RECENT TRENDS IN SCRAMJET ENGINE TECHNOLOGY

Problem Areas

4.1 The main difficulty in scramjet operation is the short residence time of air

in the engine i.e. a few thousandths of a second. So the task of burning the fuel

in the scramjet resembles lighting a match inside a tornado and keeping it alight

at any cost. The trick to work make a scramjet work lies in the extremely

sophisticated shaping of the tubes inner geometry and deciding at which part of

the tube the combustion is to be done. A scramjet generates stable thrust by

precisely controlling the speed and pressure of air flowing through the engine

and by metering the fuel into the combustor so that it burns fully and releases its

energy exactly as needed. Careful control of the relation between the flow area

and the heat release negates the need for mechanical choke of the Ramjet and

enables the scramjet to maintain supersonic flow through the combustor.

4.2 Managing the thermal energy generated in the engine is a Herculean task

for the researchers. Heat generation is primarily due to friction, from combustion

process as well as from the heat flux due to internal shock waves impinging on

the engine wall as wall of the fore body. “Active cooling” is the concept adapted

to prevent melting down of the structure of the vehicle. Therefore studies often

plan on "active cooling", where coolant circulating throughout the vehicle skin

prevents it from disintegrating from the fiery atmospheric friction. Active cooling

could require more weight and complexity. There is also safety concern since it's

an active system. Often, however, the coolant is the fuel itself, much in the same

way that modern rockets use their own fuel and oxidizer as coolant for their

engines. Both scramjets and conventional rockets are at risk in the event of a

cooling failure.

4.3 Another complexity involving the scramjet is that the given airflow

geometry is optimized for only a single and a particular set of flight conditions i.e.

speed, altitude etc. Successful scramjet operation is thus a delicate balancing act

of choosing the right material and design complication.

FUEL FOR SCRAMJET ENGINE

4.4 The primary difference between the rocket engine and a turbojet engine is

that rockets carry their own supply of oxygen internally while turbojet engines

suck in oxygen from the external atmosphere. The scramjet is an air breather,

meaning that it gets its oxygen from the surrounding air. However, the scramjet is

significantly different from other kinds of jet engines, like turbojets and ramjets, in

one key way. In most jets, the air sucked into the engines is slowed below Mach

1 and is combusted at subsonic speeds. The air within the scramjet combustion

chamber, however, remains “supersonic”.

4.5 In order to make a scramjet work, a fuel that can burn rapidly and

generate a large amount of thrust must be chosen. Hydrogen meets these

criteria. One way to illustrate the differences between various fuels and their

energy content is a measurement called the Lower Heating Value (LHV). The

LHV describes the amount of energy released when a fuel is combusted and all

of the remaining combustion products remain in gaseous form. The LHV for

hydrogen is 119,600 kJ/kg. JP-8, another fuel commonly used in military aircraft,

has a LHV of only 43,190 kJ/kg, less than half that of hydrogen. Hence, hydrogen

provides more "thrust” per kilogram than JP-8, or any other hydrocarbon fuel for

that matter. There are also other advantages to using hydrogen as a fuel. First

of all, hydrogen is extremely flammable, it only takes a small amount of energy to

ignite it and make it burn. Hydrogen also has a wide flammability range, meaning

that it can burn when it occupies anywhere from 4% to 74% of the air by volume.

Since hydrogen is a gas, it mixes very easily with air allowing for very efficient

combustion. Another advantage over hydrocarbon-based fuels like JP-8 or

gasoline is that hydrogen does not produce any harmful pollutants like carbon

monoxide (CO), carbon dioxide (CO2), or particulate matter during the

combustion process.

