 + 2O2 + 5CO26CH4 + 2O2 + 2H22.6CH4 + 2O2 + 5.8N2

30

Appendix B: Pressure-Time Data Experiments HS1664 – HS1685

31

Superdet Stage Pressure Data: HS1685 6CO2

-1

0

1

2

3

4

5

6

7

8

9

10

50 100 150 200 250 300 350

Microseconds

Volts

P37P35M35P32M32P30M30P28M28P27M27P25M25P23M23

Subdet Stages Pressure Data: HS1685

0

1

2

3

4

5

0 50 100 150 200 250 300 350

Microseconds

Volts

P22P13M13P12M12P1M1Plt

32


-1

0

1

2

3

4

5

6

7

8

9

10

50 100 150 200 250 300 350

Microseconds

Volts

P37P35P32M32P28M28P27M27P25M25P23M23


0

1

2

3

4

5

6

50 100 150 200 250 300 350

Microseconds

Volts

P22M22P13M13P12M12P1M1Plt

33


-1

0

1

2

3

4

5

6

7

8

9

10

11

50 70 90 110 130 150 170 190 210 230 250

Microseconds

Volts

P37M37P35M35P32M32P30P28M28P27M27P25M25P23M23


0

1

2

3

4

5

50 100 150 200 250 300 350

Microseconds

Volts


34


-1

0

1

2

3

4

5

6

7

8

9

10

11

50 100 150 200 250 300 350

Microseconds

Volts

P37P35M35P32M32P30M30P28M28P27M27P25M25P23M23


0

1

2

3

4

5

50 100 150 200 250 300 350

Microseconds

Volts


35


-1

0

1

2

3

4

5

6

7

8

9

50 100 150 200 250 300 350

Microseconds

Volts

P37P35P32M32P30M30P28M28P27M27P25M25P23M23


0

1

2

3

4

5

6

50 100 150 200 250 300 350

Microseconds

Volts


36


-1

1

2

3

4

5

6

7

8

9

10

11

50 100 150 200 250 300 350

Microseconds

Volts

P37P35P32P30P28M28P27M27P25M25P23M23


0

1

2

3

4

5

6

50 100 150 200 250 300 350

Microseconds

Volts


37


0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

50 100 150 200 250 300 350

Microseconds

Volts P27

P23


0

1

2

3

4

5

6

7

50 100 150 200 250 300 350

Microseconds

Volts

P22P13P12P1Plt

38

Single Stage Pressure Data: HS1670

-1

1

2

3

4

5

6

7

8

9

10

11

12

13

50 100 150 200 250 300 350

Microseconds

Volts

P37P35P32P30M30P28M28P27M27P25M25P23M23P22M22P13M13P12M12P1M1Plt

39


-1

0

1

2

3

4

5

6

7

8

50 100 150 200 250 300 350

Microseconds

Volts

P35P32M32P28M28P27M27P25M25P23M23


-1

0

1

2

3

4

5

50 100 150 200 250 300 350

Microseconds

Volts

P22M22M13P13P12M12P6M6P1M1

40


-1

0

1

2

3

4

5

6

7

50 100 150 200 250 300 350

Microseconds

Volts

P35M35P32M32P28M28P27M27P25M25P23M23


-1

0

1

2

3

4

5

6

50 100 150 200 250 300 350

Microseconds

Volts

P22M22P13M13P12M12P6M6P1M1

41


-1

0

1

2

3

4

5

6

7

8

50 100 150 200 250 300 350

Microseconds

Volts

P37P35P32M32P30P28P27M27P25P23M23


-0.5

0.5

1.5

2.5

3.5

4.5

5.5

50 100 150 200 250 300 350

Microseconds

Volts

P22M22P13P12P6M6P1M1

42

Appendix C: Experiment Shot Log HS1650 – HS1685

UNCLASSIFIED SECURITY CLASSIFICATION OF FORM (highest classification of Title, Abstract, Keywords)

DOCUMENT CONTROL DATA (Security classification of title, body of abstract and indexing annotation must be entered when the overall document is classified)

1. ORIGINATOR (the name and address of the organization preparing the document. Organizations for who the document was prepared, e.g. Establishment sponsoring a contractor's report, or tasking agency, are entered in Section 8.)

Reactive Energetics 1678 Alexis Nihon Montreal, Quebec H4R 2W2

2. SECURITY CLASSIFICATION (overall security classification of the document, including special

warning terms if applicable)

Unclassified

3. TITLE (the complete document title as indicated on the title page. Its classification should be indicated by the appropriate abbreviation (S, C or U) in parentheses after the title).

Experimental Investigation of Combustion and Propulsion for Shock-Induced Combustion Ramjets

4. AUTHORS (Last name, first name, middle initial. If military, show rank, e.g. Doe, Maj. John E.)

Higgins, Dr. Andrew

5. DATE OF PUBLICATION (month and year of publication of document)

December 2006

6a. NO. OF PAGES (total containing information, include Annexes, Appendices, etc) 45

6b. NO. OF REFS (total cited in document)

0

7. DESCRIPTIVE NOTES (the category of the document, e.g. technical report, technical note or memorandum. If appropriate, enter the type of report, e.g. interim, progress, summary, annual or final. Give the inclusive dates when a specific reporting period is covered.)

