Supersonic Now

When the Concorde was gearing up to fight off environmentalists, Lockheed and Boeing were in the middle of trying to construct a supersonic transport for the United States. Due to budget issues and noise concerns, the projects both were abandoned and we’ve spent the rest of our lives tooling along at the universal atmospheric speed limit of Mach 0.80. While I have no problem with people lobbying against having to live with 15-30 booms per day (for those who live under jet routes, it would have been very disturbing), I do have an issue with the “oh well, forget it” mentality that plagues our society after the passing of regulations. We collectively said forget it, nobody needs to fly that fast because supersonic flight was banned via federal law.

Or was it?

FAR 91.817 states that:

(a) No person may operate a civil aircraft in the United States at a true flight Mach number greater than 1 except in compliance with conditions and limitations in an authorization to exceed Mach 1 issued to the operator under appendix B of this part.
(b) In addition, no person may operate a civil aircraft for which the maximum operating limit speed MM0exceeds a Mach number of 1, to or from an airport in the United States, unless—
(1) Information available to the flight crew includes flight limitations that ensure that flights entering or leaving the United States will not cause a sonic boom to reach the surface within the United States; and
(2) The operator complies with the flight limitations prescribed in paragraph (b)(1) of this section or complies with conditions and limitations in an authorization to exceed Mach 1 issued under appendix B of this part.


Now, I for one would like it if FAA regulations were written for regular people instead of law students, but it doesn’t take many mental cartwheels to understand this rule. Nowhere in the entire rule did it say that you could not fly past Mach 1 under any circumstance. Further more, subsection (b)(1) has given operators a stipulation that allows them to break Mach. What the FAA has stated is that supersonic flight that creates a sonic boom on the surface is prohibited. Technically, if one were able to fly supersonic without creating an audible boom, it is allowed. That means you can imitate Chuck Yeager all the way from TEB to DFW provided that your boom does not reach the surface.

If my decidedly non-legal interpretation is true, why have we as a nation stopped all development on civil supersonic flight? Other than the persistent efforts of Aerion Corporation and a few attempts by Gulfstream, Dassault and Lockheed Martin, there has been very little interest in going faster. I shudder to think that manufacturers were limiting their efforts based on misinterpretation of a rule (or maybe they had it right and it’s my misinterpretation). What is more likely is that faster speeds like Mach 2.0 and Mach 3.0 were what designers wanted to achieve. Engine efficiency goes way up due to ram compression and range can actually be improved with an increase in speed (to a certain limit). Going that fast in the stratosphere will always generate an audible boom, so there was no point in conducting research. The end result is that by taking a government rule and connecting it with a physical rule (flight beyond Mach 1 will always produce some sort of boom), manufacturers in effect put themselves out of the supercruise business.

But for every law of nature, there is a workaround (you may not get exactly what you wanted but what you get is better than nothing). There is a way to go fast without scaring the crap out of citizens. It takes advantage of the different temperatures in our atmosphere and uses them as a muffler. For years, scientists, engineers and pilots have known about a phenomenon called Mach Cutoff. In layman’s terms, it is a certain Mach speed that if exceeded, will result in an audible sonic boom on the ground. Below this speed, no sonic boom will be heard by people on the surface. This speed varies based on air temperature, weight of the aircraft and of course the elevation of the ground one happens to be flying over. For practical purposes, most aerodynamicists use the range between Mach 1.15 and Mach 1.2 as a standard cutoff.

Hidden by obscurity, Mach Cutoff did not do anything to spur development of faster aircraft until fairly recently. In those intervening years there was also very little research in how to fly with a quiet, muffled or ground-level silent sonic boom. The conventional wisdom said that all sonic booms were created equal and that you couldn’t avoid or mitigate them in any way. However, in the late 1990s and early 2000s there was a resurgence in high speed flight testing. NASA tested a modified F-5 with what appeared to be a boat hull fuselage. They also tested an extendable spike in the nose of an F-15 to see if creating a series of weak oblique shocks would be less offensive acoustically. Aerion tested a natural laminar flow wing section on the bottom of an F-15 to study the effects of supersonic forces on a straight wing. It appears that there is at least some interest in going fast if it can kept quiet.