4.6 There are some disadvantages to using hydrogen as a fuel in aerospace

vehicles. Hydrogen is not a dense fuel. At standard pressure and temperature, it

has a density of only 0.09 kg/m3. Compared to the density of gasoline at 750

kg/m3 or JP-8 at 800 kg/m3. While this low density is an advantage in terms of

saving weight, hydrogen requires a large volume in order to store an adequate

amount of chemical energy for practical use. Hydrogen gas is typically stored

under pressure to increase its density, but even at 10,000 psi (68,950 KPa) it will

contain only a quarter of the chemical energy stored in an equivalent volume of

JP-8. The density of hydrogen can be further increased by cooling and

pressurizing the substance to the point that it becomes a liquid, but even in this

form it will need a tank approximately twice the size of that required by JP-8. In

addition, the cost and safety issues involved in manufacturing and storing

cryogenically-cooled fuel is another major drawback. Despite the clear

advantages of hydrogen described earlier, more energy can often be stored in

smaller volumes using denser fuels. As a result, vehicles burning denser

hydrocarbon fuels can usually fly longer distances than those using hydrogen.

4.7 Even given these limitations, however, hydrogen has been a clear choice

for many scramjet researchers due to its versatility and performance. One of the

first hydrogen-fueled scramjets ever flown was the X-43A launched on 27 March

2004. The X-43A is part of NASA's Hyper-X program to develop new air-

breathing propulsion systems for use in hypersonic flight. A milestone in scramjet

research, the X-43A achieved positive acceleration while climbing at Mach 7 for

approximately 10 seconds.

Fig 4.1 COMPARISION OF HYDROGEN AND HYDROCARBON PERFORMANCE

4.8 Fuel Injection: The most difficult task in the design of a supersonic

combustion Ramjet engine is encountering the problems associated with

combustion of fuel in the combustion chamber. In order to ensure proper mixing

of fuel and air in the combustion chamber between Mach 2 and Mach 3, it is

imperative that the chamber has to be elongated in length. Several concepts are

being studied on this issue. There are various ways of injecting fuel into the

combustion chamber. They can be broadly classified into two categories: -

(a) Transverse injection. Transverse injection provides good

penetration and near field mixing but is inevitably accompanied by shocks

that reduce the total pressure. The total pressure is directly linked to

thrust, so it is important to ascertain the conditions under which these

serious losses occur.

(b) Parallel injection. Parallel injection way of injecting fuel in the

combustion chamber is injecting it parallel to the flow. This is not an

effective way since improper mixing occurs. So a compromise researched

which has solid or aerodynamic ramps. Oblique jets produced by angled

injection does exhibit less near field mixing in comparison with transverse

injection, but they resulted in good far- field mixing through the generation

of large vertical structures.

4.9 Injection from the side walls would produce very long mixing lengths;

therefore the fuel injection using the struts is also used by NASA Langley

research Center extensively. Three struts configuration as show in the Fig below

are used to provide six planes of in stream fuel injection. This feature shortens

the combustor well reduces the inlet length, since these struts provide a

significant part of the inlet flow compression.

4.10 In the recent years, cavity flame holders and integrated fuel injection

/flame holding approaches have been proposed as a way of flame holding and

stabilization in supersonic combustor. It is known that the growth rate of mixing

layer between supersonic air and gaseous fuel in a scramjet combustor

decreases as the convective Mach number increases due to compressibility

effect. The design of combustion chamber geometry and of fuel injection

elements as well as the combustion process is some of the major challenges

when developing Ramjet/ Scramjet engines.

4.11 Intake design considerations: In hypersonic air breathing engines, the

design and performance characteristics of the inlet play a vital role in the total

performance. The purpose of the intake is to ensure total pressure recovery and

mass flow regulation. The design of the intake influences the structural pressure

and cooling loads on the engine other than generation of thrust and impulse.

Intake is also associated with aerodynamic heating of the internal surfaces and

passages. Two types of intake design are in vogue. They are:

(a) Fixed Geometry inlet

(b) Variable geometry inlet

It is desired that a scramjet engine should operate at high inlet contraction ratios

at high flight speeds in order to keep the velocities in the combustor at the low

levels required for low momentum loss and high thrust. However, in contrast

fixed geometry inlets are limited to maximum contraction ratios because the inlet

must have the capability of starting of establishing supersonic flow in the inlet at

the low end of Mach number range.