Final Contract Report

8. SPONSORING ACTIVITY (the name of the department project office or laboratory sponsoring the research and development. Include the address.)

Defence R&D Canada – Suffield, PO Box 4000, Station Main, Medicine Hat, AB T1A 8K6

9a. PROJECT OR GRANT NO. (If appropriate, the applicable research and development project or grant number under which the document was written. Please specify whether project or grant.)

Project 3ea26 (Pulse Detonation Engine TIF)

9b. CONTRACT NO. (If appropriate, the applicable number under which the document was written.)

W7702-05-R063/001/EDM

10a. ORIGINATOR'S DOCUMENT NUMBER (the official document number by which the document is identified by the originating activity. This number must be unique to this document.)

DRDC Suffield CR 2007-249

10b. OTHER DOCUMENT NOs. (Any other numbers which may be assigned this document either by the originator or by the sponsor.)

11. DOCUMENT AVAILABILITY (any limitations on further dissemination of the document, other than those imposed by security classification)

( x ) Unlimited distribution ( ) Distribution limited to defence departments and defence contractors; further distribution only as approved ( ) Distribution limited to defence departments and Canadian defence contractors; further distribution only as approved ( ) Distribution limited to government departments and agencies; further distribution only as approved ( ) Distribution limited to defence departments; further distribution only as approved ( ) Other (please specify):

12. DOCUMENT ANNOUNCEMENT (any limitation to the bibliographic announcement of this document. This will normally corresponded to the Document Availability (11). However, where further distribution (beyond the audience specified in 11) is possible, a wider announcement audience may be selected).

Unlimited UNCLASSIFIED SECURITY CLASSIFICATION OF FORM

UNCLASSIFIED SECURITY CLASSIFICATION OF FORM

13. ABSTRACT (a brief and factual summary of the document. It may also appear elsewhere in the body of the document itself. It is highly desirable that the abstract of classified documents be unclassified. Each paragraph of the abstract shall begin with an indication of the security classification of the information in the paragraph (unless the document itself is unclassified) represented as (S), (C) or (U). It is not necessary to include here abstracts in both official languages unless the text is bilingual).

An investigation of the shock-induced-combustion ramjet propulsive cycle was carried out in the 38-mm-bore ram accelerator facility at the University of Washington. Titanium-alloy shcramjet projectiles were launched into reactive propellants at Mach numbers greater than 5.5 to determine if the combustion process could be shock initiated and stabilized, what levels of thrust can be generated, the reactivity of the projectile material in hypersonic flow, and to evaluate the efficacy of investigating the operating characteristics of hypersonic propulsive cycles using gun-launched projectiles. Experiments with four different projectile configurations were carried out in methane- and ethane-based propellants with and without carbon dioxide diluent. Positive acceleration was observed in CH4/O2/CO2 and C2H6/O2 propellants in the Mach range of 5.5 - 7 (1.7 - 2.1 km/s) for distances of up to 6 meters. In the majority of cases, the acceleration process was terminated by either unstart, cruise at constant velocity (i.e., thrust equal drag), or wave fall off. Sustained accelerations greater than 9000 g and average specific thrust 150 N*s/kg were achieved in these experiments. The hypersonic projectile drag coefficient (based on projected frontal area of the projectile) in the Mach range of 5 – 6 was determined to be CD = 0.09 from experiments in which the propellant did not ignite (wave fall-off).

Details of the experimental apparatus and procedures used in this investigation are described. Key results of 18 pertinent hypersonic shcramjet experiments are presented and suggestions for follow on research are discussed. Velocity- and Mach-distance data plots and corresponding tube wall data from pressure and electromagnetic sensors for these experiments are provided. An abbreviated shot log having pertinent details of the each testing configuration and outcome for all experiments conducted in this program is provided.

14. KEYWORDS, DESCRIPTORS or IDENTIFIERS (technically meaningful terms or short phrases that characterize a document and could be helpful in cataloguing the document. They should be selected so that no security classification is required. Identifies, such as equipment model designation, trade name, military project code name, geographic location may also be included. If possible keywords should be selected from a published thesaurus, e.g. Thesaurus of Engineering and Scientific Terms (TEST) and that thesaurus-identified. If it is not possible to select indexing terms which are Unclassified, the classification of each should be indicated as with the title.) Ram accelerator Shcramjet Titanium projectiles Hypersonic Hypersonic projectiles Hypersonic propulsion Thrust Shock initiation

UNCLASSIFIED SECURITY CLASSIFICATION OF FORM

Defence R&D Canada R & D pour la défense Canada Canada's Leader in Defence

and National Security Science and Technology

Chef de file au Canada en matière de science et de technologie pour la défense et la sécurité nationale

www.drdc-rddc.gc.ca

Experimental Investigation of Combustion and Propulsion ...

Documents

Transcript of Experimental Investigation of Combustion and Propulsion ...