What does this mean for a well-organized and equally well-funded designer? It means that there is no reason that designer can’t build something fast today. Of course this is very easy for me to say sitting here with exactly $17.81 in my pocket and not a single airplane built to my name. But a person doesn’t need to run an aerospace firm to know that there is a lot to be gained from building a proof-of-concept vehicle. Investors like something they can see, touch and sit in better than CAD generated images. Other developers become inspired and improve upon another firm’s work. Sometimes breakthroughs are made when idiosyncrasies that simply cannot be predicted in computer models are experienced and rectified in a flight test program.

It’s 2015 and we are well into the future that was predicted when most of us were children. I don’t have to say that while some things are an improvement, other things are complete letdowns. Flying at the same speeds that we attained in the 1950s is not a limitation of physics or economics, but of our own desires. We’ve become complacent and comfortable with what is essentially six decade old technology. Sure we’ve refined it and eked out far more efficiency than we ever imagined, but is that it? Are we supposed to be excited over another 1% fuel savings? Are we supposed to look at an aircraft with awe because it features a self contained 5G network?

The last 60 years were nice, but it’s time to go meet the future. Let’s move beyond where we are to where we should be.

CJ Sunset

Is this as fast as we’ll ever go?


Supersonic Engines

(This is not meant to be an exhaustive text on gas turbine theory and application, plus I’m trying to refrain from being my usual aero-nerd self. Some things are not mentioned like combustion chamber design and the varieties of turbine one could opt for. I know! This is meant to focus on the really critical parts of cobbling together a supercruise engine.)


An adequate powerplant for supersonic flight is a massive sticking point for anyone hoping to develop a civilian transport. Efficiency at high speed without excessive fuel consumption is a balancing act that becomes more difficult the less one is able to spend on exotic materials. For a civilian designer faced with a budget that resonates more with FBO than DOD, we should explore what is required with regards to creating a supersonic powerplant for non-military aircraft.

The most important thing for any aircraft flying faster than the Mach is to have an exhaust speed which is equal to (in actuality, slightly faster) than the true airspeed. The drawback is very high noise levels from exhaust shearing (like I care) and very high fuel consumption for a given thrust setting (this I really do care about). The first solution most people would consider is to increase the bypass ratio. The complicated reality is that there is not a single bypass ratio that is optimal for all phases of flight, especially for an aircraft with a 750+ knot speed range. What may help at Mach 0.60 will probably hurt at Mach 1.15. Furthermore, by increasing the fan size or fan rotational speed, the core of the engine has to work harder and either give up thrust or increase turbine temperatures to compensate. Any increase in fan diameter has to be weighed against these factors.

We can rule out large bypass ratios of greater than 5, such as those used for modern airliners. The fan diameter is too large and creates too much frontal area drag. In addition, the massive volume of cool exhaust flow is not able to move fast enough to push the aircraft to supersonic speeds. Theoretically one could take a large diameter fan and spin it faster (many large civilian turbofan engines spin their fans at speeds below 3000RPM) but this would require a lot of extra energy from the core. On the other hand, a military style bypass ratio of 0.2 to 0.8 is not enough to produce the required TSFC for a civilian aircraft that cannot refuel in-flight or carry external tanks. Thus somewhere between a bypass ratio of 1.0 and 5.0 is the optimal choice for our engine.

Way too wide for our purposes, but perfect for the C-17.

Way too wide for our purposes, but perfect for the C-17.