4.12 It is also evident that the variable geometry inlet will offer only a 16 %

increase in the performance compared to fixed geometry. This comes with a

penalty in terms of increase in the system complexity, sealing and other

problems. It may also be apparent that high contraction ratios would result in

heating of the inlet thus cooling requirement of the engine. In turn the variable

geometry comes with higher weight and cost. Therefore studies often plan on

"active cooling", where coolant circulating throughout the vehicle skin prevents it

from disintegrating from the fiery atmospheric friction. Active cooling could

require more weight and complexity. There is also safety concern since it's an

active system. Often, however, the coolant is the fuel itself, much in the same

way that modern rockets use their own fuel and oxidizer as coolant for their

engines. Both scramjets and conventional rockets are at risk in the event of a

cooling failure

CURRENT PROJECTS

Recent advances in scramjet technology

5.1 There are several claims as to which group were the first to demonstrate a

"working" scramjet, where "working" in this case can refer to

(a) Demonstration of supersonic combustion in a ground test.

(b) Demonstration of net thrust in a ground test.

(c) Demonstration of supersonic combustion or net thrust in a ground test

with realistic fuels and/or realistic wind tunnel flow conditions.

(d) Demonstration of supersonic combustion in a flight test.

(e) Demonstration of net thrust in a flight test.

5.2 The problem is complicated by the release of previously classified material

and by partial publication, where claims are made, but specific parts of an

experiment are kept secret. Additionally experimental difficulties in verifying that

supersonic combustion actually occurred, or that actual net thrust was produced

mean that at least four consortiums have legitimate claims to "firsts", with several

nations and institutions involved in each consortium (For a further listing see

Scramjet Programs). On June 15, 2007, the US Defense Advanced Research

Project Agency (DARPA) and the Australian Defense Science and Technology

Organization (DSTO), announced a successful scramjet flight at Mach 10 using

rocket engines to boost the test vehicle to hypersonic speeds, at the Woomera

Rocket Range in Central Australia. No scramjet powered vehicle has yet been

produced outside an experimental program.

5.3 In recent years, significant progress has been made in the development of

hypersonic technology, particularly in the field of scramjet engines.US efforts are

probably the best funded, and the Hyper-X group has claimed the first flight of a

thrust-producing scramjet with full aerodynamic maneuvering surfaces. The first

group to demonstrate a scramjet working in an atmospheric test was a project by

an Australian team at the University of Queensland. The university's HyShot

project demonstrated scramjet combustion in July 30, 2002. The scramjet engine

worked effectively and demonstrated supersonic combustion in action, however

the engine was not designed to provide thrust to propel a craft, it was designed

more or less as a technology demonstrator. On Friday, June 15, 2007, the US

Defense Advanced Research Project Agency (DARPA), in cooperation with the

Australian Defence Science and Technology Organization (DSTO), announced a

successful scramjet flight at Mach 10 using rocket engines to boost the test

vehicle to hypersonic speeds. The following nations have active scramjet

programs: -

(a) UNITED STATES OF AMERICA

(i) NASP (CLOSED)

(ii) HYPER-X

(iii) HyTECH/HYSET

(iv) FASST

(v) FALCON

(vi) HyFLY

(b) AUSTRALIA

HYSHOT

(c) RUSSIA

(i) GLL Kholod,Igla (SA-5 MOUNTED)

(ii) GELA/RADUGA,ORYOL

(d) JAPAN

(e) GERMANY

SANGER

(f) CHINA

(g) UK

HOTOL(CLOSED)

(h) FRANCE

(i) PROMETHEE

(ii) JAPHAR

5.4 Detailed reports regarding the various projects being pursued by various

countries are given in references [4] to [10] including CFD analysis data and flight

data analysis (HyShot) data.