On the front end, the fan pressure ratio affects the specific thrust and indirectly, speed of the air through the engine. Specific thrust is the thrust divided by the inlet airflow in pounds per second. Low bypass engines tend to have very high specific thrust values while large high bypass turbofans have a very low specific thrust. Civilian turbofans usually use one large fan whereas high performance military turbofans typically use 3 or more fan stages for this exact reason (as a reference, the F100-PW-229 has a specific thrust of 71.77, the F101-GE-102 ranks at 48.84 and the GE90-B4 comes in at only 28.78). The now “hardened” bypass air is able to move through the engine with enough energy (combined with the core flow) to provide a choked nozzle condition at the exhaust orifice (Mach 1 flow at the narrowest point). The flow can then be exploited by a variable system aft of the choke point to accelerate the air to the required supersonic speed (I have simplified so much that it annoys me, but otherwise, this article would go on for days).

We also must mention overall pressure ratio, which is the amount of total compression achieved by the engine/inlet combination. A lower pressure ratio means fewer exotic materials required in the hot section, a lower engine weight, less extravagant methods of cooling and as a result, a lower manufacturing cost. On the downside, the thrust level is lower for a given engine as compared to the same engine with a higher pressure ratio. Another negative effect is that a lower overall pressure ratio raises TSFC for a given thrust setting. Depending on the aircraft budget and expected operating environment, trading extra fuel burn for lower initial cost may be acceptable. A word of note is that as Mach number increases, TSFC increases as well. This may be offset by ram pressure recovery (mentioned later) but it is important to know that the TSFC at Mach 0.80 is not going to be the same at Mach 1.3.

The intake setup is important even though strictly speaking it is a part of the airplane, not the engine. As an aircraft with the proper fan pressure ratio moves faster, it is able to delay thrust decay and in some cases, reverse the process thanks to ram recovery. But this is if and only if the intake is designed properly. At low supersonic velocities, a simple normal shock inlet is sufficient to allow adequate pressure recovery at the fan/compressor face. As velocities increase past roughly Mach 1.5, the losses mount exponentially and thrust will degrade accordingly. A multiple shock inlet can reduce these losses significantly, however there may be issues with shockwave placement at off-design speeds. In the quest for low cost and low weight, a fixed position normal shock inlet is probably the best choice for a civilian supersonic jet. If one wishes to engage in what the military terms “carefree handling”, then intake design must be given far more attention to ensure that certain flight conditions do not lead to disturbed flow. Turbulent air at the fan/compressor face can lead to surges and stalls of varying severity.

If it's good enough for the Viper, it should be good enough for us...fixed normal shock inlets are lightweight, inexpensive and have no moving parts.

If it’s good enough for the Viper, it should be good enough for us…fixed normal shock inlets are lightweight, inexpensive and have no moving parts.

Finally, the most important piece of the puzzle is the exhaust nozzle. Ideally, a jet engine exhausts air at ambient pressure to produce a stable column of thrust. A given engine can force air out at a higher pressure than ambient, but this flow will simply overexpand, collapse in upon its now low pressure core and possibly re-expand. This is very inefficient and can be hazardous to the aircraft’s operation. To allow the higher than ambient pressure flow to expand under control so that it’s energy is translated aft rather than radially, a divergent section of nozzle is required. Every angle made with respect to the convergent and divergent sections has a particular Mach number and pressure ratio associated with it. Knowing this, for an aircraft to have maximum efficiency across a wide range of Mach numbers, a variable convergent-divergent nozzle would be required. However, a variable exhaust nozzle is extremely complex to build and requires a system to activate it (oil or bleed air in most cases). A fixed nozzle will have far less efficiency but an attendant lower cost.

Controlled by bleed air, these F100-PW-220 nozzles are very complex.

To recap, we are in need of a low bypass turbofan probably between 1.0 and 5.0 with a high fan pressure ratio, moderate overall pressure ratio, adjustable exhaust and fixed inlet. Starting at the front of the engine, we can assume a bypass ratio of 2.5 for no other reason than it’s halfway (and back-of-the-napkin wit). With this we are assured of an acceptable frontal drag penalty, while still having a fan small enough to stage if required. To keep core requirements within a reasonable range, a specific thrust range of 50-70 allows us to move air fast enough without taxing the core too much. An added bonus is that multiple fan stages can eliminate the need for a low-pressure compressor altogether. At the rear of the engine, the exhaust nozzle should be adjustable to a certain extent. A dual-position nozzle may be the best alternative to a fully articulated iris nozzle when cost and complexity is considered.

As of now, there are several civilian engines that qualify with minimal modifications and many that could fit the bill with more extensive changes. In the interest of rapid development, low cost and minimal risk to the aircraft, a minimal-change option is the best choice for a civilian budget. The Williams FJ44-2 and Pratt & Whitney Canada JT15D are both contenders for small aircraft. The medium to large designs could be well served by the Rolls Royce Tay 611-8, Spey 511-8 or Pratt & Whitney JT8D with very few changes (certainly for less effort and expense than a cleansheet design). While some of these engines are no longer in production, there are enough examples to support a test program and even limited run manufacturing of aircraft.

This is a classic chicken vs egg issue. Engines will not be produced unless there is an airframe that requires them and airframes will not be built unless there is a reliable engine available to power it. Somebody has to blink first.

Delta Arrow Wings: Advantages For Civil Supersonic Flight

There is more than one way to skin a cat; this cat happens to be supersonic drag. There are many theories on how to achieve efficient civil supersonic flight, each with distinct advantages. One method that I have always favored is the simple and reliable delta wing. No complex construction, boundary layer control systems, or reliance on laminar flow. Just a big triangle without the need for high lift devices. There are tradeoffs but if your goal is to go supersonic without complexity, a good place to start is with the delta.

It is a fact that sweeping a wing delays supersonic drag rise and raises the critical mach number. It is also a fact that sweeping a wing with no special treatments will cause all kinds of hellacious stability problems at low speeds. While I could go into at least 4 pages of descriptions, case studies and NACA test data, I’m trying to write less like a mad scientist this year. So I’ll limit the following formulas to the basics in trying to get the point across.

An object moving faster than the speed of sound in air will produce a shockwave. The angle created by this shock cone (it’s a three dimensional wave) is dependent on the speed of the object. The faster the object, the smaller the angle created by the shock cone. With low supersonic Mach numbers, it is entirely possible to sweep a wing enough to contain it within the subsonic wake of the cone. The formula for determining the half-angle is:

1 / Mach#  = sin * cone angle

As stated, this is the half angle formula. To get what the entire cone would look like if drawn whole and not bisected, simply multiply the result by 2. For example, say that my aircraft is going Mach 1.3 and the wing is swept 54 degrees. With a cone half-angle of  50.3 degrees, my wing is definitely within the confines of the wake. This has a significant effect on reducing wave drag.

In addition to reducing wave drag, critical Mach number can also be reduced from sweep. Simply put, air is tricked into thinking that the wing has a longer chord and accelerates over the top of the wing at a slower rate. The formula for this effect is:

Vmach (cos * sweep angle) = Effective Vmach

This formula determines the Mach velocity over the wing when sweep is introduced. This may or may not be lower than the critical Mach number for that airfoil section. To determine what the new critical Mach number is when sweep is accounted for, the following formula is appropriate:

Mcrit / cos * sweep angle = Effective Mcrit

In effect, a highly swept wing can delay the critical Mach number to supersonic velocities. A straight wing with a relatively low Mcrit of Mach 0.7 would have an effective Mcrit of Mach 1.19 when swept to 54 degrees. This holds a lot of promise for reducing drag in the low supersonic speed range.

It’s not all free soda and candy for the swept wing. As mentioned before, there are very serious effects to consider aerodynamically. Swept wings stall at the tips first, creating unstable rolling and pitching moments during the stall. In other situations, pitch-up may occur when the horizontal stabilizer gets caught in the wing’s flowfield. Slow speed handling is degraded by spanwise flow, the same phenomenon that helps to reduce drag at higher speeds. Maximum lift coefficient for a given angle of attack is also reduced, leading to sometimes extreme pitch attitudes at slow speeds. There are other issues but these are the most critical to control and stability.