5.5 India has become force to reckon with its recent advancements in space

technology and the experience gained from its indigenous guided missile

development programme. As compared to the US, Indian efforts started very

recently in developing a scramjet engine/re-usable launch vehicle and we have

made rapid progress. As per reports, the ISRO has ground-tested scramjet

combustion for 6 seconds. The various projects in India for scramjet research

are: -

(a) ISRO

(i) CHANDRAYAN-RLV

(ii) AVATAR

(b) DRDL

HSTDV

(c) NAL

5.6 The “Hyper-X” X-43

Fig. 5.1 X-43 HYPERSONIC EXPERIMENTAL VEHICLE

In addition to the HyTech program the research efforts in scramjet technology

extended towards the Hyper-X vehicles. This project began in 1996 at NASA.

The hyper-X flew small test vehicle to demonstrate hydrogen fuelled scramjet

engine. In one of the test flight the X43A reached nearly Mach 10 speed.

Figure 5.2 HYPER-X FLIGHT TEST PROFILE

Since the ramjet/scramjet engine could not power itself to the engine operating

speeds, the plan for flight-testing was to attach the Hyper-X aircraft to a rocket

powered launch vehicle, and air launch the “stack” from underneath a B-52 wing.

The rocket powered launch vehicle was the Pegasus rocket, developed and built

by Orbital Sciences Corporation Launch Systems Group. However, the Pegasus

rocket to be used by the Hyper-X program was heavily modified. Initially a three-

stage launch system, the second and third stages was removed, leaving only an

Orion 50S solid rocket motor. The payload fairing normally used to protect

satellites was replaced with a ballast assembly and payload adapter for the

Hyper-X vehicle to be mounted in front. Furthermore, since the assembly would

spend all its time in the atmosphere, a new thermal protection system was added

to protect the booster’s composite structures (Orbital, 2002).

Indian efforts

5.7 The ‘Aerobic vehicle for hypersonic Aerospace Transportation’ (AVATAR)

is a hyper plane concept from India. This conceptual vehicle would be almost the

size of a MiG25 fighter airplane and is expected to deliver a payload of 500 Kg to

1000Kg to low earth orbit. It is expected to weigh 25 tonnes and is capable of

entering into a 100 Km orbit in a single stage and launching satellites which may

weigh up to 1 tonne. It is still in a conceptual stage and the estimated initial

development budget is around $5 million. BDL Hyderabad is involved in this

project.

5.8 Defense Research and Development Laboratory (DRDL) has also begun

the research on supersonic combustion with hydrogen and kerosene fuel. A test

facility has been designed, developed and successfully tested. The facility

simulates the test conditions prevailing at the entry of typical scramjet combustor

operating at flight Mach number of 5 and at an altitude of 25 Km. The facility

mainly consists of a Hydrogen burner for generating hot gases to simulate

stagnation temperature, a supersonic nozzle and a supersonic combustor with

fuel injection systems and flame holding mechanism. In the hydrogen burner,

gaseous hydrogen and air are burnt to achieve the required temperature while

oxygen is added to the gases to keep its volume fraction in the vitiated air same

as that of air. The supersonic combustor is made up of an isolator, constant area

combustor and divergent section. Between the isolator and constant area

combustor, rearward-facing steps are provided for improvement of ignition and

flame holding characteristics. Provision was made to inject the gaseous fuel

transversely into the supersonic vitiated air from the walls. The fuel injection is

carried out in a staged manner. A very little amount of fuel known as pilot fuel is

allowed into the combustor ahead of the rearward-facing step. Most of the fuel

named as primary fuel is injected behind the steps. The combustor is tested for

duration of 15 seconds.

5.9 The National Aerospace laboratory (NAL) Bangalore has also carried out

the self ignition studies of hydrogen. The ignition and performance

characteristics of hydrogen fueled stepped combustor have been determined. It

was observed that the hydrogen ignition temperature was substantially constant

in the stagnation pressure range tested and was also found to be independent of

number of ports through which hydrogen was injected. In general, it has been

recognized that hydrogen fuel ignition in supersonic combustion testing is

frequently facility dependent.

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16. WEBSITE www.aerospaceweb.org18. WEBSITE INTENTBLOG.COM