A straight wing that is swept may be troublesome, but delta wings have distinct advantages  that make them attractive for our purposes. The double delta design is a derivative that is configurable to nearly any range of speeds. The two main variants of the double delta are the “shovel”, with the low sweep segment in front, and the “arrow” with the high sweep segment in front. In either configuration, the forward segment produces low-pressure vortices that drift over the aft segment and delay the stall to a much higher angle of attack (non-linear lift). If properly balanced, a double delta will display highly favorable stability characteristics as well.  For this discussion, we will deal with the delta arrow variant since it is customized for supersonic speeds.


LoFlyte test vehicle at USAF Museum in Dayton, OH. Designed for hypersonic waveriding flight, this basic design is applicable to low supersonic flight as well.

A delta arrow consists of a highly swept leading edge section and a less severely swept aft wing section. This ensures that the leading edge of the wing remains behind the shockwave at moderate supersonic speeds while retaining adequate lift reserves for slow speed operations. As mentioned earlier, swept wings have a reduced lift curve slope with the degree of sweep directly correlated to the reduction (provided airfoil section and thickness remain constant). This disadvantage becomes an advantage in supersonic flight. All aircraft experience a rearward shift in the center of lift which reduces maneuverability and increases drag. Rather than rely on large control surface deflections to correct this situation, the delta arrow’s forward wing segment provides the lift required with minimal drag.

Airfoil thickness is a strong modifier of total drag at high subsonic speeds. A thin airfoil has far less drag than a thick one, but in trade, it has a lower lift curve slope, and less space for structure and fuel. A way around this is to sweep a moderately thick airfoil so that the effective thickness is reduced while retaining actual space inside for the structure and fuel. This phenomenon is more pronounced with large amounts of sweep, so the volume inside the forward segment of a delta arrow is quite extensive.

Maneuvrability is closely tied to the wing loading and center of gravity location. A heavily loaded wing will have a larger angle of attack in 1G flight, reducing the amount of lift available for aggressive maneuvering, regardless of aspect ratio. A reduced angle of attack can be achieved by moving the CG aft but within certain limits. Locating the CG too far aft would render the aircraft uncontrollable, even with fly-by-wire. A forward CG will reduce the control response and increase static angle of attack, but enhance stability. Balance between the two extremes will be determined by the aircraft’s purpose and desired handling capabilities.

The reduction in lift coefficient for a given angle of attack is subject to the same cosine formula that was applied to the critical Mach number earlier. Therefore the formula is as follows:

Cl ( cos * sweep angle) = Effective Cl

Assume an aircraft requires a Cl of 0.2 to sustain level flight at a given speed. If the aft wing is swept at 54 degrees, the effective Cl would be 0.12, demanding the aircraft increase its angle of attack to create 0.2 Cl. An alternative is to reduce wing loading, reducing the required Cl and by association, angle of attack. This reduction in lift is beneficial for flying in turbulence as the reaction of the aircraft to disturbances will be markedly reduced. While some people may not consider turbulence reaction to be a reason to reduce Cl slope, those who fly low-level, high-speed profiles, especially over warm areas or near mountains may have differing viewpoints. Not all civilian designed turbine aircraft have to be business jets.

So what does this mean in plain English? It means that economical and safe civil supersonic flight is possible. The reason we have not been able to achieve this so far is that industry has been focused on improving efficiency of existing designs. To integrate these advantages, a radical departure in construction has to be undertaken. The wing must be blended with the fuselage to keep drag and weight to a minimum without sacrificing strength. Thrust to weight has to be increased to ensure adequate acceleration is available at high altitudes. Low wing loading will not only improve induced drag numbers while subsonic, but reduce the impact of sonic booms at higher speeds.

The aviation industry has been hearing for at least 15 years about proposed civil supersonic aircraft. In each case the designs were business jets. With the tumultuous world economy of the early 21st century, no builder or prospective buyer has seen fit to invest in such a jet for understandable reasons. To date, no one has proposed building a manned research aircraft of much smaller size to investigate actual performance, handling, environmental effects and integration with the current ATC environment. The cost of a purpose-built test aircraft would be far less than attempting to build a full size business jet requiring full Part 25 certification. Something to think about for the frugal mavericks among us.